T.O.1F-16A-l

FLIGHT MANUAL
PUBNR
USAF/EPAF SERIES AIRCRAFT

036802

F-16A/B
BLOCKS 10 AND 15/J.q:'~:-,

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IF-16A-l-0001X

Commanders are responsible for bringing this publication to the attention of all Air Force personnel cleared for
operation of subject aircraft.

Published under authority of the Secretary of the Air Force.

14 AUGUST 1995
CHANGE 14

15 AUGUST 2003

10. 1 F- 16A- 1

DISTRIBUTION STATEMENT

Distribution is authorized to Department of Defense (DOD) components only (Direct Military Support).
Other requests for this document shall be referred to OO-ALCIYPVT, 6080 Gum Ln., Hill AFB, UT
84056-5825.
EXPORT CONTROL NOTICEWARNING: This document contains technical data whose export is restricted by the Arms Export Control
Act (Title 22, U.S.C. Sec. 2751 et seq) or the Export Administration Act of 1979 as amended (Title 50, U.S.C.
app. 2401 et seq). Violations of these export laws are subject to severe criminal penalties. Disseminate inaccordance with provisions of AFI 61-204.
HANDLING AND DESTRUCTION NOTICEFor unclassified, limited documents, destroy by any method that will prevent disclosure of contents orreconstruction of the document.
SUPPLEMENTAL NOTICEThis manual is incomplete without T.O. 1F-16A-1-1, T.O. 1F-16A-1-3 (Block 10), or T.O. 1F-16A-1-4 (Block 15
and l@]).
For RNLAF, when operating with the Orpheus Reconnaissance System, this manual is incomplete without T.O.
NE1F-16A-1.

SUPERSEDURE NOTICESee Technical Order Index, T.O. 0-1-CD-1, CAT-1-4, for current status of Flight Manuals, Safety Supplements,
Operational Supplements, and Flight Crew Checklists.
This manual supersedes T.O. 1F-16A-1SS-421.

T-2

Change 14

T. O. 1F- 16A- 1
Reproduction for nonmilitary use of the information or illustrations contained in this publication is not permitted. The policy for military
use reproduction is established for the Army in AR380-5, for the Navy and Marine Corps in OPNAVINST 5510.1B, and for the Air Forcein Air Force Instruction 31-401.

INSERT LATEST CHANGED PAGES. DESTROY SUPERSEDED PAGES.

LIST OF EFFECTIVE PAGES
NOTE:Changes to the current issue are indicated by a vertical line in the outer margins of the page.

Dates of issue for original and changed pages are:Original .. . o '" 14 Aug 95
Change ... . 1 ... 25 Mar 96
Change ... . 2 .... 10 Feb 97
Change ... . 3 .... 15 Sep 97
Change ..
4 .... 2 Mar 98

Change ... 10 ... Not Issued
Change .... 11 .... 15 Jul 02
Change .. , 12... 15 Nov 02
Change . .. 13 .... 15 Apr 03
Change ... 14... 15 Aug 03

Change ... . 5 , .... 4 Jan 99
Change ... . 6 '" 15 May 99
Change ... . 7 , .. 15 Mar 00
Change ... . 8 ' ... 15 Dec 00
Change .. ,. 9 ... Not Issued

TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS 584, CONSISTING OF THE FOLLOWING:
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:3-129 ... .

3-130.
3-131 ...
;:3-132 ... .
3-132.1 ............ .
3-132.2 Blank ...... .
3-13:3.. . .......... .
3-134 .............. ,3-135 ................ .
3-136 ......... , .
3-137 ......... , ....... .
3-138 ................. .
3-139 ................ .
3-140 ......... , ...... .
3-140.1 ....... , ...... .
3-140.2 .............. .
3-141 .... .
3-142 ... .
3-143 .. .
3-144 .. .
3-145 .. .
3-146 .. .

*Change
No.
1

1

1
8

8
13
8
13

o
5

o
o
o

o

5
11
11

:3
12
7

Page
No.

4-1
4-2 Blank ............. .
5-1 ................... .
5-2 ................... .
5-3 . .
. .......... .
5-4.....
. ....... .
5-4.1 ................. .
5-4.2 Blank ........... .
5-5 ................... .
5-6 ................... .
5-7 ........... .
5-8 .................. .
5-9 ................... .
5-10 ................... .
5-11 .................. .
5-12 .......... , ...... .
5-12.1 ................ .
5-12.2 ........ , ....... .
5-13 ............... '" .
5-14 .................. .
6-1 ................... .
6-2 ., ................ .

22

6-3 .,
6-4 ..
6-5 ., ................ .

o

6-6 .................. .

2

6-7 ........... .
6-8 ... .
6-9 ... .
6-10 .. .
6-11 ........... .
6-12 ................. .
6-13 ................. .
6-]4 .. .
6-15 ................ .
6-16 Blank ......... .
7-1 ................. .
7-2 ..
7-3 ..
7-4 .'
7-5 ., ...
7-6 ..
7-7 ..

11

13
12

o
11
11
11
11
11
4
4
4
4

3
4
7
8

8
11
11

11
11
6

7
8

*ChangeNo.

7-8 ................. .

8-1 .................. .
8-2 .................. .
8-3 ..
8-4 .'
8-5 . .
. ............ .
8-6 ............. .
8-7 ............. .
8-8 .. .

8-9 ..
8-10 ....
8-11 ..

o
o
3
8
11

11
8
2

o
o
13
8

1

o
o
8

11
8
13
3
8

o
o
o
3

o
4

4
4
4
4
4
8
8
8

8
12
13

1

o
8
8
8

13
13
13
1:3
1:3
13
13
13
13
1:3
13

* Zero in this column indicates an original page.

o

Change 14

Page
No.

*Change
No.

13
13
13
13
13
13
13
13
8-20.. . .... .. ....... .. .
13
8-21 . . . . . . . . . . . . . . . . . . .
138-22 ..................... 13
8-23.. . .... .. ....... . . .
13
8-24 . . . . . . . . . . . . . . . . . . .
13
8-25 ................. , .
138-26 . . . . . . . . . . . . . . . . . . .
138-27 . . . . . . . . . . . . . . . . . . .
138-28 ............ , . .... .
13
8-29 ...... , .. .... . .... .
13
8-30 ...... , .......... , .
138-31 . . . . . . . . . . . . . . . . . . .
13
13
8-32 ...... , ..... , .. . ....
8-33 .................. '" 13
8-34 ............ , ... .. .
13
8-35 ...... , .. " ., ..... .
13
8-36. . . . . . . . . . . . . . . . . . .
138-37 . . .. . . . . . . . . . . . . . . .
138-38 Blank. . . . . . . . . . . . .
13
Glossary_1 ............
13
Glossary_2 ............
13
Glossary_3 ...........
13
Glossary_4 ...........
13
Glossary_5 ............
13
Glossary _6 Blank ..
0

8-12 ..... .
8-13 . . . . . . . . . . . . . . . . . . .
8-14 . . . . . . . . . . . . . . . . . . .
8-15 . . . . . . . . . . . . . . . . . . .
8-16 . . . . . . . . . . . . . . . . . . . .
8-17 ...................
8-18 . . . . . . . . . . . . . . . . . . .
8-19 . . . . . . . . . . . . . . . . . . .

TO. 1F- 16A-l
FLIGHT MANUAL, SAFETY SUPPLEMENT, AND
OPERATIONAL SUPPLEMENT STATUS

This supplement status page is based on information available as of 15 August 2003. It is not an official status
page.
Flight Manual

Basic Date

Change No. and Date

T.O. IF-16A-l

14 Aug 1995

14

Supplemental Flight Manuals

Basic Date

Change No. and Date

T.O. IF-16A-l-1

28 Nov 199410

15 Feb 2003

T.O.IF-16A-I-3

15 Nov 1999

6

15 Jun 2003

T.O. IF-16A-I-4

15 Nov 1999

6

15 Jun 2003

Flight Crew Checklist

Basic Date

Change No. and Date

T.O.IF-16A-ICL-l

7 Mar 1994

14

15 Aug 2003

15 Aug 2003

INCORPORATED SAFETY AND OPERATIONAL SUPPLEMENTS

T.O. Number

Date

Short Title

IF-16A-lSS-421

03 Jul 03

Nozzle Failure Logic

Flight Manual
Sections Affected

II, IIIOUTSTANDING SAFETY AND OPERATIONAL SUPPLEMENTS

T.O. Number

Date

Short Title

Flight Manual
Sections Affected

NoneuI.'·.,/.:~

Change 14

a/(b blank)

T.O. 1F-16A- 1

TABLE OF CONTENTSSECTION I

Description and Operation

1-1

SECTION II

Normal Procedures

2-1

SECTION III

Emergency Procedures

3-1

SECTION IV

Crew Duties

4-1

SECTION V

Operating Limitations

5-1

SECTION V

Operating Limitations (Stores) - Block 10

T.O. IF-16A-I-3

Operating Limitations (Stores) - Block 15

T.O. IF-16A-I-4

SECTIONVand~SECTION VI

Flight Characteristics

6-1

SECTION VII

Adverse Weather Operation

7-1

SECTION VIII

Air Refueling Procedures

8-1

GLOSSARYStandard and Nonstandard Abbreviations

APPENDIX A

Performance Data - FI00-PW-200

T.O.IF-16A-l-l

APPENDIXB(Not Applicable)

T.O. IF-16A-l-l

APPENDIXC

Performance Data - FI00-PW-220 or
FI00-PW-220E

T.O. IF-16A-l-l

Glossary 1

Change 13

T.O. 1F-16A-l

BEFOREYOU
TAKE OFF,READ THIS!

1 F-16X-1--0002X@

SCOPE

This manual contains the necessary information forsafe and efficient operation of the aircraft. These
instructions provide a general knowledge of theaircraft and its characteristics and specific normal
and emergency operating procedures. Pilot experience is recognized; therefore, basic flight principles
are avoided. Instructions in this manual are prepared
to be understandable to the least experienced pilot
who can be expected to operate the aircraft. Thismanual provides the best possible operating instructions under most conditions. Multiple emergencies,
adverse weather, terrain, etc., may require modification of the procedures. This manual must be usedwith one or more of the following manuals to obtain
information necessary for safe and efficient operation:
T.O. IF-16A-l-l

Supplemental Flight
Manual, F-16AIB Aircraft

T.O. IF-16A-5-2
T.O. IF-16A-6CF-l

T.O. IF-16A-25-1

T.O. IF-16A-25-10

T.O. IF-16A-34-1-1

T.O.I-IM-44-1FD

T.O. IF-16A-34-1-3

T.O. IF-16A-1-3

Supplemental Flight
Manual, F-16AIB Aircraft (Block 10)

T.O. IF-16A-34-1-4

T.O. IF-16A-1-4

Supplemental Flight
Manual, F-16AIB Aircraft (Block 15 and

T.O.IF-16A-39~)
II

Change 13

T.O. l-lC-l

Loading Data
Acceptance and FUllctional Check Flight
Procedures Manual,
F-16AIB Aircraft
Nuclear Weapons
Delivery Manual(SECRET) (Title
Unclassified)
Aircrew PracticeBomb Delivery Procedures
Avionics and Nonnuclear Weapons Delivery Flight Manual
Combat Weapons
Delivery Software(CWDS)
Avionics and Nonnuclear Weapons Delivery Flight Manual ~Avionics and Nonnuclear Weapons Delivery Flight Manual M
Aircraft Battle Damage Repair
Basic Flight Crew Air
Refueling ProceduresI

T. O. 1F- 16A- 1•

The flight manual takes a positive approach andnormally states only what can be done. Unusual
operations or configurations are prohibited unlessspecifically covered herein. Clearance must be
obtained before any questionable operation which is
not specifically permitted in this manual is
attempted.

instructions in the supplement. \Vhen lengthy
additions are required, the formal supplement
provides one-sided insert page(s) to the flight manualand checklist. This supplement page(s) is attached tothe original page(s). The original page(s) remains in
the manual or checklist in case the supplement is
rescinded and the page(s) is needed. Added page(s)(e.g., 3-48.1) are inserted in proper numericalsequence and may be printed on both sides.

HOW TO BE ASSURED OF HAVING LATEST DATASAFETY SUPPLEMENTS

PERMISSIBLE OPERATIONSI Refer to T.O. 0-I-CD-l, CAT-I-4 for a listing of all
current flight manuals, safety supplements, operational supplements, and checklists. Also, check theflight manual title page, the title block of each safety
and operational supplement, and all status pages
contained in the flight manual or attacood to formal
safety and operational supplements. Clear all
discrepancies before flight.

Information involving safety is promptly forwarded
in a safety supplement. Urgent information is
published in interim safety supplements andtransmitted by electronic message. Formal supplements are mailed. The supplement title block andstatus page (published with formal supplementsonly) should be checked to determine the effect of this
supplement on this manual and other outstanding
supplements.

ARRANGEMENTI The manual is divided into eight sections and three
supplemental appendices.
ILLUSTRATIONS

Cockpit arrangement, cockpit console, and cockpit
instrument panel illustrations display the delivered
configuration plus the approved equipment modifications. For details of equipment modification, see theindividual equipment illustration.

OPERATIONAL SUPPLEMENTS

Information involving changes to operating procedures is forwarded by operational supplements. The
procedure for handling operational supplements isthe same as for safety supplements.
CHECKLISTThe checklist contains itemized procedures without
all of the amplification. Primary line items in theflight manual and checklist are identical.

SUPPLEMENT INFORMATION AND GUIDELINESHOW TO GET PERSONAL COPIESSupplements are safety or operational and are
indicated -ISS or -IS, respectively. Supplements areissued as interim electronic messages or formal
printed copies. All interim supplements are assigned
odd numbers, such as -ISS-195. When an interim
supplement is formalized, it will be assigned the next
following even number, such as -ISS-196. Formal
supplements not preceded by an interim supplement
are also assigned even numbers. If an interim
supplement is not to be formalized, a statement
cancelling the next assigned even supplement
number is included in the REMARKS section of theinterim supplement. If a formal supplement is notpreceded by an interim supplement, a statement
cancelling the previous odd supplement number is
included on the status page of the formal supplement.
Occasionally, a supplement has dual references in the
instructions; this is because the supplement applies
to the present and subsequent manual. Minor
text/illustration changes or deletions are given as

Each pilot is entitled to a personal copy of the flight
manual, safety supplements, operational supplements, and a checklist. The required quantities
should be ordered before needed to assure theirtimely receipt. Check with the publication distribution officer whose job is to fulfill T.O. requests.
Basically, the required quantities must be orderedfrom the appropriate T.O. Index. T.O. 00-5-1 and T.O.
00-5-2 give detailed information for properly ordering
these publications. Insure a system is established ateach base to deliver the publications to the pilots
immediately upon receipt.
FLIGHT MANUAL BINDERS

Looseleaf binders and sectionalized tabs are available for use with the manual. They are obtained
through local purchase procedures and are listed in
the Federal Supply Schedule (FSC Group 75, Office

Change 13iii

T.O. 1F-16A-1

Supplies, Part n. Check with supply personnel forassistance in procuring these items.
CHANGE SYMBOLI

The change symbol, as illustrated by the black line in
the margin of this paragraph, indicates changes
made to the current issue.
WARNINGS, CAUTIONS, AND NOTESThe following definitions apply to Warnings,
Cautions, and Notes found throughout the manual.

PILOT'S RESPONSIBILITY - TO LET US KNOW

Every effort is made to keep the flight manualcurrent. Review conferences with operating personnel and a constant review of safety investigation andflight test reports assure inclusion of the latest datain the manual. Comments, corrections, and questions
regarding this manual or any phase of the flight
manual program are welcomed. These should be
forwarded on AF Form 847 in accordance with AFI
11-215 through command headquarters to OO-ALCIYPVT, 6080 Gum Lane, Hill AFB, UT 84056-5825.
PUBLICATION DATEI WARNING I

The date appearing on the title page represents thecurrency of material contained herein.

Operating procedures, techniques,etc., which could result in personal
injury or loss of life if not carefully
followed.

AIRCRAFT AND COCKPIT DESIGNATION CODES

System and/or component effectivity for a particularaircraft version/cockpit is denoted by a letter code
enclosed in a box located in the text or on an
illustration. The symbols and designations are as
follows:
No code - F-16A and F-16B aircraft

Operating procedures, techniques,etc., which could result in damage toequipment ifnot carefully followed.

~

F -16A aircraft[ID

F-16B aircraft

NOTE

rnfl

F -16B aircraft, forward cockpit~

F -16B aircraft, rear cockpit~

Air Defense~

Belgium

USE OF WORDS SHALL, WILL, SHOULD, AND MAY~

Denmark

The word shall or will is used to indicate a mandatory
requirement. The word should is used to indicate anonmandatory desired or preferred method ofaccomplishment. The word may is used to indicate an
acceptable or suggested means of accomplishment.

~

Netherlands~

Norway

[g§]

USAF

~

USAFIUSN

An operating procedure, technique,etc., which is considered essential toemphasize.

USE OF WORDS AS DESIRED AND AS REQUIREDENGINE DESIGNATION CODESAs desired allows pilot preference in switch/control
positioning.
As required indicates those actions which vary basedon mission requirements.
AIRSPEED REFERENCES

All references to airspeed quoted in knots refer toindicated airspeed.
iv

Change 12

System and/or component effectivity for a particularengine version is denoted by an engine code enclosed
in a box located in the text or on an illustration. The
symbols and designations are as follows:
No code - Any engine

I PW200 I

FIOO-PW-200

I PW2201

FIOO-PW-220 or F100-PW-220E•

IO.lF-16A-l
HAVE QUICK RADIO DESIGNATION CODES

System and/or component effectivity for a particular
HAVE QUICK radio version is denoted by a HAVEQUICK radio code enclosed in a box located in the text
or on an illustration. The symbols and designations
are as follows:
No codeHAVE QUICK II or HAVE QUICK IIPHASE III HQ II IHAVE QUICK III HQ 11.2 tHAVE QUICK II PHASE IIBLOCK DESIGNATION CODES/SERIAL NUMBER/
TAIL NUMBER CROSS-REFERENCE

Because of differences in configuration between
aircraft and to avoid repetitious use of aircraft serial
numbers, a block effectivity system is used. The block
effectivities reflect the aircraft block, the aircraft
serial number, and the tail number. Attrited aircraft
are removed from the listing. This system is usedthroughout the manual, both in text and illustrations.

T. O. 1F- 16A- 1
AIRCRAFT MODIFICATION/RETROFIT INFORMATIONThis list includes the applicable T.O.IECP effectivities. It is not an official status page. Refer to T.O.
0-1-1-4 for the complete listing of TCTO's. Throughout this manual, black TV screen symbols containing
white numerals (D) are used to distinguishinformation related to aircraft which are modified by
TV CODE/
T.O.NO.
1F-16-1751

SHORT TITLE
Integration ofPressure Breathingfor G (PBG)(ECP 1677)

a specific T.O.IECP. Information pertaining tomodified aircraft is identified by an appropriate
effectivity symbol. Information which is not identified
by an effectivity symbol is considered common to all
aircraft. Information pertaining only to unmodified
aircraft has the appropriate effectivity symbol
preceded by LESS. For example, LESS 0 indicatesthat the information is only applicable to aircraft notmodified by a specific T.O.IECP.
EFFECTIVITYRETROFITPRODUCTION
~ NA

~ All

~

~ All

NA

~ NA

rrrn NA
INQ] NAI

Summary:• Retrofit authorized• Est start date - In work
• Installs a new oxygen system which provides pressure breathing for g (PBG).

@)

1F-16-1791

I

RLG Capability(ECP 1707A)

~ NA

~ TBD
~ TBD
INQ] All
~ All

~All

~ Complete

~ NA

~All

INQ] NA~ Complete
MAll
1M AllSummary:• Retrofit - As required• Est start date - In work
• The INS gimbal INU is interchangeable with the ring laser gyro (RLG) INU. The only cockpit indications that a RLG INU is installed are the shorter alignment times, greater accuracy, and different memory addresses.

VI

Change 1210. 1 F- 16A-l

TVCODEI
T.O. NO.
NA

EFFECTIVITYPRODUCTIONRETROFITSHORT TITLE
Early Provisions for
Improved Capability (ECP 0350)~

F-16A 80-0541
and on; F-16B
80-635 and onNA~ F-16A 80-35471

NAFA-56 and on;
F-16B 80-35881
FB-13 and on~ F-16A 80-3596 andNAon; F-16B 80-3612
and on[N;J F-16A 78-0258 andon; F-16B 80-3649
and onM F-16A 78-0300 andon; F-16B 80-3689
and onI8Nl All

NA
NA

I

Summary:• Retrofit NA• Multinational Staged Improvement Plan (MSIP) installs improved avionics and provides for CARA,AMRAAM, and LANTIRN.
BE1F-16-6007
NE1F-16-6001
N01F-16-6003

Incorporate Increased Area Horizontal Tail (ECP0425) (~mJ M
ECP 4003)~ F-16A 80-0541

NAand on; F-16B
80-635 and on~ F-16A 80-35471~ F-16A 78-0116/FA-01

FA-56 and on;
F-16B 80-3588/
FB-13 and on

thru 80-3546/FA-55;
F-16B 78-0162/FB-01
thru 78-0173/FB-12[Q§ F-16A 80-3596 andon; F-16B 80-3612
and onI8Nl AllNA

I~ Complete

M Complete
Summary:•••~ mJ I!iQl Retrofit authorized
Est start date - In work
Increased horizontal tail area to reduce rotation speed, increase departure resistance, and makeaircraft less susceptible to deep stall in the event of a departure.

Change 12VIIT. O. 1F- 16A- 1

TV CODE/
T.O.NO.

®

1F-16-1977

SHORT TITLE
Improved Switch
Guard for Fuel
Master Switch(ECP 2120)

EFFECTIVITYPRODUCTIONRETROFIT
All

NA

Summary:• Retrofit authorized• Est start date - In work
• Rescission date - TBD• Modification replaces existing in-line switch guard with a side guard configuration.

ff)

1F-16-1958

Z2 Software
ProvidesF100-PW-220E
Engine With PFL
Capability (OCP
20007)

~ NA

~ All

~ NA

~ Complete

IQ§ NA[N;] NA[NQ] NA

ml All[N;] All

rrn Complete

Summary:• Retrofit authorized• Est start date - In work
• Modification adds Pilot Fault List reporting capability to aircraft equipped with F100-PW-220E engine.

~

1F-16-1404
BE1F-16-6004

Modify WOW
Switch for Interruption of ChafflFlare
Dispense System(ECP 1212)~NA[!J§] F-16A 78-0001 thru
80-0540; F -16B All~NA

rn;J All

IQ§ NA[Q§ F-16A 78-0174 thru
78-0203; F-16B All[N;] NA~ F-16A 78-0212 thru

78-0257; F-16B All!NQ] NAI[fill Complete

Summary:~ All• Retrofit authorized• Est start date - In Work• Rescission date - Jul 01• ~ Rescission date - TBD• Incorporate an NLG WOW safety mechanism to prevent inadvertent chaffi'flare dispensing on the
ground.

VIII

Change 12

T.O. 1F-16A- 1
TVCODEI
T.O.NO.

®

IF-16-1404

EFFECTIVITYPRODUCTIONRETROFITSHORT TITLE~NA

Modify WOW
Switch for Interruption of ChaffIFlare
Dispense System(ECP 1212)~ F-16A 80-0541 and on;

F-16B NA
~NA

~NA

~NA

[Qg] F-16A 80-3596 and on;
F-16B NA~NA

~ F -16A 78-0258 and on:~NA

Irm Complete

F-16B NAI

Summary:~ All• Retrofit authorized• Est start date - In Work
• Incorporate a right MLG WOW safety mechanism to prevent inadvertent chaff/flare dispensing on
the ground.
~IF-16-1365

Provide Cockpit Indication of Probe
Heater Malfunction(ECP 1193)

~

NA

~

F-16A 87-0004
and on;F-16B 86-0197
and on1m All

~

F-16A 78-0001 thru
83-1117;
F -16B 78-0077 thru
83-1173~

Complete~

F-16A 78-0174 thru
80-3611;
F -16B 78-0204 thru
80-3615~ Complete~ Complete

I

Summary:• Retrofit authorized• Est start date - In Work• Modification installs probe heat monitoring system and PROBE HEAT caution light to provide
cockpit indication of possible probe heater failure or malfunction.

00 2J-F100(l)-594
2J-F1 00(II)-649

Redesigned 3rd
Stage Fan and New
4th Stage Compressor Blades

NA

All F100-PW-200 Engines

Summary:••••

Retrofit - TBD
Est start date - TBD
Rescission date - TBD
Modification reduces potential for 4th stage compressor blade failure. Completion of TCTO's removes the 600 KCAS limitation.

Change 12ix

T.O. 1F-16A-l
TVCODEI
T.O.NO.
(FI)

2J-FIOO-930

SHORT TITLE
Incorporation of
Arctic Trim to
PW200 Engines

EFFECTIVITYPRODUCTIONRETROFITNAAll FIOO-PW-200 Engines

Summary:•••

Retrofit authorizedStart date - Jun 98
Modification retrims engines to reduce occurrences of AB initiation stalls during takeoff.

00 IF-16-2170

Modification of theEjection Mode Selector Valve Console(ECP 2362)

NA

All F-16B

Summary:•••~

Retrofit authorizedStart date - Feb 99
Modification enlarges the hole in the console to allow the EJECTION MODE SEL handle to seatproperly in the NORM position.

5El-2-15-510

Incorporation of
EDU 2.4.0. OFP to
PW220 Engines

NA

All FIOO-PW-220 Engines

Summary:• Retrofit authorized• Start date - Feb 98• Modification installs EDU 2.4.0. OFP software for FIOO-PW-220 engines. Mter incorporation of
EDU 2.4.0A OFP (TCTO 5El-2-15-5110), ~ actions are not required.

00 IF-16-2077

Installation of
Improved Antiskid
Braking System

NA

~ All
~ All

(OCP 5003B)

Summary:•••~

Retrofit authorizedStart date - Jan 99
Modification incorporates an improved Brake Control/Anti-skid Assembly to replace the current
Brake Control and Antiskid Control Boxes.

IF-16-2348

Summary:
•
•
•
•

x

Replacement of
Cabin Low Pressure
Switch

NA

~
~

NA
All
[g] All
INEI AllINOI All

Retrofit authorizedStart date - Mar 02
Rescission date - TBD
Modification replaces the cabin low pressure switch, lowering CABIN PRESS caution light illumination from 27,000 to 22,500 feet.

Change 13

10. 1F-16A-l

Block Designation Codes/Serio'
Number/Tail Number Cross-Reference
BLOCK 10
Serial and Tall
Numbers Same for~[A]**78-0001##**78-0002##**78-0003#**78-0005**78-0007#**78-0008**78-0010#**78-0011#**78-0013##**78-0015**78-0017#**78-0019#**78-0020#**78-0021##*78-0022#*78-0024#*78-0025##*78-0026*78-0027##*78-0038#*78-0039##*78-0040##*78-0042##*78-0043##*78-0044*78-0047#*78-0049#*78-0050##*78-0051#*78-0052##*78-0053##*78-0054#*78-0056*78-0057##*78-0058#*78-0059##*78-0060#*78-0061*78-0062*78-0063#*78-0064#*78-0065##*78-0066##*78-0068#*78-0069*78-0070#*78-0073#*78-0074##*78-0076#*79-0290##*79-0294#*79-0296*79-0300#*79-0303*79-0306#*79-0307*79-0308#*79-0309##*79-0310#*79-0311#*79-0312#*79-0314#*79-0317#*79-0322#*79-0324*79-0326*79-0327*79-0329#*79-0330*79-0331#
79-0332
79-0334
79-0335##
79-0336#
79-0337##
79-0340##
79-0341#
79-0342#
79-034479-0345#
79-0346#
79-0348#
79-0349##
79-0351##
79-0352
79-0353#
79-0354#
79-0355#
79-0357
79-0359#
79-0360

79-0362
79-0363#
79-0364
79-0365#
79-0366##
79-0368
79-0370
79-0371
79-0373##
79-0375
79-0376#
79-0380##
79-0381#79-0382##
79-0383#
79-0384#
79-0387##
79-0388##
79-0389
79-0393#
79-0394##
79-0395#
79-0396#
79-0399##
79-0401##
79-0402##
79-0403##
79-0404
79-0405#
79-0406##
79-0407##
79-0408
79-0409#
80-0474#
80-0475#
80-0476#
80-0479#
80-0480#
80-048180-0482#
80-0483#
80-0485#
80-0487#
80-0488##
80-0489#
80-0492
80-0493#80-0494#
80-0495

80-0496#80-0497#
80-0498#
80-0499
80-0500#
80-0504
80-0505
80-0506#
80-0507#
80-0508#
80-0509#
80-0510#80-051180-0512#80-0513#80-0515#80-0518#80-0519#80-0520#80-0521#80-052280-0523#80-052480-0525#80-052680-052780-0528#80-0529#80-0530#80-0531#80-0533#
80-0535##
80-0537#80-0538#
80-0539#80-0540#~OO**78-0077#**78-0079#**78-0080**78-0081**78-0082**78-0083**78-0084**78-0085**78-0087**78-0088**RETROFITrED FROM BLOCK 1 *RETROFITrED FROM BLOCK 5 #STORAGE**78-0089**78-0090##**78-0091**78-0094**78-0096**78-0097##**78-0098*78-0099#*78-0100*78-0101*78-0102*78-0103#*78-0104#*78-0107#*79-0411*79-0412*79-0413##*79-0414*79-0415#*79-0417*79-0418#
79-042079-042179-042279-0426#
79-0427
79-0428
79-042979-0430
79-043179-0432
80-0623#80-0625#80-062880-0629#80-0630
80-063180-0633
80-0634~[AJ
Serial Number/
Tall Number**78-0 116/FA-O1#**78-0117/FA-02#**78-0118/FA-03### INACTIVE
Change 8xi

T. o. 1F- 16A- 1

Block Designation Codes/Serial
Number/Tail Number Cross-Reference**78-0 119/FA-04#**78-0 120/FA-05#**78-0 124/FA-09#**78-0125/FA-10#**78-0131/FA-16#**78-0 132/FA-17#*78-0 133/FA-18#*78-0 134/FA-19#*78-0 135/FA-20#*78-0 136/FA-21#*78-0 137/FA-22#*78-0 138/FA-23 #*78-0 140IFA-25#
78-0 141/FA-26#
78-0 142/FA-27#
78-0 143IFA-28#
78-0 145/FA-30#
78-0 146IFA-31#
78-0 147/FA-32#
78-0 149/FA-34
78-0151/FA-36#
78-0 152/FA-37#
78-0 153/FA-38#
78-0 155/FA-40#
78-0 158/FA-43 #78-0 159/FA-44#
78-0 160/FA-45 #78-0 161/FA-46#
80-3538/FA-47
80-3539/FA-48
80-3540/FA-49#
80-3541/FA-50#
80-35421FA-51#
80-3544/FA-53
80-3546/FA-55~oo
Serial Number/
Tall Number**78-0162IFB-01**78-0 163/FB-02**78-0 164IFB-03#**78-0 165/FB-04**78-0166IFB-05*78-0 168/FB-07*78-0 169/FB-08*78-017 01FB-09*78-0 171/FB-10
78-0173/FB-12**78-0 1741E-174**78-01761E-176*78-01771E-177*78-01781E-178*78-0 1801E-180*78-0181/E-181*78-0 1821E-182*78-0 1831E-183*78-0 1841E-184*78-0 1871E-187*78-0 1881E-188
78-0 1891E-189
78-01901E-190
78-0191/E-191
78-01921E-192
78-0 1931E-193
78-0 1941E-194
78-0 1951E-195
78-0 1961E-196
78-0 1971E-197
78-01981E-198
78-01991E-199
78-02001E-200
78-0202lE-202
78-02031E-203**78-0221/J-221**78-0222/J-222**78-0223/J-223*78-0226/J-226*78-0228/J -228
*78-0229/J-229*78-0230/J-230*78-0231/J-231*78-0232/J-232*78-0234/J-234*78-0235/J-235*78-0236/J-236
78-0238/J-238
78-0239/J-239
78-0240/J-240
78-0241/J-241
78-0242/J-242
78-0243/J -243
78-0245/J-245
78-0246/J-246
78-0247/J-247
78-0248/J-248
78-0249/J-249
78-0250/J-250
78-0251/J-251
78-0253/J-253
78-0254/J -254
78-0255/J-255
78-0256/J-256
78-0257/J-257

~oo

~oo

~oo

**78-0204IET-204*78-0206IET-206*78-0207IET-207*78-0208IET-208
78-0210IET-210
80-0626IET-626~oo**78-0212/J-212**78-0213/J-213**78-0214/J-214**78-0215/J-215**78-0218/J-218**78-0219/J-219**78-0220/J-220**78-0259/J-259**78-0260/J-260**78-0261/J-261**78-0262/J-262**78-0263/J-263**78-0264/J-264*78-0265/J-265*78-0266/J -266
78-0267/J-267
78-0268/J-268
78-0269/J-269
78-0270/J-270

MOO**78-0272/272**78-0273/273*78-0275/275*78-02761276*78-0277/277*78-0279/279*78-0281/281*78-0282/282*78-0284/284
78-0285/285
78-02861286
78-02881288
78-02891289
78-0291/291
78-0292/292
78-0293/293
78-0295/295
78-0297/297
78-0298/298
78-0299/299

MOO** 78-0302/302*78-0304/304
78-0305/305
78-0306/306

BLOCK 15
Serial and Tall
Numbers Same for~OO***80-0541#***80-0542#***80-0543#***80-0545#***80-0548#***80-0549#***80-0550
80-0551#***80-0552#***80-0553#***80-0554***80-0556#
80-0557***80-0558***80-0559**RETROFITTED FROM BLOCK 1 *RETROFITrED FROM BLOCK 5 #STORAGE***AIR DEFENSE FIGHTER
xii

Change 8***80-0560#***80-0561#***80-0562#***80-0563***80-0565***80-0567#***80-0568#***80-0569#***80-0570#***80-0571#***80-0572#
80-0573##***80-0575***80-0576#***80-0577#***80-0578***80-0579#***80-0580#***80-0581***80-0583#
80-0584***80-0587#***80-0588#***80-0589***80-0591#***80-0593#***80-0594***80-0596***80-0598#***80-0601***80-0602***80-0603***80-0604***80-0605***80-0607#***80-0608
80-0609***80-0611#***80-0612***80-0613###INACTIVE

f

T.O. 1F-16A-l

Block Designation Codes/Serial
Number/Tail Number Cross-Reference***80-0614#***80-0615***80-0616***80-0619#***80-0620#***80-0621#***80-0622
81-0663***81-0665***81-0666
81-0667***81-0668***81-0669#
81-0670***81-0673***81-0674***81-0675
81-0676
81-0677
81-0678
81-0679***81-0680***81-0681***81-0682#
81-0683***81-0685***81-0686#
81-0687
81-0688#***81-0690#***81-0691#***81-0693***81-0694#***81-0695***81-0696#***81-0698#***81-0699***81-0700#***81-0701#***81-0702#***81-0703***81-0705#***81-0707#***81-0708#***81-0709***81-0710#***81-0711#***81-0712***81-0713#***81-0715#***81-0716#***81-0718#***81-0719#***81-0720#***81-0721##***81-0722#***81-0723***81-0725#***81-0726#***81-0727#***81-0728#***81-0729#***81-0731***81-0732***81-0733#***81-0734#***81-0735#***81-0736#***81-0737#***81-0738#***81-0739#***81-0740#***81-0741#***81-0742#***81-0743#***81-0744#***81-0746#***81-0748#***81-0749#***81-0751***81-0752#***81-0753#***81-0754#***81-0755***81-0756#***81-0757***81-0759***81-0760#***81-0761***81-0762#***81-0763***81-0764#***81-0765#***81-0767#***81-0768#***81-0769#***81-0771#***81-0772***81-0773***81-0774***81-0775***81-0776***81-0777***81-0778***81-0780#***81-0781#***81-0782***81-0783***81-0784#***81-0785***81-0786***81-0787
81-0788***81-0789
81-0790***81-0791
81-0792#***81-0793
81-0794#***81-0795
81-0796***81-0797***81-0799
81-0800#***81-0801
81-0802***81-0803
81-0804***81-0805
81-0806***81-0807***81-0809#
81-0810***81-0811#
82-0900***82-0901
82-0902***82-0903
82-0904***82-0905
82-0906***82-0907
82-0908***82-0910
82-0911***82-0913
82-0914*** AIR DEFENSE FIGHTER *RETROFITTED FROM BLOCK 5***82-0915#***82-0916#***82-0917#
82-0918#***82-0919***82-0921
82-0922***82-0923#
82-0924***82-0926
82-0928***82-0929***82-0930
82-0931***82-0932
82-0933***82-0935
82-0936
82-0938
82-0941***82-0942#
82-0944#***82-0945#
82-0946***82-0947#
82-0948#
82-0949***82-0950***82-0951
82-0952#***82-0953***82-0955***82-0956
82-0957***82-0958#***82-0960#***82-0961
82-0962#***82-0963#
82-0964***82-0966#***82-0967
82-0968#***82-0969#
82-0970***82-0972***82-0973#***82-0974#
82-0975##STORAGE

82-0976
82-0977#***82-0978#***82-0979
82-0980***82-0981
82-0982#***82-0983***82-0984
82-0986#***82-0987
82-0988#***82-0989
82-0991***82-0992
82-0993***82-0995
82-0996***82-0997
82-0999#***82-1000***82-1001
82-1002#82-1004#***82-1005***82-1006
82-1007#***82-1008
82-1009***82-1010#***82-1012
82-1013***82-1014***82-1016#
82-1017#***82-1019
82-1020***82-1021
82-1022#***82-1023
82-1025
83-1066#
83-1068###INACTIVE
Change 8xiii

T.O. 1 F-16A-l

Block Designation Codes/Serial
Number/Tail Number Cross-Reference
83-1069#
83-1072#
83-1073#
83-1076#
83-1077#
83-1079
83-1080#
83-1081#
83-1083#
83-1084#
83-1085
83-1087
83-1088#
83-1090#
83-1091
83-1092#
83-1093
83-1094#
83-1095
83-1096#
83-1097
83-1098
83-1099
83-1100
83-1101#
83-1103
83-1104#
83-1105
83-1106
83-1108
83-1109
83-1110
83-1111
83-1112#
83-1113#
83-1114~[ID
80-0635***80-0636***80-0637
80-0638***81-0812
81-0813#81-081581-0816***81-0817***81-0818***81-0819***81-0820
81-082181-0822#***82-1026***82-1027***82-1031***82-1032#***82-1033***82-1034***82-1035***82-1036***82-1039***82-1041
82-1043***82-1046
82-1047***82-1049
83-1166
83-1167#
83-1168#
83-1169
83-1170
83-1171#
83-1172~OO

Change 887-0047IFA-103

87-0048/FA-104
87-0050/FA-106
87 -005l1FA-107
87-0052/FA-108
87-0053IFA-109

Serial Number/
Tall Number

80-3547/FA-56
80-3548/FA-57
80-3549/FA-58
80-3551/FA-60
80-3552/FA-61
80-3556IFA-65

80-3557/FA-66
80-3558/FA-67
80-3559/FA-68
80-3560IFA-69

80-3561/FA-70
80-3562/FA-71
80-35631FA-72

80-3564/FA-73
80-3565/FA-74
80-35661FA-7 5
80-3567IFA-76
80-3568IFA-77*** AIR DEFENSE FIGHTER
xiv

80-3569/FA-78
80-35721FA-81
80-3573/FA-82
80-3574/FA-83
80-3575/FA-84
80-3577/FA-86
80-3578/FA-87
80-3579/FA-88
80-3580/FA-89
80-3581/FA-90
80-35821FA-91
80-3583/FA-92
80-3584/FA-93
80-3585/FA-94
80-3586/FA-95
80-3587/FA-96
86-0073/FA-97
86-0074/FA-98
86-0075/FA-99
86-0076/FA-100
86-0077/FA-101
87-0046/FA-102

87-0054/FA-110
87-0055IFA-111
87-0056IFA-112

88-0039/FA-114
88-0040/FA-115
88-0041/FA-116
88-0042IFA-117

88-0043/FA-118
88-0044/FA-119
88-0045/FA-120
88-0046/FA-121
88-004 71FA-122
89-000 1/FA-123
89-0002/FA-124
89-00031FA-125
89-00041FA-126
89-0005/FA-127
89-00061FA-128
89-00071FA-129
89-00081FA-130
89-00091FA-131#STORAGE

89-0010/FA-132
89-001l1FA-133
90-0025/FA-134
90-0026/FA-135
90-0027/FA-136~[ID
80-3589/FB-14
80-3590/FB-15
80-3592/FB-17
80-3593/FB-18
80-3594/FB-19
80-3595/FB-20
87 -000 lIFB-21

88-00 171E-0 17
88-00 181E-0 18~[ID
80-3612IET-612
80-3613IET-613
80-3614IET-614
80-3615IET-615
86-0 1971ET-197
86-0 1981ET-198
86-0 1991ET-199
87 -0022IET-022[fOOOO

88-0048IFB-22

88-0049/FB-23
89-00 12/FB-24~OO
80-35961E-596
80-35971E-597
80-35981E-598
80-35991E-599
80-36001E-600

80-360 lIE-60 1
80-36021E-602
80-36031E-603
80-36041E-604
80-36051E-605

80-3606/E-606
80-3607/E-607
80-36081E-608

80-3609/E-609
80-3610/E-610
80-361l1E-611
82-10 ll1E-O 11
82-1024/E-024
83-10701E-070
83-107 41E-07 4
83-10751E-075
83-11071E-107
87-00041E-004

87 -OO05/E-005
87 -OO061E-006
87 -OO071E-007
87 -00081E-008
88-00 161E-0 16

80-3616/J-616
80-3617/J-617
80-3619/J -619
80-3620/J -620
80-3622/J -622
80-3623/J-623
80-3624/J -624
80-3627/J-627
80-3628/J -628
80-3630/J-630
80-3631/J -631
80-3632/J -632
80-3633/J-633
80-3635/J-635
80-3636/J-636

80-3637/J-637
80-3638/J -638
80-3640/J -640
80-3641/J-641
80-3642/J-642
80-3643/J -643
80-3644/J-644
80-3646/J -646
80-3647/J-647
80-3648/J -648
81-0864/J -864
81-08661J -866

T. O. 1F- 16A- 1

Block Designation Codes/Serial
Number/Tail Number Cross-Reference
81-0867/J-867
81-0868/J-868
81-0869/J-869
81-0870/J-870
81-08711J-871
81-0872/J-872
81-0873/J-873
81-087 4/J-87 4
81-0875/J-875
81-0876/J-876
81-0877/J-877
81-0878/J-878
81-0879/J-879
81-0881/J-881
83-1192/J-192
83-1193/J-193
83-1194/J-194
83-1196/J-196
83-1197/J-197
83-1198/J-198
83-1199/J -199
83-1201/J-201
83-1202/J-202
83-1203/J-203
83-1204/J-204
83-1205/J-205
83-1206/J-206
83-1207 /J-207
84-1360/J-360
84-1362/J-362
84-1363/J-363
84-1364/J-364
84-1365/J-365
84-1366/J-366
84-1367/J-367
85-0135/J-135
85-0 136/J-136
85-0137/J-137
85-0138/J-138
85-0139/J-139
85-0140/J-140
85-0141/J-141
85-0142/J-142
85-0143/J-143
85-0144/J-144
85-0145/J-145
85-0146/J-146
86-0055/J-055
86-0057/J-057

86-0058/J-058
86-0060/J -060
86-0061/J-061
86-0062/J-062
86-0063/J -063
87-0508/J-508
87 -0509/J-509
87-0510/J-510
87-0511/J-511
87-0512/J-512
87-0513/J-513
87-0514/J-514
87-0515/J-515
87-0516/J-516
88-0001/J-001
88-0002/J -002
88-0003/J -003
88-0004/J-004
88-0005/J-005
88-0006/J -006
88-0008/J -008
88-0009/J -009
88-0010/J-010
88-0011/J-011
89-0013/J-013
89-00 14/J-014
89-0015/J-015
89-0016/J-016
89-0017/J-017
89-0018/J-018
89-0019/J-019
89-0020/J-020
89-00211J-021~[ID
80-3649/J -649
80-3650/J-650
80-3651/J-651
80-3652/J-652
80-3653/J-653
80-3654/J-654
80-3655/J-655
80-3656/J-656
80-3657/J-657
81-0882/J-882
81-0884/J-884
81-0885/J-885
83-1208/J-208

83-1209/J-209
83-1210/J-210
83-1211/J-211
84-1368/J-368
84-1369/J-369
86-0064/J-064
86-0065/J -065
87 -0066/J -066
87-0067/J-067
87 -0068/J -068

MOO
80-3658/658
80-3659/659
80-3660/660
80-36611661
80-3662/662
80-3663/663
80-3664/664
80-3665/665
80-3666/666
80-3667/667
80-3668/668
80-3669/669
80-3670/670
80-36711671
80-3672/672
80-3673/673
80-3674/674
80-3675/675
80-3677/677
80-3678/678
80-3680/680
80-36811681
80-3682/682
80-3683/683
80-3686/686
80-3687/687
80-3688/688

87 -07121712~OO
90-0942/900942
90-0943/900943
90-0944/900944
90-0945/900945
90-0946/900946
90-0947/900947
92-0404/920404
92-0405/405
92-0406/406
92-0407/407
92-0408/920408
92-0409/920409
92-0410/920410~[ID
90-0948/948
90-0949/949
90-0950/950
90-0951/951
90-0952/952
92-0452/452
92-0453/453
92-0454/454
92-0455/455
92-0456/456
92-0457/457
92-0458/920458
92-0459/920459
92-0460/920460
92-04611920461

MrnJ
80-3689/689
80-3690/690
80-3691/691
80-3692/692
80-3693/693
87-07111711

Change 12xv

10. 1 F- 16A- 1

F-1 EiFighting Falcon
AB

1 F-16A-1-0003X @xvi

10. 1F- 16A- 1

SECTION I
DESCRIPTION AND OPERATIONTABLE OF CONTENTS

Page

The Aircraft . . . . . . . . . . . . . . . . . . . . .
Aircraft General Arrangement
Aircraft General Data ...........
Aircraft Gross Weight (NC GW) .
Cockpit Arrangement. .
Engine IPW200 I .. . . . . . . . . . . . . . . . . . . . . .
Engine General Description IPW200 I
Engine Fuel System I PW200 I ..........
Unified Fuel Control (UFC) I PW200 I
Electronic Engine Control (EEC)

1-6
1-6
1-6
1-6
1-6.1
1-42
1-42
1-42
1-42

I PW200 I . . . . . . . . . . . . . . . . . . . . . . . . .

1-42

Backup Fuel Control (BUC) SystemI PW200 I .. . . . . . . . . . . . . . . . . . . . . . .

Main Fuel Pump IPW200 I .....
Afterburner (AB) Fuel Pump

IPW200 I ........................
Compressor Inlet Variable Vanes(CIVV's) I PW200 I . . . . . . . . . . . . . . .
Rear Compressor Variable Vanes(RCVV's) IPW200 I ............
Compressor Bleed Air I PW200 I ..
Pressurization and Dump Valve

IPW200 I .................... _Exhaust Nozzle IPW200 I . . . . . . . . . . . . . . . .
Convergent Exhaust Nozzle Control(CENC) I PW200 I ................
Engine Oil System IPW200 I . . . . . .
Engine Anti-Ice System I PW200 I .....
Engine ANTI ICE Switch I PW200 I ....
Engine and Accessory Drive Gearboxes

IPW200 I ...........................
Engine Alternator I PW200 I ......
Engine Ignition System IPW200 I
Jet Fuel Starter (JFS) I PW200 I • . .
ENG & JET START Control
Panel I PW200 I ............
JFS Switch [8] mEl IPW200 I .....
JFS RUN Light [8] mEl IPW200 I
JFS Operation I PW200 I .......
Engine Controls and Indicators

IPW200 I ........................
EEC BUC Switch [8] mEl I PW200 I
EEC Caution Light IPW200 I . . . . .
BUC Caution Light IPW200 I . .

STARTING FUEL Switch [8] mEl
I PW200 I ...........•.......•.• _MAX POWER Switch I PW200 I . . .
RPM Indicator I PW200 I .............
NOZ POS Indicator I PW200 I . . . . . . . . .
FTIT Indicator I PW200 I .............
FUEL FLOW Indicator I PW200 I
OIL Pressure Indicator IPW200 I ..
HYD/OIL PRESS Warning LightI PW200 I . . . . . . . . . . . . . . . . . . . . . . .

ENGINE Warning Light I PW200 I . .
REDUCED IDLE THRUST (RIT)
Switch IPW200 I ................
Throttle I PW200 I ...............
Engine Operating Characteristics

1-44
1-44

IPW200 I .......................
Ground Operation I PW200 I . . . .

1-44

AB Operation in Flight IPW200 I
BUC Operation I PW200 I .............
Engine I PW2201 ... . . . . . . . . . . . . . . . . . . . . . . .
Engine General Description I PW2201 ....
Engine Fuel System I PW220 I ..........
Engine Control System I PW220 I . . . . . . . .
Main Fuel Control (MFC) I PW2201 ...
Afterburner (AB) Fuel Control

1-44
1-44
1-44
1-44
1-45
1-45
1-45
1-45
1-45
1-45
1-45
1-45
1-46
1-46
1-46
1-47
1-47
1-47
1-47
1-50
1-50

1-50
1-50
1-50
1-50
1-50
1-50
1-50
1-50
1-51
1-51
1-51
1-51
1-51

Non-AB Operation in Flight
I PW200 I . . . . . . . . . . . . . • • . . . . . .

1-51
1-53
1-53
1-54
1-54
1-54
1-54
1-54

PW2201 . . . . . . . . . . . . . . . . . . . . . . .

1-54

Digital Electronic Engine Control(DEEC) I PW2201 ...............
Secondary Engine Control (SEC)

1-56

I

I PW2201 . . . . . . . . . . . . . . . . . . . . . . . . .

Main Fuel Pump I PW2201 ...........
Afterburner (AB) Fuel PumpI

PW2201 . . . . . . . . . . . . . . . . . . . . . .

Compressor Inlet Variable Vane
(CIVV) Controll PW2201 .......
Rear Compressor Variable Vanes(RCVV's) I PW2201 ............. _Compressor Bleed Air I PW2201 ...
Pressurization and Dump ValveI PW2201 . . . . . . . . . . . . . . . . . . . . .

Exhaust Nozzle I PW2201 . . . . . . . . . . . . . . . .
Convergent Exhaust Nozzle Control(CENC) I PW2201 .................
Light-Off Detector (LO D) I PW220 I
Engine Diagnostic Unit (EDU)I PW2201 ...................
Engine Oil System I PW2201 . .
Change 2

1-56
1-56
1-56
1-56
1-56
1-56
1-57
1-57
1-57
1-57
1-57
1-57
1-1

T.O. 1 F-16A-l
EnbTine Anti-Ice System I PW220 I
Engine ANTI ICE Switch I PW2201 ....
Engine and Accessory Drive Gearboxes
I PW2201 ......................... .
Engine Alternator I PW2201 ..... .
Engine Ignition System I PW220 I .
Jet Fuel Starter (JFS) I PW2201 ..
ENG & JET START Control
Panel I PW2201 ........... .
JFS Switch [8J mEl I PW2201 .... .
JFS RUN Light [8J
I PW2201 .
JFS Operation I PW2201 ....... .
Engine Controls and Indicators
I PW2201 ................... .
ENG CONT Switch I PW2201 ...... .
AB RESET Switch [8J
I PW220 I .... .
ENGINE FAULT Caution Light~, rIDemem

I PW2201 .......................... .

1-57
1-58
1-58
1-58
1-58
1-59
1-59
1-59
1-59
1-59
1-60
1-60
1-60
1-60

ENGINE FAULT Caution LightLESS ~, LESS f1) I PW2201 ... .

SEC Caution Light I PW2201 ......... .
MAX POWER Switch[8J mEl I PW2201 ..
RPM Indicator I PW2201 ............ .
NOZ POS Indicator I PW2201 ..
FTIT Indicator I PW2201 ........ .
FUEL FLOW Indicator I PW220 I .
OIL Pressure Indicator I PW2201 ..
HYD/OIL PRESS Warning LightI PW2201 ....................... .
ENGINE Warning Light I PW2201 .. REDUCED IDLE THRUST (RIT)
Switch I PW2201 .............. .
Throttle I PW220 I .............. .
Engine Operating Characteristics
I PW2201 ...................... .
Ground Operations I PW2201 ..
Non-AB Operation in Flight
I PW2201 .................... .

AB Operation in Flight I PW2201 .
SEC Operation I PW2201 ........ .
Fire and Overheat Detection System
FIRE & OHEAT DETECT Test
Button [A] mEl ............ .
Fuel System ........ .
Fuel Tank System ..... _Fuel Transfer System ............. .
Fuel Tank Vent and Pressurization
System ....................... _Engine Fuel Supply System ..... .
Fuel Quantity Indicating System .... .
Reservoir Fuel Level Sensing System .
Fuel Low Caution Lights ...... .
HUD FUEL Low/Bingo Indication ... .
HUD TRAP FUEL Warning LESS ~1-2

Change 12

1-60
1-60
1-60
1-60
1-60
1-60
1-63
1-63
1-63
1-63
1-63
1-63
1-63
1-63
1-63
1-65
1-65
1-67
1-67
1-67
1-67
1-67
1-69
1-69
1-69
1-78
1-78
1-78
1-78.1

FUEL HOT Caution Light ...... .
Fuel Tank Explosion Suppression
System ..................... .
Refueling System ... .
Ground Refueling ......... .
Air Refueling (AR) System .... .
N\VS AIR DISC MSL STEP
Button ................ .
Air Refueling (AR) Status
Indicator ................. .
Environmental Control System (ECS)Electrical Failures .
Air-Conditioning .. _Pressurization .............. .
AIR SOURCE Knob [8J mEl "TEMP Knob [8J rntJ ...... .
DEFOG Lever [8J mEl ....... .
EQUIP HOT Caution Light ..... , .
Cockpit Pressure Altimeter [8J mEl .
CABIN PRESS Caution Light ... , ,Anti-G System .. - - - ...... Electrical System .......... .
Main AC Power System ......... .
Overcurrent Protection Units
Emergency AC Power System .
DC Power System [BQ] ........ DC Power System LESS ~ .
FLCS Power Supply ....... .
External Power Provisions ..... .
Electrical System Controls andIndicators ................ , . _Electrical Power Distribution .
Normal Operation .. Hydraulic System ..................... HYD PRESS Indicators and Warning
Light ........................... .
HYD PRESS Indicators ........... .
HYD/OIL PRESS Warning Light ... .
Emergency Power Unit (EPU) ....... .
EPU Controls and Indicators . , ... .
EPU Ground Safety Switch , ....... .
EPU Switch .............. - - - - - - - - EPU Run Light .. .
HYDRAZN Light .
AIR Light ............ .
EPU/GEN Test Switch ........ .
EPU FUEL Quantity Indicator[8J mEl ................ .
Hydrazine Leak Detector ..
EPU Fired Indicator ... EPU Operation ........ .
Landing Gear (LG) System ... .
Main Landing Gear (MLG) ..
Nose Landing Gear (NLG) .

1-79
1-79
1-79
1-79
1-79
1-80
1-80
1-80
1-80
1-80
1-80
1-80
1-83
1-84
1-84
1-85
1-85
1-85
1-85
1-85
1-85
1-85
1-85
1-86
1-86
1-90
1-90
1-90
1-90
1-101
1-101
1-101
1-101
1-101
1-104
1-104
1-104
1-106
1-106
1-106
1-106
1-106
1-106
1-106
1-106
1-107
1-107
1-107

10. 1 F- 16A- 1

Landing Gear Controls and Indicators ..
Landing Gear Handle ....... .
Landing Gear Handle Down
Permission Button ...... .
DN LOCK REL Button .
ALT GEAR Handle
LG Warning Horn ........ _HORN SILENCER Button .... .
T.O.ILAND CONFIG Warning
Light .....................
WHEELS Down Lights .....
Landing Gear Weight-on-Wheels(WOW) Switches ............ .
Landing Gear Operation ........ .
Nosewheel Steering (NWS) System ..
NWS Controls and Indicators ........ .
NWS NR DISC MSL STEP Button ..
NWS Light .............. .
NWS FAIL Caution Light ........ .
Nosewheel Steering (NWS) System [ID .
Wheel Brake System ..
Toe Brake System .......... .
Spin Down Braking System ... .
Brakes Channel Switch lID rnfJ ..
Parking Brake ~ ~ ........ .
Antiskid System ........... .
ANTI-SKID Switch lID mIl .. .
ANTI SKID Caution Light ..
Speedbrake System .......... .
SPD BRK Switch ................ .
SPEED BRAKE Position Indicator
Drag Chute System m1 ~ ........ .
DRAG CHUTE Switch m1 ~ ..
Arrestment System.
HOOK Switch ....... .
HOOK Caution Light .
Wing Flap System ................ _Leading Edge Flaps (LEF's) .......... .
LE FLAPS Switch ................ .
LE FLAPS Caution Light ..
Trailing Edge Flaps (TEF's)(Flaperons) ............ .
ALT FLAPS Switch ..... .
Flight Control System (FLCS)
FLCS Limiters ....... .
AOAlG Limiter ... .
Roll Rate Limiter ........ .
Rudder Authority Limiter
Yaw Rate Limiter .
FLCS Gains ....
Cruise Gains .
<

<

<

<

<

,

••••

••••••••••

1-107
1-107
1-107
1-107
1-107
1-109
1-109
1-109
1-109
1-109
1-109
1-113
1-113
1-113
1-113
1-113
1-113
1-113
1-113
1-114
1-114
1-114
1-114
1-115
1-118
1-118
1-118
1-118
1-118
1-118
1-120
1-120
1-120
1-120
1-120
1-120
1-120
1-120
1-120
1-121
1-121
1-121
1-121
1-121
1-121
1-127
1-127

Takeoff and Landing Gains
1-127
Standby Gains ..........
1-127
FLCS Data Recorder ~ mEl .
1-127
Gun Compensation. . . . . . . . . . . .
1-127
Flight Control System (FLCS)
Controls .................
1-127
Stick .......... _1-127
Rudder Pedals. . . . . . . . . . . . . . .
1-130
1-130
MANUAL TRIM Panel ~ illf] ......
MANUAL PITCH Override (MPO)
Switch ........................
1-130
STORES CONFIG Switch lID illf] .
1-130
1-130
Low Speed Warning Tone. . .. . .. . . .
FLIGHT CONTROL Panel (FLCP)~ ~ ............................ 1-130
FLCS Warning, Caution, and Indicator
Lights ............................
1-130
T.O.ILAND CONFIG Warning
Light. . . . . . . . . . . . . . . . . . . . . . .
1-130
FLT CONT SYS Caution Light. .
1-130
DUAL FC FAIL Warning Light .
1-130
Trim .............. " . .. . . . . . . . .. . ..
1-132
1-133
Autopilot ............. _1-133
Autopilot Operation . . _Stick Steering ....
1-133
AOA System ..................
1-138
AOA Indicator . .
1-138
AOA Indexer .......
1-138
HUD AOA Display. .
1-138
1-139
Air Data System. . . . . . . . . . .. ...........
1-139
Air Data Probes ...................
1-139
AOA Transmitters ...... .
1-139
Total Temperature Probe .
1-139
Static Pressure Ports ...
1-139
Probe Heat Monitor ~ ., .......... .
1-139
PROBE HEAT Switch. . . . .. . ..... .
1-139
Air Data Converter (ADC) .
1-141
ADC Caution Light ............... .
Central Air Data Computer (CADC) .... . 1-141
1-141
CADC Caution Light .............. .
1-141
Warning, Caution, and Indicator Lights ..
1-141
Voice Message System (VMS) ........ .
1-143
VOICE MESSAGE Switch rru rn£l ... .
1-143
MASTER CAUTION Light ......... .
1-143
Caution Light Panel ................. .
1-152
MAL & IND LTS Test Button.. 1-152
Pilot Fault List (PFL) ...
1-152.1
Lighting System ..... .
Exterior Lighting .......... ,1-152.1
1-152.1
Anticollision Strobe Light ,1-152.1
PositionIFormation Lights .. _1-152.1
Air Refueling Lights ...... - ... .
<

••

_

<

Change 6

1-3

10.lF-16A-1

Vertical Tail-Mounted
Floodlight (illj ......... .
1-152.1
Landing and Taxi Lights ... .
1-154
Identification (ID) Light ~ ...... .
1-154
Identification (ID) Light ~ M ~ ..
1-154
Interior Lighting ................... .
1-154
PRIMARY CONSOLES Knob .
1-155
PRIMARY INST PNL Knob ... .
1-155
NAV/FREQ DISP Knob ~ ~ .. .
1-155
FREQ DISP Knob ~ ........ .
1-155
MAL & IND LTS Switch .. .
1-155
FLOOD INST PNL Knob .. .
1-156
FLOOD CONSOLES Knob .
1-156
Utility Light ...... .
1-156
Cockpit Spotlights .
1-157
Escape System ...... .
1-157
Canopy ....................... .
1-157
Canopy Controls and Indicators .... - 1-157
MANUAL CANOPY CONTROL
Handcrank ................. _1-157
Canopy Handle ~ ~ .
1-157
Canopy Jettison ......... , ........ .
1-157
Ejection Seat ............... .
1-157
Ejection Seat Controls andIndicators .............. ______ _1-163
Ejection Handle .................. .
1-163
SHOULDER HARNESS Knob ..... 1-163
EMERGENCY MANUAL CHUTE
Handle ...................... .
1-163
Ejection Safety Lever ............. .
1-163
SEAT NOT ARMED Caution Light ..
1-163
SEAT ADJ Switch ................. ' 1-163
Shoulder Harness StrapsfParachute
Risers ......................... .
1-163
Inertia Reel Straps .
1-163
Lapbelt .......................... .
1-163
Survival Kit ........................ . 1-163
RADIO BEACON Switch ~ ~ ~~~ ......................... .
1-163
Emergency Oxygen ............... .
1-164
Ejection Seat Operation ............. .
1-164
Ejection Mode Envelopes .
1-164
Ejection Sequence Times ...... .
1-164
Canopy Jettison/Seat Ejection ..
1-164
Ejection Seat Performance .1-164
Oxygen System .......... .
1-173
Liquid Oxygen System ...... .
1-173
Pressure Breathing for G
(PBG)
1-173
Oxygen System Controls andIndicators ................ .
1-173
Oxygen Duration ....... .
1-173
Emergency Oxygen System .
1-173
Oxygen System Schematic
1-173

m ................ .

1-4

Change 12

Communications System .
Antenna Locations ............... .
Have Quick (HQ) System LESS ~ .
Single Word of Day (WOD) ... .
Single WOD Transfer ......... .
Multiple Word of Day (MWOD) .
Automatic M\VOD Loading
I HQ 11.2 I...............•....

Time-of-Day (TOD) Transmission .
Time-of-Day (TOD) Reception ..
Net Number ..................... ,Antijamming (AJ) Mode Operation
HAVE QUICK (HQ) UHF
Radio LESS ~ .......... .
UHF Radio Control Panel
Function Knob ..... - - - l\1ode Knob I HQ I I I ... .
Mode Knob I HQ I I . 2 I ..
CHAN Knob ............. .
Manual Frequency Knobs
A-3-2-T Knob I HQ II I .. _A-3-2 Knob I HQ 11.2 I ..
VOL Knob .......... .
SQUELCH Switch. . . . . .
TONE Button I HQ II I . . . . .
CHAN Display I HQ II. 2 I .
Frequency/Status Display

1-180
1-180
1-180
1-180
1-180
1-180
1-182
1-182
1-182
1-182
1-182
1-182
1-182
1-183
1-183
1-183
1-183
1-184
1-184
1-184
1-184
1-184
1-184
1-184

I HQ I I . 2 I •.••••.•.•..•.
•• 1-184
STATUS Button I HQ 11.2 I •,. 1-184.1
T-TONE Switch I HQ 11.2 I ........... 1-184.1
TEST DISPLAY Button I HQ 11.2 I ... 1-184.1
ZERO Switch I HQ II. 2 I ........... " 1-184.1
FILL Connector I HQ 11.2 I .
,. 1-184.1
UHF RADIO ~ ............ ,,. 1-184.1
UHF Radio Control Panel .
,. 1-184.1
Function Knob . __ ... __ .. ___.. 1-184.1. ........... 1-184.1
Mode Knob.....
CHAN Knob...... .... .. ..
" 1-184.1
Manual Frequency Knobs .
1-185
VOL Knob....... . . . . . . . .
1-185
SQUELCH Switch. ,1-185
TONE Button ...............
1-185
UHF ANT SEL Switch~ ~ .
1-185
VHF Radio ....................
1-185
VHF Radio Control Panel. .
1-185Function Knob ........
1-186
Preset CHAN Knob.
1-186
1-186
Mode Knob. . . . . . . . . . . . . . .
Manual Frequency Knobs .
1-187
VOL Knob ............ 1-187
SQ DIS TONE Switch. . . .
1-187
LOAD Button crfj ~ ~ . . . . .
1-187
1-187
WB MEM LOAD Switch [Illi] .

10. 1 F- 16A- 1

HF Radio I@l [A] . . . . . . . . . . . . . . . . .
HF Radio Control Panel ~ ~ .,....
OFF VOL Control Knob ............
OFF SQL Control Knob ..............
USB AM LSB Select Switch ..........
Preset Channel Select Knob .

1-187
1-187
1-188
1-188
1-188
1-188

CHAN Display ......... .
Frequency Select Switches ..

1-188
1-188

FREQ Display ..... .
LOAD TEST Switch ..

1-188
1-188

PRE Tune Light .

1-188

Change 121-4.1/(1-4.2 blank)

10. 1F- 16A- 1

XMIT Light ..... .
1-188
1-188
MAN Tune Light ..
UNTUNED Light .......... .
1-188
Built-In Test (BIT) Function ..... .
1-189
Radio Channel/Frequency Indicator(RCFI) ....................... ..
1-189
COMM Control Panel ~ rnfJ .. .
1-189
HOT MIC CIPHER Switch .
1-189
INRCOM/SV Knob .......... .
1-190
THREAT Tone Knob . ..
1-190
MSL Tone Knob .......... .
1-190
Master Zeroize Switch ~ rnfJ .,1-190
Secure Voice System LESS IBNI ....... .
1-190
1-190
Secure Voice Control Panel ~POWER Switch ............... .
1-190
ZEROIZE Switch
1-190
DELAY Switch .
1-191
FILL Switch ....
1-191
MODE Switch ........ .
1-191
1-191
PLAIN Cipher Switch. " .
Inertial Navigation Set (INS) ........... .
1-191
Tactical Air Navigation (TACAN) System .
1-192
1-192
TACAN Control Panel ............... .
TACAN Function Knob .. .
1-192
VOL Knob ........ .
1-192
1-192
XIY Channel Ring '"Units Channel Knob .............. .
1-192
Hundreds-Tens Channel Knob
1-192
1-192
CHANNEL Display .' .....
1-192
TEST Pushbutton ..
1-193
TEST Light ................. .
Instrument Landing System (lLS)
1-193
ILS Control Panel ....... .
1-193
ILS Frequency Knobs .
1-193
1-193
ILS PWR Switch ..
1-193
ILS VOL Ring .................... .
MRK BCN Light .................. .
1-193
IFF System ........................... .
1-194
IFF Advisory Function ............... .
1-194
IFF Control Panel ..................... . 1-194
1-194
Mode 4 CODE Knob ~ ........... .
Mode 4 CODE Knob LESS ~ .
1-194
REPLY Light IBNI ............ , .... .
1-194
1-195
REPLY Light LESS ~ ............ .
TEST Light ...................... .
1-195
IFF MASTER Knob ..... .
1-195
RAD TEST MON Switch .
1-195
1-195
IFF IDENT Switch .... '" ....... .
1-195
Mode TEST Switches .
1-196
Code Wheels ....... .
1-196
MODE 4 SwitchM ...... .
1-196
MODE 4 Switch LESS IBNI .... .
1-196
MODE 4 Monitor Switch ~ ..

em ..

MODE 4 Monitor Switch LESS ~ ..
IFF ANT SEL Switch ~ rn.El ..... .
IFF IDENT Button ....... .
IFF Caution Light
AIFF System ~ .....
Transponder ~Interrogator!BQ] .......... .
IFF Advisory Function [@]AIFF Control Panel!BQ] ....
Mode 4 CODE Knob !BQ]
M-4REPLYLight~ ....
AIFF MASTER Knob !BQ]
M-4 Switch!BQ] ........... .
Mode 4 Monitor Switch !BQ]
M-l Switch!BQ]
M-2 Switch!BQ] ...
M-3/A Switch [@]M-C Switch ~ ................ .
IFF ANT SEL Switch !BQ] ~ rnfJ .
IFF IDENT Button ~ ........ .
IFF Caution Light ~ ..
INSTR MODE Select Panel ....
INSTR HDG Knob ~ rn£l .
INSTR MODE Knob ........ .
Navigation Aids and Display ..
Instrument Modes . . ..... .
Flight Instruments ..
Altimeter ............... .
AirspeedlMach Indicator ......... .
Standby Attitude Indicator (SAl) .
Vertical Velocity Indicator (VVI) ..
Magnetic Compass .............. ,Attitude Director Indicator (ADI) ..... .
Horizontal Situation Indicator (HSI) ..
Clock ..................... .
Standard Flight Data Recorder(SFDR) [BQ] ................. .
F-16B Aircraft ............ .
Engine Controls IPW200 I ..
Throttle IPW200 I ~ ....... .
BUC Switch IPW200 I ~ .. .
Engine Controls I PW2201 .. .
Throttle I PW220 I ~ ........... .
ENG CONT Switch I PW2201 ~Fuel System .................... .
Fuel Control Panel [ID!J
Speedbrake System ..... .
SPD BRK Switch [ID!] .... .
Drag Chute System [N;] IBQ] ......... .
DRAG CHUTE Switch [frn ~ ~Landing Gear (LG) System ........ .
DN LOCK REL Button [ID!JChange 12

1-196
1-196
1-196
1-196
1-196
1-196
1-197
1-197
1-197
1-198
1-198
1-198
1-198
1-198
1-198
1-198
1-198
1-198
1-199
1-199
1-199
1-200
1-200
1-200
1-200
1-200
1-207
1-207
1-207
1-207
1-209
1-209
1-209
1-209
1-209
1-209
1-212
1-212
1-212
1-212
1-212
1-212
1-212
1-212
1-212
1-212
1-212
1-212
1-212
1-213
1-213

1 -5

lO. 1 F- 16A-1

Flight Control System
Stick Control Operation ........ .
STICK CONTROL Switch rnfJ ..
Stick Indicator ~ ........... .
Stick OVRD Lights [ID ..
Escape System rnJ ................... .
EJECTION MODE SEL Handle~ .
Communications System [ID .......... .
INSTR MODE Select Panel [ID ..
Flight Instruments ~ . - ..... Accelerometer ~ ........ .
RadarlElectro-Optical (REO)
Display~ ................. .
Stores Management Subsystem(SMS)~ .................. _Fire Control Radar (FCR) ~ ..
Throttle and Stick Switches/
Controls ~ .............. .
ARMT CONSENT Switch ~ .
Servicing Diagram

1-213
1-213
1-213
1-213
1-213
1-213
1-213
1-214
1-214
1-214
1-214
1-214
1-214
1-214
1-214
1-214
1-214

THE AIRCRAFTThe F -16A is a single-engine, single-seat, multirole
tactical fighter with full air-to-air and air-to-surface
combat capabilities. The F-16B is a two-seat (tandem)
version and performs the secondary role of a trainer.
The fuselage is characterized by a large bubble
canopy, forebody strakes, and an under fuselage
engine air inlet. The wing and tail surfaces are thinand feature moderate aft sweep. The wing has
automatic leading edge flaps which enhance
performance over a wide speed range. Flaperons aremounted on the trailing edge of the wing and combine
the functions of flaps and ailerons. The horizontal
tails have a small negative dihedral and provide pitch
and roll control through symmetrical/differential
deflection. The vertical tail, augmented by twin
ventral fins, provides directional stability. All flightcontrol surfaces are actuated hydraulically by two
independent hydraulic systems and are directed bysignals through a fly-by-wire system.
The fire control system includes a fire control radarwith search and tracking capability, a radar electro-optical (REO) display, and a head-up display (RUD). A
stores management system (SMS) presents a control
panel and visual display for inventory, control, andrelease of all stores. Basic armament includes afuselage-mounted multibarrel 20 mm gun and an
air-to-air missile on each wingtip. Additional stores ofvarious types can be carried on pylons mounted under
the wings and on the fuselage centerline.

1-6

Change 2

AIRCRAFT GENERAL ARRANGEMENT

Refer to figure 1-1 for general arrangement andoverall view of the aircraft.
AIRCRAFT GENERAL DATA

Refer to figure 1-2. The aircraft approximate
dimensions are:• Span - including missile
fins

32 feet 10 inches• Length - including noseprobe

49 feet 6 inches• Height - top of vertical tail16 feet 8.5 inches• Height - top of canopy

9 feet 4 inches• Tread• Wheelbase= 7 feet 9 inches

13 feet 2 inches

Refer to TURNING RADIUS AND GROUND
CLEARANCE, Section II.
AIRCRAFT GROSS WEIGHT (A/C GW)[.B] IPW200 I Block 10 The GW of the aircraft including
pilot, oil, two tip AIM -9 missiles, and a full load of 20mm ammunition is approximately 17,500 pounds andwith full internal JP-8 fuel 24,800 pounds. I PW2201
Aircraft is approximately 100 pounds heavier.

[ID IPW200 I Block 10 The GW of the aircraft including
two pilots, oil, two tip AIM-9 missiles, and a full loadof 20 mm ammunition is approximately 18,300
pounds and with full internal JP-8 fuel 24,400
pounds. I PW2201 Aircraft is approximately 100
pounds heavier.
~ I PW200 I Block 15 The GW of the aircraft including

pilot, oil, two tip AIM-9 missiles, and a full load of 20mm ammunition is approximately 17,900 pounds andwith full internal JP-8 fuel 25,200 pounds. I PW2201
Aircraft is approximately 100 pounds heavier.

[ID IPW200 I Block 15 The GW of the aircraft including
two pilots, oil, two tip AIM-9 missiles, and a full loadof 20 mm ammunition is approximately 18,700
pounds and with full internal JP-8 fuel 24,800
pounds. I PW2201 Aircraft is approximately 100
pounds heavier.

T. O. 1F- 16A- 1

These GW's are approximate and shall not be used forcomputing aircraft performance. For maximum GW
limitations, refer to Section V, OPERATING LIMITATIONS. For detailed information, refer to T.O.
IF-16A-l-l, Part 1.

COCKPIT ARRANGEMENT
Refer to figure 1-3. The cockpit arrangement isconventional except for the seat, which is reclined 30
degrees, and the stick, which is mounted on the rightconsole. The cockpit contains no circuit breakers.

Change 2

1-6.1 /(1-6.2 blank)

General Arrangement (Typical)
11

26

15

27

~28
~30
31 32

~-~

rn
For antenna locations, refer
to ANTENNA LOCATIONS,
this section.

33_~~~17 35
')c

34

47
64
1_ Air Data IPitot) Probe
59 57
20 Nose Radome
1 R LEF Drive Unit
3. AOA Probe
19. Hydraulic Reservoir
4. AOA Transmitter
20. AR Slipway
5. Forward Electronic Equipment Bay 21. FLCS Accumulators
6. Cockpit Pressure Regulator
22. Anticollision Strobe Light Power Supply
7. Cockpit Pressure Safety Valve
23. Vertical Tail Attachment Fittings
R Head--Up Display (HUD)
24. Vertical Tail
90 Instrument Panel
25. BLOCK 15 Vertical Tail-Mounted Floodlight
10. Canopy (Movable)
26. Anticollision Strobe Light
11> Ejection Seat
27. Rudder
28. Position Light
12. Canopy IFixed)
29. Rudder ISA
13. AIM--9 Missile
14. Missile Launcher
30. [N£] INOI Drag Chute Fairing
15. AR and Formation Light
31. Turbofan Engine
16. Position/Formation Light
32. Speed brake
17. Static Discharger
33. Speedbrake Actuator

44 43

51. Air-Conditioning Package
34. ChaffIFlare Dispenser
35. Horizontal Tail ISA
52. M61A1 20MM Gun
36. Horizontal Tail
53. Ammunition Drum
37. Formation Light
54. EPU Nitrogen Bottle
38. Flaperon ISA
55. Gun Port
39. Hook
56. NLG Door
40. Flaperon
57. BLOCK 10 Total Temperature Probe
41. Wing Structural Box
58. Position Light
42. Fixed Trailing Edge Panel 59. NLG
43. Asymmetry Brake
60. Engine Air Inlet
44. LEF Rotary Actuator
61. Left Console
45. LEF Torque Shaft
62. Throttle
63. Strake
46. LEF
47. Wing Attachment Fittings 64. Lower Equipment Compartment
48. MLG
65. IAOland IOEllNOlrn ID Light
49. MLG Door
50. LEF Angle Drive Gearbox

1F-16A-1-1004X®

:-I

9

-;n
......
~~

I........

Figure 1-1.

10. 1 F-16A-l

General Datam (Typical)
MISSILE LAUNCHEROUTBOARD STORES LOCATIONS

18 FT 6 IN.
,.....,...._-......
-_.1

SPEEDBRAKE HINGE

49 FT 6 IN.
----------------------OVERALL LENGTH--------------------~

1o+-~~47.50

r

15.35

1 1
95.0 101.0

1 1
16 FT 8.5 IN.
OVERALL
HEIGHT~-......-49.81

NOTE: Dimensions are in inches unless specified otherwise.
1F-16A-1-1 005-1 X®

Figure 1-2. (Sheet 1)
1-8

10. 1F- 16A- 1

WINGS
Area ...................•... 300 Sq Ft
Span ......•...........•... 30 Ft
Aspect Ratio ............... 3.0
Taper Ratio ................ 0.2275
Sweep (LE) ................ 40 0
Dihedral ................... 0 0
Airfoil ...................•. NACA 64A204
Incidence ..•....•...•..•.•. 0 0
Twist
At BL 54.0 ............. 0 0
At BL 180.0 ............ 3 0
Flaperon Area ............. 31.32 Sq Ft
LEF Area .................. 36.71 Sq Ft
HORIZONTAL TAILSArea ...................... 63.70 Sq Ft
Aspect Ratio ............... 2.114
Taper Ratio ................ 0.390 (Theo)
Sweep (LE) ................ 40 0
Dihedral ................... -10 0
AirfoilAt Root ................. 6% Biconvex
At Tip .................. 3.5% Biconvex
VERTICAL TAIL
Area ...................... 54.75 Sq Ft
Aspect Ratio ............... 1.294
Taper Ratio •........•..•.•. 0.437
Sweep (LE) ................ 47.5 0
AirfoilAt Root. ................ 5.3% Biconvex
At Tip .................. 3.0% Biconvex
Rudder Area ....•........•. 11.65 Sq Ft

VENTRAL FIN (EACH)
Area ....................... 8.03 Sq Ft
Span ....................... 23.356 In. Theo (27.5 In. Actual)
Aspect Ratio ............... 0.472 (Theo)
Taper Ratio ................. 0.760 (Theo)
Sweep (LE) ................. 30 0
Dihedral (Cant) .............. 15 0 Outboard
AirfoilAt Root ..........•...... 3.886% Modified Wedge
At Tip ................... Constant 0.03R
SPEEDBRAKES
Area (4 Element Clamshell) .. 14.26 Sq Ft (3.565 Sq Ft Ea)
LANDING GEAR (LG)Main Gear (MLG)
Tire Size ................. 25.5 x 8-14 20 Ply
Tire Size ................. 25.5 x 8-14 18 Ply
Stroke ............•..••.. 10.5 In.
Static Rolling Radius ..... 11.0 In.
Nose Gear (NLG)
Tire Size ................ 18 x 5.7-8 18 Ply
Tire Size ................ 18 x 5.5 14 Ply
Stroke .................. 10.0 In.
Static Rolling Radius ..... 7.5 In.

I

I

ENGINE
F100-PW-200/220
Thrust ..........•........ 25,000 Lb Class
Engine Compressor
Face Diameter ..•••.... 34.8 In.
Engine Length ........... 191.16 In.

32 FT 10 IN.
14-----0VERALL SPAN W/MISSILES----~
1 + - - - - OVERALL

31 FT
SPAN W/O MISSILES ---~

1F-16A-1-1 005-2X ®

Figure 1-2. (Sheet 2)

1-9

T.O. 1 F-16A-1

General Data LESS GD (Typical)

MISSILE LAUNCHEROUTBOARD STORES LOCATIONS

24.05

tI

f

I

_--=__

I

18 FT 0.34 IN.
SPEEDBRAKE HINGE

49 FT 6 IN.
~--------------------OVERALL LENGTH--------------------~
t+----t~ 47.50

15.35

r 1
95.0 101.0

jj
16 FT 8.5 IN.
OVERALL
HEIGHTSTATIC GROUND LINE
t+---~-49.81

NOTE: Dimensions are in inches unless specified otherwise.
1 F-16A-1-1 006-1 X ®

Figure 1-2. (Sheet 3)
1-10

T.0.1F-16A-1

WINGS
Area .. _________ .. __ . ___ .. _. 300 Sq Ft
Span __ ... ____ . _. ____ .. _. _. 30 Ft
Aspect Ratio ..... _... _. _... 3.0
Taper Ratio ................ 0.2275
Sweep /LE) .....•.......... 40 0
Dihedral ................... 0 0
Airfoil ..................... NACA 64A204
Incidence ..•............... 0 0
Twist
At BL 54.0 ............. 0 0
At BL 180.0 ............ 3 0
Flaperon Area •............ 31.32 Sq Ft
LEF Area .................. 36.71 Sq Ft
HORIZONTAL TAILSArea ...................... 49.0 Sq Ft
Aspect Ratio ............... 2.598
Taper Ratio ................ 0.3 (Theo)
Sweep (LE) ................ 40 0
Dihedral ................... -10 0
AirfoilAt Root. ................ 6% Biconvex
At Tip ...........•...•.. 3.5% Biconvex
VERTICAL TAIL
Area ...................... 54.75 Sq Ft
Aspect Ratio ............... 1.294
Taper Ratio .•.....•........ 0.437
Sweep (LE) ................ 47.5 0
AirfoilAt Root .•..•............ 5.3% Biconvex
At Tip .................. 3.0% Biconvex
Rudder Area ............... 11.65 Sq Ft

VENTRAL FIN (EACH)
Area ....................... 8.03 Sq Ft
Span ....................... 23.356 In. Theo /27.5 In. Actual)
Aspect Ratio ............... 0.472 (Theo)
Taper Ratio ................. 0.760 (Theo)
Sweep (LE) ................. 30 0
Dihedral (Cant) .......•...... 15 0 Outboard
AirfoilAt Root ................. 3.886% Modified Wedge
At Tip ................... Constant 0.03R
SPEEDBRAKES
Area (4 Element Clamshell) .. 14.26 Sq Ft (3.565 Sq Ft Ea)
LANDING GEAR (LG)Main Gear (MLG)
Tire Size ................. 25.5 x 8-14 20 Ply
Tire Size ................. 25.5 x 8-14 18 Ply
Stroke ................... 10.5 In.
Static Rolling Radius ..•.. 11.0 In.
Nose Gear (NLG)
Tire Size ................ 18 x 5.7-8 18 Ply
Tire Size ................ 18 x 5.5 14 Ply
Stroke ....•......•....•. 10.0 In.
Static Rolling Radius ..... 7.5 In.

I
I

ENGINE
F100-PW-200
Thrust ...•.•.•.••••.••••. 25,000 Lb Class
Engine Compressor
Face Diameter .••...... 34.8 In.
Engine Length ........... 191.16 In.

32 FT 10 IN.
14-----0VERALL SPAN W/MISSILES-------+t
31 FT
WIO MISSILES ---~

1 4 - - - - OVERALL SPAN

1F-16A-1-1006-2X®

Figure 1-2. (Sheet 4)

1- 11

10. 1 F- 16A-l

Cockpit Arrangement~ IBFI (Typical)
Block to

FOR DETAILS OF OXYGEN REGULATORPANEL DIFFERENCES, SEE THE INDIVIDUAL
EQUIPMENT ILLUSTRATION.

RIGHTAUXILIARYCONSOLE

STICK

LEFT
CONSOLERIGHTCONSOLELOCATED ON SHELF
BEHIND SEAT.

1 F-16A-1-1007X@

Figure 1-3. (Sheet 1)

1-12

T.O. 1F-16A- 1

Cockpit Arrangement [8J IBFI (Typical) Block 10
LEFT CONSOLE

20
/

/

) 2

19
3

4

5

11

25

1. RADAR Control Panel
2. UHF Radio Control Panel
3. FC/NAV Panel
4. TACAN Control Panel
5. MANUAL TRIM Panel
6. AVTR Control Panel
7. ECM Pod Control Panel
8. G-Suit Hose Connection
9. ANTI G TEST Button
10. TEST Switch Panel
11. DEFOG Lever
12. FLIGHT CONTROL Panel
13. Fuel Controi Panel
14. CANOPY JETTISON T-Handle
15. COMM Control Panel
16. EPU Control Panel
17. ELEC Control Panel
18. Throttle FRICTION Control
19. ENG & JET START Control Panel
20. MANUAL PITCH Override Switch
21. CHAFF/FLARE Dispense Button
22. Throttle
23. REDUCED IDLE THRUST Switch
24. Stowage
25. I SF I STICK CONTROL Switch

1F-16A-1-1008X®

Figure 1-3. (Sheet 2)

1- 13

T.O. 1 F-16A-1

Cockpit Arrangement ~ I BF I (Typical) Block to
LEFT AUXILIARY CONSOLE

20RIGHT AUXILIARY CONSOLE

1 23
L_\-h 4

""r ~
~~.~~
l71G~· ~I
19
18

16

l

I

:~~kd
14 11

89

5

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. ANTI-SKID Switch
5. LANDING TAXI Lights Switch
6. DN LOCK REL Button
7. LG Handle Down Permission Button
8. LG Handle Warning Light (Red)
9. LG Handle
10. IFF Control Panel
11. THREAT WARNING AUX (DIM) Knob
12. THREAT WARNING AUX Controls andIndicators
13. ALT GEAR Reset Button
14. ALT GEAR Handle
15. SPEED BRAKE Position Indicator
16. LE FLAP POSITION Indicator (Deactivated)
17. STORES CON FIG Switch
18. HORN SILENCER Button
19. GND JETT ENABLE Switch (Lever Lock)
20. BRAKES Channel Switch

1. Magnetic Compass
2. Clock
3. OXY FLOW Indicator
4. System A HYD PRESS Indicator
5. System B HYD PRESS Indicator
6. EPU FUEL Quantity Indicator
7. Cockpit Pressure Altimeter
8. LIQUID OXYGEN Quantity Indicator
9. Caution Light Panel
10. FUEL QTY SEL Knob
11. EXT FUEL TRANS Switch

Figure 1-3. (Sheet 3)
1-14

IF-16A-1-1009X®

T. O. 1F- 16A- 1

Cockpit Arrangement ~ IBFI (Typical) Block 10
INSTRUMENT PANEL

L;]~~

• t:]

~

~

c

c~~~~~~~c=~~OO~

2
1. HUD Combiner Glass
2. Television Sensor
3. AR Status/NWS Indicator
4. HUD Control Panel
5. Standby Attitude Indicator
6. FUEL FLOW Indicator
7. DUAL FC FAIL Warning Light (Red)8. HYD/OIL PRESS Warning Light IRed)
9. CANOPY Warning Light (Red)
10. ENGINE Warning Light (Red)
11. Radio Channel/Frequency Indicator
12. Vertical Velocity Indicator
13. OIL Pressure Indicator
14. RPM Indicator
15. NOZ pas Indicator
16. FTIT Indicator
17. FUEL Quantity Indicator
18. Altimeter
31. MASTER ARM Switch
19. MRK BCN Light
32. [A] ALT REL Button
20. Attitude Director Indicator
33. SMS PWR Switch
21. Horizontal Situation Indicator
34. [QID ~ mJ IFF IDENT Button;
22. Radar/EO Display~ [NQ] Drag Chute Switch
23. Rudder PEDAL ADJ Knob
35. TO/LDG CONFIG Warning Light (Red)
24. AOA Indicator
36. ALT LOW and ENG FIRE Warning Lights (Red)
25. INSTR MODE Select Panel
37. THREAT WARNING Controls and Indicators
26. Airspeed/Mach Indicator
38. MASTER CAUTION Light (Amber)
27. Stores Control Panel
39. THREAT WARNING Azimuth Indicator
28. Autopilot ROLL Switch
40. Spotlight
29. Autopilot PITCH Switch
41. AOA Indexer
30. AUTOPILOT Switch
42. ~ OVRD Light

42
28
1 F-16A-l-l 010X®

Figure 1-3. (Sheet 4)

1-15

10. 1 F- 16A- 1

Cockpit Arrangement [AJ IBFI (Typical) Block 10
RIGHT CONSOLE

20

16

7 6

~

8
13

12

9

11

1. ILS Control Panel
2. NUCLEAR CONSENT Switch (Guarded)
3. VHF Radio Control Panel
4. RDR ALT Switch
5. Interior LIGHTING Control Panel
6. ECS Control Panel
7. Secure Voice Panel
8. ENGINE ANTI ICE Switch
9. ANT SEL Panel
10. Utility Light
11. OXYGEN REGULATOR Panel
12. Data Transfer Unit
13. CHAFF/FLARE Control Panel
14. Stowage
15. Oxygen/Communications Hookup
16. EXT LIGHTING Control Panel
17. VOICE MESSAGE Switch,
18. MASTER ZEROIZE Switch
19. Stick
20. ~ TAKE CONTR Panel
21. rn::EJ NWS Control Button/
Indicator (Green)
LOCATED ON SHELF
22. Chaff/Flare Programmer
BEHIND SEAT.
23. SEAT ADJ Switch
1F-1SA-1-1011X@

Figure 1-3. (Sheet 5)

1-16

T. O. 1F- 16A- 1

Cockpit A~
IBRI (T~. rangement.,plcal)Block 1.0

INSTRUMENT
PANEL

"'"

RIGHT
CONSOLE1 F-16A-1-1 012X@

Figure 1-3• \l/Sh eet 6)

1-17

10. 1F-16A- 1

Cockpit Arrangement IBRI (Typical) Block 10
LEFT CONSOLE-TAKECONTfU..-

t--r===~---irnJl ~ ~ ~9~~B\OB\ :'U'

1~'I'C ~11'C~11'C(ifff
j

CF
F PWR

I LS

~® [@f@

®
VOL

11

d:J

12

7

13

6

5

1. RADIO SEL Panel
2. ILS Control Panel
3. TACAN Control Panel
4. UHF Radio Control Panel
5. G-Suit Hose Connection
6. ANTI G TEST Button
7. MAL & IND LTS Test Button
8. CANOPY JEniSON T-Handle
9. ENG CONT Switch
10. MANUAL PITCH Override Switch
11. Throttle
12. FUEL MASTER Switch
13. Stowage

1 F-16A-1-1013X®

Figure 1-3. (Sheet 7)

1-18

T.O. 1F- 16A- 1

Cockpit Arrangement IBRI (Typical) Block 10
LEFT AUXILIARY CONSOLERIGHT AUXILIARY CONSOLE

UP

AOC

flTaM
SYS

fWD
fUEllilW
AIT
FUEl lOW

CAOC

1

ANTI-$(~

(@) I

I

WOOl
NWSfAll- HYD PRESS A
B

O'/EIIOTOXYlilW

AYI)Ij(SEQUI'HOT

Noo.EM

5[ATNOT ARMED

RADAR
AlT
STO<IES[NQNE'AlA.T@) @lEFlAPS

CA8IN~SSElECI't$

5[CCOlrflCj

EJECTION MODE SEL
NORM
I

SOLO
/

1 2 3

rttlrtlJ
15~ I
16

/~ •

17
5

1~ --{j!l] 14 12

lLI

OJ

I6lITJ

I

LJo~:

8

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. DN LOCK REL Button
5. LG Handle Down Permission Button
6. LG Handle Warning Light (Red)
7. LG Handle
8. ALT GEAR Handle
9. ALT GEAR Reset Button
10. THREAT WARNING AUX (DIM) Knob
11. THREAT WARNING AUX Controls
and Indicators
12. SPEED BRAKE Position Indicator
13. ALT FLAPS Switch (Lever Lockl
14. HORN SILENCER Button
15. GND JETT ENABLE Switch (Lever Lock'
16. LE FLAP POSITION Indicator (Deactivated:
17. [NID [NQJ DRAG CHUTE Switch

5

1. Clock
2. OXY FLOW Indicator
3. System B HYD PRESS Indicator
4. System A HYD PRESS Indicator
5. EJECTION MODE SEL Handle
6. Caution Light Panel

1 F-16A-l-1 0 14X@

Figure 1-3. (Sheet 8)

1-19

T.O. 1F·16A·l

Cockpit Arrangement IBRI (Typical) Block 10
INSTRUMENT PANEL

1. MASTER CAUTION Light (Amber)
2. AR Status/NWS Indicator
3. Radar/EO Display
4. DUAL FC FAIL Warning Light (Red)6. HYD/OIL PRESS Warning Light (Red)
6. CANOPY Warning Light (Red)
7. ENGINE Warning Light (Redl
8. Standby Attitude Indicator
9. FUEL FLOW Indicator
10. Vertical Velocity Indicator
11. Radio Channel/Frequency Indicator
12. RPM Indicator
13. OIL Pressure Indicator
14. NOZ POS Indicator
15. FUEL Quantity Indicator
16. FTIT Indicator
17. Altimeter
18. MRK BeN Light
19. Attitude Director Indicator
20. Horizontal Situation Indicator
21. AOA Indicator
22. INSTR MODE Select Panel
23. Rudder PEDAL ADJ Knob
24. Airspeed/Mach Indicator
25. Reduced Idle Thrust Indicator
26. Stores Control Panel
27. IFF IDENT Button
28. ARMT CONSENT Switch (Guarded)
29. Accelerometer

30. Stick Indicator
31. OVRD Light
32. VIDEO SEL Switch
33. THREAT WARNING Controls & Indicators
34. THREAT WARNING Azimuth Indicator
35. Spotlight
36. TO/LOG CONFIG Warning Light (Red)37. ALT LOW and ENG FIRE Warning Light (Red)38. AOA Indexer
39. Handgrip, F-' 6A-' -1 01 5X ~Figure 1-3. (Sheet 9), .20

T. o. 1F- 16A-1

Cockpit Arrangement \BRI (Typical) Block 10
RIGHT CONSOLE

SUIT

PRESS

@
(n

\

--PRlMARY-\lHD.ES

O~

9

7

m PNl

O~

fREQ DISP

O~

6

1. SUIT PRESS Switch (Nonfunctional)
2. Interior LIGHTING Control Panel
3. OXYGEN REGULATOR Panel
4. Utility Light
6. Oxygen/Communications Hookup
6. Stowage
7. NWS Control Button/lndicator (Green)
8. Stick
9. SEAT ADJ Switch

1F-16A-1-1016xe

Figuree 1-3. (Sheet 10)

1-21

T. O. 1F- 16A- 1

Cockpit Arrangement~~[AJ ~ (Typical)

FOR DETAILS OF OXYGEN REGULATORPANEL OR IpW200i ENGINE DIFFERENCES,
SEE THE INDIVIDUAL EQUIPMENTILLUSTRATION.

INSTRUMENT
PANEL

RIGHT
AUXILIARYCONSOLELEFT
AUXILIARYCONSOLERIGHT - - - - . 4
CONSOLELEFTCONSOLELOCATED ON SHELF
BEHIND SEAT,

Figuree 1-3. (Sheet 11)

1-22

Change 1

1F-1SA-1-10"xeT. O. 1F- 16A- 1

Cockpit Arrangement [AJIBFI (Typical) Block 15
LEFT CONSOLE

1415 ---+l"i-+!
11

24
9

8

25
1. RADAR Control Panel
2. UHF Radio Control Panel
3. FC/NAV Panel
4. TACAN Control Panel
6. MANUAL TRIM Panel
6. AVTR Control Panel
7. ECM Pod Control Panel
8. G-Suit Hose Connection
9. ANTI G TEST Button
10. TEST Switch Panel
11. DEFOG Lever
12. FLIGHT CONTROL Panel
13. Fuel Control Panel
14. CANOPY JETTISON T-Handle
16. COMM Control Panel
16. EPU Control Panel
17. ELEC Control Panel
18. Throttle FRICTION Control
19. ENG & JET START Control Panel
20. MANUAL PITCH Override Switch
21. CHAFF/FLARE Dispense Button
22. Throttle
23. REDUCED IDLE THRUST Switch
24. Stowage
26. (If] STICK CONTROL Switch

IFF~

1F-18A-1-101aX<il

Figure 1-3. (Sheet 12)

Change 1

1-23

T.O. 1F-16A-1

Cockpit Arrangement 00 IBF I (Typical) Block 15
LEFT AUXILIARY CONSOLERIGHT AUXILIARY CONSOLE

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down lights (Green)
3. HOOK Switch (lever lock)4. ANTI-SKID Switch
5. LANDING TAXI Lights Switch
6. ON lOCK REl Button
7. lG Handle Down Permission Button
8. lG Handle Warning Light (Red)
9. lG Handle
10. IFF Control Panel
11. THREAT WARNING AUX (DIM) Knob
12. THREAT WARNING AUX Controls andIndicators
13. AlT GEAR Reset Button
14. AlT GEAR Handle
15. SPEED BRAKE Position Indicator
16. STORES CON FIG Switch
17. HORN SilENCER Button
18. GND JETT ENABLE Switch (lever lock)
19. BRAKES Channel Switch

1. Magnetic Compass
2. Clock
3. OXY FLOW Indicator
4. System A HYD PRESS Indicator
5. System B HYD PRESS Indicator
6. EPU FUEL Quantity Indicator
7. Cockpit Pressure Altimeter
8. LIQUID OXYGEN Quantity Indicator
9. Caution light Panel
10. FUEL QTY SEl Knob
11. EXT FUEL TRANS Switch
1F-18A-1-1019X.

Figure 1-3. (Sheet 13)

1-24

T.O. 1F-16A- 1

Cockpit Arrangement [AJ IBF I (Typical) Block 15
INSTRUMENT PANEL

L;](@[@]
IfflOENTco

LJ!ERNI/

1. HUD Combiner Glass
2. Television Sensor
3. AR Status/NWS Indicator
4. HUD Control Panel
5. Standby Attitude Indicator36
6. FUEL FLOW Indicator
32
7. DUAL FC FAIL Warning Light (Red)31 I
8. HYD/OIL PRESS Warning Light (Red)
9. CANOPY Warning Light (Red)29~_
10. RDR ALT LOW Warning Light (Red)
11. ENGINE Warning Light (Red)
12. Radio Channel/Frequency Indicator
13. Vertical Velocity Indicator26
14. OIL Pressure Indicator
15. RPM Indicator
30. [8J ALT REL Button
16. NOZ POS Indicator
31. LASER ARM Switch
17. FTIT Indicator
32. ~~~ IFF IDENT Button;
MJ~ DRAG CHUTE Switch
18. FUEL Quantity Indicator
19. AOA Indicator
33. Stores Control Panel
20. INSTR MODE Select Panel
34. Threat Warning Azimuth Indicator
21. Airspeed/Mach Indicator
35. THREAT WARNING Controls and Indicators
36. ENG FIRE Warning Light (Red)
22. Attitude Director Indicator
23. Horizontal Situation Indicator 37. T.O./LAND CON FIG Warning Light (Red)38. MASTER CAUTION Light (Amber)
24. Altimeter
25. Radar/EO Display
39. Spotlight
40. AOA Indexer
26. Autopilot PITCH Switch
27. Autopilot ROLL Switch
41. MRK BCN Light
42. Rudder PEDAL ADJ Knob
28. AUTOPILOT Switch
29. MASTER ARM Switch
43. ~ OVRD Light

1

27~J."

33

~,\30
28

1F-16A-1-1020X®

Figure 1-3. (Sheet 14)

1-25

T. O. 1F- 16A- 1

Cockpit Arrangement [8J IBF I (Typical) Block 15
RIGHT CONSOLE

4 24

6
17

18 7

16

13

~

25~

8
10
12

11

1. ILS Control Panel
2. VHF Radio Control Panel
3. NUCLEAR CONSENT Switch (Guarded)
4. SUIT PRESS Switch (Nonfunctional)
5. Interior LIGHTING Control Panel
6. SNSR PWR Control Panel
7. ECS Control Panel
8. ENGINE ANTI ICE Switch
9. Utility Light
10. ANT SEL Panel
11. OXYGEN REGULATOR Panel
12. Data Transfer Unit
13. CHAFF/FLARE Control Panel
14. Stowage
15. Oxygen/Communications Hookup
16. AVIONICS POWER Panel
17. EXT LIGHTING Control Panel
18. Secure Voice Control Panel
19. VOICE MESSAGE Switch
20. MASTER ZEROIZE Switch
21. Stick
22. [AJ Chaff/Flare Programmer
23. SEAT ADJ Switch
24. ~ VHF Control Button/
Indicator
25. ~ TAKE CONTR Panel
26. [[8 NWS Control
Button/Indicator (Green)

LOCATED ON SHELF
BEHIND SEAT.
1I~e:::........:~.-=...J

1F-1SA-1-1021X@

Figure 1-3. (Sheet 15)

1-26

T. O. 1F- 16A- 1

Cockpit Arrangement
IBRI (Typical)
Block 15

. . . - - LEFTAUXILIARYCONSOLERIGHT - - . - I
AUXILIARY
I-~~~:;-l
CONSOLE~~,[~~tQ"f"W'Ul,.1
"0

'.

STICK

LEFT
CONSOLERIGHTCONSOLEFOR DETAILS OF OXYGEN REGULATORPANEL OR IpW2001 ENGINE DIFFERENCES,
SEE THE INDIVIDUAL EQUIPMENTILLUSTRATION.

1F-16A-1-1022X@

Figure 1-3. (Sheet 16)

1-27

T.O. 1 F- 16A-l

Cockpit Arrangement IBRI (Typical) Block 15
LEFT CONSOLE- - TAKE CD'JTROL-VHF

TACAN

ILS

MS,

THREAT,~'ERCOMOODmO\C O~O~ID
OfT PWR

®

I LS

VOl

®

[108.IOJ

CHAN~[l(~l~Q~~~ ® ID'/~A/AR[C

orf

eM r

CHAN

,I

.
I

1

AlAT/R

U

0
11

U
11

2
3

12

4

s-H
7
13

1. RADIO SEL Panel
2. ILS Control Panel
3. TACAN Control Panel
4. UHF Radio Control Panel
5. G-Suit Hose Connection
6. ANTI G TEST Button
7. MAL & IND LTS Test Button
8. CANOPY JETTISON T-Handle
9. ENG CONT Switch
10. MANUAL PITCH Override Switch
11. Throttle
12. FUEL MASTER Switch
13. Stowage

<ZI--

C!l

1F-16A-1-1023X@

Figure 1-3. (Sheet 17)

1-28

T.O. 1F- 16A- 1

Cockpit Arrangement IBRI (Typical) Block 15
LEFT AUXILIARY CONSOLE
9EMERSTGRESJETTISOli

RIGHT AUXILIARY CONSOLE

~
~P
0
U e
~

~En°
~

.

L

UP

S

GNO JETTENABLE~OffFLT CONT
SYS
ADC

FWD
FUEL LOW
AFT
fUEL LOWLE
fLAPS

OVERHEAT

CAOC

ENelNE
fAULT

ELEC
SYS

SEC
fUEL
HOT
SEAT NOT
ARWEO

AVIONICS
ATf NOTENGAGED
RAOAR
ALT
EQUIP
HOT

ANTI
SKID

OXY FLOW

HOOK

Iff

NWS
fAIL
CABIN
PRESS
OXY
LOW

ECM

NUCLEARSTORESCONFIGPROBE HEAT

~

(@)

~

- HYD PRESS A
B@ @EJECTION MODE SEL
NORM
i

SOLO

/

1 2 3

rt-};ill
15~JI
4 II
~. I
16

14-r, f-\

II

~o1B

5
6

7

8
1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. DN LOCK REL Button
5. LG Handle Down Permission Button
6. LG Handle
7. LG Handle Warning Light (Red)
8. ALT GEAR Handle
9. ALT GEAR Reset Button
10. THREAT WARNING AUX (DIM) Knob
11. THREAT WARNING AUX Controls
and Indicators
12. SPEED BRAKE Position Indicator
13. HORN SILENCER Button
14. ALT FLAPS Switch (Lever Lock)
15. GND JETT ENABLE Switch (Lever Lock)
16. ~DRAG CHUTE Switch

M

1. Clock
2. FUEL Quantity Indicator
3. OXY FLOW Indicator
4. System B HYD PRESS Indicator
5. System A HYD PRESS Indicator
6. EJECTION MODE SEL Handle
7. Caution Light Panel

1 F-16A-1-1 024X@

Figure 1-3. (Sheet 18)

1-29

10. 1 F-16A-l

Cockpit Arrangement IBRI (Typical) Block 15
INSTRUMENT PANEL

1. MASTER CAUTION Light (Amber)
2. AR Status/NWS Indicator
3. Radar/EO Display
4. DUAL FC FAIL Warning Light (Red)
5. HYD/OIL PRESS Warning Light (Red)
6. CANOPY Warning Light (Red)
7. RDR ALT LOW Warning Light (Red)
8. Standby Attitude Indicator
9. ENGINE Warning Light (Red)
10. FUEL FLOW Indicator
11. OIL Pressure Indicator
12. MRK BCN Light
13. NOZ POS Indicator
14. FTIT Indicator
15. RPM Indicator
16. AOA Indicator
17. Vertical Velocity Indicator
18. Attitude Director Indicator
19. Horizontal Situation Indicator
20. Altimeter
21. Airspeed/Mach Indicator
22. Accelerometer
23. Radio Channel/Frequency Indicator
24. INSTR MODE Select Panel
25. ARMT CONSENT Switch (Guarded)
26. IFF IDENT Button
27. Reduced Idle Thrust Indicator
28. Stick Indicator
29. OVRD Light
30. THREAT WARNING Controls & Indicators
31. Threat Warning Azimuth Indicator
32. Stores Control Panel

33. Spotlight
34. VIDEO SEL Switch
35. ENG FIRE Warning Light (Red)
36. T.O.ILAND CONFIG Warning light (Red)37. ADA Indexer
38. Rudder PEDAL ADJ Knob
39. Handgrip
1F-16A-1-1025X®

Figure 1-3. (Sheet 19)

1-30

10. 1F-16A- 1

Cockpit Arrangement IBRI (Typical) Block 15
RIGHT CONSOLE

9

3

8

: - - PRIMARY - -~ cnro.ES

INST PNL:FREQ DISP

I~ O~ ~~ O~

~
!

G
H

T

-FLOOD -Q ~~ '~

~

r®

@

tJAL&

(Jffl}llllT

®

CD

ICD
I

I ~D~ 6

MrD

Sw

.

1. SUIT PRESS Switch (Nonfunctional)
2. VHF Radio Control Panel
3. Interior LIGHTING Control Panel
4. DSPL POWER Switch
5. OXYGEN REGULATOR Panel
6. Utility Light
7. Oxygen/Communications Hookup
8. Stowage
9. NWS Control Button/Indicator (Green)
10. Stick
11. SEAT ADJ Switch

~rE
Cff
\ B

OXYC£NREGULATOR

0flOW

["['B

NCO"

~

~

.•

.llIml~lllllmlmIIMIII

1F-16A-1-1026X®

Figure 1-3. (Sheet 20)

1-31

T. O. 1F- 16A- 1

Cockpit Arrangement !~~[8J IBFI (Typical) IAOIFOR DETAILS OF OXYGEN REGULATORPANEL OR IpW200i ENGINE DIFFERENCES,
SEE THE INDIVIDUAL EQUIPMENTILLUSTRATION.

INSTRUMENT
PANEL

LEFT
AUXiLiARy ............
CONSOLE.........

D
RIGHT - - - - . 1
CONSOLELOCATED ON SHELF
BEHIND SEAT.

Figure 1-3. (Sheet 21)

1-32

1F-16A-1-1027X@

10. 1 F- 16A-l

Cockpit Arrangement [ruIBFI (Typical) ~LEFT CONSOLE

/

) 2
3
14
15----i-i-i~

5~-'--'------'..---~

11

24-

25
1. RADAR Control Panel
2. UHF Radio Control Panel
3. FC/NAV Panel
4. TACAN Control Panel
5. MANUAL TRIM Panel
6. AVTR Control Panel
7. ECM Pod Control Panel
8. G-Suit Hose Connection
9. ANTI G TEST Button
10. TEST Switch Panel
11. DEFOG Lever
12. FLIGHT CONTROL Panel
13. Fuel Control Panel
14. CANOPY JETTISON T-Handle
15. COMM Control Panel
16. EPU Control Panel
17. ELEC Control Panel
18. Throttle FRICTION Control
19. ENG & JET START Control Panel
20. MANUAL PITCH Override Switch
21. CHAFF/FLARE Dispense Button
22. Throttle
23. REDUCED IDLE THRUST Switch
24. Stowage
25. [!E] STICK CONTROL Switch

I

~@

I

~
1F-1SA-1-1028X@

Figure 1-3. (Sheet 22)

1-33

10. 1 F- 16A- 1

Cockpit Arrangement [KJIBFI (Typical) IAOILEFT AUXILIARY CONSOLE

14 11

RIGHT AUXILIARY CONSOLE

5

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. ANTI-SKID Switch
5. LANDING TAXI Lights Switch
6. DN LOCK REL Button
7. LG Handle Down Permission Button
8. LG Handle Warning Light (Red)
9. LG Handle
10. IFF Control Panel
11. THREAT WARNING AUX (DIM) Knob
12. THREAT WARNING AUX Controls
and Indicators
13. ALT GEAR Reset Button
14. ALT GEAR Handle
15. SPEED BRAKE Position Indicator
16. STORES CONFIG Switch
17. HORN SILENCER Button
18. GND JETT ENABLE Switch (Lever Lock)
19. BRAKES Channel Switch

Figure 1-3. (Sheet 23)

1-34

1. Magnetic Compass
2. Clock
3. OXY FLOW Indicator
4. System A HYD PRESS Indicator
5. System B HYD PRESS Indicator
6. EPU FUEL Quantity Indicator
7. Cockpit Pressure Altimeter
8. LIQUID OXYGEN Quantity Indicator
9. Caution Light Panel
10. FUEL QTY SEL Knob
11. EXT FUEL TRANS Switch
1F-16A-1-1029X@

T. O. 1F- 16A- 1

Cockpit Arrangement [8J [if] (Typical) lAD I
INSTRUMENT PANEL

3839

1. HUD Combiner Glass
37
2. Television Sensor
36 ~~-,....-_--,.-.-e=~
3. AR Status/NWS Indicator
4. HUD Control Panel
32
5. Standby Attitude Indicator
31 r~F~---'
6. FUEL FLOW Indicator
7. DUAL FC FAIL Warning Light (Red)29
33
8. HYD/OIL PRESS Warning Light (Red)
9. CANOPY Warning Light (Red)
10. RDR ALT LOW Warning Light (Red)
11. ENGINE Warning Light (Red)26
28
12. Radio Channel/Frequency Indicator
13. Vertical Velocity Indicator
14. OIL Pressure Indicator
15. RPM Indicator
16. NOZ POS Indicator
30. IAJ ALT REL Button
31. LASER ARM Switch
17. FTIT Indicator
32. IFF IDENT Button
18. FUEL Quantity Indicator
19. AOA Indicator
33. Stores Control Panel
34. Threat Warning Azimuth Indicator
20. INSTR MODE Select Panel
35. THREAT WARNING Controls and Indicators
21. Airspeed/Mach Indicator31}j32
36. ENG FIRE Warning Light (Red)
22. Attitude Director Indicator
23. Horizontal Situation Indicator 37. T.O./LAND CONFIG Warning Light (Red)43/
38. MASTER CAUTION Light (Amber)
24. Altimeter
29~
25. Radar/EO Display
39. Spotlight
40. AOA Indexer
26. Autopilot PITCH Switch
27. Autopilot ROLL Switch
41. MRK BCN Light
42. Rudder PEDAL ADJ Knob
28. AUTOPILOT Switch
29. MASTER ARM Switch
43. !IEJ OVRD Light

I

2711

~,\30

[@

1F-16A-1-1030X@

Figure 1-3. (Sheet 24)

1-35

T.O. 1 F-16A-l

Cockpit Arrangement [AJ IBF I (Typical) lAD I
RIGHT CONSOLE

15
17

16

13

12

1. ILS Control Panel
2. VHF Radio Control Panel
3. NUCLEAR CONSENT Switch (Guarded)
4, lAl HF Radio Control Panel
6, RDR ALT Switch
6. Interior LIGHTING Control Panel
7. ECS Control Panel
8, ENGINE ANTI ICE Switch
9, Utility Light
10. ANT SEL Panel
11. OXYGEN REGULATOR Panel
12. Data Transfer Unit
13, CHAFF/FLARE Control Panel
14, Stowage
15, Oxygen/Communications Hookup
16, AVIONICS POWER Panel
17. EXT LIGHTING Control Panel
18. VOICE MESSAGE Switch,
19, MASTER ZEROIZE Switch
20. Stick
21, Chaff/Flare Programmer
22. SEAT ADJ Switch
23. ~ VHF Control
Button/Indicator
24. lIE! TAKE CONTR Panel25, [[EJ NWS Control
Button/Indicator (Green)
LOCATED ON SHELF
BEHIND SEAT.
1F-16A-1-1031X@

Figure 1-3. (Sheet 25)

1-36

T. O. 1F- 16A- 1

Cockpit Arrangement
IBRI (Typical) IAOIINSTRUMENT
PANEL

LEFT
AUXILIARY ___CONSOLE---...

LEFT
CONSOLEFOR DETAILS OF OXYGEN REGULATORPANEL OR IpW200! ENGINE DIFFERENCES,
SEE THE INDIVIDUAL EQUIPMENTILLUSTRATjON

1 F-1 6A-'1-1 032X ®

F£gure 1-3, (Sheet 26)

1-37

T. O. 1 F- 16A- 1

Cockpit Arrangement IBRI (Typical) [8QJLEFT CONSOLE,fii'\WI®

- - TAKE CeNTRaL-ClH<

VHf

TAGPI'<

r----;:::::==::;---1I~D sE [J []
ID 6L
CD

i

C

OFF

ieS

D D

MSL

THREATm

INTERCOM

O~ O~O~m
I LS(® !108.101 ®

OfT PWIl

VOl

DVlNNEl(CIQM]

~Q~
~~
1<
II
II
17

9

12

CHAN

U

II
II

"

2
3
4

L--

sH,
13

7

1. RADIO SEL Panel
2. ILS Control Panel
3. TACAN Control Panel
4. UHF Radio Control Panel
5. G-Suit Hose Connection
6. ANTI G TEST Button
7. MAL & IND LTS Test Button
8. CANOPY JETTISON T-Handle
9. ENG CONT Switch
10. MANUAL PITCH Override Switch
11. Throttle
12. FUEL MASTER Switch
13. Stowage

cCZI--

C!l

1F-16A-1-1033X@

Figure 1-3. (Sheet 27)
1-38

T. O. 1F- 16A- 1

Cockpit Arrangement IBR] (Typical) lAD I
LEFT AUXILIARY CONSOLERIGHT AUXILIARY CONSOLE

UP

flTS;~NT
AoC

rWD
rUEL LOW
ArTrUEL LOW

ANTI
SKID

AVIONICS
ATr NOTENGAGEDRADARAlTOXY FLOW

HOOK~ (@) ~lE
HAPS

OVERHEAT

CAoC

ENGINE
rAU T

E~~iP

ElEC
SYS

SEC

Iff

NWS
rAil
CABIN
PRESS
DXY
lOW

ECM

NUCLEARSTORESCONFIePROBE HEAT

-

-

rUEL
HOT
SEAT NOT
ARr,jED--- HYD PRESS-

A

B

-

EJECTION MODE SEL
NORM
I

SOLO

/

1 2 3

157i.r:1-i00/114ff. i-f II4

I~0-1r£J

5
6
7

OJ®
10m

LJO~4

8
1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. DN LOCK REL Button
5. LG Handle Down Permission Button
6. LG Handle
7. LG Handle Warning Light (Red)
8. ALT GEAR Handle
9. ALT GEAR Reset Button
10. THREAT WARNING AUX (DIM) Knob
11. THREAT WARNING AUX Controls
and Indicators
12. SPEED BRAKE Position Indicator
13. HORN SILENCER Button
14. ALT FLAPS Switch (Lever Lock)
15. GND JETT ENABLE Switch (Lever Lock)~:
1. Clock
2. FUEL Quantity Indicator
3. OXY FLOW Indicator
4. System B HYD PRESS Indicator
5. System A HYD PRESS Indicator
6. EJECTION MODE SEL Handle
7. Caution Light Panel

1F-16A-1-1034xe

Figure 1-3. (Sheet 28)

1-39

T.O. 1F- 16A- 1

Cockpit Arrangement IBFI (Typical) IAOIINSTRUMENT PANEL

1. MASTER CAUTION Light (Amber)
2. AR Status/NWS Indicator
3. Radar/EO Display
4. DUAL FC FAIL Warning Light (Red)
5. HYD/OIL PRESS Warning Light (Red)
6. CANOPY Warning Light (Red)
7. RDR ALT LOW Warning Light (Red)
8. Standby Attitude Indicator
9. ENGINE Warning Light (Red)
10. FUEL FLOW Indicator
11. OIL Pressure Indicator
12. MRK BCN Light
13. NOZ pas Indicator
14. FTIT Indicator
15. RPM Indicator
16. AOA Indicator
17. Vertical Velocity Indicator
18. Attitude Director Indicator
19. Horizontal Situation Indicator
20. Altimeter
21. Airspeed/Mach Indicator
22. Accelerometer
23. Radio Channel/Frequency Indicator
24. INSTR MODE Select Panel
25. ARMT CONSENT Switch (Guarded)
26. IFF IDENT Button
27. Reduced Idle Thrust Indicator
28. Stick Indicator
29. OVRD Light
30. THREAT WARNING Controls & Indicators
31. Threat Warning Azimuth Indicator
32. Stores Control Panel

7
9

28

11

"

26

10

~12

o-~ \

15

~~
14

13

38
33. Spotlight
34. VIDEO SEL Switch
35. ENG FIRE Warning Light (Red)
36. T.O./LAND CON FIG Warning light (Red)37. AOA Indexer
38. Rudder PEDAL ADJ Knob
39. Handgrip
1f-1SA-1-1035X@

Figure 1-3. (Sheet 29)

1-40

10. 1 F- 16A- 1

Cockpit Arrangement IBRI (Typical) IAOI
RIGHT CONSOLE

SUIT
PRESS

<&
0'1

9
3
8

1. SUIT PRESS Switch (Nonfunctional)
2. VHF Radio Control Panel
3. Interior LIGHTING Control Panel
4. DSPL POWER Switch
5. OXYGEN REGULATOR Panel
6. Utility Light
7. Oxygen/Communications Hookup
8. Stowage
9. NWS Control Button/Indicator (Green)
10. Stick
11. SEAT ADJ Switch

1F-16A-l-1036X@

Figure 1-3. (Sheet 30)
Change 121-40.1

T.O.1F-16A-1

Cockpit Arrangemen~t__
L8J iEJ (Typical) I AN I ,.,
PANEL

RIGHT
AUXILIARYCONSOLE-LEFTCONSOLERIGHT - - _ I
CONSOLELOCATED ON SHELF
BEHIND SEAT.
1 F-1SAN-l-0006X (g)Figure 1-3. (Sheet 31)

1-40.2

Change 1210. 1F-16A- 1

Cockpit Arrangement[AJ IEJ (Typical) IAN I
LEFT CONSOLE

20
/

/

) 2

19

23 18

6

3

14-Q

I ~ 13
~ ~

15

4

5

f-L--.L.......I...-----.J.,-,.-l--ll

6

24-

t-~1

~~}

9

8

25
1. RADAR Control Panel
2. UHF Radio Control Panel
3. FC/NAV Panel
4. TACAN Control Panel
5. MANUAL TRIM Panel
6. AVTR Control Panel
7. ECM Pod Control Panel
8. G-Suit Hose Connection
9. ANTI G TEST Button
10. TEST Switch Panel
11. DEFOG Lever
12. FLIGHT CONTROL Panel
13. Fuel Control Panel
14. CANOPY JETTISON T-Handle
15. COMM Control Panel
16. EPU Control Panel
17. ELEC Control Panel
18. Throttle FRICTION Control
19. ENG & JET START Control Panel
20. MANUAL PITCH Override Switch
21. CHAFF/FLARE Dispense Button
22. Throttle
23. REDUCED IDLE THRUST Switch
24. Stowage
25. ~ STICK CONTROL Switch"~~" i~);;;~cg.~~~'"~~"w""m

5~UJcn ....

rnIJ
1F-16AN-1-0007X @Figure 1-3. (Sheet 32)

Change 121-40.3

10. 1F-16A-1

Cockpit Arrangement [AJ IBFI (Typical) IANILEFT AUXILIARY CONSOLERIGHT AUXILIARY CONSOLE

13

1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)
4. ANTI-SKID Switch
5. LANDING TAXI Lights Switch
6. ON LOCK REL Button
7. LG Handle Down Permission Button
8. LG Handle Warning Light (Red)
9. LG Handle
10. IFF Control Panel
11. THREAT WARNING AUX (DIM) Knob
12. THREAT WARNING AUX Controls andIndicators
13. ALT GEAR Reset Button
14. ALT GEAR Handle
15. SPEED BRAKE Position Indicator
16. STORES CONFIG Switch
17. HORN SILENCER Button
18. GND JETT ENABLE Switch (Lever Lock)
19. BRAKES Channel Switch

Figure 1-3. (Sheet 33)

1-40.4

Change 12

1. Magnetic Compass
2. Clock
3. OXY FLOW Indicator
4. System A HYD PRESS Indicator
5. System B HYD PRESS Indicator
6. EPU FUEL Quantity Indicator
7. Cockpit Pressure Altimeter
8. LIQUID OXYGEN Quantity Indicator
9. Caution Light Panel
10. FUEL QTY SEL Knob
11. EXT FUEL TRANS Switch
IF-16AN-l-0008X @10. 1F- 16A- 1

Cockpit Arrangement [A] IBFI (Typical) IANIINSTRUMENT PANEL

1. HUD Combiner Glass
2. Television Sensor
3. AR Status/NWS Indicator
4. HUD Control Panel
5. Standby Attitude Indicator
6. FUEL FLOW Indicator
7. DUAL FC FAIL Warning Light (Red)8. HYD/OIL PRESS Warning Light (Red)
9. CANOPY Warning Light (Red)
10. RDR ALT LOW Warning Light (Red)
11. ENGINE Warning Light (Red)
12. Radio Channel/Frequency Indicator
13. Vertical Velocity Indicator
14. OIL Pressure Indicator
15. RPM Indicator
30. []U ALT REL Button
16. NOZ POS Indicator
31. LASER ARM Switch
17. FTIT Indicator
32. IFF IDENT Button
18. FUEL Quantity Indicator
33. Stores Control Panel
19. AOA Indicator
34. Threat Warning Azimuth Indicator
20. INSTR MODE Select Panel
35. THREAT WARNING Controls and Indicators
21. Airspeed/Mach Indicator
36. ENG FIRE Warning Light (Red)
22. Attitude Director Indicator
23. Horizontal Situation Indicator 37. T.O./LAND CONFIG Warning Light (Red)38. MASTER CAUTION Light (Amber)
24. Altimeter
39. Spotlight
25. Radar/EO Display
40. AOA Indexer
26. Autopilot PITCH Switch
41. MRK BCN Light
27. Autopilot ROLL Switch
42. Rudder PEDAL ADJ Knob
28. AUTOPILOT Switch
43.
OVRD Light
29. MASTER ARM Switch

m

32

31)
43+
29~

1 F-16AN-1-0009X @Figure 1-3. (Sheet 34)

Change 121-40.5

T.o. 1 F-16A-1

Cockpit Arrangement [8J IBF I (Typical) IAN I
RIGHT CONSOLE
1

21~ 69h='\[ill rd I

17

Off PWRW

ILS~I [108 01

'All

®

-0 f----6-----1
1_1

14

23 ~

16

15

12
9

11

10

1. ILS Control Panel
2. VHF Radio Control Panel
3. HUD Control Panel
4. SUIT PRESS Switch (Nonfunctional)
5. Interior LIGHTING Control Panel
6. SNSR PWR Control Panel
7. ENGINE ANTI ICE Switch
8. Utility Light
9. ANT SEL Panel
10. OXYGEN REGULATOR Panel
11. Data Transfer Unit
12. CHAFF/FLARE Control Panel
13. Stowage
14. Oxygen/Communications Hookup
15. ECS Control Panel
16. EXT LIGHTING Control Panel
17. VOICE MESSAGE Switch
18. MASTER ZEROIZE Switch
19. Stick
20. Chaff/Flare Programmer
21. SEAT ADJ Switch
22. ~ VHF Control
Button/Indicator
23. ~ TAKE CONTR
Panel
24. ~ NWS Control
Button/Indicator (Green) ~~ ~~ §r'7:n~ \:§ ~ ~~Y.J)

~r&\ ~<;' 9.Q~\:::;y c§~ ~LOCATED ON SHELF
BEHIND SEAT.
1 F-16AN-1-0010X ®

Figure 1-3. (Sheet 35)

1-40.6

Change 1210. 1F- 16A- 1

Cockpit Arrangement IBRI (Typical) IANIINSTRUMENT
PANEL

LEFT
AUXILIARYCONSOLERIGHT=-:--1-_------;:=;:;;;;;;;::;;;;::;;;~AUXILIARYCONSOLE

STICK

LEFT
CONSOLERIGHT---+l
CONSOLE

IF-16AN-l-00llX@

Figure 1-3. (Sheet 36)

Change 121-40.7

T.o. 1F- 16A-1

Cockpit Arrangement lBRl (Typical) [8NJLEFT CONSOLE
uHF

WF

TACA~

ILS

TMR'AT

I'JTERCOMOOD"SL

orr PWR~(@ [108.10] ®
I LS

VOl

CHAi'ML

CH

I

2

,

1.

.
11

15

CHAN

(7),

lUl W

I'
17
11
11

,.

m
CD

2

3

10

4
~

__________________-400

OJ

ID

8

7

1. RADIO SEL Panel
2. ILS Control Panel
3. TACAN Control Panel
4. UHF Radio Control Panel
5. G-Suit Hose Connection
6. ANTI G TEST Button
7. MAL & IND LTS Test Button
8. Stowage
9. CANOPY JETTISON T-Handle
10. FUEL MASTER Switch
11. MANUAL PITCH Override Switch
12. ENG CONT Switch
13. Throttle

«ZI--

C!J

1 F-16AN-1-001 2X@

Figure 1-3. (Sheet 37)

1-40.8Change 12

T. O. 1F- 16A- 1

Cockpit Arrangement IBRI (Typical) IAN I
LEFT AUXILIARY CONSOLERIGHT AUXILIARY CONSOLE:::::::::::::::;::::===~eeEMERSTORES*~

~

~

.~

L

o~:
UP

S
GNDJ,'T

ENABLEo~OFF

FLl CONISYS

AL T FLAPS

HOR~

SPeED

NORM

SILENCER

BRAKE

Q

EXTEND

© ~

LE
flAPS

FWD
FUEL LOW
AFT
FUEL LOW
OVERHEATENGINEfAUl I

FIFe

SYS

m

ATF "lOT

ENGAGED
AlT
EQUP
HOT

OXY FLOW

HOOK

@

NWS

flADAf(

SEC
fUEL
HOT
SEAT NOI
ARMED

ANII
SKID

AVIONICS

FAIL

CAB"
PRESS
DXY
LOW

~

~ HYO PRESS

~

-

A

EC~

STORESCONFIG~B

GD GD
2':1,000

BE]

~

0

EJECTION MODE SEL
NORM
I

SOLJ

/

7

1. EMER STORES JETTISON Button ~Covered)
2. WHEELS Down Lights ~Green)
3. HOOK Switch ~Lever Lock)
4. ON LOCK REL Button
5. LG Handle Down Permission Button
6. LG Handle
7. LG Handle Warning Light ~Red)
8. ALT GEAR Handle
9. ALT GEAR Reset Button
10. THREAT WARNING AUX ~DIM) Knob
11. THREAT WARNING AUX Controls
and Indicators
12. SPEED BRAKE Position Indicator
13. HORN SILENCER Button
14. ALT FLAPS Switch ~Lever Lock)
15. GND JETT ENABLE Switch ~Lever Lock)

1. Clock
2. FUEL Quantity Indicator
3. OXY FLOW Indicator
4. System B HYD PRESS Indicator
5. System A HYD PRESS Indicator
6. EJECTION MODE SEL Handle
7. Caution Light Panel

1 F-16AN-1-0013X@

Figure 1-3. (Sheet 38)

Change 121-40.9

IO.lF-16A-1

Cockpit Arrangement IBRI (Typical) IANIINSTRUMENT PANEL

1. MASTER CAUTION Light (Amber)
2. AR Status/NWS Indicator
3. Radar/EO Display
4. DUAL FC FAIL Warning Light (Red)
5. HYD/OIL PRESS Warning Light (Red)
6. CANOPY Warning Light (Red)
7. RDR ALT LOW Warning Light (Red)
8. Standby Attitude Indicator
9. ENGINE Warning Light (Red)
10. FUEL FLOW Indicator
11. OIL Pressure Indicator
12. MRK BCN Light
13. NOZ POS Indicator
14. FTIT Indicator
15. RPM Indicator
16. AOA Indicator
17. Vertical Velocity Indicator
18. Attitude Director Indicator
19. Horizontal Situation Indicator
20. Altimeter
21. Airspeed/Mach Indicator
22. Accelerometer
23. Radio Channel/Frequency Indicator
24. INSTR MODE Select Panel
25. ARMT CONSENT Switch (Guarded)
26. IFF IDENT Button
27. Reduced Idle Thrust Indicator
28. Stick Indicator
29. OVRD Light
30. THREAT WARNING Controls & Indicators
31. Threat Warning Azimuth Indicator
32. Stores Control Panel

37

36
35

2

34

33. Spotlight
34. VIDEO SEL Switch
35. ENG FIRE Warning Light (Red)
36. TO./LAND CON FIG Warning light (Red)37. AOA Indexer
38. Rudder PEDAL ADJ Knob
39. Handgrip
1 F-16AN-l-0014X@

Figure 1-3. (Sheet 39)

1-40.10

Change 1210. 1F- 16A- 1

Cockpit Arrangement IBRI (Typical) IANI
RIGHT CONSOLE

93
80Jl CONSUS
L
I
G
H

~~1

~~~~~~

O~

iN
FLOOD
G n
~~T
:(lIISOLES

@

~

~ ~T
.... -

1. SUIT PRESS Switch (Nonfunctional)
2. VHF Radio Control Panel
3. Interior LIGHTING Control Panel
4. DSPL POWER Switch
5. OXYGEN REGULATOR Panel
6. Utility Light
7. Oxygen/Communications Hookup
8. Stowage
9. NWS Control Button/Indicator (Green)
10. Stick
11. SEAT ADJ Switch

INST P~L

FREO CISP

O~

m

WlL&
INO LTS

0 ID
DIM

MII~~II~lm~m~~llmll~1111•

1 F-16AN-l-0015X®

Figure 1-3. (Sheet 40)
Change 12

1-41

T.O. 1F- 16A- 1

ENGINE IPW200 I

and start bleed controls, and the convergent nozzle
area control.

ENGINE GENERAL DESCRIPTION I PW200 I

The UFC schedules basic engine fuel flow according
to throttle position, rear compressor rotor speed,
FTIT, and combustion chamber pressure. The UFC isfurther trimmed by the EEC which uses fan speed,engine inlet temperature, and FTIT to adjust fuelflow. AB fuel flow is scheduled by the UFC but may be
reduced by the EEC through the segment 5 ABlockout and stall recovery logic features.

Refer to figure 1-4. The aircraft is powered by a single
FIOO-PW-200 afterburning turbofan engine. Maximum thrust is approximately 25,000 pounds.
ENGINE FUEL SYSTEM I PW200 I

Refer to figure 1-5. The engine fuel system delivers
the required fuel to the engine for combustion and foruse by the control system for scheduling the engine
variable geometry.
Unified Fuel Control (UFC) I PW200 I

Primary engine control is provided by a hydromechanical UFC.
The UFC contains the engine fuel control, the AB fuelcontrol, the rear compressor variable vanes (RCVV)Electronic Engine Control (EEC) I PW200 I

The EEC is an engine-mounted, solid-state digitalcomputer which is fuel cooled.
The EEC commands UFC trims of engine fuel flow
and convergent nozzle area. The EEC controls thescheduling of the compressor inlet variable vanes(CIVV's) and provides signal inputs to the UFC forengine stall recovery, segment 5 AB lockout, AB
ignition inhibit, and idle area reset.

F1.00-PW-200 Engine
FAN MODULE

CORE ENGINE MODULEFAN DRIVE
TURBINE
MODULEAB MODULE

CIVVHIGH PRESSURE TURBINEUNIFIED FUEL CONTROLENGINE DRIVENGEARBOXELECTRONIC ENGINE CONTROL
MAIN FUEL PUMP
1 F-16X-l-0003X@

Figure 1-4.

1-42

1 0 .1 F-16A-1

Engine Fuel/Control System Schematic(Typical)ENGIN E Fl 0 0-PW-2 0 0--... --.----_.
FUEL PRESSURIZATION AND."~:II.

r

FROMUFCIICENe~ CENe ~ AIR-MOTOR

CENe

EXHAUSTSHUTOFF VALVEDUMP VALVE

1JW~ _ START BLEED ACTUATOR (UFC)_REAR COMPRESSOR VARIABLE:
VANES IUFC)

r.~

••••

_ START BLEED ACTUATOR (BUC)

FUEl/_ REAR COMPRESSOR VARIABLE
VANES (BUC)

TO ••
CIW

BACKUPUNIFIED

FUEL CONTROL(BUC)

FUEL CONTROL

~

n
r"-

I RTN

TO
CENC

(UFC)

ENGINE: 'AIRFRAME~ __ -,- + -----:-----..,Iro'

1 RSVRS1

1----1------------ ------1

ENGINEELECTRONICCONTROL
COOLINGFUELSHUTOFF
VALVE

1

1

FUEL LINESTRAINERfUEL
TRANSMITTER
L ___
_FLOW
__
__________

~

ER

I l
I

MAIN
PUMP
_~F~u~'~L~~~~:II..

~

MAIN FUEL PUMP

- - ELECTRICAL~ PRESSURE, BUC___ _ MECHANICAL__J COOL-J~

I

LEGEND:FUELITO OIL l Oll
SYSTEM

1! FUEl./OIL HEAT EXCHANGER
MAIN FUEL SHUTOFF VALVE__REMOVED
ON SOMEENGINESTHROTTLE

BOOST PUMP PRE SSURE_

MAIN FUEL PUMPPRESSURE, UFC OR BUC
AB FUEL PRESSURE

MAIN FUEL PUMPPRESSURE~

AB FUEL PRESSURE

MAIN FUEL PUMPPRESSURE, UFC_

AB PUMPCONTROLLER~ (FILTEREDf

ENGINE BLEED AIR

ENGINEFAN DUCT
IF- 16A- l - l038A.

Figure J-5_

1-43

T. O. 1F- 16A- 1

The EEC closed-loop idle control trims the UFC idle
fuel flow to maintain scheduled fan speed. This
results in constant, repeatable idle thrust in flightand on the ground. To reduce the idle thrust level, thenozzle is commanded open when the throttle is at ornear IDLE and the LG handle is DN.
The EEC limits minimum engine operation throughout the flight envelope to maintain stable operation.
At high altitude, low airspeed conditions, the EEC
protects against low thrust engine stalls. Duringtransonic and supersonic conditions, the EEC limits
minimum engine operation as a function of machnumber (from the CADC) to provide sufficient engine
airflow.
To minimize the possibility of stalls during AB
operation at high altitude and low airspeed, the EEC
commands termination of segment 5 AB. If an engine
stall does occur in AB, the EEC automaticallycommands the engine to minimum AB fuel flow
regardless of throttle position. Below AB, the EEC
stall recovery logic is most effective with the throttle
at MIL; at MIL, the EEC commands the UFC to open
the nozzle to relieve excessive back pressure createdby a stall. All these features are deactivated when the
EEC is turned off.

An engine fan overspeed condition causes the EEC toautomatically turn off and illuminates the EEC
caution light. The EEC cannot be reset in this case.

The EEC receives power directly from the engine
alternator. In the event of engine alternator or engine
gearbox failure indicated by rapid decrease to zero
percent rpm and illumination of the ENGINE
warning light, the EEC is inoperative and the EEC
caution light may not illuminate. The engine isoperable without the EEC within certain limits (refer
to Section III).
Backup Fuel Control (BUC) System I PW200i

The BUC is a hydromechanical system which
provides engine control in the event of a UFC
malfunction. The BUC is selected by the EEC BUC
switch. Fuel flow is scheduled only by throttle
position and engine fan duct static pressure which
compensates for altitude.
Main Fuel Pump IPW200 I

The gearbox-mounted main fuel pump provides
pressurized fuel to the engine and boosts pressure to
the AB fuel pump.

1-44

Change 1

Afterburner (AB) Fuel Pump I PW200i

The AB fuel pump is driven by engine bleed air andprovides pressurized fuel to the AB section of theUFC. The pump also provides pressurized fuel forhydraulic actuation of the ClVV's, the RCW's, andthe start bleed actuator.
Compressor Inlet Variable Vanes (CIW's) I pW200i

The CIVV's are located immediately forward of thefirst fan stage. They are positioned by signals fromthe EEC using pressurized fuel from the AB fuelpump. In BUC, the ClVV's are in a fixed position.
Rear

Compressor Variable

Vanes(RCW's)

IPW2001

The first three stages of the rear compressor are
equipped with RCVV's. RCVV's are positioned by the
UFC using pressurized fuel from the AB fuel pump.
In BUC, the RCW's are controlled by the throttle
position.
Compressor Bleed Air IPW200i

Low-pressure (seventh-stage) bleed air is directedfrom the bleed strap into the fan duct to increase thecompressor stall margin during starting. The bleed
valve is actuated by the UFC using pressurized fuelfrom the AB fuel pump. In BUC, it is activated bythrottle position. Low-pressure bleed air is also used
for engine inlet anti-icing.
High-pressure (thirteenth-stage) bleed air is supplied
to the EPU and engine nacelle ejectors. It is also used
to drive the AB fuel pump and to drive the CENC
motor.
Either low-pressure or high-pressure air is provided
to the ECS depending on engine bleed pressure levels.
Pressurization and Dump Valve IpW200i

A pressurization and dump valve is located in theengine fuel manifold line between the fuel/oil cooler
and fuel nozzles. It provides a minimum fuel pressure
for UFC operation at low rpm, enhances quick
starting, and dumps the engine fuel manifold when
the throttle is retarded to OFF.

I

T. O. 1F- 16A- 1
EXHAUST NOZZLE I PW200 I

controlled by a three-position ANTI ICE switch. The
anti-ice system can be activated manually or
automatically by an ice detector in the engine inlet.

The exhaust nozzle is variable and consists of two
sections. The divergent nozzle floats freely and moves
in conjunction with the convergent nozzle. The
convergent nozzle is controlled by the convergent
exhaust nozzle control.

The inlet strut is electrically heated to prevent ice
buildup. The heater is activated by the engine ANTIICE switch.

Convergent ExhaustEngine ANTI ICE Switch IPW200 I

Nozzle Control(CENC)

IpW2001

The CENC is activated by a high-pressure bleed air
motor. The nozzle schedule is primarily controlled by
the throttle position input to the UFC and trimmed by
the EEC. With the LG handle in DN and EEC on, thenozzle is approximately 70-95 percent open at IDLE
(idle area reset). As the throttle is advanced, thenozzle closes. With the LG handle in UP, the nozzle
position is 35-50 percent at throttle settings below
midrange with the EEC on or off. Nozzle position
decreases as the throttle is advanced. The nozzle isnear minimum area at throttle settings above
midrange except when approaching MIL or when in
AB. At MIL and above, the EEC trims the nozzle toregulate engine back pressure to control fan speed. As
the throttle is advanced in the AB range, the UFC
opens the nozzle to compensate for increasing AB fuelflow. With the EEC off or inoperative, the nozzle isnearly closed at throttle settings above midrange andAB operation is not recommended. In BUe, the nozzle
is not scheduled by the CENC but is aerodynamicallyloaded toward the closed position; therefore, AB
operation is prohibited.

The engine ANTI ICE switch is located on the rightconsole.
Functions are:
• ON - Low-pressure bleed air is directed to the fixed
inlet guide vanes and nose cone. The inlet strut
electrical heater turns on.
• AUTO - When engine icing is detected by the icing
detector (automatic detection may not occur on the
ground or above 7 degrees AOA), low-pressure bleed
air is directed to the fixed inlet guide vanes and nose
cone and the inlet strut electrical heater turns on.
• OFF - Electrical power closes the engine anti-ice
valve (loss of power opens the valve). The inlet strut
heater and ice detector are turned off.
ENGINE AND ACCESSORY DRIVE GEARBOXESI PW200 I

ENGINE OIL SYSTEM I PW200 I

Refer to figure 1-6. The engine gearbox drives themain fuel pump, the oil pump assembly, the engine
alternator, and the PTO shaft, which powers theaccessory drive gearbox (ADG).

The engine is equipped with a self-contained oil
system to lubricate the engine and gearbox. Systempressure is nonregulated and varies with rpm, oiltemperature, and altitude.

The ADG powers the main generator through theconstant-speed drive (CSD), system A and B
hydraulic pumps, and FLCS PMG. The JFS is also
mounted on the ADG.

• Below approximately 35,000 feet MSL, oil pressure
should increase approximately 15 psi from IDLE to
MIL. At very high altitudes (50,000 feet), the oilpressure increase is approximately 5 psi from IDLE
to MIL. At all altitudes, however, a definite oilpressure increase should be evident when the rpm isincreased. Refer to SERVICING DIAGRAM, thissection for servicing/specifications information.
ENGINE ANTI-ICE SYSTEM I PW200 I

The anti-ice system routes low-pressure bleed air toand through the fixed inlet guide vanes and the nose
cone to prevent ice formation. The system is

ENGINE ALTERNATOR I PW200 I

The engine alternator is driven by the engine gearbox
and provides sale power for the EEC, engine and AB
ignition, and the rpm signal to the rpm indicator.
ENGINE IGNITION SYSTEM I PW200 I

The ignition system is powered by the engine
alternator and contains three igniter plugs (two forthe engine and one for the AB). With the throttle at or
above IDLE and rpm at 15 percent or above, engine
ignition is continuous. When the throttle is moved

Change 111-45

10. 1F- 16A- 1

Engine and Accessory Drive Gearboxes
ENGINE F100-PW-200

2

1. Tower Shaft (From Engine)
2. Main Fuel Pump
3. Engine Gearbox
4. Engine Alternator
5. JFS Exhaust Duct
6. Jet Fuel Starter
7. PTa Shaft
8. System B Hydraulic Pump
9. Accessory Drive Gearbox
10. FLCS PMG
11. System A Hydraulic Pump
12. Constant-Speed Drive
13. Main Generator
14. Engine Oil Pump

7
1F-16A-1-0007X@

Figure 1-6.

into AB, AB ignition is activated for approximately 1
second. For subsequent AB ignition, the throttle must
be retarded to MIL or below for a minimum of 1.5
seconds and then returned to AB.

used to start the engine on the ground and to assist inengine airstart. Refer to JET FUEL STARTER
LIMITS, Section V.
ENG & JET START CONTROL PANEL I PW200 I

JET FUEL STARTER (JFS) I PW200 I

The JFS is a gas turbine which operates on aircraft
fuel and drives the engine through the ADG. The JFS
is connected by a clutch to the ADG and only provides
torque when required to maintain engine rpm. If the
ADG is not able to rotate (i.e., seized engine), the JFS
runs, but the clutch prevents it from rotating theADG. The JFS receives fuel at all times regardless ofthe FUEL MASTER switch position. The JFS isstarted by power from two brake/JFS accumulators
used either singly or together. The brake/JFS
accumulators are charged automatically by hydraulicsystem B or manually by a hydraulic hand pumplocated in the left wheel well. Automatic recharging
takes between 40 seconds (hot ambient conditions)and 60 seconds (cold ambient conditions). The JFS is

1-46

Change 1

Refer to figure 1-7. The ENG & JET START control
panel is located on the left console.
JFS Switch [BJ mEl I PW200 I

Functions are:• OFF - Normal switch position. The JFS can be shut
down at anytime by selecting OFF. The switch
returns to OFF automatically during a normal
ground start at 50 percent rpm.
• START 1 - Vents one of the brake/JFS accumulators to the hydraulic start motor.
• START 2 - Vents both brakelJFS accumulators to
the hydraulic start motor.

I

T.O. 1F- 16A-1

ENG & JET START Control Panel (Typical)
1

ENGINE F100-PW-200

2
6
3
J

OFF

E
T

S
T

~

Bue
STARTING FUEL

4T

1. RUN Light (Green)
2. JFS Switch
3. EEC BUC Switch
4. STARTING FUEL Switch
5. MAX POWER Switch
6. BUC Switch

NOTE:~ For BUC switch, referto F-16B AIRCRAFT, thissection.

5

1F-16A-1-1042X~

Figure 1-7.

JFS RUN Light ~ ~ I PW200l

shut down until the JFS switch is manually
positioned to OFF.

The green JFS RUN light illuminates within 30
seconds after initiating JFS start to indicate that the
JFS has attained governed speed.

ENGINE CONTROLS AND INDICATORS I PW200i

JFS Operation I PW200i

Refer to figure 1-8. The engine instruments arelocated on the right side of the instrument panel.
Refer to ENGINE LIMITATIONS I PW2001, Section V.

During a ground engine start, the brakelJFS
accumulators begin to recharge after the engine
accelerates through 12 percent rpm. As the engine
accelerates through 50 percent rpm, a sensor causes
the JFS to shut down automatically and the JFS RUN
light goes off. During in-flight operation, thebrake/JFS accumulators begin to recharge (provided
system B hydraulic pressure is available) when the
JFS reaches 70 percent of governed speed (3-4seconds before the JFS RUN light illuminates). If the
JFS RUN light does not illuminate within 30 seconds
or the JFS RUN light goes off once illuminated, theJFS START switch will not reengage and the JFS
cannot be restarted until the JFS has spooled down.
JFS spooldown takes approximately 17 seconds fromfull governed speed. Once running, the JFS does not

EEC BUC Switch ~ [Ill IPW200 I
The EEC BUC switch (guarded out ofBUC) is locatedon the left console. ~ For EEC BUC switch
differences, refer to F-16B AIRCRAFT, this section.
Functions are:
• EEC - EEC in operation (normal position).
• OFF - EEC not in use (basic UFC operation).
• BUC - BUC operation. Transfer occurs whenthrottle is in OFF or at or above BUC IDLE. The
throttle may have to be advanced past BUC IDLE
to have transfer occur.

Change 1

1-47

IT. O. 1F- 16A- 1

Engine Controls and Indicators (Typical)
Block 10

E

N
G

&
J
E

Bue

EEC
A
NI
Te
!E

T

s

T

suc

R

STARTING FUEL
AUTO LEAN

A
T

A
F
R
I
C
T
I
0
N

2

~~ I: sENGINETE
L

CFF

FUEL HOT11

OVERHEAT
EEC

12

Q~
LEAN

v

OVERHEAT
EEC

1. Throttle
2. 1All!El EEC BUC; [IB] BUC Switch
3. ENG FIRE Warning Light (Red)
4. FUEL FLOW Indicator
5. HYD/OIL PRESS Warning Light (Red)
6. ENGINE Warning Light (Red)

Figure 1-8. (Sheet 1)

1-48

7. OIL Pressure Indicator
8. RPM Indicator
9. NOZ POS Indicator
10. FTIT Indicator
11. Caution Lights (Amber)
12. ANTI ICE Switch

1F-16A-1-1039X8

T.O. 1F-16A-l

Engine Controls and Indicators (Typical)
ENGINE F100-PW-200

4

&

J

E

T

s

EEC

N

A
R

T

BUC
STARTING FUEl

T

AUTO LEAN

;;
F
R
I

OVERHEAT
EEC

2

QRICH

A
NI
TC
IE

LEAN

C

ENGINE

~:"
OFF

A
NS
TE
L

11

Bue
FUEL
HOT

T

I
0

12

N

\1

OVERHEAT
EEC

Bue
FUEL
HOT
1. Throttle
2. [K] ~ EEC SUC; ~ SUC Switch
3. ENG FIRE Warning light (Red)
4. FUEL FLOW Indicator
5. HYD/OIL PRESS Warning Light (Red)
6. ENGINE Warning light (Red)

7. OIL Pressure Indicator
8. RPM Indicator
9. NOZ pas Indicator
10. FTIT Indicator
11. Caution Lights (Amber)
12. ANTI ICE Switch

1F-16A-1-1040X@

Figure 1-8. (Sheet 2)

1-49

T.O. 1 F-16A-1

EEC Caution Light I PW200i
The EEC caution light, located on the caution light
panel, indicates an EEC self-failure or input failure.
The EEC caution light also illuminates in conjunction
with the CADC caution light if the malfunction
affects the mach number signal. The EEC caution
light illuminates if the EEC BUC switch is in OFF or
BUC or if an engine overspeed condition occurs.

BUC Caution Light I pW200i
The BUC caution light, located on the caution light
panel, illuminates when the engine is operating in
BUC or main fuel pump pressure is low.

STARTING FUEL Switch [!] ~ I PW200i
The STARTING FUEL switch is located on the left
console.
Functions are:
• AUTO LEAN - In UFC (EEC on or om, lean fuelflow is provided during the engine start cycle until
30 seconds after the main generator comes on line.
Fuel flow then increases by 100 pph (rich fuel flow).
In BUC, rich starting fuel flow is provided.
• RICH - Fuel flow is rich at all times.
• LEAN - Fuel flow is lean at all times.
A lean fuel mixture is required during both UFC
ground starts and airstarts. Rich fuel mixture isrequired for BUC starts.

MAX POWER Switch IPW200 I
The MAX POWER switch, located on the left console,
is solenoid held in the MAX POWER position when
the throttle is at MAX AB and airspeed is 1.1 mach or
greater. Refer to ENGINE LIMITATIONS I PW2ool,
Section V.
Functions are:
• MAX POWER - Delivers maximum thrust byallowing maximum FTIT to increase by 22°C.
• OFF - Normal (deenergized) position.

RPM Indicator IpW200i
The RPM indicator has a pointer display and the rpm
signal is supplied by the engine alternator. RPM is

1-50

Change 2

expressed in percent from 0-100. The indicator is
powered by~ battery bus No. 2, LESS~ the battery
bus.

NOZ POS Indicator I PW200 I
The NOZ POS indicator displays the position of the
CENC exhaust nozzle drive shafts which arecalibrated from 0 percent (closed) to 100 percent (fullyopen). The indicator accurately reflects exhaust
nozzle position in UFC and BUC unless both drive
shafts are failed. The indicator is powered byessential ac bus No.2.

FTIT Indicator I pW200i
The FTIT indicator displays an average FTIT indegrees C. The indicator has a range of 200°-1 200°C
in major increments of 100°C and is powered by IAQ]
battery bus No.1, LESS IAQ] the battery bus.

FUEL FLOW Indicator I pW200i
The FUEL FLOW indicator is a pointer-counter
indicator which displays the total fuel flow to theengine, including AB, in pph. The indicator has arange of 0-80,000 pph and is powered by essential ac
bus No. 1.

OIL Pressure Indicator I pW200i
The OIL pressure indicator displays engine oilpressure from 0-100 psi and is powered by essential
ac bus No.2.

HYD/OIL PRESS Warning Light I PW200i
The HYD/OIL PRESS warning light, located on theedge of the right glareshield, serves as a monitor ofengine oil pressure and hydraulic system pressure.
For engine oil pressure, the warning light illuminates when oil pressure has been below approximately 10 psi for 30 seconds (time delay minimizes
warning light illuminating during maneuvering).
The light goes out when oil pressure exceeds
approximately 20 psi. For hydraulic pressure, thewarning light illuminates when either A or B systempressure decreases below 1000 psi. The light goes out
when both system A and B pressures are above 1000
psi. During engine start, the warning light usually
goes off before reaching idle rpm; however,acceptable operation is indicated if the light goes off
before exceeding 70 percent rpm and remains off
when the throttle is retarded to IDLE. The warning
light is powered by~ battery bus No. 1, LESS~ thebattery bus.

T.O.1F-16A-1

ENGINE Warning Light I PW200i

I

The ENGINE warning light, located on the edge ofthe right glareshield, illuminates when RPM and
FTIT indicator signals indicate that an engine
overtemperature, flameout, or stagnation has
occurred. Illumination also occurs for an engine
alternator failure and may occur as a result of an
RPM or FTIT indicator failure. The warning lightilluminates when the rpm decreases to subidle (below
55 percent), when engine stagnates (determined from
rpmIFTIT rates), or approximately 2 seconds after
FTIT indication exceeds 1000°C. The warning lightgoes off when the condition that turned it on iseliminated. The warning light is powered by ~battery bus No.1, LESS ~ the battery bus.

REDUCED IDLE THRUST (RIT) Switch IPW200i
The RIT switch, located on the left sidewall just aft of
the throttle, is deactivated.

ThroHle I PW200 I
Refer to figure 1-9. The engine is controlled by athrottle mounted above the left console with detents
at OFF, IDLE, MIL, and MAX AB. The throttle ismechanically connected to the UFCIBUC. The OFF
position terminates engine ignition and fuel flow. The
IDLE position commands minimum UFC thrust andis used for all ground starts and airstarts. From IDLE
to MIL, the throttle controls the output of the engine.
Forward of the MIL position, the throttle modulates
the operation of the AB (through five segments) while
maintaining constant basic engine operation. When
BUC is selected, a BUC IDLE detent drops into placeforward of IDLE. Throttle travel from IDLE to BUC
IDLE provides a manual BUC starting schedule.
Then the BUC IDLE detent is used to command
minimum BUC thrust.
!A] ~ The throttle must be rotated outboard to allow

advancement from OFF to IDLE and from MIL to AB.
Retarding the throttle from AB to MIL automaticallyrotates the throttle. At IDLE, a cutoff release at thebase of the throttle must be actuated to allow thethrottle to be rotated outboard and retarded to OFF.
~ For throttle differences, refer to F -16B AIRCRAFT, this section.

A single white reflective stripe is located ~ [If) on
both the upper surface of the throttle foot and on the
sidewall fairing, lIB] on both the lower throttle radius
next to the console and on the panel outboard of thethrottle radius. Alignment of the two stripes aids in
identifying the IDLE position.
Six switches are located on the throttle. A throttle
friction control is located inboard at the base of thethrottle. ~ The throttles are mechanically linked
together.

ENGINE OPERATING CHARACTERISTICS I PW200i
Ground Operation I PW200 I
During ground operation, closed-loop idle can beverified by observing rpm and FTIT indications when
the EEC is cycled on and off. When the EEC is turnedoff, rpm should decrease and FTIT should increase
approximately 30°C. When the EEC is turned backon, it may take up to 1 minute for rpm and FTIT toreturn to the previous closed loop levels. Since theEEC maintains constant IDLE thrust, rpm will varywith temperature and pressure altitude (higher
temperature or pressure altitude results in higher
rpm).

Non-AB Operation In Flight I PW200i
After a MIL takeoff, engine FTIT is usually 900
950°C with rpm of 89-94 percent for any outside airtemperature above -7°C. FTIT and rpm are lower for
temperature below _7 C. Regardless of temperature,
nozzle area should not exceed 30 percent at MIL.
0-0Engine operation is continually optimized as flight
conditions change. This is evident by slight changes
in the NOZ POS, RPM, and FTIT indicator
indications.
At low altitudes (below approximately 10,000 feet),idle rpm should always be equal to or slightly higher
than the ground idle rpm. As altitude increases, idle
rpm increases to provide the engine sufficient stallmargin during throttle transients. At 1.4 mach andabove, the minimum thrust level is MIL even though
the throttle may be retarded below MIL. Typically,
the minimum thrust level increases from IDLE to
MIL between 0.84-1.4 mach. All of the minimum
operating level features are deactivated when the
EEC is turned off.

Change 1

1-51

T.O. 1F-16A-1

Throttle (Typical)
ENGINE F100-PW-200

13

6

10
11

9

8

1. [A]~ Throttle Cutoff Release
2. UHF VHF Transmit Switch (2-Way, Momentary Rocker)
3. ~[A] UHF HF VHF Transmit Switch (3-Way, Momentary Rocker)
4. MAN RNG/UNCAGE Knob/Switch (Rotate, Depress)
5. ANT ELEV Knob (Rotate, Center Detentl
6. DOG FIGHT Switch (3-Position, Slide)
7. SPD BRK Switch (3-Position, Aft Momentary)
8. RDR CURSOR/ENABLE Switch (Depress Multidirectional)
9. Throttle Foot
10. Throttle
11. OO~ Throttle Friction Control
12. OO~ IDLE Stripe
13. OO~ Throttle Stripe

BUC IDLE (SOLENOID ACTIVATED; DETENT SHOWNIN STOWED POSITION)NOTE:[jBJ For throttle differences, refer to THROTILE and F-16B AIRCRAFt this section.
Figure 1-9.

1-52

1F-16A-1-1 043X ®

T.O. 1 F-16A-l

I A low frequency engine vibration may be sensed inflight or on the ground primarily at or near idle, but
may also occur at higher thrust settings. The
vibration has no adverse effect on engine or aircraft
structure and should disappear if engine rpm is either
increased or decreased. Vibrations that change in
intensity with throttle movement and are present
across the throttle/rpm range may indicate apotential engine malfunction.

I

AB Operation in Flight I PW200 I

In AB, FTIT, rpm, and oil pressure vary with altitude
and airspeed. Nozzle position at MAX AB duringsubsonic operation will be approximately 75 percent.
During supersonic acceleration, the nozzle opens tomaintain the proper fan speed.
Refer to figure 5-3. When climbing or decelerating
into the segment 5 AB lockout region, nozzle position
decreases from approximately 75-60 percent with anassociated thrust and fuel flow decrease. Whendescending or accelerating from the segment 5
lockout zone, nozzle position increases from approximately 60-75 percent with a corresponding increasein thrust and fuel flow.
Refer to figure 5-3. AB rumble may occur in segment
5 AB in an area just below the segment 5 lockout line.
If rumble is experienced, an AB blowout and/or stall
may result. Retarding throttle should eliminate therumble.
Methods to improve AB operation in flight include:• Stabilizing at MIL prior to initiating AB.
• Delaying throttle advance into AB for a minimumof 1.5 seconds after an AB cancellation.

• Cancelling AB with throttle snaps. This results incontrol logic opening the nozzle area 10 percent
larger than during nonsnap cancellations, thus
providing more stall margin.
• Initiating throttle transients to AB before maneuvering aircraft rather than during maneuvers.
BUC Operation I PW200 I

The engine transfers to BUC when the EEC BUC
switch is placed to BUC and the throttle is in OFF or
if it is at or above BUC IDLE.
When the engine transfers to BUC, the BUC caution
light illuminates.
When starting in BUC, the starting fuel schedule isrich with the STARTING FUEL switch in either
AUTO LEAN or RICH. LEAN is available but
degrades airstart capability and safe idle operation.
When transferring to BUC with the engine running,
rpm and FTIT may either increase or decreasedepending on throttle setting and flight conditions.
The simplicity of the BUC restricts its operatingenvelope and requires smooth and slow throttle
movements (5 seconds through the BUC IDLE to MIL
or MIL to BUC IDLE range). Refer to BUCOPERATIONAL ENVELOPE, Section V. Since theengine is not trimmed by the EEC when in BUC, MIL
rpm can exceed 96 percent at low altitude on a hot day
and must be monitored to prevent the rpm fromexceeding 96 percent.
The nozzle is aerodynamically loaded toward theclosed position when operating in BUC; therefore, AB
selection results in an engine stall and should not be
attempted. Above 25,000 feet MSL or if the nozzle ismore than 50 percent open, the nozzle may not close
due to insufficient aerodynamic loading .
MIL thrust in BUC is less than that in UFC. Thrustin BUC IDLE is higher than that in UFC.

Change 2

1 -53

T.O.1F-16A-l

ENGINE IpW2201

Main Fuel Control (MFC) I PW2201

The F100-PW-220 and F100-PW-220E engines havethe same operating limitations, normal operatingprocedures, and emergency procedures. Both engines
are referred to as I PW220\'

The MFC operates in both the PRI and SEC modes.
During PRI control, the MFC receives throttle inputs,
fuel from the main fuel pump, and electricalcommands from the DEEC. It controls main ignition,
start bleed strap position, main engine fuel flow, andrear compressor variable vane (RCVV) position. The
MFC also provides actuation pressure to thecompressor inlet variable vane (CIVV) control, theconvergent exhaust nozzle control (CENC), and both
the AB fuel control and AB pump controller.

ENGINE GENERAL DESCRIPTION I PW2201
Refer to figure 1-10. The aircraft is powered by asingle F100-PW-220 or F100-PW-220E afterburning
turbofan engine. Maximum thrust is approximately
25,000 pounds.

In SEC, the MFC receives throttle inputs, fuel fromthe main fuel pump, and static pressure and total
temperature signals from the fan inlet case. The MFC
controls main engine fuel flow, start bleed strap
position, RCVV's, and engine ignition.

ENGINE FUEL SYSTEM I PW2201
Refer to figure 1-11. The engine fuel system delivers
the required fuel to the engine for combustion and foruse by the control system for scheduling the engine
variable geometry.

Afterburner (AB) Fuel Control I PW2201

Engine Control System I PW2201

During primary operation, the AB fuel control
receives fuel from the AB fuel pump and electricalcommands from the DEEC. It provides AB ignition,
AB segment sequencing, and fuel flow to the AB
segments. During SEC control, AB fuel flow is
inhibited.

The engine control system is composed of three major
components: the main fuel control (MFC), the AB
fuel control, and the digital electronic engine control(DEEC). The engine has two modes of operation: primary (PR!) and secondary (SEC).

Fl00-PW-220 Engine
FAN MODULEFAN DRIVE
TURBINE
CORE ENGINE MODULE MODULEREAR COMPRESSOR
10 STAGESFAN DUCTCOMBUSTIONCHAMBERAB MODULE
AS SPRAY RINGSAND FLAME HOLDERCONVERGENTNOZZLEAB LIGHT-OFF DETECTORHIGH PRESSURE TURBINEAB FUEL CONTROL
MAIN FUEL CONTROLENGINE DRIVEN GEARBOXENGINE DIAGNOSTIC UNIT

DIGITAL ELECTRONIC ENGINE CONTROL

Figure 1-10.

1-54

1F-16X-1-0004X @T.O. 'F-'6A-'

Engine Fuel/Control System Schematic
ENGINE Fl00-PW-220(Typical),. LEGEND:;:- - ELECTRICAL- - - - - MECHANICAL_ENG BLEED AIR~ FUEL (COOLINGI~ AS FUEL FLOWATNTOFUEL PRE S~~URI;­ZATION ANDOUMP VALVE~ SERVO PRESSURE

RSVR$
FUEl/OIL

COOLER

MAIN
FUEL

AIRFRAME
SIGNALS

CONTROLIMFC)DIGITAL
ELECTRONICENGINECONTROL
IDEEelavec>-----,ENGINEIBLEED AlR,ENGINE IFOR ACTUATIONI :- !- -THROTTlE• AS RESET~----------------

FUEL FLOW
TRANSMITTER

-,
I
I
I
I
I
I
I
I

SHUTOFF VALVEENGINE SIGNALS• COCKPIT ANTI-ICE• ENGINE FAULT• ENGINE CONTROL° FTlT• ICE DETECTOR SIGNAL

MAIN FUEL• NOZ POS
ENGINE• OIL
CONTROL CaDUNG______________
o_RP_M____________-"e=,-",c___.:..:FU!!:. SHUTOF~ALVE.J• MACH

IF_ HS"'_1_104SX.

Figure I-II .

, -55

T. O. 1F- 16A- 1
Digital Electronic Engine Control (DEEC) ! PW2201

The DEEC is an engine-mounted, fuel-cooled, solidstate digital computer. It controls the scheduling ofengine fuel flow in PRI, nozzle position, CIVV's,
RCVV's, start bleed strap position, and AB fuel flow
sequencing. The DEEC provides electrical signals to
the MFC, CENC, and AB fuel control for engine stallrecovery, segment 5 AB fuel flow redistribution,
segment 1 AB limiting, and AB ignition.
The DEEC closed-loop idle control schedules MFC
idle fuel flow to maintain a constant temperaturecorrected fan speed which results in constant idle
thrust in flight and on the ground. To reduce the idle
thrust level, the nozzle is commanded open when thethrottle is at or near IDLE and the LG handle is DN.
The DEEC limits minimum engine rpm throughout
the flight envelope to maintain stable operation. At
high altitude, low airspeed conditions, the DEEC
protects against engine stalls. During transonic andsupersonic conditions, the DEEC limits minimum
idle rpm as a function of mach number (from CADC)
to provide sufficient engine airflow.

To minimize the possibility of stalls during AB
operation at high altitude and low airspeed, theDEEC commands termination of segment 5 AB. At
extremely high altitude and low airspeed, the DEEC
limits AB operation to segment 1 AB. When a stall issensed, the DEEC cancels the AB (if throttle is in AB
range) and opens the nozzle until the stall clears. For
subsequent AB operation, the throttle must be
retarded below AB before AB can be reinitiated.
An engine overspeed or overtemperature condition
causes the DEEC to automatically transfer to SEC
and illuminate the SEC caution light.

The DEEC receives power directly from the engine
alternator. In the event of an engine alternator orengine gearbox failure indicated by rapid decrease tozero rpm and illumination of the ENGINE warning
light, the DEEC loses power and an automatictransfer to SEC occurs.
Secondary Engine Control (SEC) ! PW2201

The SEC is a hydromechanical system which
provides engine control in the event of a DEEC
system malfunction. In SEC, the CIVV's move to afIxed (cambered) position, nozzle position is closed,

1-56

the RCYVs are positioned by a hydromechanicalcontrol in the MFC, and AB operation is inhibited.
SEC is selected manually with the ENG CONT switch
or automatically by the DEEC. During SEC
operation, the SEC caution light illuminates.
Main Fuel Pump! PW2201

The gearbox-mounted main fuel pump provides
pressurized fuel to the MFC and boosts pressure to
the AB fuel pump.
Afterburner (AB) Fuel Pump! PW2201

The AB fuel pump is driven by engine bleed air andprovides pressurized fuel to the AB. The pump
operates only during AB operation.
Compressor Inlet Variable Vane (CIW) Control!PW2201

The CIVV control positions the CIYVs using MFC
fuel pressure in response to an electrical signal fromthe DEEC. In SEC, the CIYVs are in a fixed(cambered) position.
Rear

Compressor

Variable

Vanes(RCVV's)!PW2201

The first three stages of the rear compressor are
equipped with variable geometry vanes. RCYVs arecontrolled by the DEEC and are positioned usingpressurized fuel from the main fuel pump. In SEC,
the RCYVs are positioned by a hydromechanicalcontrol in the MFC.
Compressor Bleed Air! PW2201

Low-pressure bleed air is directed from the bleed
strap into the fan duct to increase the compressor
stall margin during starting. Pressurized fuel fromthe main fuel pump is used to drive the start bleed
actuator. The bleed valve is scheduled as a function ofengine rpm by the DEEC when starting in PRI and as
a function of time and engine inlet pressure in SEC.
High-pressure bleed air is supplied to the EPU andengine nacelle ejectors. It is also used for engine inlet
anti-icing, to drive the AB fuel pump, and to drive the
CENC motor.
Either low-pressure or high-pressure air is provided
to the ECS depending on engine bleed pressure levels.

T.o. 1F- 16A- 1
Pressurization and Dump Valve I PW2201

A pressurization and dump valve is located in theengine fuel manifold line between the fuel/oil cooler
and fuel nozzles. It provides a minimum fuel pressure
for MFC operation at low rpm and dumps the engine
fuel manifold when the throttle is retarded to OFF.
EXHAUST NOZZLE I PW2201

The exhaust nozzle is variable and consists of two
sections. The divergent nozzle floats freely and moves
in conjunction with the convergent nozzle. The
convergent nozzle is controlled by the convergent
exhaust nozzle control.
Convergent

Exhaust Nozzle

Control(CENC)I PW2201
The CENC is actuated by a high-pressure bleed air
motor. The nozzle schedule is controlled by the DEEC
as a function of throttle input to the MFC. In PRI with
the LG handle down, the nozzle is approximately
70-95 percent open at IDLE (idle area reset). As thethrottle is advanced, the nozzle closes. With the LG
handle up, the nozzle is near minimum area except
when approaching MIL or above. At MIL and above,
the DEEC schedules the nozzle to control enginepressure ratio as a function of fan speed. When thethrottle is advanced in the AB range, the DEEC
commands the nozzle open to compensate forincreasing AB fuel flow. In SEC, the nozzle ispositioned to the closed position and AB operation is
inhibited.
Light-Off Detector (LOD) I PW2201

The engine incorporates an AB LOD, which, whencombined with the DEEC logic, provides AB no-lightand blowout detection. When the LOD senses an AB
no-light or blowout, the DEEC automaticallyterminates AB fuel flow. If the throttle is left in AB,
the DEEC attempts AB light-off up to three times. If
these attempts are unsuccessful, the throttle must be
retarded to MIL or below and then advanced into AB
for further AB attempts.
ENGINE DIAGNOSTIC UNIT (EDU) I PW2201•

The EDU operates in conjunction with the DEEC toautomatically acquire and record diagnostic datawhenever the engine is operating. 1M!] , ~ In addition,

the EDU communicates with the avionic MUX BUS
for real time reporting of faults. All EDU operation
takes place automatically. Approximately 10 seconds
of engine data is recorded during certain predetermined events. Data may also be recorded by placing
the AB RESET switch to ENG DATA. In either case,data is recorded 8 seconds prior to and 2 seconds
following the event or switch movement. Use of the
AB RESET switch overwrites any previously
recorded data.
ENGINE OIL SYSTEM I PW2201

The engine is equipped with a self-contained oil
system to lubricate the engine and gearbox. Systempressure is nonregulated and varies with rpm, oiltemperature, and altitude.
Below approximately 35,000 feet MSL, oil pressure
should increase approximately 15 psi from IDLE to
MIL. At very high altitudes (50,000 feet), the oilpressure increase is approximately 5 psi from IDLE
to MIL. At all altitudes, however, a definite oilpressure increase should be evident when the rpm isincreased. Refer to SERVICING DIAGRAM, thissection for servicing/specifications information.
ENGINE ANTI-ICE SYSTEM I PW2201

The anti-ice system routes high-pressure engine
bleed air to and through the fixed fan inlet guidevanes, the CIVV's, and the inlet pressure probe
support cone to prevent ice formation. Additionally,
the inlet probe is continuously heated electrically toprevent ice formation. The system is controlled by the
DEEC and a three-position ANTI ICE switch. The
anti-ice system can be activated manually by placing
the ANTI ICE switch to ON or automatically, if the
ANTI ICE switch is in AUTO and a sensor located in
the inlet senses the accumulation of ice. Activation
can also occur if the detection system fails or if
essential dc bus No.1 power is lost (unless inhibitedby the DEEC).
The inlet strut is electrically heated to prevent ice
buildup. This heater is also controlled by the ANTIICE switch for manual or automatic operation.
The DEEC prevents anti-ice operation above 30,000
feet MSL and when engine inlet or bleed airtemperatures are high. In addition, a DEEC
malfunction may result in loss of bleed air for engine
anti-icing.

Change 121-57

10. 1F- 16A- 1
Engine ANTI ICE Switch I PW2201

The engine ANTI ICE switch is located on the rightconsole.

The ADG powers the main generator (through theconstant-speed drive (CSD», system A and B
hydraulic pumps, and FLCS PMG. The JFS is also
mounted on the ADG.

Functions are:
ENGINE ALTERNATOR I PW2201• ON - Engine anti-ice system is activated if notinhibited by the DEEC. The inlet strut electricalheater turns on.
• AUTO - Ice detector is operating. When an ice
accumulation is sensed (automatic detection may
not occur on the ground or above 7 degrees AOA),
the engine anti-ice system activates (unlessinhibited by the DEEC) and the inlet strut electricalheater turns on.
• OFF - Electrical power closes the engine anti-ice
valve (loss of power opens the valve). The inlet strut
heater and ice detector are turned off.
ENGINE AND ACCESSORY DRIVE GEARBOXESI PW2201
Refer to figure 1-12. The engine gearbox drives themain fuel pump, the oil pump assembly, the engine
alternator, and the PTO shaft, which powers theaccessory drive gearbox (ADG).

The engine alternator is driven by the engine gearbox
and provides sale power for the DEEC, engine and AB
ignition, inlet pressure probe heater, and the rpm
signal to the RPM indicator.
ENGINE IGNITION SYSTEM I PW2201

The ignition system is powered by the engine
alternator and contains four igniter plugs (two for theengine and two for the AB). With the throttle at or
above IDLE and engine rpm at 12 percent or above,engine ignition is continuous. When the throttle is
moved into AB, AB ignition is activated by the DEEC
for up to 3 seconds or until the LOD detects an AB
light. In the event of an AB blowout or no-light
condition with the throttle left in AB, AB ignition is
automatically resequenced by the DEEC up to three
additional times. For subsequent AB ignition, thethrottle must be retarded to MIL or below and thenreturned to AB.

Engine and Accessory Drive Gearboxes
ENGINE F100-PW-220

2

1. Tower Shaft (From Engine)
2. Main Fuel Pump
3. Engine Gearbox
4. Engine Alternator
5. JFS Exhaust Duct
6. Jet Fuel Starter
7. PTO Shaft
8. System B Hydraulic Pump
9. Accessory Drive Gearbox
10. FLCS PMG
11. System A Hydraulic Pump
12. Constant-Speed Drive
13. Main Generator
14. Engine Oil Pump

8

1F-16A-1-0013X@

Figure 1-12.

1-58

7

T.O. , F- , 6A-1
JFS Switch ~ mEl I PW2201

JET FUEL STARTER (JFS) I PW2201

Functions are:
The JFS is a gas turbine which operates on aircraft
fuel and drives the engine through the ADG. The JFS
is connected by a clutch to the ADG and only provides
torque when required to maintain engine rpm. If the
ADG is not able to rotate (i.e., seized €ngine), the JFS
runs, but the clutch prevents it from rotating theADG. The JFS receives fuel at all times regardless ofthe FUEL MASTER switch position. The JFS isstarted by power from two brake/JFS accumulators
used either singly or together. The brake/JFS
accumulators are charged automatically by hydraulicsystem B or manually by a hydraulic hand pumplocated in the left wheel well. Automatic recharging
takes between 40 seconds (hot ambient conditions)and 60 seconds (cold ambient conditions). The JFS isused to start the engine on the ground and to assist inengine airstart. Refer to JET FUEL STARTER
LIMITS, Section V

ENG & JET START CONTROL PANEL I PW2201

I

Refer to figure 1-13. The ENG & JET START control
panel is located on the left console.

• OFF - Normal switch position. The JFS can be shut
down at anytime by selecting OFF. The switch
returns to OFF automatically during a normal
ground start at 50 percent rpm.
• START 1 - Vents one of the brake/JFS accumulators to the hydraulic start motor.
• START 2 - Vents both brake/JFS accumulators to
the hydraulic start motor.
JFS RUN Light ~ mEl I PW2201

The green JFS RUN light illuminates within 30
seconds after initiating JFS start to indicate that the
JFS has attained governed speed.
JFS Operation I PW2201

During a ground engine start, the brake/JFS
accumulators begin to recharge after the engine
accelerates through 12 percent rpm. As the engine
accelerates through 50 percent rpm, a sensor causes
the JFS to shut down automatically and the JFS RUN
light goes off.

ENG & JET START Control Panel (Typical)
1

ENGINE F100-PW-220~ETFUEl~

0 ~START
rID
lOFF

N

G
&

1
T
L;:
F

R
I

C
T
I
0

N

\J

3

START?

2

ENG CONT
PRI

A~3
NORM

ENG DATA

f~~

4

1. RUN Light (Green)

2. JFS Switch
3. ENG CaNT Switch
4. AS RESET Switch
5. MAX POWER Switch

5

NOTE:

~ For ENG CaNT switch,refer to F-16S AIRCRAFT,
this section.

1F-16A-1-1049A@

Figure 1-13.

1-59

T.O. 1F-16A-1

During in-flight operation, the brake/JFS accumulators begin to recharge (provided system B hydraulicpressure is available) when the JFS reaches 70
percent of governed speed (3-4 seconds before the JFS
RUN light illuminates). If the JFS RUN light does notilluminate within 30 seconds or the JFS RUN lightgoes off once illuminated, the JFS START switch will
not reengage and the JFS cannot be restarted until
the JFS has spooled down. JFS spooldown takes
approximately 17 seconds from full governed speed.
Once running, the JFS does not shut down until the
JFS switch is manually positioned to OFF.

ENGINE FAULT Caution Light LESS ~, LESS fI!)
, PW2201

The ENGINE FAULT caution light, located on thecaution light panel, indicates an AB malfunction orloss of mach signal to the DEEC. When the lightilluminates for an AB malfunction, AB operation may
be partially or completely inhibited. If the CADC
caution light also illuminates, loss of mach signal is
the probable cause. In this case, AB operation is notinhibited; however, supersonic flight should be
avoided due to potential inlet buzz.

ENGINE CONTROLS AND INDICATORS I PW2201

SEC Caution Light' PW2201

Refer to figure 1-14. The engine instruments arelocated on the right side of the instrument panel.
Refer to ENGINE LIMITATIONS I PW2201, Section V.

The SEC caution light, located on the caution light
panel, indicates that the engine is operating in SEC
or that main fuel pump pressure is low.

ENG CONT Switch' PW2201

MAX POWER Switch ~ rnfJ' PW2201

The ENG CONT switch (guarded out of SEC) is
located on the left console. ~ For ENG CONT switch
differences, refer to F-16B AIRCRAFT, this section.
Functions are:

The MAX POWER switch, located on the left console,
is solenoid held in the MAX POWER position when
the throttle is at MAX AB and airspeed is 1.1 mach or
greater. Refer to ENGINE LIMITATIONS , PW2201,
Section V.

• 00 [!ffi PRI - DEEC in operation (normal position).

Functions are:

• SEC - SEC operation. Transfer occurs when theswitch is moved to the SEC position.

• MAX POWER - Delivers maximum thrust byallowing maximum FTIT to increase by 22°C .

AB RESET Switch 00 mEl' PW220 I• OFF - Normal (deenergized) position.

The AB RESET switch, located on the left console, isa three-position toggle switch, spring-loaded to center(NORM) position.

RPM Indicator I PW2201

Functions are:• AB RESET - When operating in primary, thisposition is used to attempt to clear the DEEC of an
AB fault or to reestablish a mach signal to the
DEEC.
• NORM - Normal (deenergized) position.
• ENG DATA - This position may be used to record
engine data in the EDU.
ENGINE FAULT Caution Light~, fI!)' PW2201

The ENGINE FAULT caution light, located on thecaution light panel, indicates that an engine PFL
item was detected. The ENGINE FAULT caution
light goes off when the fault is acknowledged. Refer
to T.O. 1F-16A-34-1-1 or ~ T.O. 1F-16A-34-1-4 for a
detailed description of fault recall.

I

1-60

Change 13

The RPM indicator has a pointer display and the rpm
signal is supplied by the engine alternator. RPM isexpressed in percent from 0-100. The indicator is
powered by ~ battery bus No. 2, LESS~ the battery
bus.
NOZ POS Indicator I PW2201

The NOZ POS indicator displays the position of the
CENC exhaust nozzle drive shafts which arecalibrated from 0 percent (closed) to 100 percent (fullyopen). The indicator accurately reflects exhaust
nozzle position in PRI and SEC unless both drive
shafts are failed. The indicator is powered byessential ac bus No.2.
FTIT Indicator I PW2201

The FTIT indicator displays an average FTIT indegrees c. The indicator has a range of200o-1200°C
in major increments of 100°C and is powered by [B:Q]
battery bus No.1, LESS ~ the battery bus.

T. O. 1F- 16A- 1

Engine Controls and Indicators (Typical)
Block 10ENGINE F100-PW-220

4

&

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T
A
R
T

t;
F

R

I
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I

ENG CONT

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FUEL HOTENGINE

SEC

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TE

OFF

OVERHEATENGINEFAULT

L

AS RESETgNORIot

11

12

ENG DATA

0

N

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SEC
FUEL HOT11

OVERHEATENGINE
FAULT
1. Throttle
2. ENG CONT Switch
3. ENG FIRE Warning Light (Red)
4. FUEL FLOW Indicator
5. HYD/OIL PRESS Warning Light (Red)
6. ENGINE Warning Light (Red)

7. OIL Pressure Indicator
8. RPM Indicator
9. NOZ POS Indicator
10. FTIT Indicator
11. Caution Lights (Amber)
12. ANTI ICE Switch

1F-16A-1-1 046X ®

Figure 1-14. (Sheet 1)

1-61

10. 1 F-16A-l

Engine Controls and Indicators (Typical)
ENGINE F100-PW-220

4START 2ENG CONT
PRI

T
A

u-tSECOVERHEATTAS RESETENGINE
FAULT

S

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ENG DATA

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ENGINE
FAULT

11

SEC
FUEL
HOT
1. Throttle
2. ENG CONT Switch
3. ENG FIRE Warning Light (Red)
4. FUEL FLOW Indicator
5. HYD/OIL PRESS Warning Light (Red)
6. ENGINE Warning Light (Red)

Figure 1-14. (Sheet 2)

1-62

7. OIL Pressure Indicator
8. RPM Indicator
9. NOZ POS Indicator
10. FTIT Indicator
11. Caution Lights (Amber)
12. ANTI ICE Switch

1F-16A-1-1047X®

T.O.1F-16A-1
FUEL FLOW Indicator I PW2201

The FUEL FLOW indicator is a pointer-counter
indicator which displays the total fuel flow to theengine, including AB, in pph. The indicator has arange of 0-80,000 pph and is powered by essential ac
bus No.1.
OIL Pressure Indicator I PW2201

The OIL pressure indicator displays engine oilpressure from 0-100 psi and is powered by essential
ac bus No.2.
HYD/OIL PRESS Warning Light I PW2201

The HYD/OIL PRESS warning light, located on theedge of the right glareshield, serves as a monitor ofengine oil pressure and hydraulic system pressure.
For engine oil pressure, the warning light illuminates when oil pressure has been below approximately 10 psi for 30 seconds (time delay minimizes
warning light illuminating during maneuvering).
The light goes out when oil pressure exceeds
approximately 20 psi. For hydraulic pressure, thewarning light illuminates when either A or B systempressure decreases below 1000 psi. The light goes out
when both system A and B pressures are above 1000
psi. During engine start, the warning light usually
goes off before reaching idle rpm; however,acceptable operation is indicated if the light goes off
before exceeding 70 percent rpm and remains off
when the throttle is retarded to IDLE. The warning
light is powered by ~ battery bus No.1, LESS ~ thebattery bus.

REDUCED IDLE THRUST (RIT) Switch I PW2201

The RIT switch, located on the left sidewall just aft of
the throttle, is inoperative.
Throttle IPW2201

Refer to figure 1-15. The engine is controlled by athrottle mounted above the left console with detents
at OFF, IDLE, MIL, and MAX AB. The throttle ismechanically connected to the MFC. The OFF
position terminates engine ignition and fuel flow. The
IDLE position commands minimum thrust and isused for all ground starts and airstarts. From IDLE
to MIL, the throttle controls the output of the engine.
Forward of the MIL position, the throttle modulates
the operation of the AB (through five segments) while
maintaining constant basic engine operation.
~rm

The throttle must be rotated outboard to allow
advancement from OFF to IDLE and from MIL to AB.
Retarding the throttle from AB to MIL automaticallyrotates the throttle. At IDLE, a cutoff release at thebase of the throttle must be actuated to allow thethrottle to be rotated outboard and retarded to OFF.
~ For throttle differences, refer to F -I6B AIRCRAFT, this section.
A single white reflective stripe is located IAl rntl on
both the upper surface of the throttle foot and on the
sidewall fairing, g on both the lower throttle radius
next to the console and on the panel outboard of thethrottle radius. Alignment of the two stripes aids in
identifying the IDLE position.
Six switches are located on the throttle. ~ mE A
throttle friction control is located inboard at the baseof the throttle. 00 The throttles are mechanicallylinked together.
ENGINE OPERATING CHARACTERISTICS I PW2201

ENGINE Warning Light I PW2201

The ENGINE warning light, located on the edge ofthe right glareshield, illuminates when RPM and
FTIT indicator signals indicate that an engine
overtemperature, flameout, or stagnation has
occurred. Illumination also occurs for an engine
alternator failure and may occur as a result of an
RPM or FTIT indicator failure. The warning lightilluminates when the rpm decreases to subidle(below 55 percent), when engine stagnates (determined from rpmlFTIT rates), or approximately 2
seconds after FTIT indication exceeds IOOO°C. The
warning light goes off when the condition thatturned it on is eliminated. The warning light ispowered by ~ battery bus No. 1, LESS~ the batterybus.

Ground Operations I PW2201

Since the DEEC maintains constant idle thrust, rpm
varies with temperature and pressure altitude(higher temperature or pressure altitude results inhigher rpm).
At MIL, the DEEC controls fan speed and enginepressure ratio to maintain consistent thrust. RPM
and FTIT vary as a function of flight conditions.
Non-AI Operation In Flight I PW2201

After a MIL takeoff, FTIT is usually 890°-960°C withrpm of 89-94 percent for any outside air temperature
above 2°C. FTIT and engine rpm are lower for
temperatures below 2°C. Regardless of temperature,
nozzle position should not exceed 30 percent at MIL.
Change 2

1-63

T.o. 1 F-16A-l

Throttle (Typical)
ENGINE F100-PW-220

13

6

9

11

8

1. [AJ[jfJ Throttle Cutoff Release
2. UHF VHF Transmit Switch (2-Way, Momentary Rocker)
3. 18:Q][AJ UHF HF VHF Transmit Switch (3-Way, Momentary Rocker)
4. MAN RNG/UNCAGE Knob/Switch (Rotate, Depress)
5. ANT ELEV Knob (Rotate, Center Detent)
6. DOG FIGHT Switch (3-Position, Slide)
7. SPD BRK Switch (3-Position, Aft Momentary)
8. RDR CURSOR/ENABLE Switch (Depress Multidirectional)
9. Throttle Foot
10. Throttle
11. [AJ~ Throttle Friction Control
12. [AJ~ IDLE Stripe
13. [AJ~ Throttle Stripe

NOTE:~ For throttle differences, refer to THROTILE and F-16B AIRCRAFT, this section.
1F-16A-1-1050X®

Figure 1-15.

1-64

10. 1F-16A- 1
Engine operation is continually optimized as flight
conditions change. This is evident by slight changes
in the NOZ POS, RPM, and FTIT indicator
indications.
At low altitudes (below approximately 10,000 feet),idle rpm should always be equal to or slightly higher
than the ground idle rpm. As altitude increases, idle
rpm increases to provide the engine sufficient stallmargin during throttle transients.
At 1.4 mach and above, the minimum thrust level is
MIL even though the throttle may be retarded below
MIL. Typically, the minimum thrust level increases
from idle to MIL between 0.84-1.4 mach. All of theminimum operating level features are deactivatedduring SEC operation.
A low frequency engine vibration may be sensed inflight or on the ground primarily at or near idle, but
may also occur at higher thrust settings. The
vibration has no adverse effect on engine or aircraft
structure and should disappear if engine rpm is either
increased or decreased. Vibrations that change in
intensity with throttle movement and are present
across the throttle/rpm range may indicate apotential engine malfunction.
AS Operation in Flight I PW2201

I Refer to figures 1-16 and 1-16.1. The DEEC monitors
AB operation and takes appropriate action to prevent
engine stalls. In AB, the DEEC provides thefollowing:• Fast acceleration capability: The AB has nolimitations during throttle transients from IDLE to
MAX AB. Near sea level, AB operation occurs
immediately after AB is selected. At high altitude,
a higher fan speed must be attained prior to AB
operation. For example, during an IDLE-to-MAX
AB throttle transient at low altitude, the AB lights
just above idle thrust and the total time from idle
thrust to MAX AB thrust is approximately 4
seconds. In contrast, at high altitude, the time from
idle thrust to MAX AB thrust is approximately 11
seconds.
• AB fuel flow redistribution: Flight at high altitude
and low airspeed results in the redistribution ofsegment 5 fuel flow to segment 3 to maintain ABstability.

• AB segment sequencing limiting: When AB isselected at extremely high altitudes and low
airspeeds, only segment 1 AB is scheduled. However,if this area is entered with AB above segment 1,
there will be no change. If AB segment 2 or greater
is selected while in this area, the engine
automatically sequences up to the requested throttle
position as the aircraft exits the area. A self-recoverable AB stall may occur during this automaticsequencing if the engine is operating on approved
fuels other than JP-4, NATO F-40, or JET B.
• AB recycle capability: The DEEC, in conjunction
with the LOD, provides automatic AB recycle
capability in the event of an AB blowout or no-light
condition (if the throttle is left in AB). In that event,the DEEC automatically resets the control system
to MIL, performs a control system check, andreattempts to light the AB up to three additionaltimes before returning the engine to MIL. If the
LOD is failed, the DEEC attempts one AB relightusing a duct pressure signal to verify AB lightoff. No
caution lights result from unsuccessful AB recycles.
Additional AB attempts can be made by moving thethrottle to MIL or below and then back into AB.
SEC Operation I PW2201

The engine transfers to SEC when the ENG CONT
switch is manually switched to the SEC position.
Transfer to SEC also occurs automatically if the
DEEC senses a major engine control system
malfunction or ifloss of electrical power to the DEEC
occurs.
When the engine transfers to SEC, the SEC caution
light illuminates and AB operation is inhibited. RPM
may increase or decrease slightly except at highaltitude where rpm and FTIT decrease significantly
if the transfer occurs with the throttle at or near
IDLE.
While subsonic in SEC, throttle movement isunrestricted between 15,000 and 40,000 feet MSL.
The throttle may be moved in the AB range; however,the AB is inhibited. Refer to ENGINE - OPERATIONAL ENVELOPE, Section V for transfer andthrottle movement restrictions.
SEC provides 80-100 percent of normal MIL thrust.
This level provides a measure of protection against
exceeding engine operating limits and provides
sufficient thrust for safe flight operations. SEC idle
thrust is approximately twice that in PRI with anormal nozzle during landing approach and ground
operations because the nozzle is closed.

Change 13

1-65

IT.O. 1 F-16A-l

Light-Off

AS Envelope

ENGINE F100-PW-220
NOTES:• Throttle movement is unrestricted throughout the aircraft flight envelope.
• Region 1 - Unlimited 5 segment AB operation.
• Region 2 - AB segments 1 through 4 available.
• Region 3 - AB segment 1 available.
• Region 4 - AB inhibited.

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IF-16X-l-0006X@

Figure 1-16,

1-66

10. 1 F-16A-l

AB -Operating Envelope
ENGINE F 1 OO-PW-220• Region 1 - Unlimited AB Operation•

AB recycles not expected but occasional recycles may occur during AB
selection and operation .

• Region 2 - Unlimited AB Operation•

AB recycles may increase with decreasing airspeed and increasing altitude.

•

The DEEC will attempt up to three recycles if the initial AB
attempt is not successful or if the AB blows out. If the recycles are
unsuccessful, the throttle must be placed to MIL and back to AB
for additional attempts.

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NOTE:

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IF-16X-1-3QQ6X@

Figure 1-16.1.

Change 13

1-66.1/(1-66.2 blank)

T. O. 1F- 16A- 1

FIRE AND OVERHEAT DETECTIONSYSTEM
The fire and overheat detection system consists of two
separate parallel loop sensing systems, one for fire
and the other for overheat. The fire detection loops
are routed through the engine compartment. The
overheat detection loops are routed through theengine compartment, MLGwheel wells, ECS bay, andEPU bay. Activation of the overheat detection loops
occurs approximately 100°C below the activation
temperature of the fire detection loops. The firewarning signal causes the ENG FIRE warning lightto illuminate. The overheat signal causes the
OVERHEAT caution light to illuminate. When thetemperature of the element drops below the criticaltemperatures, the signal ceases, allowing the ENG
FIRE warning or the OVERHEAT caution light to go
off. The detection circuit is powered by the essential
ac bus No.2 and the ~ battery bus No.1, LESS ~the battery bus.
FIRE & OHEAT DETECT TEST BUnON ~ [If!

Refer to figure 1-17. The FIRE & OHEAT DETECT
test button, located on the TEST switch panel, checks
continuity of both systems and illuminates the ENG
FIRE warning light and the OVERHEAT caution
light.

TEST Switch Panel~ IBFI (Typical)

I
1. MAL & IND LTS Test Button
2. OXY QTY Indicator Test Switch
3. FLCS PWR Lights (Green)
4. FLCS PWR TEST Switch
5. [!f] STICK CONTROL Switch
6. EPU/GEN Test Switch
7. PROBE HEAT Switch
8. FIRE & OHEAT DETECT Test Button
1F-18A-1-1052Xct

Figure 1-17.

FUEL SYSTEM
Refer to figure 1-18 for a simplified system diagramand figures 1-19 and 1-20 for system schematics. The
fuel system is divided into seven functional
categories. These are the fuel tank system, fueltransfer system, fuel tank vent and pressurization
system, engine fuel supply system, fuel quantity/fuellevel sensing system, fuel tank explosion suppression
system, and refuelingldefueling system.
FUEL TANK SYSTEM

Refer to figure 1-21. The aircraft has seven internalfuel tanks located in the fuselage and wings that are
integral to the structure. There are provisions forcarrying three external tanks on the wings and thecenterline station. Five of the internal tanks arestorage tanks: the left and right wing tanks, two
forward fuselage tanks (F-1 and F-2), and the aftfuselage tank (A-1). The two internal reservoir tanks(forward and aft) supply fuel directly to the engine. 00
The F -1 fuel tank is reduced in size to allow room forthe rear cockpit.
FUEL TRANSFER SYSTEM

Fuel is transferred by two independent methods. The
primary method provides a siphoning action throughstandpipes connecting the fuel tanks. Siphoning
action depends on the absence of air in the bays
receiving fuel. Air ejectors in each reservoir tank
automatically expel air. In case of failure of thesiphoning system, powered fuel pumps work
continually to pump fuel from the internal tanks to
the reservoirs. The powered transfer system also
scavenges tanks to minimize unusable fuel by using
electrically driven pumps and pumps powered bybleed fuel pressure from the engine manifold. Bothmethods operate simultaneously and independently
to transfer fuel through the system.
The transfer system is divided into two separate tanksystems, the forward and the aft. The forward system
consists of right external tank (if installed), rightinternal wing tank, F-1, F-2, and forward reservoir.
The aft system consists of left external tank (ifinstalled), left internal wing tank, A-1, and aftreservoir. If a centerline tank is installed, it isconsidered to be part of both forward and aft systems.
The wing external tanks empty into the respective
internal wing tanks. Fuel flows from the internal
wing tank to the fuselage tanks and then to theforward and aft reservoirs. Fuel is pumped to the
Change 1

1-67

T.0 .1 F- 16A-1

Fuel Tank System (Typical)
Q;NTERUNE
FUEL TANKLEFT EXTERNALRIGHT EXTERNAL
FUEL TANKFUEL TANKLEGEND:_

BOOST PUMP ..... SSr . . J FUEL TRAHSF£R_

AFT/lEFT FUEL TANK SYSTEM_

FWD/RIGHT FUEL TANK SYSTEM_

EXTERNAl.. FUEL TANK

::~:

•

•

ENGINEELECTRlCAl

C ) FUEL TANK COhHECTING STANDPIPE()J TRANSfER PUMP FUEl PRESSURE DRMN)TRAHSffRlREFVEl PRESSURlZAnON
RfFUEl../TRANSfERoom DISCONNECT

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TRANSFER SHUTOFF VAlVE<I>

fUEL FlOW DIRECTIONSTANDPIPE OPENING AT TOP OF TANKSTANDPIPE OPENING AT 8OnOM Of TANKBOOST PUMP !ELECTRICAl)
REFUEL SHUTOff VAlVENOTE: This is • simplified (AI diagram. Not .11 Iv-' systemitems are shown Icroufeod. mouve Row. etd. Ref.r
to FUEL SYSTI:M SCHEMATIC for ctet..ils.
'F-16A-, -106lJt.

Fi6urc 1-18.
1-68

TRANSFER PUMP IELECnUCAU

T.O. 1F- 16A-1

engine from the reservoirs. 1b automaticallymaintain the CG, fuel is transferred through theforward and aft systems simultaneously.

high altitude. An external tank vent and pressurization valve regulates pressure supplied to the externaltanks.

If external tanks are installed, air pressure transfers
fuel to the internal wing tanks. If the EXT FUEL
TRANS switch is in NORM, the sequence offuel flow
is from the centerline tank to the internal wing tanks.
Mter the centerline tank empties, each external wingtank flows to its respective internal wing tank.

If the combat schedule (reduced pressure) is activatedby the TANK lNERTING switch, Halon, if available,
is mixed with air and the internal tank vent andpressurization valve controls the pressure.

The external tank fuel transfer valve in each internal
wing tank shuts off fuel to prevent overfilling theinternal tanks. If one of these valves fails, a float
switch senses fuel and shuts off all external tank fueltransfer before fuel flows overboard. By placing the
EXT FUEL TRANS switch to WING FIRST, theexternal wing tanks empty before the centerline tank,and the float switch does not prevent fuel fromspilling overboard if a transfer valve fails.

I

The automatic forward fuel transfer system supplements the function of the FFP by preventing
undesirable aft CG. The automatic forward fueltransfer system operates only when the FUEL QTY
SEL knob is in NORM and the total forward fuselage
fuel quantity indication is less than 2800 ([I] 1500)
pounds. In the~, forward fuel transfer starts when
the forward heavy fuel differential drops below 300
pounds and stops when the forward heavy fueldifferential reaches 450 pounds. In the [1], forward
fuel transfer starts when the aft heavy fueldifferential exceeds 900 pounds and stops when theaft heavy fuel differential reaches 750 pounds. Thissystem does not correct a forward fuel imbalancesince it only transfers fuel from aft to forward.
For proper operation, the automatic forward fueltransfer system depends on a properly functioning
fuel quantity indicating system. Fuel is transferredthrough a solenoid-operated trim valve powered fromessential dc bus No.2. The automatic system isdeactivated if electrical power is lost through failure,by moving the FUEL QTY SEL knob out of NORM, orduring gravity feed conditions.
FUEL TANK VENT AND PRESSURIZATION SYSTEM

The fuel tank vent and pressurization system
supplies cooled pressurized air from the ECS to force
fuel from the external tanks to the internal wingtanks and to power the air ejector pumps whenever
the AIR SOURCE knob is in NORM or DUMP. It also
prevents fuel in internal tanks from vaporizing at

If the AIR SOURCE knob is placed in OFF or RAM or
if the ECS is inoperative, tank pressurization is notavailable and external fuel cannot be transferred.
With multiple generator failures, fuel tank pressurization continues and external fuel still transfers.
ENGINE FUEL SUPPLY SYSTEM

Refer to figure 1-22. When the ENG FEED knob is in
NORM, boost pumps in the forward and aft reservoirs
pump the fuel through the engine feedline to the fuelflow proportioner (FFP). In the FFP, twin constantdisplacement pumps, powered by hydraulic system A,
supply equal amounts of fuel from each reservoir tomaintain CG. Two fuel lines with check valves can
bypass the FFP in case it fails so that fuel flow will not
be interrupted. After fuel flows through the FFP, asmall amount of cooling fuel is routed to the IPW200i
EEC, IPW2201 DEEC and then returned to thereservoirs. The remainder of the fuel passes througha fuelloil heat exchanger to cool hydraulic systems A
and B, the main generator CSD, and the ADG. Thenfuel flows through an electric main fuel shutoff valve
which has a full travel time of 2-4 seconds and iscontrolled by the FUEL MASTER switch. (The JFS
receives fuel at all times regardless of the FUELMASTER switch position.) After passing through themain fuel shutoff valve, fuel passes through the fuelflow transmitter (which operates the FUEL FLOW
indicator) to the engine.
FUEL QUANTITY INDICATING SYSTEM

Refer to figure 1-23. The fuel quantity indicating
system displays the amount and location of fuelremaining. The totalizer shows all fuel in theinternal and external tanks in pounds. The AL and
FR pointers show the fuel quantity in the tanks as
selected by the FUEL QTY SEL knob. Erroneous
fuel indications may occur during or immediately
after maneuvering flight. The selected tanks should
normally be the fuselage tanks (FUEL QTY SEL
knob in NORM). The difference between the forward
and aft tanks should remain essentially constantsince the FFP maintains an equal flow of fuel.

Change 1

1-69Ito.1F-16A-l

Fuel System [Al (Typical)

RIGHT
WING

TANK

lNERTING

SWITCHFun
QUAHTn'IfI)fCATaot

EXTERNAL TANK
(TYPtCAL)

Figuro 1-19. (SM.. 1)

1-70

LEFT
WING

T.O.1F-16A-l

ENG FEEDFUEL MASTERKNOB

SWITCHLEGEND,(F;~~ FUEl flOW
INDICATOR

PUMP
DRIVE
PRESSURE
FROM

"YO
SYS A

FUEL FLOW
TRANSMlnER

FUEL flOW
PROPORTIONERCAUTION liGHT

o

oo
(J)
o

©

9
@

TRANSfER SHUTOFF VALVE
TRANSFER PUMP IELECTR1CAL)
BOOST PUMP (ELECTRICALIREFUEL SHUTOFF VALVESOLENOlD VALVECROSSFEED VALVE
TRANSFER PUMP!FUEL PRESSURE DRIVEN)
SUCTION FEED VALVECD DISCONNECT[])

FUEL/Oil HEATEXCHANGERGROUND COOLING RECEPTACLEGFUEL EJECTOR PUMPOIT]

ELECTRIC MOTOR OPERATEDSHUTOFF VALVE[!]CHECK VALVEORIFICE

9

TEMPERATURE SENSOR
STANDPIPE OPENING AT TOP
OF TANK
STANDPIPE OPENING AT
BOTTOM OF TANK
0:::) FUEL. TANK CONNECTING
STANDPIPE():::]

A-I

o

•
••

TO RIGHT WING EXT TANKTO CENTERLINE TANKAHALON/AIR MIXING VALVEEXT TANK VENT & PRESS VALVEINT TANK VENT & PRESS VALVEREMOTE SENSING PRESSURE
RELIEF VALVEAR RECEPTACLE

B
C

o

E

ELECTRICAL----

rz:zI OVERBOARD VENT
c::::::J FUEL SUPPLYGROUND REFUEliNG

RECEPTACU:BOOST PUMP PRESSURE
PUMP DRIVE PRESSURETANK INERTING PRESSURE~ FUEL TRANSFERr::=-

CAUTION LIGHTS

TRANSFEA/REFUEl
PRESSURIZATIONTANK PRESSURIZATIONREFUEVTRANSFERECS PRESSURIZATION SUPPLYTRANSFER BLEED

IF·HIA-'-1054·2)(.

Figure 1·19. (Sheet 2)

1-71

IO. 1F-1M-1

Fuel System rn:J (Typical)

RIGHT
WING

*

F-2

TAlI<
ItuTING

SW"'"

LEFT
WING
FWD , AfT fl.ll
QUANTITY INOICATOfiS

I F·l6A-'_ l OSS- 1)(eFigu.re 1·20. (Sheet 1)

1-72

IO.IF-16A-l
FWD fUEl
ArT MLfWD FUELMASTER SWITCH MASTER SWITCH flOW INDICATORAFT FUEl
flOW INDICATOR

fROM
SYS

HY.

®

oo

''''' FEED
KNOO

PUMP
DRIVE
PRESSlJIE

LEGEND:
TRANSFER SHUTOFF VALVE

TRANSFER PUMP (ELECTRICAl!
BOOST PUMP {ELECTRICAl!

(J) REFUEL SHUTOFF VALVE®

FUEL flOW
PftOPORTlOfD

AFT CAUTION©

SOLENOID VALVECAOSSFEED VALVEg

TRANSFER PUMP(FUEL PRESSURE DRIVENIrn

DISCONNECT~

FUEL EJECTOR PUMP

@ SUCTION FEED VALVE.iI.~ TO ENGINEFLOW TRANSMITTER

rI~..:::~~;t£AT

EXCHANGER

CD GROUND COOLING RECEPTACLEOOJELECTRIC MOTOR OPERATEDSHUTOFF VALVE[!]CHECK VALVETEMPERATURE SENSOR
STANDPIPE OPENING AT TOP
OF TANK
STANDPIPE OPENING AT
BOnOM OF TANKFUEL TANK CONNECTING
STANDPIPE

'VI

Q:::l

o

A-I

0::::>ORIFICE•TO RIGHT WING EXT TANK

••
A
B
C

TO CENTERLINE TANKHALON/AIR MIXING VALVEEXT TANK VENT & PRESS VALVEINT TANK VENT & PRESS VALVEREMOTE SENSING PRESSURE
RELIEF VALVEAR RECEPTACLEoE

ELECTRICAL~ OVERBOARD VENT

C:=J FUEL SUPPLYGROUND REFlElINGRECEPTACL£FWD ..

STATUS

INDICATOR_BOOST PUMP PRESSURE_

PUMP DRIVE PRESSURE~ TANK INERTING PRESSURE~ FUEL TRANSFER

AFT..

1}

STATUS
INDICATc:J

iii

B

_

-=_

TRANSFER/REFUELPRESSURIZATIONTANK PRESSURIZATIONREFUEl./TRANSFER_ECS PRESSURIZATION SUPPlY_

TRANSFER BLEED

FWD

fWD CAUTION
LIGHTS

AFT CAUTION
LIGHTS

FUEl LOW
AFT
FUEL LOW

If 161\ 1 10551X.

Figure 1-20. (Slu!et 2)

1-73

10. 1F- 16A- 1

Fuel Quantity Indication and Tank
Arrangement (Typical)

FUEL QUANTITY INDICATIONS (JP-8) WITHNORM SELECTED AND FULL INTERNALTANKS WITH TWO FULL 370-GALLON
EXTERNAL WING TANKS.
1 F-16A-1-0019B@[ID~
TANK LOCATIONFUEL QTY SELKNOB SETTINGS POINTER

1. LEFT INTERNAL WING
2. RIGHT INTERNAL WING

I

AL
FR

FUEL
QTY (LB) JP-5/8

525 ± 100525±100550± 100550± 100525±100525±100550± 100550± 100

3100± 1003250± 1001800± 100

1890± 100

3. F- 1 FUSELAGE
4. F-2 FUSELAGE
5. FWD RESERVOIR

}

NORM

FR

6. AFT RESERVOIR
7. A-l FUSELAGE} NORM

AL*2800± 100**2940± 100*2800± 100**2940± 100

RSVR
RSVR

FR
AL

460±30460±30480±30480±30460±30460±30480±30480±30EXT CTR
EXT WING
EXT WING

FR
AL
FR

1800± 1002300± 1002300± 100

1890± 1002420±1002420±1001800± 1002300± 1002300± 100

1890± 1002420± 1002420±100

TOTAL INTERNAL FUEL

*6950±300**7290±300*5650±300**5930±300

TOTAL EXTERNAL FUEL

6400±300

6730±300

6400±300

6730±300

5. FWD RESERVOIR
6. AFT RESERVOIR
8. CENTERLINE
9. LEFT EXTERNAL WING
10. RIGHT EXTERNAL WINGII

INTWING
INTWINGFUEL
FUEL
QTY (LB) JP-4 QTY (LB) JP-5/8

FUEL
QTY (LB) JP-4

NOTES:
1. Tolerances are due to indication errors with the variations in density resulting from temperatures, additives, etc.
2. The quantity of wing fuel varies depending upon aircraft attitude during refueling.
3. Usable fuel and indicated fuel quantities are approximately equal.
4. Indications are approximate and shall not be used for computing weight and balance data. Refer to 10. 1F-16A-1-1 for detailed
information.
5. ~ */** Subtract 125/130 pounds.

Figure 1-21.

1-74

Change 12

T. O. 1F- 16A- 1

Fuel Control Panel (Typical)
ENG FEEDAIR REFUEL 7TN

N~~OO~ WU

Off\0 @GSE
LESS

LESS

Em

m

I:t
o

1. FUEL MASTER Switch
2. TANK INERTING Switch
3. ENG FEED Knob
4. AIR REFUEL Switch

1

lEm

4

3

1F-16A-1-1057B@

1.

2.

CONTROLPOSITIONFUNCTION

FUEL MASTER Switch(lever lock)

MASTER (guarded)

Opens main fuel shutoff valve which thenopens the engine electronic control coolingfuel shutoff valve

OFF

Closes main fuel shutoff valve which thencloses the engine electronic control coolingfuel shutoff valve

TANK INERTING

Reduces internal tank pressurization. If
Halon is available, allows 20 seconds of initial
Halon flow; thereafter, allows a small meteredflow of Halon to the F-1, A-I, and internal
wing tanks

OFF

Stops Halon flow. Returns internal tank pressurization to normal schedule

TANK INERTING Switch(lever lock to OFF)

Figure 1-22. (Sheet 1)

Change 2

1-75

T. O. 1F- 16A- 1

Fuel Control Panel (Typical)
CONTROL
3.

4.

ENG FEED Knob

AIR REFUEL Switch

POSITIONFUNCTION

OFF

Deenergizes all electric-driven pumps. Engine
supplied by FFP

NORM

Energizes all pumps. CG maintained automatically

AFT

Energizes pumps in aft tanks and opens
crossfeed valve. Fuel is transferred from afttanks to the engine and forward tanks. CG
moves forward

FWD

Energizes pumps in forward tanks and opens
crossfeed valve. Fuel is transferred from forward tanks to the engine and aft tanks. CG
moves aft

OPEN

Opens slipway door. Places FLCS in takeoff
and landing gains
Enables slipway light. Turns on AR floodlightand vertical tail-mounted floodlight
Reduces internal tank pressurization, depressurizes external tanks, and allows the refuelvalve in each reservoir to open when a centerline tank is installed and refuel pressure isapplied

CLOSE

Reverses the OPEN actions

Figure 1-22. (Sheet 2)

1-76

10. 1 F- 16A- 1

FUEL Quantity Indicator and Select Panel(Typical)

BLOCK 10BLOCK 15

2

~ RS~'"~C~SNORM
YE
L

WING

EXT WINGTEST"XT CTR,%3n I

EXT FUEl~

WING FIRST

1. FUEL Quantity Indicator
2. FUEL QTY SEL Knob
3. EXT FUEL TRANS Switch

CONTROUINDICATOR
1.

2.

FUEL Quantity
Indicator

FUELQTYSEL
Knob

1 F-16A-1-0021 A ®

FUNCTIONPOSITION
AL and FR
pointers

Display fuel quantities as determined by the FUEL QTY
SEL knob

Totalizer

Displays total fuel in all fuel tanks (fuselage + wing + external)

Red portion of
AL pointershowing

Indicates fuel imbalance between forward and aft fuselage
tanks

TEST

AUFR pointers drive to 2000 (± 100) pounds
Totalizer drives to 6000 ( ± 100) pounds
Both fuel low caution lights illuminate

NORM

AL pointer displays sum of fuel in the aft (left) reservoir and A-I
fuselage tanks
FR pointer displays sum offuel in the forward (right) reservoir and F-I, F-2 fuselage tanks
Enables automatic forward fuel transfer system, LESS ~trapped fuel warning, and bingo fuel computation based on fuselage fuelFigure 1-23. (Sheet 1)

Change 12

1-77

IT. O. 1F- 16A- 1

FUEL Quantity Indicator and Select Panel(Typical)
CONTROLIINDICATOR
2.

FUELQTYSEL
Knob - continued

POSITIONFUNCTION

RSVR

AUFR pointers display fuel in aft/forward reservoir tanks

INTWING

Al1FR pointers display fuel in left/right internal wing tanks

EXT WING

Al1FR pointers display fuel in left/right external wing tanks

EXTCTR

AL pointer drops to zero
FR pointer displays fuel in centerline tank3.

EXT FUEL
TRANS Switch

NORM

Centerline tank transfers first and then external wing tanks

WING FIRST

External wing tanks transfer first and then centerline tank
Figure 1-23.

00 Normally, the forward tank fuel quantity is 0-600
pounds greater than the aft tank quantity. [IDNormally, the aft tank fuel quantity is 700-1350
pounds greater than the forward fuel quantity. If
these values are exceeded in either direction, a red
portion of the AL pointer becomes visible. Fueldistribution can be changed by rotating the ENG
FEED knob to the FWD or AFT position until theimbalance is corrected.
RESERVOIR FUEL LEVEL SENSING SYSTEM

Fuel level sensors in the reservoir tanks are used toturn on/off the air ejectors and the fuel low caution
lights. When a reservoir tank is not full, the air ejector
in that tank is operating. The reservoir tank sensors,
associated sensor circuitry, and fuel level sensing unit
operate independently of the fuel quantity indicating
system.
Fuel Low Caution Lights

The fuel low caution lights, located on the caution
light panel, indicate either a low fuel quantity in thereservoir tanks or a reservoir fuel level sensing
system malfunction. The lights function independently of the fuel quantity indicating system. The
FWD FUEL LOW caution light illuminates when fuelquantity in the forward reservoir drops below 400 ([ID250) pounds. The AFT FUEL LOW caution lightilluminates when aft reservoir fuel quantity drops
1-78

Change 13(Sheet 2)

below 250 ([ID 400) pounds. The caution lights are
powered by the essential dc bus No.2.
HUD FUEL LOW/BINGO INDICATION

In addition to the fuel low caution lights, a fuel low
condition may be indicated by the word FUEL in the
HUD in conjunction with the home mode of the FCC
or the previously entered bingo fuel value.
With the FUEL QTY SEL switch in NORM, the bingo
computation is based on the lesser of fuselage fuelweight or total fuel weight. That is, with the FUELQTY SEL switch in NORM, bingo fuel warning will be
triggered when either fuselage fuel or total fueldecreases below the bingo fuel value.
With the FUEL QTY SEL switch out of NORM, thewarning will only be triggered when total fueldecreases below the bingo value. With trapped
external fuel, this could lead to fuel starvation before
the bingo warning is triggered.
In addition to the HUD FUEL low indication, theVMS provides a BINGO-BINGO message in theheadset when the entered bingo fuel value isreached.
For a more detailed description of the home modeand the bingo fuel option, refer to T. O.
IF-16A-34-1-1, ~ T.O. IF-16A-34-1-3 or lM!I T.O.
IF -16A-34-1-4.

I

T. O. 1F- 16A- 1•

HUD TRAP FUEL WARNING LESS ~A false TRAP FUEL warning may occur after thefollowing:

A trapped external fuel condition is indicated byflashing TRAP FUEL in the HUD. Three
conditions must be met for a TRAP FUEL warning
to occur.

• A fuel leak which exceeds the transfer rate of theexternal tank(s).

Conditions are:

• Receiving a partial fuel load during air refuelingwith an external tank(s).

• FUEL QTY SEL knob is in NORM.
• Fuselage fuel is 500 pounds less than fuselage
capacity.

The TRAP FUEL warning clears automatically after
the condition is corrected; the warning may be
manually reset by placing the DRIFT C/O switch to
TEST.

• Total fuel is 500 pounds greater than fuselage
fuel.

For a more detailed description of the HUD TRAP
FUEL warning, refer to T.O. IF-16A-34-1-1 or ~T.O. IF-16A-34-1-3.

• Prolonged AB use if fuel flow to the engine exceeds
the transfer rate from the external tank(s).

Change 12

1-78.1/(1-78.2 blank)

T. O. 1F- 16A- 1
FUEL HOT CAUTION LIGHT

The FUEL HOT caution light, located on the caution
light panel, illuminates when the temperature offuelto the engine becomes excessive.

When a partial fuel load is required, fuel distribution
should be corrected prior to flight by selective
operation of the fuel transfer pumps controlled by the
ENG FEED knob.
Air Refueling (AR) System

FUEL TANK EXPLOSION SUPPRESSION SYSTEM

The fuel tank explosion suppression system places
the fuel tank vent and pressurization system on a
reduced pressure schedule and inerts the fuel vapors
inside the tanks (if serviced with Halon). The system,
intended for use only in combat or during
emergencies, is controlled by the TANK INERTING
switch on the fuel control panel. The system uses
Halon as an inerting agent which prevents
combustion when mixed with air. For the agent
specification and reservoir location, refer to SERVICING DIAGRAM, this section. The Halon
reservoir has a heater, controlled by a thermostatic
switch, which assures sufficient operating pressure.
The RMLG WOW switch prevents operation of theheater while the aircraft is on the ground.
When the TANK INERTING switch is placed to
TANK INERTING, the fuselage and internal wingtanks are placed on a reduced pressure schedule anda valve at the Halon reservoir is opened. At eachactivation of the TANK INERTING switch, Halon (ifavailable) is released into the F-l, A-I, and internal
wing tanks for 20 seconds for initial inerting.
Thereafter, a continuous metered flow of Halon ismixed with the pressurization air to maintain theinert condition. The metered flow continues until thesystem is turned off or until the MAIN PWR switch is
positioned to OFF.
Because of limited Halon supply, the system shouldbe activated after the external tanks have emptied,but before half of the internal fuel is depleted. Since
the 20 seconds of initial inerting occurs each time the
TANK INERTING switch is placed to TANKINERTING, do not cycle the switch. The fuel tank
explosion suppression system does not protect theexternal fuel tanks.

The AR system consists of a hydraulically actuatedreceptacle and slipway door, a signal amplifier, andthe associated controls and indicators. Hydraulicsystem B provides pressure for operation of the door
and latch mechanism. The receptacle is located on the
top fuselage centerline aft of the canopy. When theslipway door is opened, a mechanical linkage retracts
the aft end of the slipway door into the fuselage,
forming a slipway into the receptacle.
When the AIR REFUEL switch is placed to OPEN,
the external tanks are depressurized, external fueldoes not transfer, and the FLCS is placed in takeoff
and landing gains.
When closed, the slipway door is flush with thefuselage skin. The AR receptacle is equipped withfour lights, two located on each side. An AR floodlight
is located on the top fuselage centerline immediately
aft of the canopy. A light on the upper leading edge ofthe vertical tail floods the AR receptacle area and theupper fuselage.
During AR operations, the AR boom enters thereceptacle and is automatically latched in place by ahydraulic actuating mechanism. The HOT MIC
switch allows intercom communications with compatible tankers through the AR boom. When the lastrefuel shutoff valve closes, a pressure switch
automatically provides a signal to unlatch the boom
from the receptacle. A disconnect signal can be
manually initiated at anytime during AR by thereceiver or by the tanker boom operator.

lID Disconnect from the boom may occur before all

Ground Refuelingtanks are full if the external fuel tank configurationconsists of only a centerline fuel tank. Such a
disconnect typically occurs when refueling with an
initial internal fuel load of 4000 pounds or more andthe centerline tank empty. At disconnect, the aircraft
total fuel may be up to 1600 pounds less than full,with many occurrences resulting in approximately
1000 pounds less than full.

All external and internal fuel tanks can be pressure
filled from a single-point ground refueling receptacle
located on the lower left side of the fuselage justforward of the wing trailing edge. Electrical power is
not required to refuel the aircraft unless fuel quantity
is to be monitored. Terminating refueling withpartially filled tanks could result in fuel imbalance.

Fuel venting from under the left wing can occur
during AR, particularly when the aircraft isconfigured with external tank(s). Terminating the AR
operation in a partially filled condition could result infuel imbalance. When a partial fuel load is required,fuel distribution should be monitored and corrected
as required by use of the ENG FEED knob.

REFUELING SYSTEM

Change 11

1-79

T. O. 1F- 16A- 1
NWS AIR DISC MSL STEP BUnONThe NWS AIR DISC MSL STEP button is located on
the outboard side of the stick. The AIR DISC functionof the switch is activated when the aircraft is airborne
and the AIR REFUEL switch is positioned to OPEN.
The button provides a means of manually disconnecting the AR boom. Depressing the switch causes theboom latching mechanism to unlatch and release theboom.
AIR REFUELING (AR) STATUS INDICATOR

The AR status indicator, located to the right of the
HUD, contains three lights.

There is no interruption of bleed air flow due to a lossof electrical power. All pressurization functions (servo
PBG, and fuel tank •air, canopy seal, g-suit,pressurization) operate normally since the bleed air
function continues to operate.

m

AIR-CONDITIONING
Engine bleed air is directed through a turbine
compressor and air-to-air heat exchangers where it iscooled by ram air. The conditioned air is then used forthe functions shown in figure 1-24.

• ARINWS - Illuminates green when the boom islatched in place.

A cockpit temperature controller receives signals
from temperature sensors and from a manuallyoperated control panel to automatically control the
cockpit temperature. Conditioned air enters the
cockpit on both sides, the top rear of the seat, throughthe angle vent on the instrument panel, and throughthe canopy defogger. In the event of an ECS
malfunction, emergency ram air operation can beselected for ventilation and cooling.

• DISC - Illuminates amber when a disconnect
occurs. After the disconnect, the system automatically recycles to ready, and the RDY lightilluminates after a 3-second delay.

A ground cooling cart can be connected to the ground
cooling receptacle on the lower left side of the fuselage
just above the nosewheel area to provide cooling air
to the cockpit and avionic equipment.

A lever for dimming the three lights is located on theright side of the unit.

PRESSURIZATION

Functions are:• RDY - Illuminates blue when the AR slipway door
is open and the system is ready.

ENVIRONMENTAL CONTROL SYSTEM(ECS)Refer to figure 1-24. The ECS combines air-conditioning and pressurization functions to provide
temperature-controlled, pressure-regulated air forheating, cooling, ventilating, canopy defogging,cockpit pressurization, canopy sealing, g-suitpressurization, fuel tank pressurization, electronicequipment cooling,
and pressure breathing for g(PBG). Most of these functions are lost when the AIR
SOURCE knob is placed to OFF or RAM. Refer to
AIR SOURCE KNOB, this section.

mRefer to figure 1-25 for the automatically controlled
cockpit pressure schedule. Air pressure is provided by
the pressurization system for controVoperation of
PBG, fuelsome of the ECS, canopy seal, g-suit,tanks, and radar. A cockpit pressure safety valve
relieves pressure anytime the cockpit pressure
exceeds ambient pressure by 5.4 psi.

m

The canopy seal is inflated/deflated with themechanical locking/unlocking of the canopy.

AIR SOURCE Knob ~ rnrl
Refer to figure 1-26. The AIR SOURCE knob IS
located on the ECS panel.

ELECTRICAL FAILURES

Functions are:

When the ECS loses electrical power, cabintemperature control and water separator anti-icing
become inoperative. The cockpit receives cold air only.
The water separator coalescer sock freezes ataltitudes where there is enough moisture in the air(below 30,000 feet) and the built-in bypass valve
opens allowing air to flow to the cockpit and avionics.

• OFF - Engine bleed aIr valves close. All
air-conditioning, cooling, and pressurizing functions shut off, including g-suit, m PBG, canopy seal,
and fuel tank pressurization.

1-80• NORM - Air-conditioning system set for automatic
temperature and pressure regulation.

IT.O. 1F-16A-l

Environmental Control System (Typical)
Block 10CAUTION LIGHT

COCKPITMECHANICALLY OPERATED
FlAPPER VALv[- - - - - - - - - - - - - --I

ANTI-G VALV£_ _. . . AND CANOPY
SEAl.

,

,""''''''..:,> SERVO AIR'\o""''''''~ FUEL TANKPRESSURIZATION

lolli""'liE./ __ ,-.. E.lCTOR''''~::::;ENGINECOMPRESSORBlEED AIRLEGEND:®

SOLENOID VALVE@MODULATING VAlVE®

PRESSURE REGULATOR

l OW

HIGH

PRESSURE

PRESSURE- - - MECHANICAL

ws WATER SEPARATOR- - ELECTRICALTIC TURBINE/COMPRESSOR~ CHECK VALVE

H.

9 TEMPtpRESS SENSOR~

AIR -lO-AIR
HEAT EXCHANGERSENSOR/CONTROLLER
IF_16A_l _1059X.

Figure 1-24. (Sh«et 1)

1-81T.0 . 1F-16A_l

Environmental Control System (Typical)
Block JSCAUTION LIGHT

COCKPITMECHANICALLY OPERAltD
FlAPPEII VAlVE,- ~ -----------I

I

ANTI-G VALVE_ _. . AND CANOPY
SEAl,,=-=-~ ""VO AIR~<::iII<:>-t> FUEL TANKPRESSURIZATION

NACEllE

EPIJ

ENG""

COMPRESSOR
SUEO AIR

lOW

HIGH

PRESSURf

PRESStJI:E

GROUNOCOOLING
RECEPTACLE

TOTAL
TEMP
FROM

n PROBELEGEND:_BLEED AIR

~ WARMAIR
COOL AIRRAM AIRoGo- - - MECHANICAL- - ELECTRICALSOLENOID VALVEMODULATING VAlVE
PRESSURE REGULATOR~CHECK VALVE

Ii

TEMP/PRESS SENSOR~SENSOR/CONTROLLER

WS WATER SEPARATORTIC TURBINE/COMPRESSOR

H,

AIR-TO-AIR
HEAT EXCHANGER~ELECTRONIC FLOWCONTROL SET
IF_16A _l _l060X.

Figure 1·24. (Shat 2)

1-82

T. O. 1F- 16A- 1

Cockpit Pressure Schedule60...... ~
l/~

,
t-

w
w

50

...... ~....... ~000... 40......
......

w

Ct-~~\'-~"\' "30......

20

,

~

0

0
0

"

10

~

'"

.......

'"....... ~.......'"
5 PSI DIFFERENTIAL PRESSURE"....... ~......

~

ii:.......

......

5t-~...... ~u.

;:).......

>- -

SL

5

.......

'"

.....

1/

_10-:' ~.'.-'!"'~ .... ,:;. ~-.......

~~ ~.'"" "" ... ::,:... ..... :"": "=': .'.
~.'j:;a: ~-~""1" •...... ".

- ...... .
"."

", ""
iOo'"

~foI""t-~~-..-:,:;.~......

'"\)~I?~~A- r- CONSTANT PRESSURE RANGE

-~~II
II
SL ~
I

"

....... ~'"\1..'-\) 1\\

c::r:;·.J~~......

I
I

I I

I

I

I

10

15

20

~F"

I I
I I

25

30

35

40

AIRCRAFT ALTITUDE - 1000 FEET

45

50

55

60

1F-1eX-1-0001X@

Figure 1-25.

- DUMP - Cockpit pressure dump valve opens toatmospheric pressure. Cockpit pressure altitude
increases if DUMP is selected above approximately
8000 feet MSL. Conditioned air ventilates cockpit
and performs all other system functions.

I-

RAM - Engine bleed air valves close and the
cockpit pressure dump valve opens to atmosphericpressure. Cockpit pressure altitude increases if
RAM is selected above approximately 8000 feet
MSL. All air-conditioning, cooling, and pressurizing functions shut off, including g-suit, m PBG,
canopy seal, and fuel tank pressurization. The ram
air valve opens to admit ram air to ventilate the
cockpit and avionic equipment.

Functions are:
- AUTO - Cockpit temperature is automaticallymaintained (60°-80°F) relative to the setting of theknob.
- MAN - The temperature control drives the air
modulating valve to a set position. Cockpit
temperature varies according to throttle setting,
OAT, and cockpit heat load. If WARM is selected,
the cockpit supply air temperature may exceed the
maximum allowable limit of approximately 177°F.
This causes the warm air valve to cycle on and off.
This is a normal occurrence and can be stopped by
selecting a cooler setting.
- TEMP OFF - Hot air mixing is shut off. Only air at
approximately 35°F is delivered to cockpit.

TEMP Knob ~ mEl

Under extreme temperature conditions, system
performance on the ground can be improved byadvancing the throttle 1-3 percent above idle rpm.

Refer to figure 1-26. The TEMP knob, located on the
ECS panel, only controls cockpit temperature.

Ground operation with the radar in OFF improves
cockpit cooling and ground operation with the radarin STBY improves cockpit heating.

Change 3

1-83

T.O. 1 F-16A-l

Environmental Control System Control
Panel [8J IBF I (Typical)

Defog Control Panel[A] IBF I (Typical)
L

1I
1. DEFOG Lever

1. TEMP Knob
2. AIR SOURCE Knob
, F-'6A-'~026A.

'F-'6A-1~025A.

Figure 1-26.

DEFOG Lever ~ [lEI

Refer to figure 1-27. The DEFOG lever, located on thefar aft portion of the left console, mechanicallycontrols a flapper valve in the cockpit air supply line.
Functions are:• MIN - Minimum airflow toward the canopy forward
area and air vent in center pedestal; maximum
airflow to outlets behind seat.
• MAX - Most of the cockpit air supply is diverted tothe canopy forward area for defogging and to the air
vent in the center pedestal. A partial opening of thecen ter pedestal air outlet allows a more balanced
defogging of the right and left sides of the canopy.
When placed in the full forward defog (MAX)position and with the TEMP knob in AUTO, thelever activates a switch which shifts the cockpit air
supply control to full warm. The full warm air
supply automatically terminates 3 minutes after
activation. The lever may be cycled to restart thefull warm, 3-minute period.

1-84

Figure 1-27.

Under extremely humid conditions or after initialengine start, fog may form at the cockpit air outlets as
the cold air mixes with the moist cockpit atmosphere.
This condition can be eliminated by selecting MAN
and moving the TEMP knob toward WARM until thefog stops forming. In flight, while operating in AUTO,
the most rapid method of eliminating air outlet
fogging is by selecting the MAX position with the
DEFOG lever. Fog may form on the interior surface ofthe canopy as a result of moisture in the cockpit air
condensing on the cold surface. 1b warm the canopy
surface above the dewpoint and permit the cockpit air
to retain more moisture during cold weather
operation, the DEFOG lever should be placed in a
forward position and the TEMP knob positioned to
MAN WARM.
EQUIP HOT Caution Light

The EQUIP HOT caution light, located on the caution
light panel, illuminates when the avionic equipment
cooling air temperature/pressure is insufficient.

T.O. 1F- 16A-1

Degraded equipment performance and/or damage
can result from overheating. Therefore, when the
EQUIP HOT caution light illuminates, the electronicequipment should be turned off unless it is essential
for flight. Illumination of the EQUIP HOT caution
light automatically interrupts electrical power to theradar. Turning the radar to OFF in flight does not
close the radar cooling air shutoff valve.
A short duration or intermittent EQUIP HOT
caution light may occur when ground cooling air isdisconnected.
Cockpit Pressure Altimeter ~ mEl

The cockpit pressure altimeter, located on the right
auxiliary console outboard of the stick, is labeled
CABIN PRESS ALT.
CABIN PRESS Caution LightI

MAIN AC POWER SYSTEM

AC power is normally supplied by a 40 kva main
generator located on and driven by the ADG. The
main generator supplies power to the overcurrent
protection panels and LESS ~ ECM power panel andnonessential and essential ac buses. The main
generator contains a permanent magnet generator(PMG) which provides one source of start power forthe emergency power unit (EPU) if the main
generator fails but is still rotating.
Overcurrent Protection Units

The ~ eight, LESS ~ five overcurrent protection
units, located on ~ overcurrent protection panels
No.1 and No. 2, LESS~ overcurrent protection panel
and ECM power panel, protect certain ac buses;
stations 3, 5, and 7; ~ and inlet stations fromovercurrent. The ELEC CAUTION RESET buttonon the ELEC control panel is used to reset a tripped
overcurrent protection unit for the nonessential ac
bus No. 1 ~ and the nacelle nonessential ac bus. The
unit may not remain reset if the fault persists.

The CABIN PRESS caution light, located on thecaution light panel, illuminates when the cockpit
pressure altitude is above W 22,500, LESS 00 27,000
feet.

The items with nonresettable overcurrent protection
units are the radar ac bus; stations 3, 5, and 7; and left
and right inlet stations.

ANTI-G SYSTEMEMERGENCY AC POWER SYSTEM

The anti-g system includes the ANTI-G panel/valve,
the g-suit, m and PBG equipment.
The g-suit connector and TEST button are located on
the ANTI-G panel at the aft end of the left console.
The ECS delivers cooled bleed air to the g-suit andto the oxygen regulator as a control pressure for
PBG. Airflow is proportional to the positive g forces
sensed. If an ECS shutdown occurs, g-suit m and
PBG protection are not available.

m

If the main generator fails, emergency ac power is

supplied automatically by a 5 kva EPU generator
driven by the EPU. The system supplies power to theessential ac buses. The overcurrent protection units
are not functional if the EPU generator is supplying
power. The EPU generator has a PMG which
supplies de power through an ac to dc converter tothe four FLCS branches. Refer to EMERGENCY
POWER UNIT, this section, for further discussion of
the EPU.
DC POWER SYSTEM ~The system can be manually tested by depressing theanti-g TEST button to inflate the g-suit m and tocheck the PBG function. The system incorporates anautomatic pressure relief valve.

ELECTRICAL SYSTEM
Refer to figure 1-28. The electrical system consists ofa main ac power system, an emergency ac powersystem, a de power system, an FLCS power supply,and provisions for external ac power.

DC power is supplied to the essential dc buses by ac
to dc converters. With the main generator operating,one of these buses also powers the nonessential dc bus
and nacelle dc bus. When the essential dc buses are
powered, they power the battery buses; then theaircraft battery is disconnected and charged by abattery charger which receives power from essential
dc bus No. 1. With the main generator failed and the
EPU generator operating, only the essential de buses
and battery buses are powered. If the main and EPU
generators both fail, none of the dc buses is powered
and the aircraft battery supplies power to the battery
buses.

Change 13

1-85

T.O. 1 F-16A-1

DC POWER SYSTEM LESS [BQIDC power is supplied to the essential dc buses by ac
to dc converters. With the main generator operating,one of these buses also powers the nonessential dc bus~ and nacelle dc bus. When the essential dc buses are
powered, they power the battery bus; then theaircraft battery is disconnected and charged by abattery charger which receives power from essential
dc bus No. 1. With the main generator failed and the
EPU generator operating, only the essential dc buses
and battery buses are powered. If the main and EPU
generators both fail, none of the dc buses is powered
and the aircraft battery supplies power to the battery
bus.
FLCS POWER SUPPLYThe primary FLCS power supply includes a dedicated
FLCS PMG, two dual-channel converter/regulators,
and four inverters.
Other FLCS power sources are the main generator,
the EPU generator, the EPU PMG, the aircraft
battery, and the FLCS batteries.
The FLCS PMG is the primary power source for the
FLCS during normal operations. The FLCS PMG is
located on the ADG and shares the hydraulic system
A pump shaft and generates power whenever the
ADG is rotating. The PMG has four outputs, one foreach branch of the FLCS, and generates sufficient
power to operate the FLCS at 40 percent engine rpm
or greater.
~ Two converter/regulators, having two channels

each, provide a separate channel for each inverter.
Both converter/regulators receive power from theFLCS PMG and, if the EPU is running, the EPU
PMG. The branch A and B converter/regulator also
receives power from essential dc bus No. 1 and1-86

battery bus No. 1 and the branch C and D
converter/regulator also receives power from essential dc bus No. 1 and battery bus No.2. Eachconverter/regulator channel converts ac power fromthe FLCS PMG to dc, selects the power source withthe highest voltage (within limits), and provides dc
power to the respective inverter. Converter/regulator
output voltages are regulated to prevent overvoltage
to the FLCS inverters.
LESS ~ Two converter/regulators, having twochannels each, provide a separate channel for each
inverter. The converter/regulators normally receive
power from the FLCS PMG, dc essential bus No.1,
and the battery bus. If the EPU is running, the EPU
PMG also provides power to the converter/regulators.
Each converter/regulator channel converts ac powerfrom the FLCS PMG to dc, selects the power sourcewith the highest voltage (within limits), and provides
dc power to the respective inverter. Converter/regulator output voltages are regulated to prevent
overvoltage to the FLCS inverters.

Each inverter then changes dc to ac power for the
FLCS. If the converter/regulator output voltage isless than the FLCS battery voltage, the FLCS battery
powers the FLCS branch until balance is reached or
the battery is depleted. The converter/regulators also
provide fault indications for display on the ELEC
control panel and provide test indications to the
TEST switch panel.
The primary function of the FLCS batteries is toprovide temporary emergency power to the FLCS; thebatteries are not intended to be a continuous
emergency power source. The FLCS batteries arecontinually charged by power from the converter/regulators when the MAIN PWR switch is in BATT or
MAIN PWR. When the MAIN PWR switch is OFF, a
trickle charge circuit maintains the FLCS battery
charge with power from the aircraft battery.

10.lF-16A-l

Electrical Power Distribution Diagram
Block to~

MAIN P\NR CONTACTORLEGEND:
ELECTRICAL!

I- - - - - MECHANICALRESETTABLP
00

MONITOR UNIT

I

EXTERNAL PWA =CABLE RECEPTACLENONRESrnABLE·

I - ~~ c

,

n n n

1

ADG

I

~ Cen be powered by EPU-

uPROT!ANEL

with EPU Running
IIIIlIIID Powered
and Main Goo on Una
Powered thru NONESS AC
BUI No. 2.~

thru Overcurrent
CJ Powered
Prot Panel0~Powered thru ECM Power

Panel

'PU
ON

A
F

,

I~
.

'm~

~ I AC:BU; NO.1

IDCBJI :

H

T P
S W
T R
R P
A N
K L

G'N

I R
P
P
B N
A L

R
I A
G C

I B~~;iJ ,•Le; f=': m c-

;,A]

~-

I BUS

;R - JW··OVERCURRENT PROTECTION UNIT

T 0
C

IRAOAR AC BUsl

F~LI
1 0CBU

~

:

V

I

I

I ~'~s~BOX BUS IIYBAITERY CHARGER! I
CONTROL ASSEMBLV

R
I 0
G C
H

T P
S W
T R
R P
A N

:

r ~

B~S :] ,K L

BRANCH'S{
B. C,

FlCS PMG

~ '-

-

-- --

~

t::

'--+l 'PU

AND 0

BRANCH A
TYPICAL

I CONY~.r.'" .Tno

: BAIT Lr-

L...... FLCSCOMPUTER

1

rl

PMG

I

'PU
'--+lAC TO DC

Figure 1·28. (Sheet 1)

Change 2

1·87

T.O. 1F·16A·l

Electrical Power Distribution Diagram
Block 1.5

Powered with EPU Running
IiITIliIll andMain Gen on LinePowered 1hru NONESS AC

BUI No. 2

ttlro Overcurrent
c:::J Powered
Prot Penal No. 1
Powered lhru Overcurrent

Prot PeMI No. 2·OVERCURRfNT PAOTECT10N UNIT

F

T 0
C

,o

P
U W
, R
p
P
0 N
L

•

111 1 DC:-----1

0

NONESS AC
BUS NO. 1

I~

-

I~
Iii

EPU
ON

I

t,J

I~ I[.c·ou ;NO ]
Ii:

~
,

j

I DCESSENTIALI
BUS NO. 1 r

I• ~ .c'~

,
,

0 P
U N
L
P

~-

R

~

-m·

,

A
G c
H
T P
W
5 R
T
R P
A N
K L

~r~~ ,,,,,,
r

CHARGER!

ASSEMBLY'W~io
ORANCH A {

,

R

:

I~Oo':: ~~ 21 G C0
H
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W
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~t~

11111:oc,.~t.

71j'BOX OUS IJtI AC-etrssi

, 6

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I

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I

0
G
H Y
T
A
N C
A /
C 0
C
L
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R

PRO~-"~N':-N(

,,
,,
,,
,,
,,

TO.

m~

I

!,,

W·

I

,

,,
,,

I

ADO

,,

I:. ~

ou~lllll-

? W£f

y

PWR

L;.;; ,~ Can be powered by EPU•.rJ 1m .

f- - ELECTRiCAl----- MECHANICALRESETIABLP
0
NONReSETIABlE·
I2l-

MAIN PWR CONTACTOR, UNI1LEGEND:-- - -

FlCS PMG

---

:

,

-

,PU

,,i
,,,
,, Y

~us ' al ,f...--lLO~T

I C( NVTYPICAL

L- FlCS COMPUTER,I-

1::PMG'PU

I

L,.fAC TO DC
IF·l'lA-l_108''''.

FigUT'f! 1·28.

, -88

Change 2(Sheet 2)

IO. l F·16A-l

Electrical Power Distribution Diagram
Air Defense Fighter- - ELECTRICAl
MECHANICALMONITOR UNITI-- --00MAIN PNR CONTACTORILEGEND:RESETTABLPt
EXTERNAl f'INR .E

CABLE RECEPTACLE-

Powered thru NONESS AC

Bus No.2

c::::::J Prot Panel No. 1

Powered thru Overcurr 001

Powered ttlru Overeurr onl

Prot Panel No. 2· OVERCURRENT PROTECTION U NITI~,~NU ~ J

I~

-

W ·

I"
I~
I L IL'C-BU , ."'o~, lAF

rn I. ""

p

p
N
B
A L
Y

;

BU'

~

~~Aii.

~~I

ESSENTIAl
DC BUS NO. 1

1

,1
I

1
R
I B
G A
H Y
T
A
N C
A /
C a
C

,

,

1

I

H

~Ai~
~t ~
'W
I

AC

Bu~SSI

AND 0

::t

1

,,
,,
,
!
,,

BRANCH B

B. c,

FlCS PMG ~

--

!,

1---

P

PROT P~NE~ N ~ 1

I

I

I

",,6&71~-11111oC~

BRANCH A
TYPICAL

R
I A
G C

>~

H

T

K

~

LBAT' II CONV\I

L - FLCS COMPUTER

,1---

L

R
I a
G C

,

:

P

W
S R
T
R p
A N

'~ ,

_OA~ '~~< ;~

H

T

P

s WR

T
R p
A N

,
K

L

" 'A

~~

'PU

PMG

a P
U N
I L

I
OVERCURRENT;BATTERYCONTROL

1

P

, W
R

' BOX BUS

:,
,,
,
,,,
,,,

I

ADO

- M~

:TO
a
C

a p
UW
I R

L
L

,,
,,
,,
,,
:

I

NONESS AC
BUS NO.1

,

M ·

,,

I;0 I ~
1

~IJ;-------1

NONRESETTABLE-IWZJ Can be powered by EPU
with EPU Run nlng
IIIIIIIIIl Powered
and Main Gan on Une

T

I

-M

'PU

I
I

~ACTQOC

lr·15A-l-,oeu..
Figure 1·28. (Shftt 3)

Change 2

1-89

T. O. 1F- 16A- 1
EXTERNAL POWER PROVISIONSThe external power provisions include a standard
external power cable receptacle and a monitor unit
which are part of the airframe. The monitor unit
allows external power to be connected to the aircraft
buses if the phasing, voltage, and frequency of theexternal power are correct. When connected, theexternal power provides the same power as the main
generator.
ELECTRICAL SYSTEM CONTROLS AND INDICATORS

Refer to figure 1-29.
ELECTRICAL POWER DISTRIBUTION

held in TEST, the ACFT BATT TO FLCS and FLCS
PMG lights go off. The FLCS BATT lights come on andthe FLCS PWR lights remain on to indicate that the
FLCS batteries are good.
With the FLCS PWR TEST switch in NORM, when
the MAIN PWR switch is moved to MAIN PWR forstart on battery power, the lights do not change. If
external power is used, the MAIN GEN, ACFT BATTTO FLCS, and the FLCS PWR lights go off when the
MAIN PWR switch is moved to MAIN PWR.
During engine start, the FLCS PMG light goes off
before the engine reaches idle and the main
generator comes on line at approximately 45 percent
rpm. External power, if used, is disconnected fromthe aircraft buses when the main generator comes online.

Refer to figures 1-30 and 1-31.
NORMAL OPERATION

Prior to engine start, the MAIN PWR switch is placed
to BATT to permit a check of the aircraft and FLCS
batteries. The ELEC SYS, FLCS PMG, and MAIN
GEN lights come on. In addition, the ACFT BATT TO
FLCS light comes on to indicate that the aircraft
battery is powering the FLCS, and the FLCS PWRlights come on to indicate that the FLCS inverter
outputs are good. With the FLCS PWR TEST switch

1-90

Anytime after selecting MAIN PWR, including inflight, the FLCS PWR TEST switch may be held
momentarily in TEST to check inverter output.
During the EPU test, the FLCS PWR lights come on
to indicate that EPU PMG power is available to the
FLCS.
During engine shutdown, the ELEC SYS caution
light and FLCS PMG and MAIN GEN lights come on
as the engine spools down. The ACFT BATT TO
FLCS light may also illuminate.

10.lF-16A-l

Electrical System Controls and Indicators~ IBFI (Typical)
23

1. MAIN PWR Switch
2. FLCS PMG Indicator Light
3. MAIN GEN Indicator Light
4. STBY GEN Indicator Light
5. EPU GEN Indicator Light
6. EPU PMG Indicator Light
7. FLCS BAIT Indicator Lights
8. ACFT BATT Indicator Lights
9. ELEC SYS Caution Light
10. ELEC CAUTION RESET Button
11. FLCS PWR TEST Switch
12. FLCS PWR Indicator Lights

4

5
6

12

11

8

7

1F-16A-1-1067X@

1.

CONTROUINDICATORPOSITIONIINDICATION

MAIN PWR Switch

MAINPWR

Connects external power or the main generator to the electrical system and detenninesfunction of FLCS PWR TEST switch. If ac
power is not available, connects aircraft battery to the battery bus

BATT

Connects aircraft battery to the battery bus,
disconnects main generator or external
power, determines function of FLCS PWRTEST switch, and is used to reset the main
generator

OFF

In flight - inoperative

NOTE:

During ground operation, if
the MAIN PWR switch is
moved from MAIN PWR to
OFF without a delay of one
second in BAIT, the EPU
does not activate and electrical power for braking,
NWS, hook, and radios islost.

FUNCTION

On ground - disconnects all power from aircraft electrical system. Canopy operation and
FLCS battery trickle charge are available

Figure 1-29. (Sheet 1)

1-91

T. O. 1F- 16A- 1

Electrical System Controls and Indicators[8] [1!E] (Typical)
2.

CONTROUINDICATORI POSITIONIINDICATIONFLCS PMG Indicator Light

FLCS PMG (amber)

FUNCTION
In flight- None of the FLCS branches arereceiving power from the FLCS PMG
On ground - FLCS PMG power is not available at one or more FLCS branches. Light isdelayed 60 seconds after initial NLG WOW

3.

MAIN GEN Indicator Light

MAIN GEN (amber)

Indicates external power or main generator
not connected to the electrical system

4.

STBY GEN Indicator LightSTBY GEN (amber)

Not operational

5.

EPU GEN Indicator LightEPU GEN (amber)

Indicates the EPU has been commanded on
but the EPU generator is not providing power
to both essential ac buses. The light does not
function with the EPU switch in OFF (WOW)and the engine running6.

EPU PMG Indicator LightEPU PMG (amber)

Indicates the EPU has been commanded on
but EPU PMG power is not available to allbranches of the FLCS

7.

FLCS BATT Indicator
LightsA, B, C, and D (amber)

Indicates the respective FLCS battery discharging to the indicated FLCS branch

8.

ACFT BATT Indicator
LightsTO FLCS (amber)

In flight - indicates battery bus power isgoing to one or more FLCS branches and voltage is 25V or lessOn ground - indicates battery bus power isgoing to one or more FLCS branches

FAIL (amber)

In flight - indicates aircraft battery failure(20V or less)On ground - indicates aircraft battery failure or cell voltage imbalance condition. Lightis delayed 60 seconds after MLG WOW

9.

ELEC SYS Caution LightELEC SYS (amber)

Figure 1-29. (Sheet 2)

1-92

Illuminates in conjunction with any of theabove lights. Also illuminates if one or more
FLCS batteries are not connected when themain generator is on line

T. O. 1F- 16A- 1

Electrical System Controls and Indicators[8J [IE] (Typical)
CONTROLIINDICATORPOSITIONIINDICATION

10.

ELEC CAUTION RESET
ButtonPushResets resettable overcurrent protection
units and ELEC SYS caution light andclears MASTER CAUTION light for futureindications

11.

FLCS PWR TEST Switch

TEST

When MAIN PWR switch is in:

FUNCTION

MAINPWR

BATT

Tests FLCS inverter Tests FLCS inverter
outputoutput on FLCB batteries on ground

12.

FLCS PWR Indicator Lights

NORM

Nonnal position.
Tests EPU PMG
power availability
during EPU/GEN
test on ground

Tests FLCS inverter
output on aircraft
battery

MAINT

For maintenance use on the ground. Inoperative in flight

A, B, C, and D (green)

Illuminate to indicate proper FLCB inverter
output during FLCB power tests

Figure 1-29. (Sheet 3)

Change 2

1-93/{1-94 blank)

T.O.1F-16A-l

AC Power Distribution Diagram
Block 10MAIN GENERATORQVERCURRENT PROT PANELI

SUI 3, 5, 7 - ECM, EO &.
Redar-Guided Weapons

FCR

FCC
INS Heater

Aircraft Banary HeatBr
Flight loads Recorder

Halon Haeter
Seat Adjust

Flood Console Ughts
Flood Instrumant lighll

Stations 4 &. 6 - EO &.

Interference Blanker
REO
TT Probe Hester

Formation Ughts
Fuel Boost &. Transfer
Pumps

EPU GENERATOR

Radar-Guided Weapons
Taxi Ught
TlSL

TWS

'PU

'PU

ON

-w

Air Dale Probe Heaters
Inose &. fusel.ge l
Altimeter (ELEen
ADA Indicalor
ADA Probe Heater.

ADI
AIM-9

Anticolli.ion Strobe LightAR Ught (Flood)
Data Transfer Unit
DC Converter

CAOC

INS
LEF'.
Nav/Fraq Display
Primary Console Lights
Primary Instrument Ughts

DC Converter'CA
FUEL FLOW Indicator
HUD/eTVS

Engine Ice Detector
Fire/Overheat Detect & Test
FUEL Quantity IndicatorG,"
HSI
HYD PRESS Indicators

LEGEND:otzI

ELECTRICALMECHANICALRESETTABLENONRESITTABLE

ON

IFF (mode 41
Inlet Strut HeaterLan~ Ught
UQUID OXYGEN
Quantity Indicator

1All§fJ

NOZ POS Indicator
OIL Pressure Indicator
Position Lights
TACAN~ Can be powered by EPU_

Powered thru NONESS AC Bus No.2

c::::J Powered thru Overcurrent Prot Panel_

Powered Ihru ECM Power Panel
IF·l&A·l·118 .......

Figure 1·30. (Sheet l)

Change 3

1-95

10.lF-16A-l

AC Power Distribution Diagram IADI

Block 15

MAIN GENERATORQVERCURAENT PROT PANEL NO. 1

Oal8 Control

I

Oata Trenlifer
Flight loads RecorderSia 3 , 5 , 7 - EeM, EO &Radar-Guided Weapons

Flood Console Lights
Flood Instrument Lights
Formation Lights
Fuel Boost & Transfer Pumps
Halon Heeler
Inlet Strut Heater

NONESS AC aus NO. 1
GPS
INS Heater
Interference Blanker

REO

S•• t Adjust

TT Probe Heater

Station 4 & 6 - EO &.
Radar-Guided Weapons

EPU

-

ON

-w
-----------EPU GENERATOR

Taxi Light
TISL

TWS

~
A1M-7/A1M-9/AIM- 1 20

EPU

-w

ON

."""""""""""""""",,,,,,1'1111111111111: 1111:. •:Air Data Probe Heater.

CADC

Inose & fuselage}
Ahimetaf (ELECT)DC Converter

LEF's

EC'
FCC

Nav/Freq Displav
Primary Console Lights
Primary Instrument Lights

ADA Indicator
ADA Probe Healers"""""",.""""""

FUEL FLOW Indicator
HUO/CTVS.",1.

."",..

~I
Anticollision Strobe LightAR Light !nood)
DC Converter
Engine Ice Detector

FUEL Quantity IndicatorLan~ Light

G,n
"51

[!]lIEJ LIQUID OXYGEN
Quantity Indicator

OIL Pressure Indicator
Position Lights
TACAN

HYD PRESS Indicators
Can be powered by EPU
Powered thru NONESS AC Bus No. 2
MECHANICALRESEnABLE
NONRESETTABLf

Powered thru OverCUfrent Prot Panel No. ,Powered thru Overcurrent Prot Panel No. 2
1f-1SA-1-1198A.

Figure 1·30. (SMet 2)

1-96

Change 3

T.0.1F-16A-l

AC Power Distribution Diagram LESS M
Block 1.5

MAIN GENERATOROVERCURRENT PROT PANEL NO. 1

NONESS AC BUS NO. 2

Aircraft Banery Heater

Data Control
Flight Loads RecorderSla 3, 5, 7 - ECM, EO &:RadlH-Guided Weapons

Flood Console Lights

Flood Instrument Lights
Formation lights
Fuel Boost &. Transfer Pumps
Halon Heater

NONESS AC BUS NO. 1•

Inlet Strut Heater
FCC
INS Heater
Interference Blanker

REOn

AIM-9/AIM-120

Probe Heater

Seat Adjust
Station 4 & 6 - EO &:Radar-Guided Weapons

Taxi light
nSl

EPU

-

TWS

ON

EPU GENERATOR•

AIM-9/AIM- 120

EPU-w
Air Data Probe Heaters
Inose &: fuselagel
Altimeter (ELEen

INS

CAOC
DC Converter
ECA
FUEL FLOW Indicator
HUD/ClVSADA Indicator
ADA Probe Healers

Anticollision Strobe LightAR Light [flood)
Data Transfer Unit
DC Converter

lEF's
Nav!Freq Display
Primary Console Lights
Primary Instrument Lights

Engine Ice Detector
Fire/Overheat Detect &: Test
FUEL Quantity IndicatorGoo
HSI

HYD PRESS Indicators
LESS ~ IFF (mode 4 )Landing Light[!1[1f] LIQUID OXYGEN
QuantilV Indicator~ Can be powered by EPU- - - ELECTRICALMECHANICAL

c:::::J Powered thru NONESS AC Bus No. 2
CJ Powered Ihru Overcurrent Prot Panel No. 1

RESETTABlE
NQNRESETTABLE_•OIL Pressuro Indicator
Position Lights
TACAN

LEGEND:

o
o

-

ON

Powered thru Overcurrenl Prot Panel No. 2
IF- 16A-1 - 11988.

Figure 1·30. (Sheet 3)

Change 12

1-97

T.O.1F-16A-l

DC Power Distribution Diagram
Block 10IAl AlT AEl Button•

ADA Indexer

Environmental Control Test
FCNP

Stations 1, 4, 5, 7, & 8
Arm Release
Store. Jetti.on (SEL

AR System

FLCP 'aU Hghts, reset, arm,

MASTER ARM Switch
MSl Step Switch
MLG WOW

Autopilot
Benary BUI Feeder
CAOC Caution lightand self-test capabifity

NUCLEAR CONSENT Switch

and EMER)

FlCS Power Source
IFF

NWS

TACAN

eJU

INSTR MODE SelectControlCautionfaulta]'N'~;;<E ANTI ICE Switch

LEF Motor ,1)
LG Handle DOWN Permission
ButtonlG WHEELS Down lights

lim Probe Hoat Monitor
PROBE HEAT Switch
Radar Altimeter

VHF Redio~ VVPN REL Button

SCP (18ft half)

Speedbrakes

MAIN GEN
FAil

o

W :-

Battery Charge PowerENG FEED Knob

Flight load_ Recorder
FUEL HOT Caution light
Fuel Tank lnortlng
Mal & Ind Light T08VDim

FeR

Featureena11IFtara Dispensers,eM

MAX POWER Switch
Necelle Ejector Shutoff

TISl

IF_l$A_l_1'''5-1.

Figure 1·31. (Sheet 1)

1-98

Change 5

T.O.1F-16A-l

Auto Fwd Fuel Tren.
Autopilot

AVTR
Banary Bu. Feeder
CABIN PRESS Caution
Ught
CAMERA/GUN Trigger

EngIne Bleed Air Valve.
{CION capabllityl
IPW220I ENG!NE FAULT
Caution light
EaUIP HOT Caution Ught
FUEL LOW Caution Ught

CIU

G,n
HUD

Cockpit Pressure Oump

ILS

ECA

Altimeter (PNEUI
AMi.kJd/Brake.
ANTl-SKIo Caution Ught
Sanery Ind Power Te..
IPW200I SUC Caution Light·Canopy
CANOPY Warning LightILANOING/TAXV£XT/!QlJfm[AJ 10 Light
Switches
LEF Motor (1 1
LEF at ACe Caution Lights
LG Door Close
LG Hyd Isolation
LG UP- ON Command
MASTER ARM Switch
OXY LOW Caution LightSCP (right halfl
SEAT NOT ARMED
Caution light
Seeur. Voice
Stetlon. 2. 3 . 6 at 9
Ann Release
Stick Trim
STORES CON FIG Ceution Ught
Store. Jenison (SEL at EMERITWS[!J~ WPN REL Bunon

ELEe SYS Caution Ught
~ Engine Auto Lean
~ Engine History
RecorderENGINE Warning Light

EPU
Fire/Overheat Light Bulbs
FLCC (FLT CaNT SVS and
DUAL FC FAIL light.,
FLCS Power Source
FTIT Indicator

FUEL MASTER Switch
Hook at HOOK Caution Light
HYO/OIL PRESS
Werning LightIntercom
JFS
LG Downlock Relays(WHEELS down lights)
LG Warning Light thandle,
MaIn Power Indicator
MAIN PWFI Switch
Ma.. er Caullon Light

MLG WOWNLG WOW
Parking Brake
RPM Indicator
SAl
IpW2201SEC Caution Light
TO/lOG CONFIG
UHF Radio
Utility Ught
VMS

LEGEND:~ Can be powered by EPU- - - ELECTRICAL

I!IJIIID Powered by EPU with tha Main Gan on Une

MECHANICAL_Powered thru NONESS AC Bus No. 2_

Powered thru ECM Power Panel• Power ~ailable from the battery bu. box with MAIN PWFI switch OfF.

Figun 1-31. (Shat 2)

Change 2

1-99

T.0 . 1F- 16A-l

DC Power Distribution Diagram [AD[Block J5AI FF

[KJ AlT REl Button
ANTI ICE Switch
AOA Indeller
AR System
Autopilot
Battery Bus Feeder
88ttery Charge Power
CAOC Caution light
CIU

Controlfaults)~~~~o~'~~~'~1 Control Test
FlCP lall lights, reset, armand self-test capability)
FlCS Power Source
Fuel Tank lnening
INST MODE SelectLEF Motor '1 I

lG Handle DOWN Permission
ButtonlG WHEELS Down lights
Mal & Ind light TesVDim
FeatureMASTER ARM Switch
MSl Step Switch
MlG WOWNUCLEAR CONSENT SWltdi

PROBE HEAT Switch
Radar Altimeter
SCP (left half)
Speedbrake$
Stations 1, 4 , 5 , 7 & 8
Arm Release
Station 7 Control PowerStores Jettison ISEL
& EMERJ

NWS

TACAN

Probe Heat Monitor

Cockpit Temp Control

VHF Radio

IIRI WPN AEl Button

MAIN GEN
FAIL

o

We-

" ,
' CM

Station 5 Control Power

Chaff/Flare DIspensers

FlCS Battery Heater
Inlet Stations
lIeft & rig ht)
Nacelle DC Bus Feeder

TWS

CWI

TISl
Upper/lower Antenna
Arrays (AIFF interrogation)

ENG FEED Knob

Altimeter (PNEUIAntiskid/Brakea IChan
Battery Ind Power Test
CANOPY Warning light
Cockpit Spotlightsl'~ E DU~ EEC caution light(switch position only)

ELEC SYS caution light~ Engine Auto lean
~ Engine History

RecorderENGINE Warning Ught

'PU

Firo/Overheat LIght Bulbs

FUEL HOT Caution light
MAX POWER Switch
Nacelle Ejector Shutoff

FCR

FlCC (FlT CaNT SYS end
DUAL FC FAIL lightsl
FlCS Power Source
Ibrand,es A & B)
FTlT Indicator
fUEL MASTER Switch
HYD/Oll PRESS
Warning light
JFS

Main GCU
Main Power Indicator
MAIN PWR Switch
MlG WOW (branches
A & 8)
SAl

TO/lDG CON FIG
Warning light
Utility light

VMS
1f-l$A-l_11H·lxe

Figure 1-31. (Sheet 3)

1-100

Chang e 2

T.O.1F-16A-1

Autopilot
Auto Fwd Fuel Tran.

EO.

IlS
Bleed AIr Valvas

AVTR
Battery Bu. Feeder
CABIN PRESS CautionLight
CAMERA/GUN Trigger·Canopy
CIU
Cockpit Prassure Dump

FAULT
I
C. .
light
EaUIP HOT caution Light
FUEL LOW Caution Light'o,G"[AJ HF Radio
HUDLANDING/TAXVEXTI10 LIght Switches
Swltche.
LEF Motor (1)LEF & ADC Caution Lights
LG Door Close
LG Hyd I.olatlon
LG UP-ON Command
MASTER ARM Switch
OXY LOW caution Ught
SCP (right half)

Hook & HOOK Caution LIght
IntercomLEF Aaymmetry Brake.
LG Warning Light {handlet
Maater Caution Light
NLG WOW

Antlskid/Brakes
ANTI-SKID Caution LightI PW2001 EEC Caution LightEPU
EPU Fuel Indicator
FLCS Power Source(branches C & OJ

SEAT NOT ARMED
Caution Light
Secure Voice
Station 3 Control Power
Stations 2, 3, 6 & 9
Arm Relea.e
Stick Trim
STORES CON FIG Caution LightStore. JenilOn (SEL & EMERt[!][@!] WPN REL Button

OVercurrant Trip
Brake
I
Caution LightI"""w,_~ Can be powered by EPU

LEGEND:
ELECTRICALMECHANICALIIIIIIIID Powered by EPU with MaIn Gen on Linec::::::J Powered thru NONESS AC Bus No. 2~ Powered thru OVercurrent Prot Panel No.1

• Power available from the battery bu. box wtth MAIN PWR awitch OFF.

1F-1 .... _1_11"_ZX.

Figun 1-31. (Shut 4)

Change 21-100.1

T.O.1f-16A-1

DC Power Distribution Diagram LESS Ml
Block 15•[!] ALT REL Button

LG Handle DOWN Permission

ANTI ICE Switch
ADA Indexer
AR System
Autopilot
Banery Bus Feeder
Battery Charge PowerCADC Caution LightCIU
Cockpit Temp Control

ButtonData Control

and self-test capability)
FLeS Power Source

LG WHEELS Down Ughts
Mal &. Ind Light Test/Dim
FeatureMASTER ARM Switch
MSL Step Switch

Fuel Tank Inerting

MlG WOWIFF
INSTR MODe SelectLEF Motor 11)

NWS,"",,,,"",,",1.' Control Test
FLCP (all lights, reset, arm

LESS~ NUCLEAR CONSENT
Switchm Probe Heat Monitor

PROBE HEAT Switch
Radar Altimeter
SCP Ueft halfl
Speedbrakes
Stations 1, 4, 5, 7 &. B
Arm Release
Station 7 Control PowerStores Jettison ISEl&. EMER)TACANVHF Radio
ftfil WPN REl Button

MAIN GEN
FAIL

o

•

-

FLCS Banery Heater
Inlet Stations (left &. right!

Station 5 Control Power
TISl

Chaff/Flare Dispensers
ENG FEED KnobFeRlESS ~ Flight Loads Recorder

FUEL HOT Caution light
MAX POWER Switch
Nacelle Ejector Shutoff

If 1SA 1 1197 18.

Figure 1-31_ (Sheet 5)1-100.2

Change 12

T.O.1F-16A-1

Autopilot
Auto Fwd Fuel Trans

AVTR
Banery Bus Feeder
CABtN PRESS Caution
Ught
CAMERA/GUN Trigger

Engine Bleed Air Valves(close capabilityl
IpW220lENGlNE FAULT
Caution Ught
EaUIP HOT Caution Ught
FUEL lOW Caution light

G,n

CIU

HUD

Cockpit Pressure Dump

ILS

ECA

LANDING/TAXI/EXT/~rmI00 10 Ught
SwitcheslEF Motor III
lEF &. ADC Caution Ughts
lG Door CloselG Hyd IsolationlG UP-ON Command
MASTER ARM Switch0')('( LOW Caution light
SCP Iright half)

SEAT NOT ARMED
Ceution Ught
lESS ~ Secure Voice
Station 3 Control Power
Stations 2, 3, 6 &. 9
Arm Release
Stick Trim
STORES CONFIG Caution light
Stores Jenison (SEl & EMERITWS[AJ1D.f] WPN REl Bunon

BATTERY

Altimeter jPNEUJ
Antiskid/Brskes
ANTI-5KIO Caution light
Banery Ind Power Test
Bleed Air Control
IPW200 1BUC Caution light·Canopy
CANOPY Warning light
Cockpit Spotlights~INOJ Drag ChUie~EDU~ EEC Caution light

ElEC SYS Caution Ught
~ Engine Auto lean
~ Engine History
RecorderENGINE Warning Ught
EPU
Fire/Overheat light Bulbs
FleC (FlT CONT SYS and
DUAL FC FAtl lightsl
FlCS Power Source
FTIT Indicator

FUEL MASTER Switch
Hook &. HOOK Caution Ught
HYD/OIL PRESS
Warning light
Intercom

JFS
LG Oownlock Relays(WHEELS down lightsl
LG Warning Light Ihandlef
Mltin PO)Ner IndicltlOI
MAIN P'VVR Switch
Master Caution Ught

MLG WOWNlG WOW
Parking Brake
RPM Indicator
SAl
IPW2201SEC Caution Ught
TO/lDG CONFIG
UHF Radio
Utility Ught

VMS

(switch position onlyl

lEGEND:~ Can be powered by EPU- - - ELECTRICAL[]]]]J Powered by EPU with Main Gen on Una

MECHANICAL

c::::;;:::] Powered thru NONESS AC Bus No. 2

c::::::::J Powered thru Overcurrent Prot Panel No. 1• Power available from the battery bus box with MAIN P'VVR switch OFF.

1F- 16A-1- 1197- 2A8

Figur:e 1-31. (Sheet 6)

Change 121-100.3/(1-100.4 blank)

•

o

o
o

o
o

o

10.lF-16A-1

HYDRAULIC SYSTEMHYD PRESS Indicators

Refer to figure 1-32. Hydraulic pressure is suppliedby 3000 psi hydraulic systems designated as systems
A and B. The systems are powered by two
independent engine-driven pumps located on the
ADG. Each system has a reservoir to store hydraulicfluid. The reservoirs are pressurized by theirrespective hydraulic system to insure positive
pressure at the pump. For hydraulic system cooling,refer to ENGINE FUEL SUPPLY SYSTEM, thissection,The HYD PRESS indicators, one for system A and one
for system B, are located on the right auxiliary
console. The indicators are powered by essential ac
bus No.2.

Both systems operate simultaneously to supply
hydraulic power for the primary flight controls andLEF's. If one of the systems should fail, the remaining
system provides sufficient hydraulic pressure;
however, the maximum actuation rate of the FLCS is
reduced. System A also supplies power to the FFP and
the speedbrakes. All remaining utility functions,consisting of the gun and gun purge door, AR system,
LG, brakes, NWS, and ~ ~ drag chute system aresupplied by system B. System B also charges thebrake/JFS accumulators (which provide start power
for the JFS and backup pressure for the brakes),
provided the engine is rotating at a minimum of 12
percent rpm. System B takes between 40 seconds (hotambient conditions) and 60 seconds (cold ambientconditions) to recharge the brake/JFS accumulators.
Irrn IHQ] System B also contains a drag chuteaccumulator which provides hydraulic pressure to
the drag chute system in case of hydraulic system B
failure.
The LG can be extended pneumatically in the eventof hydraulic system B failure. Should both hydraulicsystems fail, a third hydraulic pump located on the
EPU automatically provides hydraulic pressure tosystem A. (Refer to EMERGENCY POWER UNIT
(EPU), this section, for a further discussion of the
EPU.)

I

Each hydraulic system has an FLCS accumulator
which is isolated from the main system by checkvalves. These FLCS accumulators serve a dual
function. If demand exceeds the pump maximum flowrate during rapid control surface movement, theFLCS accumulators provide additional hydraulicpressure. Also, if both hydraulic systems fail, theFLCS accumulators provide adequate hydraulicpressure to the flight controls while the EPU comesup to speed. Refer to SERVICING DIAGRAM, thissection, for servicing/specifications information.
HYD PRESS INDICATORS AND WARNING LIGHT• Refer to figure 1-33.

HYD/OIL PRESS Warning LightA HYD/OIL PRESS warning light, located on theright glareshield, comes on when hydraulic system
A or B pressure drops below 1000 psi or when engine
oil pressure drops below 10 ( ± 2) psi. For the oilpressure function only, there is a 30-second timedelay in the light circuit to minimize transient lights
during negative g maneuvers. The light is powered by~ battery bus No.1, LESS ~ the battery bus.

EMERGENCY POWER UNIT (EPU)
Refer to figure 1-34. The EPU is a self-contained
system which simultaneously provides emergency
hydraulic pressure to system A and emergency
electrical power. The EPU is automatically activatedwhen both hydraulic system pressures fall below
1000 psi or when the main generator disconnects
from the bus system. The EPU may be operated
manually regardless of failure conditions.
The EPU requires battery bus (~ either batterybus) or main generator PMG power for automatic or
manual activation. When the EPU is operating, theessential ac and de buses are powered by the EPU
generator. If the main generator is on line, the EPU
also powers the nonessential dc bus, and the main
generator powers the nonessential ac buses. When
operating, the EPU augments hydraulic system A as
required. If the normal system A hydraulic pumpfails, the EPU is the only source of system A pressure.
The EPU uses engine bleed air and/or hydrazine tooperate. Normally, engine bleed air is used tomaintain operating speed. When bleed air isinsufficient, hydrazine augmentation automaticallyoccurs. Hydrazine is always used when the EPU iscommanded to start except when activated during
ground test using the EPU/GEN test switch.
On system command, hydrazine is forced by nitrogenpressure into a decomposition chamber. The gaseous
products of the reaction spin the turbine/gearbox
which then powers the EPU generator and hydraulicpump. Hydrazine exhaust is vented overboard on the

1-101

T.O.1 F-16A-l

H}'Idraulic Power Supply(Ty.pical)
SYSTEM A~

r

=.

"n

----- II FU'UOIL

II

~

:.!~:

'-

FLCS
ACCUMULATOR

LEF'SENGINEDRIVENHYDRAULICPUMPRUDDER

HORIZONTAL
TAILS

'PU

HYDRAULIC
PUMP

l.

l!

FLAPERONS,.

~FILTERSPHDBRAKES

FFP

l~

I
I

LG SYSTEMNWSWARNING LIGHT

I@~DRAGCHUTE SYSTEMBRAKE/JFSACCUMULATORS

f)r'~ '~GUN SYSTEM~HYD PRESS
INDICATOR

DUMP

AND
RELIEFMANIFOLD

cr:J
f-

'I

,.
DUMP ANDREUEF VAlVEI::JAR SYSTEMBRAKE SYSTEML@JFS SYSTEMRESERVOIRPRESSURIZATIONACCUMULATOR~RESERVOIRPRESSURIZATIONACCUMULATOR

10M-IX.

Figure 1-32. (Shut 1)

1-102

T.O.1F-16A-1

SYSTEM B
LEGEND:FILTER

ELECTRICAL§

ENGINEDRIVENISOLATION VALVEcS BLEED VALVE[] FILL CONNECTOR

FLCS

ACCUMULATOR~ QUICK DISCONNECT
8 CHECK VALVE

ffi PRESSURE SWITCHm

PRESSURE TRANSMITTERrn

FILTER

BYPASSPRESSURE

RETU RN
LG DOORSUPLOCK

SUPPlYACTUATE D[@!'lQJDRAG CHUTE
ACCUMULATORDUMP ANO_ _ _~ RELIEF VALVE

Figure 1·32. (811«, 2)

1-103

T.O. 1 F-16A-1

HYD PRESS Indicators and Warning Light(Typical)
HYD/OILHYD/OIL

PRESS

PRESS

\

1

1

1

2
1. HYO/OIL PRESS Warning Light (Red)
2. System A HYD PRESS Indicator
3. System B HYD PRESS Indicator23
1F-16A-1-0028A ®

Figure 1-33.

lower inboard side of the right strake and consists
primarily of nitrogen, hydrogen, ammonia, andwater. The temperature of exhaust gases can reach1600°F and will ignite in the presence of a flame. The
exhaust gases have an ammonia odor, are irritating
to the nose and eyes, and should be avoided to the
maximum extent possible.

• Prevents or terminates EPU operation on the
ground (WOW).

EPU CONTROLS AND INDICATORS• Prevents EPU operation in flight if switch has
remained in the OFF position since takeoff (sinceWOW).

• Refer to figure 1-35.
EPU Ground Safety Switch

A ground safety switch, located on the right side of theengine inlet, is used to disable the EPU on the ground.
With the EPU safety pin installed, the EPU does notoperate.
EPU Switch

The EPU switch is a three-position toggle switch.
Functions are:
• OFF

1-104

• Does not prevent or terminate EPU operation inflight for main generator failure if switch was
cycled or placed to NORM at anytime since takeoff(since WOW).

• Terminates EPU operation in flight except duringmain generator failure.
• NORM - The system is armed for automatic
operation except during engine shutdown on the
ground. With WOW and throttle in OFF, the EPU
does not activate when the main generator drops
offline.
• ON - Commands EPU to run regardless of failure
conditions. Operation will cease when switch is
positioned to OFF except for a main generator
failure in flight .

T.O.1F-16A-1

Emergency Power Unit Schematic (Typical)
HYD SYS A RSVRNOTE:Shown in augment mode with main generator failed.
HYDRAULICTEST RUNCOMMAND

PUMP
__
_J

, ~~"~~.D FOR

PRESSURE ,. 2000 PSIEPU RUN LIGHTI{BLEED AIR
ONLY}

EPU

SPEED

RUN
COMMANDMONITORTO{AUGMENT}CONVERTER!
.-- 1
REGULATORS +---+-----1~111
5 KVA• , GENERATOR,EPU FUELI

ESSENTIAl.~t~~A~C~PO~W~E~Rg;~~..J~~~=~~1~-=~r'!:-lr
OVERSPEED SHUTOFFAC BUSESTO

MAIN GENERATOR
FAIL INDICATION!TERTIARY CONTROLIBLEED AlR
SOLENOIDVAlVE

EPU GENERATOR FAIL INDICAnON

o+--{ "PU PMG FAIL INDICATlON FROM CONVERTER/REGULATORSHIGH PRESSUREBLEED AIR

EPU
CONTROLLERIHYDSYSAIIHYDSYSBIRUN UGHTFOR OVER OR
UND£RSPEEO
CONDITION

CLOSED FOR
PRESSURE < 1000 PSI

OPeN WITH

wow &CLOSED IF MAIN
GENERATOR OfF UNE

THROTTLEo

BELOW IDlE 0

E.
'0:"
U

_ - I:I__

~.:
.".

j:

AUGMENT
MODE

I
I

___ J

ARM!

EPU!GEN TEST

- - - -,

MAIN
GENERATOR---,BLEED AIR

CLOSED ON
START. MAIN

,

GENON

,0

EPU TEST RUN COMMAND-I==~~~~~:=~ EPU START UP AND CONTROL POWEREPU GROUNDSAFETY SWITCHoLEGEND:_

2B VDC PW. -t:=~~==:J-' ~OPEN IF EPU
SAFETY PININSTAlLED

BLEED AIRHVORAZlNE FUEL_HYDRAULIC FlUID SUPPlY_HYDRAUUC PRESSURE_GASEOUS NITROGEN

MODE ONLY

Ic L--'n___ -.Jo FIllCONNECTORQUICKOJ DISCONNECT- - ELECTRICAL- - - - MECHANICAL
1F_16o'_' _101 0)ll.

Figure 1-34.

1-105

T. O. 1F- 16A- 1

EPU Control Panel[8J IBF I (Typical)

EPU/GEN Test Switch

The EPU/GEN test switch, located on the TEST
switch panel, has positions of OFF and EPU/GEN.
The switch is spring-loaded to the OFF position. It
provides a means to test the EPU generator and EPU
PMG output to FLCS on the ground without using
hydrazine.
EPU FUEL Quantity Indicator ~ [I£]Refer to figure 1-3. The EPU FUEL quantity
indicator, located on the right auxiliary console, isgraduated 0-100 and indicates the percent ofhydrazine remaining. The indicator is powered I
by ~ battery bus No.2, LESS ~ the battery bus.

4

1. HYDRAZN Light (Amber)
2. AIR Light (Amber)
3. EPU Run Light (Green)
4. EPU Switch
1 F-16A-1-0029A<tHydrazine Leak Detector

The hydrazine leak detector is a silicone base,mustard yellow disc visible through access door 3208.
The viewing area is black on one half to provide
contrast with the mustard yellow disc. The mustard
yellow turns purple/black in the presence of
hydrazine and/or its vapors, indicating a leak in the
EPU and/or fuel tank system.

Figure 1-35.

EPU Fired Indicator

The switch has a split guard; the top half can be raised
to move the switch to ON and the bottom half can beraised to move the switch to OFF. When both sectionsof the guard are down, the switch is retained in the
NORM position.
EPU Run Light

The EPU run light, located on the EPU control panel,
illuminates when the EPU turbine speed is within theproper range and the EPU-driven hydraulic pumpdischarge pressure is above 2000 psi.
HYDRAZN Light

The HYDRAZN light, located on the EPU control
panel, illuminates when the EPU is commanding
hydrazine for operation (whether hydrazine isavailable or not) or if a primary speed control failure
has occurred.
AIR Light

The AIR light, located on the EPU control panel,
illuminates whenever the EPU has been commanded
to run with the EPU safety pin removed. It remains
on even when the EPU is augmented by hydrazine.

1-106

Change 2

The EPU fired indicator is located next to the EPU
ground safety switch on the right side of the engine
inlet. Normally, the indicator displays a gray andblack disc. If the EPU has been activated, theindicator displays six equally spaced black and white
triangles.
EPU OPERATIONThe EPU is designed to operate automatically formain generator failure, dual hydraulic system
failure, PTO shaft or ADG failure, and engine
flameout or if the engine is shut down in flight. The
EPU can also be activated manually. After receiving
any start command, the EPU requires approximately
2 seconds to come up to speed. EPU startup may not
be audible. Once operating, however, the EPU may beheard but does not sound the same as during the EPU
ground check. A lack of sound during EPU startup
does not indicate lack of EPU operation which must
be confirmed by monitoring the EPU run light. EPU
rpm is controlled by three speed controls. The
primary and secondary speed con troIs are based on
EPU rpm. The tertiary speed control is based on EPU
PMG frequency.

T. O. 1F- 16A- 1

When the EPU is operating, engine thrust settings
should be maintained to prevent using hydrazine.
This normally requires a minimum of 75-80 percent
rpm depending on pressure altitude.
If the engine fails, hydrazine alone is used to power
the EPU. With hydrazine only, operating time of thesystem is approximately 10 minutes under normalload requirements. Increased flight control movement reduces this operating time. When the EPU isthe sole source of hydraulic power, EPU loss resultsin loss of aircraft control. Refer to SERVICINGDIAGRAM, this section, for servicing/specifications
information.

LANDING GEAR (LG) SYSTEM
The LG system is normally operated by hydraulicsystem B. The NLG is extended and retracted byhydraulic pressure. The MLG's are retracted
hydraulically but are extended by free-fall assisted byairloads. All the LG doors are hydraulically activatedwith electrical sequencing during retraction andmechanical sequencing during extension. If hydraulic system B fails, the LG may be extended
pneumatically.
MAIN LANDING GEAR (MLG)

The two MLG are independent of each other andretract forward with a mechanical wheel twist into
two separate wheel wells. Each MLG wheel isequipped with three fusible (thermal pressure relief)plugs.
NOSE LANDING GEAR (NLG)

I

The NLG retracts aft with a 90-degree mechanicalwheel twist into the wheel well. A torque ann
quick-disconnect is provided so that the nosewheel
can be turned beyond the steerable range for towing.
LANDING GEAR CONTROLS AND INDICATORS

I

Refer to figure 1-36. The LG control panel is locatedon the left auxiliary console.
Landing Gear Handle

The LG handle, located on the LG control panel, has
a wheel-shaped grip. Movement of the handle
operates electrical switches (powered by essential dc
bus No.2) to command LG retraction or extension. A
warning light in the LG handle, powered by IAQJ battery bus No.2, LESS IAQJ the battery bus, illuminates
when the LG and doors are in transit or have failed tolock in the commanded position. The warning lightalso illuminates when all LG's are not down andlocked, airspeed is less than 190 knots, altitude is lessthan 10,000 feet, and rate of descent is greater than
250 feet per minute. The handle is locked in the DN
position when the aircraft is on the ground (weight onwheels), In flight, a signal from the left MLG WOW
switch automatically activates a solenoid which
unlocks the handle, allowing movement to the UP
position. The handle is locked in the UP position toprevent LG extension during high g maneuvers.
Landing Gear Handle Down Permission Buffon

The LG handle down permission button, located on
the LG handle, unlocks the handle electrically topermit movement to the DN position. The buttonenergizes an electrical solenoid which releases thespring-actuated handle lock. The button must be
depressed before downward force is applied to the LG
handle. The electrical solenoid may not unlock thehandle while any appreciable downward force isapplied.
DN LOCK REL Buffon!Al ~ The DN LOCK REL button, located on the LG
control panel, when depressed, mechanically unlocks
the spring-actuated handle lock if the electricalsolenoid should fail or not be powered. It overrides allelectrical LG control signals. Depressing this button
and raising the LG handle on the ground retracts the
LG. The DN LOCK REL button may not unlock the
LG handle while any appreciable downward force isapplied. ~ For DN LOCK REL button differences,refer to F -16B AIRCRAFT, this section.
ALT GEAR Handle

The ALT GEAR handle, located just outboard of andbelow the LG control panel, is used to extend the LG
if normal extension is not possible. Pulling the ALT
GEAR handle supplies pneumatic pressure to open
all LG doors, extend the NLG, and shut off the LG
selector hydraulic valve. The LG/hook emergency
pneumatic bottle is also used to lower the hook andcontains sufficient pneumatic pressure for one LG
extension and to hold the hook down. The bottle
cannot be recharged in flight. Since pneumaticpressure is reduced by expansion as the actuators
extend, less than the normal extending force isavailable.
An LG reset button, located on the ALT GEAR
handle, provides a means of retracting the LG afteran alternate extension if system B hydraulic pressure
is available. Refer to LANDING WITH LG UNSAFE/
UP, Section III.

Change 1

1-107

IT. O. 1F- 16A- 1

Landing Gear Control Panel (Typical)
3

4

5
6

11'=';;';"-:"'--7

8

•••

•••
•••

8
14

~

3
19

5
6
1. EMER STORES JETTISON Button (Covered)
2. WHEELS Down Lights (Green)
3. HOOK Switch (Lever Lock)4, ANTI-SKID Switch
5. DN LOCK REL Button
6. LG Handle Down Permission Button
7. LG Handle
8. LG Handle Warning Light (Red)
9. ALT GEAR Handle
10. ALT GEAR Reset Button
11. SPEED BRAKE Position Indicator
12. BLOCK 10 LE FLAP POSITION
Indicator (Deactivated)
13. LANDING TAXI LIGHTS Switch
14. HORN SILENCER Button
15. STORES CON FIG Switch
16. BRAKES Channel Switch
17. GND JETT ENABLE Switch (Lever Lock)
18. ALT FLAPS Switch
19. ~INQIDRAG CHUTE Switch

7

11

10

•••
•••
•••
9

8J
1F-16A-1-1072A@

Figure 1-36.

1-108

8

T. O. 1F- 16A- 1
LG Warning HornLanding Gear Welght-on-Wheels (WOW) Switches

The LG warning horn is an intermittent fIxed volume
signal which sounds in the headset when the NLG or
MLG is not down and locked and all the following
conditions exist:

The LG WOW switches, located on both MLG's and on
the NLG, operate as a function of LG strut extension
to allow or terminate various system functions.

• Airspeed is below 190 knots.
• Pressure altitude is less than 10,000 feet.
• Rate of descent is greater than 250 fpm.

LANDING GEAR OPERATIONHORN SILENCER Buffon

Movement of the LG handle to the UP position causes
the following events:

The HORN SILENCER button is located on the LG
control panel. Depressing the button silences the LG
warning hom. If the warning condition is eliminated,the horn resets. If it is not eliminated, subsequent LG
audio warnings do not occur.
T.O./LAND CONFIG Warning Light

The T.O.ILAND CONFIG warning light, located on
the left glareshield, illuminates in flight wheneverpressure altitude is less than 10,000 feet, airspeed isless than 190 knots, rate of descent is greater than
250 fpm, and either of the following conditions exists:• TEF's not full down.
• NLG or either MLG not down and locked(accompanied by LG warning hom).
The T.O.ILAND CONFIG warning light illuminates
on the ground if TEF's are not full down.
With TEF's full down, rapid reversals of roll command
inputs may cause the T.O.ILAND CONFIG warning
light to momentarily illuminate if the altitude,
airspeed, and rate of descent conditions outlined
above are met or WOW.
WHEELS Down Lights

I

Refer to figure 1-37 for a list of systems affected by the
WOW switches and symptoms of WOW switch
failure.

The three green WHEELS down lights, located on the
LG control panel, are arranged on the silhouette of
the aircraft. When any LG is down, its respective lightis on. A safe up and locked LG condition is indicatedwhen all three of the lights and the LG handle
warning light are off. The lights are powered by the~ battery bus or buses, LESS ~ battery bus (toenergize downlock relays) and essential dc bus No.1
(to illuminate lights).

• LG handle warning light illuminates.
• LG unlocks and retracts.
• Three WHEELS down lights go off.
• MLG wheel spin is stopped.
• LG doors close and lock.
• LG handle warning light goes off.
• Hydraulic pressure is removed from LG.
• FLCS switches to cruise gains.
• TEF's retract to streamlined position.
• Electrical power is removed from brake channell.
• LG UP nozzle scheduling is activated.
Movement of the LG handle to the DN position causes
the following events:• LG handle warning light illuminates.
• LG doors and LG unlock, extend, and lock into
place.
• Three WHEELS down lights illuminate.
• LG handle warning light goes off.
• TEF's extend.
• FLCS switches to takeoff and landing gains.
• LG DN nozzle scheduling is activated.
• Speedbrakes close to 43 degrees if not overridden.
• Electrical power is supplied to brake channell.

Change 7

1-109

T.O. 1F-16A-1

LG WOW Switch
RIGHT MLG - SYSTEMS
Aircraft Battery

EPU

Air Data Probe

FCR

Altimeter (ElECn

FLCC~ Chaff/Flare Dispenser

FlCP

AOA Probes

lG Warning

Bra kes/Antiskid

Pitot Probe

ECA@!) Probe Heat Monitor

ECS

SMS

Engine Controls

BLOCK 15 ~ VMS

Falls to Ground Position In FlightI

Falls to In-Flight Position on Ground~ Chaff/flare dispenser is inoperative~ Chaff/flare dispenser is operative

FCR can not transmit

FCR can transmit

Stores cannot be emergency jettisoned unless GND JETT ENABLE
switch is in ENABLE

Stores can be emergency jettisoned with GND JETT ENABLE switch
in OFF

BLOCK 15 ~ VMS is inoperative

BLOCK 15 ~ VMS is operative unless INHIBIT is selected

Brakes can be applied before touchdown if toe brakes are
depressed

With simultaneous failure of left and right MLG WOW switches
and ANTI-SKID switch in ANTI-SKID, toe brakes are inoperative
when groundspeed is less than 20 knots

ANTI-SKID switch holds in PARKING BRAKE with throttle in OFF to
IDLE range

With simultaneous failure of left and right MLG WOW switches,
ANTI-SKID switch must be held in PARKING BRAKE to operateparking brake

With simultaneous failure of left and right MLG WOW switches,
ACFT BATT FAil light indicates aircraft battery failure (voltage 20Vor less) or cell voltage imbalance condition

ACFT BATT FAIL light indicates aircraft battery failure only

~ LG and low speed warning tones are inoperative

EPU is commanded on during engine shutdown; operation cannot
be terminated with the EPU switch

mProbe heat monitor is inoperative unless TEST or PROBE HEAT

mProbe heat monitor is operative

is selected

All probe heaters except total temperature are on
ACC, DUAl Fe FAIL, CADC, and LE FLAPS lights may be on if AOA
probes are not equally positioned
FLCS self-test cannot be initiatedFigure 1-37. (Sheet 1)

1-110

Change 7I10. 1 F-16A-l

LG wow Switch
LEFT MLG - SYSTEMS
Aircraft Battery

Ground Test Panel (fuel pump lights)

AOA Probe (right)

JFS Ground Cutout

Brakes/Antiskid

LG Handle

ECA
LG Warning
EPU
Total Temperature Probe

FlCC
FLCP

BLOCK 15 ~ VMS

FLCS Batteries

BLOCK 10 VMS

Falls to Ground Position in Flight

Fails to In-Flight Position on Ground

AOA probe heater (right) is on
JFS shuts down automatically during engine start

JFS does not shut down automatically during engine start

LG UP position cannot be seleded unless ON LOCK REL button isdepressed

LG UP position can be selected without ON LOCK REL button
depressed

T.O./LAND CONFIG warning light is on with TEF's not down

T.O./LAND CONFIG warning light is off with TEF's up

Total temperature probe heater is inoperative

Total temperature probe heater is on

BLOCK 15BLOCK 15is selededlID and BLOCK 10 VMS is inoperativeI

rm and BLOCK 10 VMS is operative unless INHIBIT

Brakes can be applied before touchdown if toe brakes are
depressed

With simultaneous failure of right and left MLG WOW switches
and ANTI-SKID switch in ANTI-SKID, toe brakes are inoperative
when groundspeed is less than 20 knots

ANTI-SKID switch holds in PARKING BRAKE with throttle in OFF to
IDLE range

With simultaneous failure of right and left MLG WOW switches,
ANTI-SKID switch must be held in PARKING BRAKE to operateparking brake
Fuel pump lights on external ground test panel are inoperative

With simultaneous failure of right and left MLG WOW switches,
ACFT BAn FAIL light indicates aircraft battery failure (voltage 20Vor less) or cell voltage imbalance conditionrm LG and low speed warning tones are inoperative

ACFT BAn FAIL light indicates aircraft battery failure only
ADC, DUAL Fe FAIL, CADC, and LE FLAPS lights may be on if AOA
probes are not equally positioned
EPU is commanded on during engine shutdown; operation cannot
be terminated with the EPU switch
FLCS self-test cannot be initiated
FLCS batteries do not turn off after shutdown

Figure 1-37. (Sheet 2)

Change 7

1-111

T. O. 1F- 16A-1

LG WOW Switch
NLG - SYSTEMSFLCS PowerAir Data Proberm~ Chaff/Flare Dispenser(DI) 1m ~ 10 lightandAOA Probe (left)~ NWS[!] AR

Pitot Probe

FLCP

Speedbrakes

Falls to Ground Po.ltlon In .lIght

Falls to In-Flight Position on GroundAir data, pitot, and left AOA probe heaters are on~ AIR DISC button is operative[!] AIR DISC button is inoperativeI

FLCS fails self·test~ 10 light is inoperative unless TEST is selected[QEl!m2110 light is operative if 10 UGHT position is selected. TEST
position is inoperative

m[!] 10 light is inoperativeem~ 10 light is operative if TAXI position of LANDINGTAXI LIGHTS switch is selected

NWS can be engaged and follows rudder inputs with NLG down

NWS is inoperative

Speedbrakes are not limited to 43 degrees with right MlG down
and locked

Speedbrakes do not remain open more than 43 deg ....s

ACFT BAn TO FLCS light indicates aircraft battery bus is IUpplying power to one or more FLCS branches

ACFT BAn TO FLCS light indicates battery bus is supplying
power to one or more FLCS branches (bus voltage 25 vdc or less)

FLCS PMG light indicates FLCS PMG power is not available atone or more FLCS branches

FLCS PMG light indicates the FLCS PMG is not supplying power
to any FLCS branches~ Chaff/flare dispenser is inopera1ive

Gm Chaff/flore dispenser is operative

Figure 1-37. (Sheet 3)

1-112

Change 4

T.O. 1 F-16A-l

NOSEWHEEL STEERING (NWS) SYSTEMThe NWS is electrically controlled using dc bus No.1
power and is hydraulically operated using system B
pressure. Steering signals are provided through therudder pedals. Should NWS be engaged with therudder pedals displaced, the nosewheel drives to therudder pedal commanded position. NWS is limited to
32 degrees in each direction; however, turn radius canbe reduced by using inside brake. NWS is
automatically disengaged when the NLG strut is
extended. NWS is not available following analternate LG extension and may not be available
anytime the NLG WHEELS down light is notilluminated. Refer to TURNING RADIUS ANDGROUND CLEARANCE, Section II.

down and/or electrical power is removed. NWS can
only be engaged and operated from the cockpit which
has control. When NWS is selected, the NWS lightilluminates green in both cockpits. The aft stick
paddle switch allows immediate override of theforward cockpit if the STICK CONTROL switch on
the TEST switch panel is in AFT; the NWS control
button/ indicator in the aft cockpit and OVRD lightilluminates. The front cockpit paddle switch cannot
override the aft cockpit regardless of the position of
the STICK CONTROL switch. With the AIR
REFUEL switch in OPEN, NWS can be engaged ordisengaged from either cockpit without using thepaddle switch regardless of which cockpit hascontrol.

WHEEL BRAKE SYSTEMNWS CONTROLS AND INDICATORS

Refer to figure 1-3.
NWS AIR DISC MSL STEP Buffon

The NWS AIR DISC MSL STEP button, located on theoutboard side of the stick, is used to engage or
disengage NWS when the aircraft is on the ground.
Once depressed, NWS is engaged and the button may
be released. If the button is held depressed,
continuous NWS is provided.
NWS Light

The NWS light, the center element of the ARlNWS
status indicator located on the top of the glareshield,
illuminates green when NWS is engaged. NWS does
not operate even though the NWS light is illuminatedwhen the NWS FAIL caution light is on or whensystem B hydraulic pressure is unavailable. On the
ground, NWS continues to operate with the AIR
REFUEL switch in OPEN even though the NWS lightis off.
NWS FAIL Caution Light

The NWS FAIL caution light, located on the caution
light panel, illuminates when a failure in the NWS
system has caused electrical power to be switched off.
NOSEWHEEL STEERING (NWS) SYSTEM [I]Either cockpit can control the NWS. Control isaccom plished by means of an NWS con trol
button/indicator located in both cockpits just aft ofthe stick. Depressing the button/indicator transfers
control of the NWS; the indicator illuminates green
in the cockpit which has control. Control thenremains in that cockpit even if the engine is shut

Each MLG wheel is equipped with a hydraulicallypowered multiple disc brake. The brakes areelectrically controlled by conventional toe brake
pedals. The amount of braking gradually increases
as pedal pressure is applied. A parking brake is also
provided. An an tiskid system protects against blown
tires and is only available when using toe brakes.

mJ Brakes may be applied singly or simultaneously
from the forward or rear cockpits. The brake signals
from both cockpits are additive so that the total signal
to the brakes is the sum of the pedal forces from both
cockpits.
Brake hydraulic power is supplied by system B. If
system B fails or the engine is operating at less than
12 percent rpm, the toe brakes and parking brake areavailable until the brakelJFS accumulators deplete.
Continuous use of the toe brakes, even with theparking brake set, depletes brakelJFS accumulator
fluid and causes loss of all braking capability after
approximately 75 seconds (brakelJFS accumulators
initially fully charged). When holding the aircraft
stationary, use of the parking brake is preferred sincebrakelJFS accumulator fluid is not depleted.
TOE BRAKE SYSTEM

Refer to figures 1-38 and 1-39. The toe brakes use
electrical power from the four FLCS inverters and
CHAN 1 and CHAN 2 dc power sources. The brake
pedals require FLCS inverter power to operate. The
pedal signals are supplied to the (P£) brake
control/anti-skid assembly LESS ~ brake control box
which, in turn, uses both CHAN 1 and CHAN 2 dc
power sources to operate valves for controlling
hydraulic pressure to the brakes. ~ CHAN 1 and
CHAN 2 are powered by battery buses No.1 and No.
2, respectively. LESS ~ CHAN 1 and CHAN 2 arepowered by the battery bus.
Change 6

1-113

IT.O. 1F-16A-l

Toe Brake System
BRAKE
CHAN 2

BRAKE
CHAN 1, - - -____A_ _ _ __ _same except that when CHAN 1 is selected, both dc
power sources are switched off when the LG handle is
up. With CHAN 1 selected, the brakes only operatewith the LG handle down; with CHAN 2 selected, thebrakes are operable with the LG handle either up or
down. If the LG handle is stuck in the UP position,
CHAN 2 must therefore be selected to achieve braking.

SPIN DOWN BRAKING SYSTEM
FlCS

INVERTERS

c

A

o

B

The spin down braking system provides hydraulicbrake pressure to stop MLG wheel spin during LG
retraction. The hydraulic pressure is relieved whenthe LG is up and locked.

BRAKES CHANNEL SWITCH ~ [If]

TOE BRAKES

L

R

1 F-16X-1-0008X II

Figure 1-38.

The electrical power sources are grouped to provide
two redundant channels. Channel 1 uses FLCS
inverters A and C and CHAN 1 dc power. Channel 2
uses FLCS inverters Band D and CHAN 2 dc power.
If one inverter fails, one toe brake in either CHAN 1
or CHAN 2 is inoperative.
If a dual inverter failure occurs, both brakes inchannel 1 or 2 may be inoperative, or one brake ineach channel may be inoperative. Inverter failure canbe indicated by the FLCS PWR lights on the TEST
switch panel. Labels adjacent to the lights indicate
the brake and channel affected by each failure.
Regardless of which channel is selected, hydraulicpressure to three of the six pistons in each brake iscon trolled by electrical power from one de powersource and pressure to the other three pistons of eachbrake is controlled by electrical power from anotherdc power source. A loss of one dc power source when
CH \N 1 is selected results in degraded brake
operation (only one-half of the pistons are powered
and significantly more brake pedal force than normal
is required to stop). Due to redundancy features,
selecting CHAN 2 may restore full braking. If all dc
power sources fail or if all FLCS inverters are off, thetoe brakes are totally inoperative.
Channels 1 and 2 use separate redundant circuit
elements for controlling the brakes and operate the
1-114

Change 6

The BRAKES channel switch, located on the LG
control panel, has positions of CHAN 1 and CHAN 2
and allows wheel brake system switching. CHAN 1 isthe nonnal position.

PARKING BRAKE ~ em
The parking brake is activated by the ANTI-SKID
switch, located on the LG control panel, and supplies
full, unmetered pressure to three of the six pistons ineach brake. The parking brake holds the aircraft
stationary without the use of toe brakes. It C81) also be
used for emergency braking if the toe brakes areinoperative. The parking brake is poweredby ~ battery bus No.2, LESS ~ the battery bus
and system B hydraulics or one brakelJFS accumulator (the same brakelJFS accumulator which is used
for START 1).

ANTISKID SYSTEM
The antiskid system is available in either brakechannel anytime the toe brakes are powered.

&FE The antiskid system will deliver a corresponding
deceleration rate to a given pedal deflection. The
deceleration skid control will dampen brake pedal
inputs to the brakes resulting in a smoother, more
efficient stop than with previous antiskid systems. Th
optimize braking performance and reduce wear on
aircraft brakes and tires, smoothly apply brakes in asingle application.
Functions are:• Thuchdown skid control - Prevents brake application prior to wheel spinup even if brake pedals are
fully depressed.
• (PE Deceleration skid control- Active when either

brake pedal deflection is less than 85 percent of
maximum and runway surface can provide therequested deceleration.

I10. 1F- 16A-1.00 Maximum performance skid control- Active
when both brake pedal deflections are equal to or
greater than 85 percent or runway surface cannot
provide requested deceleration.

• 00 Antiskid failure detection - Detects a failure
affecting braking or in a system component.
• LESS 00 Proportional skid control- Prevents
skidding due to overbraking at 5 knots groundspeed
or greater.
• LESS mD Locked wheel skid control- Backs up the
proportional skid control and operates at 20 knots
groundspeed and greater.
• LESS 00 Antiskid failure detection - Detects anantiskid system malfunction.

I (fj) If a failure affecting braking performance is
detected while the aircraft is moving above 5 knots
groundspeed, the ANTI SKID caution light illuminates. In most cases this represents the loss of a wheel
speed sensor signal, and the system switches to analternate braking mode. In this mode, if differential
braking is applied (15 percent or greater difference
between pedals), both brakes alternate between
pedal pressure as metered and no pressure. Braking
effectiveness is reduced by 50 percent or greater. If
brake pedals are within 15 percent, the system uses
the information from the remaining good wheel speed
sensor and stopping distance is increased byapproximately 25 percent on both wet and dry
runways.
(fj) The alternate mode continues until the BRAKES

channel switch is switched to CHAN 2 and the
ANTI-SKID switch is placed to OFF. The ANTI SKID
caution light remains on and braking is manual. The
brakes then can be locked by applying too much pedal
pressure, which may result in blown tires.

ANTI-SKID switch in OFF while avoiding wheel
lockup and its associated risk of control difficulty.
Short field landing distances are increased approximately 60 percent for dry runway and 25 percent for
wet runway from those normally computed. The
amount of pulsating braking is dependent on the toepressure applied. Pulsating braking continues until
the ANTI-SKID switch is placed to OFF. At that time,
the ANTI SKID caution light remains on, and thebrake system reverts to manual control. The brakes
then can be locked by applying too much pedal
pressure, which may result in blown tires.
LESS 00 The antiskid system does not provide skid or
locked wheel protection if MLG wheels are notspinning due to hydroplaning. If WOW occurs prior to
spinup of at least one MLG wheel, wheel brakes
become operative without antiskid protection.
LESS (fj) Antiskid protection and failure detection areintentionally diminished as speed decreases below 20
knots to allow a complete stop without continuous
brake release or an ANTI SKID caution light.
Consequently, maintaining maximum toe pressure
while the aircraft comes to a complete stop may cause
wheel lockup during the last 5-10 feet before stoppingwith resultant tire flat spotting. A dragging brake notcorrectly releasing in response to antiskid signals
may also cause wheel lockup without an illuminated
ANTI SKID caution light at speeds below 20 knots.
Avoid maximum braking at speeds below 20 knots
whenever possible to prevent tire flat spotting andpossible blowout.
ANTI-SKI D Switch [!] mE

The ANTI-SKID switch, located on the LG control
panel, is not lever-locked in the ANTI- SKID positionand can be bumped to OFF.
Functions are:protection function which prevents brake application
until the wheels have spun up, even if WOW has
occurred before spinup.

• PARKING BRAKE - Full unmetered brake pressure
is applied with the throttle in the OFF to IDLE range
and WOW. Advancing the throttle more than 1 inch
beyond IDLE automatically returns the switch to
ANTI-SKID which releases the parking brake.

00 Full antiskid function becomes active at 12 knots• ANTI-SKID - Antiskid protection is available.

groundspeed when accelerating and is available tobelow 5 knots when decelerating. Maximum braking
below 12 knots groundspeed may result in tire flatspotting.

.00 OFF - Parking brake feature is deactivated andmD The antiskid system incorporates a hydroplaning

LESS mJ If a failure is detected, the ANTI SKID
caution light illuminates and the brake system
automatically switches to pulsating pressure(constant frequency pulsating on-off pressure). In
this mode, braking effectiveness is reduced approximately 50 percent; however, in most cases, braking
effectiveness is as good as can be obtained with

antiskid functions are as follows:• With BRAKES channel switch in CHAN 1Touchdown skid control is not available, but
deceleration and maximum performance skid
control remain active.
• With BRAKES channel switch in CHAN 2 - Allantiskid functions are deactivated.
• LESS (fj) OFF - Antiskid and parking brakefeatures are deactivated.
Change 8

1-115

T.O. 1F-16A-l

Wheel Brake Schematic (Typical) ii fi)
THROTTLE
POSITIONSWITCH
MLG W Ow

INEAR, 'DLE

I>:D~~H~~~~ .~~~~ VI~.
28 VOCI~BATTERY
BUS

Np~~
I~~~

I~~

I~

ANTISKIO
CAUTION
LIGHT
RELAV

1;-L

PARK
VALlIE

I'"
t--_
lr

~

,

CHANNEL 2

8RAKE CONTROLIANTI-SKID ASSV

28 VDC
8ATTERVBUS NO. 1~

CHANNEL SELECTORCHANNEL'OR
CHANNEL 2OUTPUT.w.. .....,eC~2 .,"" , .-.um·SllC~~I-L-L -LAC

I I~ ~olt-{

F

SHUTTlE
VALVE

PREBRAKE
RESTRICTOR
VAlVE

-

~

FLCS
INVERTERS

B

0

~ ~

~

WHEEL
LG UP
PRESSURE SPEEDSENSOR~-I
LESS fAQI

L...,

,

L

R

'--

'--

•

TOE
BRAKES,

26 VAC 800 Ii,
4 INDEPENDENT EXCITATION POweR
SOURCES FROM FLCS INVERTERS

Figure 1-39. (Sheet 1)

Change 6~

BRAKE
CHAN 2,.----"-----,

BRAKE
CHAN'
,.----"-----,

~~

1-116

l::::t!PM'~

28 VDC
8ATTERV
BUS
NO. 2

I~

T.O.1F-16A-1

HYO SYS B PRESSURE . . . .

LEGEND:- - ELECTRICAL- - - - MECHANICAL_HYD POWER
SUPPLYI ~ HYD RETURN

cz:D BRAKE HYD
CIRCUIT_

PARKING BRAKECIRCUIT[!] CHECK VALVELEFTMETERING

VAlVE

WHEEL
SPeED
SENSORLEFT
BRAJ(£

1F-ll1A-1_2074-lJe

Figure 1·39. (SMet 2)

Change 61-116.1

T.O. 1F-16A-l

I Wheel Brake Schematic (Typical) LESS ifO..t:1
THROTILE
POSITIONSWITCH

MLG wow

NEAR IDLE

28 VDC
BAmRV
BUS

I~"'II/I".
r I.. )..;~~~---,!},!
ANTlSKIDCONTROL
BOX

. -

N~~~

I~

r-~-h ANTISKIO-f

CAUTION'. ~I~r-o--

LIGHT
RELAY

;::

r:

"::o--J

2 8 VDC
BATIERVBUS NO. 1

BRAKE CONTROL ~
BOX

I@l

F-

g~ANNEl 1 {~~t~

CHANNEl o<o<~no

CHANNEL 2OUTPUT~ 'i",~

L....
;:;:.,;..,..,'.·-,--m-./
I -.,

1

2B VDC

SRAKe

BRAKE

CHAN'

CHAN 2~j::-

BATIERY ttt!i!i~
BUS

I
I

PAE BRAKERES TRICTOR ~;i;J

~LV E

NO. 2

C

A

o

FLCS

B

INVERTERSW HEELLG UP
PRESSURE SPEEDSENSORLESS IAQ]

-

L

R

L-

L-

TOE
BRAKES"---~---'...

26 VAC 800 Hz

4 INDEPeNDENT EXCITATION POWER
SOURCES FROM FLCS INVERTERS

Figun 1·39. (Soot 3)

1-11 6 _2

Cha nce 6

T. O.1 F-16A-l

HYD SYS B PRESSURE •

•• •

HYO

LEGEND :- - ELECTRICAL- - - - MECHANICAL_HYDPOWERSUPPLY_HYD RETURN

c:::J BRAKE HYO
CIRCUIT_

PARKING BRAKECIRCUIT8CHECK VALVE

WHEEL
SPEED
SENSORLEFT
BRAKE

IF- ,s.o.-' - 1074-2Xe

Figure 1-39. (SMet 4)

Change 6

1-117

T.O. 1 F-16A-1

ANTI SKID Caution Light

00 The ANTI SKID caution light, located on thecaution light panel, illuminates at groundspeeds
above 5 knots when a malfunction affecting braking
performance is detected. If a system malfunction notaffecting braking performance (e.g. loss of red un dancy) is detected, the light illuminates whengroundspeed is below 5 knots. The caution light is notlatched and will extinguish above 5 knots if a failurethat does not affect braking performance is present.

holding in the open position. If one switch is in theclosed position, the speedbrakes close when the otheris released from the open position.
SPEED BRAKE Position Indicator

A three-position SPEED BRAKE indicator is locatedon the LG control panel.
Positions are:
• CLOSED - Both speedbrakes closed.

00 The ANTI SKID caution light illuminates whenpower is applied to the brake controlJantiskid
assembly and goes off when power-up BIT has beensuccessfully completed (approximately 1/2 second
later). This brief illumination of the ANTI SKID
caution light may be observed when power is first
applied or after the LG handle is placed down with the
BRAKES channel switch in CHAN 1.

• Speedbrake symbol - Speedbrakes not closed.

LESS ~ The ANTI SKID caution light, located on thecaution light panel, illuminates when a malfunction
occurs with the ANTI-SKID switch in ANTI-SKID.

A drag chute is provided to minimize stopping
distance. Drag chute deployment is obtained when
hydraulic system B pressure is routed to the dragchute actuator by placing the DRAG CHUTE switch
to DEPLOY. Drag chute accumulator pressure isavailable in case of hydraulic system B failure.
Extension of the actuator closes a set of jaws onto theparachute D-ring and pulls the ripcord that releases
a spring-loaded pilot chute. The pilot chute functions
to pull the main canopy from deployment bag locatedin an aerodynamic fairing below the rudder.
Deployment below 90 knots may result in improper
deployment and damage to the drag chute.

The ANTI SKID caution light illuminates when the
LG handle is down and the switch is in OFF.

SPEEDBRAKE SYSTEM
The speedbrake system consists of two pairs ofclamshell surfaces located on each side of the engine
nozzle and inboard of the horizontal tail and is
powered by hydraulic system A. The speedbrakes
open to 60 degrees with the right MLG not down andlocked. With the right MLG down and locked,
speedbrake opening is limited to 43 degrees toprevent the lower surfaces from striking the runway
during landing. This limit can be overridden byholding the SPD BRK switch in the open (aft)position. When the NLG strut compresses on landing,the speedbrakes can be fully opened and remain fully
open without holding the SPD BRK switch.
SPD BRK Switch~ mIl The SPD BRK switch, located on the throttle,

is a thumb-activated, three-position slide switch. The
open (aft) position is spring-loaded to off (center) andallows the speedbrakes to be incrementally opened.
The closed (forward) position has a detent, allowing
a single motion to close the speedbrakes. To prevent
possible creeping, the switch should be left in theclosed position. ~ For SPD BRK switch differences,refer to F-16B AIRCRAFT, this section.

rID The speedbrake switches are connected in paralleland function so that either can override the other by

1-118

Change 11• Diagonals - Electrical power removed from theindicator. Diagonals also appear momentarily
during switching.

DRAG CHUTE SYSTEM ~ ~The drag chute system has safety provisions for
accidental deployments, both commanded anduncommanded:• Above 190 knots, the mechanical fuse section on the
D-ring fails and releases the chute.
• A deployed drag chute (or some residue thereof)
resulting from placing the DRAG CHUTE switch to
DEPLOY can be released by moving the switch
to [ru mIl NORMIREL or ~ REL at any airspeed.
• If the drag chute is deployed uncommanded (i.e.,
ripcord failure) and airspeed is above 60 knots, theD-ring pulls out of the jaw mechanism.
DRAG CHUTE Switch ~ ~Refer to figure 1-40. [ru mIl The DRAG CHUTE switch,
located on the MISC panel, is a two-position guarded
switch used to deploy and release the drag chute. The
switch is powered by the battery bus. ~ For DRAG
CHUTE switch differences, refer to F-16B AIRCRAFT, this section.

T. O. 1F- 16A- 1

Drag Chute Controls INEIINOI (Typical)

DRAG CHUTE
DEPLOY

2NOTE:

1. !AJIIEJ DRAG CHUTE Switch
2. [IB] DRAG CHUTE Switch[IB] For DRAG CHUTE switch,refer to F-16B AIRCRAFT, thissection.
1F-16A-1-1075X@

SWITCHPOSITIONNORM

FUNCTION
Chute stowed
Attachment jaws at midrange

DEPLOY

Hydraulic actuator extends
Ripcord pulled to release spring-loaded pilot chuteAttachment jaws close on D-ring

REL

Hydraulic actuator retracts for 2 seconds, releasing the D-ring from the attachment jaws
and then returns to midrange

Figure 1-40.

1-119

10. 1 F- 16A- 1

ARRESTMENT SYSTEM
The hook is electrically controlled and pneumaticallyoperated. Pneumatic pressure is supplied by the
LG/hook emergency pneumatic bottle which contains
sufficient pressure to lower the LG and hook.
When extended, pneumatic pressure holds the hookon the runway. When subsequently retracted, thehook rises enough to allow the cable to drop off thehook or to be disengaged. The hook is spring-loaded
partially up to allow taxiing over a cable. The hookmust be raised manually to reset it to the stowed
position.
HOOK Switch

The HOOK switch, located on the LG control panel,is lever-locked in the UP or DN position. Positioning
the switch to DN causes the hook to extend.
Returning the switch to UP partially retracts thehook, allowing for cable disengagement and for taxi
over the cable. 00 Either HOOK switch may be used
to extend the hook. Both switches must be positioned
to UP to raise the hook.

• When the throttle is at IDLE and MLG wheel speed
is greater than 60 knots groundspeed (the LEF's are
2 degrees up).
• LEF asymmetry brakes are locked.
• When the FLCS is operating on standby gains.
Refer to STANDBY GAINS, this section.
LE FLAPS Switch

The LE FLAPS switch is covered as a part of the
FLIGHT CONTROL panel.
LE FLAPS Caution Light

The LE FLAPS caution light, located on the caution
light panel, illuminates for dual air data signal
failures (mach number computations), dual AOA
failures, detection of LEF asymmetry, and detectionof flap command servo failures or when the LE
FLAPS switch is positioned to LOCK. The light may
also illuminate after main generator failure. Not all
LEF malfunctions result in illumination of the LE
FLAPS caution light.
TRAILING EDGE FLAPS (TEF'S) (FLAPERONS)HOOK Caution Light

The HOOK caution light, located on the caution light
panel, illuminates anytime the hook is not up andlocked.

WING FLAP SYSTEMLEADING EDGE FLAPS (LEF'S)

The LEF's consist of a spanwise flap on each wingleading edge controlled as a function of mach number,
AOA, and altitude by command signals from the ECA.
An asymmetry sensing and braking mechanism
prevents LEF asymmetry. If an asymmetry is sensed,

the LEF's lock and the LE FLAPS caution lightilluminates.
The LEF's are automatically programmed when the
LE FLAPS switch is in AUTO.
Exceptions are:• When weight is on both MLG (the LEF's are 2
degrees up).

1-120

The flaperons are located on the wing trailing edge
and function as ailerons and TEF's. The flaperons
have a maximum command deflection of 20 degrees
down and 23 degrees up. When acting as flaps, thedeflection is downward; when acting as ailerons, thedeflection is up or down, as commanded. Bothfunctions are operable whenever the FLCS ispowered. The TEF's are controlled as a function of the
LG handle position, the ALT FLAPS switch, airspeed,
and mach number. Positioning the LG handle to DN
or the ALT FLAPS switch to EXTEND causes the
TEF's to deflect downward and the FLCS to switch to
takeoff and landing gains. At all airspeeds below 240
knots, the TEF position is 20 degrees down. Above
240 knots, the TEF's reduce deflection as a functionof airspeed until nearly/fully retracted at 370 knots.
ALT FLAPS Switchrm

The ALT FLAPS switch is located on the !AIon
the FLCP, ~ on the LG control panel. With theswitch in NORM, the TEF's are controlled by the LG
handle and airspeed. Placing the switch to EXTEND
lowers the TEF's only, depending on airspeed. The
ALT FLAPS switch does not affect the operation of the
LEF's unless the FLCS is operating on standby gains.
Refer to STANDBY GAINS, this section.

T. 00 1F- 16A- 1

FLIGHT CONTROL SYSTEM (FLCS)
The FLCS is a computer-controlled, four-channel,
fly-by-wire system that hydraulically positions
control surfaces. Electrical signals are generated
through a stick, rudder pedals, and a MANUAL
TRIM panel. Redundancy is provided in electronicbranches, hydraulic systems, and power supplies.
The FLIGHT CONTROL panel (FLCP) provides
malfunction indications and controls.
Refer to figures 1-41 and 1-42. Command signals to
the FLee are initiated by applying force to the stickand rudder pedals. These signals are processed by the
FLCC along with signals from the air data system,
flight control rate gyros, and accelerometers. The
processed signals are transmitted to the ISA's of thehorizontal tails, flaperons, and rudder which are
positioned to give the commanded response.
Pitch motion is controlled by symmetrical movementof the horizontal tails. Roll motion is controlled bydifferential movement of the flaperons and horizontal
tails. Yaw motion is controlled by the rudder. Rollcoordination is provided by an ARI. The ARI function
is not available whenever MLG wheel speed exceeds
60 knots or if AOA exceeds 29 degrees. After takeoff,
ARI is activated within 2 seconds after the LG handle
is raised (spin down braking system). If the LG
handle remains down, 10-20 seconds are required for
the MLG wheels to spin down and activate ARI.
FLCS LIMITERS

Refer to figure 1-43 for limiter values. FLCS limiters
are provided in all three axes to help prevent
departures/spins,
AOA/G Limiter

Refer to figure 1-44. In cruise gains, the AOAlglimiter
reduces the positive g available as a function of AOA.
The negative g available is a function of airspeed.
Below 15 degrees AOA, the maximum positive gavailable is +9g. As AOA increases, the maximum
allowable positive g decreases. The positive g limit
and maximum AOA depend on the position of the
STORES CONFIG switch. In eAT I, positive gdecreases to a value of 19 at 25 degrees AOA.
Maximum commanded AOA is approximately 25.5
degrees. In eAT III, maximum AOA varies fromapproximately 16-18 degrees as a function ofGW andg. The negative g available above approximately 250
knots is -3g. Below 250 knots, the available negativeg decreases until below approximately 100 knots,
where the maximum negative g available is zero g.
In takeoff and landing gains, the STORES CONFIG
switch has no effect on limiting or gains. Maximum
positive g is a function of airspeed and AOA. The
negative g command limit is not a function of
airspeed. It is a fixed limit. The maximum AOA for 19
is approximately 21 degrees.
In inverted or upright departures, the AOAlg limiter
overrides stick pitch commands if the MPO is notengaged. The MPO can always override the negative
g function of the limiter. It can also override the AOA
function of the limiter when the AOA is above 29
degrees. Refer to MPO, this section.
Roll Rate LimiterIn cruise gains, the roll rate limiter reduces available
roll rate authority to help prevent roll coupled
departures. This authority is reduced as airspeed
decreases, AOA increases, or trailing edge down
horizontal tail deflection increases. In takeoff andlanding gains, roll rate limiting is available but is a
fixed value independent of AOA, airspeed, or
horizontal tail position,
Rudder Authority LimiterIn cruise gains, the rudder authority limiter
reduces the pedal commanded rudder deflection as
a function ofAOA, roll rate, and STORES CONFIG
switch position for departure protection. However,
ARI authority, stability augmentation, and trim
authority are not reduced. In takeoff and landing
gains, category I rudder authority limiting isprovided.
Yaw Rate Limiter

When AOA exceeds 29 degrees, the yaw rate limiter
overrides stick roll inputs. The yaw rate limiter
provides rudder against and flaperon with the yaw
rate until AOA is below 29 degrees to enhance spin
resistance. The yaw rate limiter is not functional for
inverted departures; below 29 degrees ADA, the yaw
rate limiter provides no protection against yaw
departure.

1-121

T. O. 1 F- 16A- 1

FLCS Functional Schematic (Typical)
FLIGHT CONTROL PANELCOMMANDS• Cockpit• Self-Test-• SWITCHES• INDICATORS• STBY GAINSMANUAL TRIM PANEL
TRIM INPUTS• Side Panel - - - -..
~• TRIM CONTROLS• Stick• TRIM/AP DISCAIRCRAFT MANUAL TRIMMANUAL PITCH..."'1 MANUAL PITCH OVERRIDE• Override ------+t
SWITCH PANELIFLCS POWER INPUTS ~DC POWERCONTROL INPUTS• Stick• Rudder Pedals

QUADRUPLE FORCESENSORS

R:T~~I• Pitch

•

..

I

~

QUADRUPLE FLIGHT CONTROL ~SYSTEM INVERTERS

3-AXIS COMMANDSTEF COMMAND--QUADRUPLE RATESTABILITY/COMMAND
AUGMENTATIONAOA LIMITER------+4 _S_E_N_S_O_R_S_ _ _ _ _-----'
L-.

• Yaw

ACCELERATION _ _ _....1 QUADRUPLEACCELEROMETERS• Normal• Lateral
AIRFLOW
.AOA

~

~~

AOA

DUAL AIRFLOW SENSORS- -CONVERTER-----------I- AIR-DATAPNEUMATIC SENSOR
ELECTRONIC COMPONENT.I~ASSEMBLY (PSA)r
ASSEMBLY (ECA)I
L -------------~-.

AIR DATA• Total Pressure• Static Pressure·AOA

1--- -~

AIR DATA• Total Pressure• Static Pressure• Total liemperature~-...,

j~-- 1 - - - ~---+CENTRAL AIR DATACOMPUTER (CADC)

I

J

I
I

~_.....J

AOA~

ALTITUDE DATANAV INPU TS• Roll Angle• Pitch Angle

1 F-16A-1-1076-1A@

Figuree 1-41. (Sheet 1)
1- 122

Change 1

T. O. 1F- 16A- 1

WARNING and CAUTION
LIGHTS,,RIGHT HORIZONTAL TAIL
INTEGRATEDSERVOACTUATOR
(lSA)

I

FLIGHT CONTROL COMPUTER
I - - - -...
~LEFT HORIZONTAL TAIL
INTEGRATEDSERVOACTUATOR(ISA)

t---"'

RIGHT FLAPERONINTEGRATEDSERVOACTUATOR(lSA)

LEFT FLAPERONINTEGRATEDSERVOACTUATOR(lSA)• QUADRUPLE ELECTRONICS• CONTROL DYNAMICS• GAIN SCHEDULING• QUADRUPLE SELECTORS• AC POWER MONITORS• AILERON-RUDDER
INTERCONNECT (ARI).

r• STRUCTURAL FILTERS• YAW RATE LIMITING• HIGH AOA ROLL RATE
LIMITING

RUDDER

AUTOPILOT COMMANDS

1------'" I NTE GRATE DSERVOACTUATOR(ISA)GAIN SCHEDULINGSENSOR POWER AND TEST . -______________________________________________~~IDUAL LEF COMMAND I. SERVOS.

AIR DATA• Static Pressure

ANGLE OFSIDESLIPISIDESLIP PRESSURE SENSORSLEF ACTUATIONSYSTEM

LEGEND:- - - ELECTRICAL- - - - - - - MECHANICAL
1F-16A-1-1076-2A@

Figure 1-41. (Sheet 2)
Change 3

1-123

T. O. 1F- 16A-1

FLCS Pitch, Roll & Yaw Schematic (Typical)

I PITCH I

I

1--FLIG~CONmmCOMPUTER--1

I.
i

I

I, -

ROLL RATEII

I...-------.1

I

Vr
4

,I

T

I ACCELERATION I L_ l
ROLL

II

I ROLL TRIM

PRESSURE

IL,-~

r-

AILERON STICK FORCE

LIMIT

I

I ROLL RATE ll-----+--L~..,-.

I
I'-----------' I
GUN
COMPENSATIONII

I GUN FIRE I t - - - - - - t - 'II
...

yAW

I

II

YAW TRIM

AOA__I----J

ISA

-

~A
I

ISA

-

ISA

--1 ~ON I

1'1

~
r-

SELECTOR

A "

I
I
I

AO~-+-+.........
SELECTOR .........-40--_~,
.~
'-0

~

~

TURN
~O~~"'_I----I
TE FLAP--+ COORD - -<>-]..0
4~
~ 29"
~~
COMMAND'1-

II....--~ll MECH LIMIT

I
I

~.

1 -.: LIMIT

11-+-____...

L _____

I
,

ISA

-

@;J
~RON

NOTE:

FLCS lains scheduled byair data inputs.
-----,LEGEND
ELECTRICAL- - - - MECHANICAL~A SERVO

L;/ AMPLIFIER

,

~

.•

..,!.. ....

I
I
I,

•

'------'~AILERON.... SELECTOR ARUDDERI
INTERCONNECTI

1..- - - -.J

LEFT HORtz
TAIL

,

~

I YAW RATE 1t---+-tI,TI--e-----.c YAW AXISLATERALI!
l ~ACCELERATION I I

I""""R-IG-HT-H-OR-IZ"""
TAIL'-------',L. -I RUDDER PEDAL FORCE 1

I

PITCH COMMAND

AUTOPILOT ROLL

tH

~I

h/j-_.
I_

MECH LIMIT :1-++-+----....

J

-

f--

.----+-_':.=====::~,.

II.. rIMPACT,.....-____---, r - . . . .1 ,I..... SELECTOR

SELECTOR ~ STAB ILlTY/
COMMAND
AUGMENTATIONII

SELECTOR J,--MANUALPITCH~OVERRIDE

I~_ _ _ _+--.-t

NORMAL......

J

~-

I _..

I

PITCH RATE

-

I
I

AOA

-

. -I INTEGRATOR

I

STICK
LIMIT

hI ,I

..J PITCH

AUTOPILOT
PITCH
I

MECH LIMIT "I HORIZ TAIL STICK FORCE I
I

I

LL

PITCH TRIM:

SELECTOR

~

r:=l
ISA

--

~

I
I
_________

~29·
~

~

1F-1SX-1-0009xe

Figure 1-42.

1-124

T. O. 1F- 16A- 1

FLCS Limiter Functions
ROLL AXISYAW AXIS

Maximum AOA=25°

Maximum roll rate commanddecreases with:

Maximum deflection (pedal command) reduced for:g command system until 15° AOA• AOA above 15°• AOA> 14° (zero roll rate)• Airspeed less than 250 knots• Roll rate>200/sec

PITCH AXIS

CAT
I

g/AOA command system above
15° AOA

Maximum AOA= 16°-18°
pending on GW)

CAT
III

(de-

g command system until 7° AOA
at 100 knots to 15° AOA at 420
knots and above• Horizontal tail deflection morethan 5° trailing edge down

Maximum roll rate command
reduced by approximately 40 percent of CAT I authority. Additional
decreases as function of AOA, airspeed, and horizontal tail position

NOTE:
Zero rudder authority available at
26° AOA.
Maximum deflection (pedal command) reduced for:
• AOA>3° (zero roll rate)• Roll rate>200/sec

NOTE:
Zero rudder authority available at
15° AOA.

g/AOA command system above
these values

NOTES

1. In takeoff/landing gains, theFLCS operates as a pitch ratecommand system until 10°
AOA and a pitch rate/AOA
command system above 10°
AOA.
2. +9g available until 15° AOA.
Maximum g decreases as afunction of AOA and airspeed.

1. In takeoff/landing gains, maximum roll rate is fixed at
approximately one-half the
maximum roll rate available incruise gains, regardless of
AOA, airspeed, or horizontal
tail deflection.
2. Above 29° AOA, the yaw rate
limiter cuts out stick roll commands and provides roll axis
antispin control inputs.

1. Above 29° AOA, the yaw rate
limiter provides yaw axis
antispin control inputs.
2. Above 29° AOA with MPO
engaged, pedal-commanded
rudder deflection is possible.
3. Maximum deflection (30°)
always available thru ARI andstability augmentation.

Figure 1-43.
Change 1

1-125

I

I

T.O. 1F-16A-1

AOAlG Limiter Function (Cruise Gains)

10

I

I

I

9

.....

.....

... r .....

CAr I SELECTED~..... 1-...

8

r- .....
I-r---t

I

II

CJ

I

I 7

a::

...0CJ
~

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toI-

If

6

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;--r--

1\
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....

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.... 5

.....

,

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II

<C
0

r--.

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II

I

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-

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=

'

I

,

I '

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I

I

1\

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I

:i

I

!\

I

<C

I

1\

CAT III SELECTED
GW
25,000 POUNDS

1

I'

i

I

\

17"-1-..

'

I
,

\
I\.

a::

0 4
2

,I II, I

J

3

CAT III SELECTEDGW = 30,000 POUNDS

I

IJ

'\

\.

I

II

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,...--

II
2

L\
:\.
""'i-- CAT III SELECTED= 35,000 POUNDSII I II I I I

II
, I I
I III

I

J

1

15

17

1\

I

I

16

,
\

GW

I I

18

1\

19

20

21

22

23

24

25

26

AOA - DEGREES

1F-16X-1-0010X®

Figure 1-44.

1-126

10.lF-16A-l
FLCS GAINS

During normal operation, the FLCS receives inputs(gains) from the ADC and provides relatively
constant aircraft response for a given stick input,
regardless of altitude or airspeed. This response
varies slightly depending on configuration. In theevent of a dual air data failure, the FLCS switches tostandby (fixed) gains.
Cruise Gains

The FLCS is in cruise gains with the LG handle in
UP, the ALT FLAPS switch in NORM, and the AIR
REFUEL switch in CLOSE. At low AOA, the pitch
axis of the FLCS is a g command system. As AOA
increases, the FLCS switches to a blended g and AOA
system to provide a warning of high AOAllow
airspeed. Roll rate limiting is available andmaximum roll rate decreases as a function of low
airspeed, high AOA, and horizontal tail position.
Takeoff and Landing Gains

The FLCS is in takeoff and landing gains with the LG
handle in DN, the ALT FLAPS switch in EXTEND,
or the AIR REFUEL switch in OPEN. In takeoff andlanding gains, the FLCS pitch axis operates as apitch rate command system until 10 degrees AOA
and a blended pitch rate and AOA command system
above 10 degrees AOA. Roll rate limiting is available
but is a fixed value independent of AOA, airspeed, or
horizontal tail position.

The STBY GAINS light cannot be reset in flight, andthe FLCS continues to operate on fixed gains even if
the failure clears. The light may reset when weightis on the MLG.
FLCS DATA RECORDER [!] ~The FLCS data recorder is attached to the ejection
seat and departs the aircraft on ejection. It retains thesame information as the FLCC and ECA including
FLCS failure data, control surface position data,
airspeed, altitude, AOA, and elapsed time fromtakeoff. A frame of data can be stored in the FLCS
data recorder anytime in flight by depressing the FCSCAUTION RESET button. This information isparticularly valuable in determining the origin ofabnormal FLCS response. The FCS CAUTIONRESET button should be used sparingly under thesecircumstances since memory capacity is limited.
GUN COMPENSATIONThe FLCS automatically compensates for the offcenter gun and the aerodynamic effects of gun gas
emissions during firing by moving the flaperons andrudder. Gun compensation is optimized for 0.7-0.9
mach range; therefore, all excursions may not be
eliminated. For example, gunfiring at low mach may
result in nose left excursions while nose right
excursions are likely at higher mach. Failure
monitoring of gun compensation circuits is notprovided and there are no caution light indications forincorrect compensation.
FLIGHT CONTROL SYSTEM (FLCS) CONTROLSStandby Gains•

In standby gains, control response is tailored for a
fixed altitude (sea level, standard day) and one of two
airspeeds. 600 knots is used when the LG handle is in
UP, the ALT FLAPS switch is in NORM, and the AIR
REFUEL switch is in CLOSE. 230 knots is used whenthe LG handle is in DN, or the ALT FLAPS switch isin EXTEND, or the AIR REFUEL switch is in OPEN.
The following lights illuminate: STBY GAINS, ADC,LE FLAPS, and FLT CONT SYS.
When operating on standby gains, the LEF's are atzero degrees with the LG handle in UP and the ALT
FLAPS switch in NORM. The LEF's deflect 15
degrees down with the LG handle in DN or the ALT
FLAPS switch in EXTEND. The operation of the
TEF's is not affected in standby gains.

Stick

Refer to figure 1-45. The stick is a force-sensing unit
which contains transducers in both pitch and rollaxes, moves approximately 114 inch in both axes, andis rotated slightly cwo
Maximum noseup and nosedown pitch commands are
generated by 25 and 16 pounds of input, respectively.
Roll commands are generated by a maximum of 17
pounds in cruise gains and by 12 pounds in takeoff
and landing gains. When using the switcheslbuttons
on the stick, inadvertent inputs to the FLCS are
possible.
The wristrest and armrest assemblies which may be
used in conjunction with the stick are located on theright side wall aft of the stick.

Change 7

1-127

T. O. 1F- 16A- 1

Stick (Typical)
NWS

AIR DISCMSL STEPCAMERAl.

GUNBLOCK 10
1. NWS A/R DISC MSL STEP Button
2. TRIM Button (4-Way, Momentary)
3. BLOCK 10 DESIG RET SRCH
Switch (2-Way, Momentary)
BLOCK 15 DSG RS DO DS
Switch (4-Way, Momentary)
4. WPN REL Button
5. Paddle Switch
6. [AJ~ Chaff/Flare Dispense Button(if installed)
7. CAMERA/GUN Trigger (2-Position)

BLOCK 101 F-16A-1-0034A @CONTROL

FUNCTIONPOSITION

1. NWS AIR DISC MSL(NWS)Depress (on ground)

Activates NWS

Depress (2nd time)

Deactivates NWS

Depress (in flight)

Disconnects boom latching. AIR REFUEL
switch must be in OPEN position

Depress (in flight)

Activates missile step function. Refer toT.O. IF-16A-34-1-1, ~ T.O. IF-16A-341-3 or ~T.O. IF-16A-34-1-4 for a detailed
description of switch functions(NOSE DOWN)

Fwd

Trims nosedown(NOSE UP)

Aft

Trims noseup

(LWD)

Left

Trims left wing down

(RWD)

Right

Trims right wing down

STEP Button(AIR DISC)(MSL STEP)I
2. TRIM Button(4- way,
momentary)

Figure 1-45. (Sheet 1)

1-'28

Change 13

10. 1F- 16A-1

Stick (Typical)
CONTROL
3. DESIG RET SRCH
Switch Block 10

POSITION(DESIG)(RETSRCH)

Refer to T.O. 1F-16A-34-1-1, ~ T.O.
1F-16A-34-1-3 or 1M T.O. 1F-16A-34-1-4 fora detailed description of switch functions

Up

I

Down

(DSG)

Up

(RS)

Down

(DO)

Left

~

(DS)

Right

1m

(DS)

Right

3. DSGRS DO DS
Switch Block 15

FUNCTION

4. WPN REL Button

Depress

Signals consent to FCC or SMS to initiate
weapon release and operates AVTRlCTVS
for 30 seconds when in AUTO

5. Paddle Switch

Depress

Interrupts the autopilot while switch IS
depressed

rn:J For stick override function, refer to
F-16B AIRCRAFT, this section
6.

~ ~ ChaffIFlare Dispense
Button (if installed)

7. CAMERA/GUN Trigger(2-position)

Depress

Activates chaff/flare dispensers. Refer toT.O. 1F-16A-34-1-1, ~ T.O. IF-16A-34-1-31
or ~T.O. 1F-16A- 34-1-4

Squeeze trigger todetentStarts operation of AVTRlCTVS with the
HUn CAMERA switch in AUTO

Squeeze trigger pastdetentFires gun (if selected and armed) and AVTRICTVS operation continues (camera operation continues for 30 seconds after trigger isreleased)

Figure 1-45.

(Sheet 2)

Change 13

1-129

T.O. 1 F-16A-l
Rudder Peda Is

The rudder pedals are force-sensing units containing
transducers. Force on the applicable rudder pedal
produces electrical yaw command signals.
The rudder pedals are also used to generate brakeand NWS signals. Rudder pedal feel is provided bymechanical springs.
MANUAL TRIM Panel ~ rnfl

I

I

• Combined airspeed and pitch angle fall on a point
within the tone on area with LG handle up and ALT
FLAPS switch in NORM.

I

The low speed warning tone has priority over the LG
warning horn. Depressing the HORN SILENCER
button silences the low speed warning tone.
The low speed warning tone is reactivated only after
the original warning condition is eliminated. The
MAL & IND LTS test button does not test the low
speed warning tone.

Refer to figure 1-46. The MANUAL TRIM panel,
located on the left console, contains trim controls
and indicators.

FLIGHT CONTROL Panel (FLCP) ~MANUAL PITCH Override (MPO) Switch

Refer to figure 1-50. The FLCP, located on the left
console, contains indicator lights and controls related I
to flight control functions.

mEl

Refer to figure 1-47. The MPO switch has two
positions, NORM and OVRD, and is spring-loaded to
the NORM position. This switch is used during a deep
stall condition to enable manual control of thehorizontal tails. Positioning and holding the switch to
OVRD overrides the negative g limiter. If AOA
exceeds 29 degrees, the OVRD position overrides the
AOA/g limiter and allows rudder inputs.

The instrument panel, right auxiliary console, the
FLCP, ELEC control panel, and TEST switch panel
contain warning, caution, and indicator lights related
to the FLCS.

STORES CONFIG Switch ~ mEl

T.O./LAND CON FIG Warning Light

Refer to figure 1-48. The STORES CONFIG switch,
located on the LG control panel, has two positions,
CAT I and CAT III. The CAT III position shall beselected when the aircraft is configured with acategory III loading. LESS rn Category II loadings
can be flown with the STORES CONFIG switch ineither CAT I or CAT III. However, with the switch in
CAT I, refer to AOA AND ROLLING LIMITATIONS,
Section V, as AOA monitoring is required to observe
the category II maneuver limits.
AOA limiting is provided. Refer to FLCS LIMITERS,
this section, for a description of categories I and III
AOA limiter.

FLCS WARNING, CAUTION, AND INDICATOR
LIGHTS

Refer to LANDING GEAR SYSTEM, this section.
FLT CO NT SYS Caution LightThe FLT CONT SYS caution light, located on thecaution light panel, illuminates when a failure occurs
in the FLCS. Associated FLCP indicator lights also
illuminate.
With WOW, if a SERVOS light illuminates with the
FLCS not in self-test, the FLT CONT SYS caution
light cannot be reset until the ISA SERVOS is reset.
DUAL FC FAIL Warning Light

Low Speed Warning Tone

Refer to figure 1-49. A low speed warning tone(steady) sounds in the headset when either of thefollowing conditions exists:·I

AOA is 15 degrees or greater with LG handle down
or ALT FLAPS switch in EXTEND.

1-130

The DUAL FC FAIL warning light, located on theright upper edge of the glareshield, illuminates toindicate that a dual malfunction has occurred in oneof the electrical control axes, in an ISA, or in the AOA
portion of the air data system. The FLT CONT SYS
caution light and associated indicator lights on the
FLCP also illuminate.

T. O. 1F- 16A- 1

AUTOPILOT
The autopilot provides attitude hold and heading
select in the roll axis and attitude hold and altitude
hold in the pitch axis. These modes are controlled by
PITCH and ROLL switches on the MISC panel. The
TRIMIAP DISC switch on the MANUAL TRIM panel
disengages the autopilot. The paddle switch on the
stick interrupts autopilot operation while the switch
is held depressed.
The AUTOPILOT switch is solenoid held in ON andreturns to OFF if any of the following conditions
occur:• AIR REFUEL switch - OPEN.
• ALT FLAPS switch - EXTEND.
• LG handle - DN.
• STBY GAINS light - ON.
• TRIMIAP DISC switch - DISC.
The ROLL and PITCH switches are two-position
switches which allow mode selection. Any combination may be selected.
AUTOPILOT OPERATION

Positioning the AUTOPILOT switch to AUTOPILOT
permits use of roll and/or pitch autopilot functions.
Autopilot options are selected by positioning thePITCH switch (ATT HOLD or ALT HOLD) and/or the
ROLL switch (ATT HOLD or HDG SEL).
Stick trim is inoperative with the autopilot engaged.
Manual trim is operable and may be used while theautopilot is engaged. However, due to the limited
authority of the autopilot, engagement of any mode inother than a trimmed flight condition degrades
autopilot performance.
There are no caution light indications of autopilot
operation, malfunction, or disconnect. The autopilot
does not include the redundancy of the FLCS so its
use must be closely monitored at low altitude or inclose formation.
Positioning the PITCH switch to ALT HOLD enables
the FLCS to use CADC information to generatecommands to the horizontal tails which result in theaircraft maintaining a constant altitude. The FLCS
limits the pitch command to +0.5g-+2g. Engagementof altitude hold at rates of climb or dive less than 2000
fpm selects an altitude within the pitch command glimits. Engagement above rates of2000 fpm causes nounsafe maneuvers; however, the engaged altitude
may not be captured. Control accuracy of ± 100 feetis provided to 40,000 feet pressure altitude for normal
cruise conditions. The altitude reference may bechanged by depressing the paddle switch, changing
altitude, and releasing the paddle switch. ALT HOLD
in the transonic region may be erratic.
Positioning the PITCH switch to ATT HOLD routes
an attitude signal from the INU to the FLCS which
results in the aircraft maintaining the selected pitch
attitude. This mode does not function if pitch angle
exceeds ± 60 degrees; however, the PITCH switch
remains engaged.
Positioning the ROLL switch to HDG SEL allows the
FLCS to use a signal from the HSI to maintain theheading set on the HSI. Adjusting the HSI headingreference marker to the aircraft heading prior to
engagement maintains the existing aircraft heading;otherwise, when the autopilot is engaged with the
ROLL switch in HDG SEL, the aircraft turns tocapture the heading indicated by the headingreference marker on the HSI. The roll command does
not exceed a 30-degree bank angle or a 20-degree/second roll rate. This mode does not function if bankangle exceeds ± 60 degrees; however, the ROLL
switch remains engaged.
Positioning the ROLL switch to ATT HOLD routes anattitude signal from the INU to the FLCS which
results in the aircraft maintaining the selected roll
attitude. This mode does not function if bank angle
exceeds ± 60 degrees; however, the ROLL switch
remains engaged.
STICK STEERING

Stick steering is operable only with the pitch and roll
attitude hold modes. Stick steering operation isaccomplished by applying force to the stick. With A'fT
HOLD selected, a force applied in the appropriate
axis large enough to activate stick steering causes theau topilot to drop the selected reference and thesystem accepts manual inputs from the stick.

1-133

10. 1F- 16A- 1

FLIGHT CONTROL Panel [8JlBFl (Typical)

;
MISC
Panel

1'+-----"'-- 6_-----;.----7~AL-T-FL-AP-S---,.,...-- 8---+!--9

1. HORIZ Tail SERVOS Status Lights
2. HORIZ Tail SERVOS ARM Switch
3. FLAPERON SERVOS Status Lights
4. FLAPERON SERVOS ARM Switch
5. RUDDER SERVOS Status Light
6. RUDDER SERVOS ARM Switch
7. FCS CAUTION RESET Button
8. SERVO ELEC RESET Switch
9. ALT FLAPS Switch (Lever Lock)
10. STBY GAINS Light
11. LE FLAPS Switch
12. P, R, and Y Status Lights
13. MON Receptacle
14. Test Number
15. Test ADV MAL Lights
16. SELF TEST Switch
17. ADV SLEW Switch
18. AUTOPILOT Switch
19. Autopilot ROLL Switch
20. Autopilot PITCH Switch
21. FLCC Data Word Dot Light
22. ECA Data Word Dot Light
23. Spare
24. Rate Gyro Speed Detect Dot Light

21

22

1 F-16A-1-1 084X @FUNCTIONCONTROUINDICATORPOSITIONIINDICATIONL HORIZ Tail SERVOS Status LightOn (amber)

Indicates first failure of left horizontal ISA

R HORIZ Tail SERVOS Status LightOn (amber)

Indicates first failure of right horizontal ISA

HORIZ Tail SERVOS ARM
SwitchLARM

Arms servo monitor. Subsequent failure in
left ISA positions left horizontal tail to neutral, locks out roll commands to the right horizontal tail, and illuminates the DUAL FC
FAIL warning light

CenterNormal position

RARM

Arms servo monitor. Subsequent failure inright ISA positions right horizontal tail toneutral, locks out roll commands to the left
horizontal tail, and illuminates the DUAL
FC FAIL warning light

L FLAPERON SERVOS
Status LightOn (amber)

Indicates first failure of left flaperon ISA

R FLAPERON SERVOS
Status LightOn (amber)

Indicates first failure of right flaperon ISA

1.

2.

3.

Figure 1-50. (Sheet 1)
1-134

10.lF-16A-l

FLIGHT CONTROL Panel [A] rn:EJ (Typical)
4.

FUNCTIONCONTROUINDICATORPOSITIONIINDICATIONFLAPERON SERVOS ARM
SwitchL ARM

AnTIs servo monitor. Subsequent failure in
left ISA positions left flaperon to neutral,locks out TEF commands to both flaperons,
and illuminates the DUAL FC FAIL warning
light

CenterNormal position

I RARM

AnTIs servo monitor. Subsequent failure inright ISA positions right flaperon to neutral,locks out TEF commands to both flaperons,
and illuminates the DUAL FC FAIL warninglight

On (amber)

Indicates first failure of rudder ISA

5.

RUDDER SERVOS Status
Light

6.

RUDDERSERVOSARM
SwitchAnTIs servo monitor. Subsequent failure of
ISA positions rudder to neutral and illuminates the DUAL FC FAIL warning light
Not Armed (center)

Normal position

7.

FCS CAUTION RESET
ButtonPushResets FLT CONT SYS caution light so that
subsequent failures can be indicated andgives consent to allow SERVO reset. Also
resets MASTER CAUTION light caused by
FLT CONT SYS caution light illumination.
Depressing FCS CAUTION RESET buttonin flight also causes the FLCS data recorderto store a frame of data

8.

SERVO ELEC RESET
Switch (momentary)

SERVO

Resets all failed I SA's when the FCSCAUTION RESET button is depressed
simultaneously (and illuminates five SERVOS lights to test bulbs)

Center (spring-loaded)

Normal position

ELEC

Resets malfunctioning P, R, Y, ADC, LE
FLAPS, and CADC electronics. Also resets
FLT CONT SYS and MASTER CAUTION
lights

NORM

TEF's operation controlled by LG handle

EXTENDTEF's extend regardless of LG handle position

9.

ALT FLAPS Switch(lever lock)Figure 1-50. (Sheet 2)

1-135

T.O. 1 F-16A-l

FLIGHT CONTROL Panel [A] mE] (Typical)
FUNCTIONCONTROUINDICATORPOSITIONIINDICATION

10.

STBY GAINS LightOn Camber)

Indicates FLCS is operating on standby
gains

II.

LE FLAPS Switch

AUTO

LEF's are automatically controlled as a function of mach number, altitude, and AOA

LOCK. Locks LEF's in present position and illuminates LE FLAPS caution light

P Camber)

Indicates signal malfunction in pitch control
electronics

R Camber)

Indicates signal malfunction in roll control
electronics

12.

P, R, and Y Status Lights

Y Camber)I Indicates signal malfunction in yaw control
electronics

P, R, and Y (amber)i

Indicate loss of power to one or more FLCS
branches13.

MON Receptacle

CappedConnection for external test equipment

14.

Test Number

00 Camber)

Indicates step number of test being performed

15.

Test ADV MAL Lights

MAL Camber)If accompanied by a dot light, indicates a malfunction in the FLCS was detected during
self-test

ADV (green)

Indicates a stop in test program which
requires manual advance

TEST

Initiates FLCS self-test

OFF

Deenergizes self-test circuits

ADV

Advances the test program after a stop in test
procedure

SLEW

Advances the test sequence approximately
one test number per second

AUTOPILOT

Engages autopilot which provides attitude
hold, altitude hold, and heading select functions

OFF

Disengages autopilot functions

16.

17.

18.

SELF TEST Switch

ADV SLEW Switch

AUTOPILOT Switch

Figure 1-50. (Sheet 3)

1-136

10.lF-16A-l

FLIGHT CONTROL Panel lAl [DE] (Typical)
19.

20.

FUNCTIONCONTROLIINDICATORPOSITIONIINDICATION

Autopilot ROLL Switch

HDGSEL

Autopilot turns the aircraft to capture andmaintain the heading selected by the heading reference marker on the HSIATTHOLD

Autopilot maintains
determined by INS

ALTHOLD

Autopilot maintains constant altitude as
determined by CADCATTHOLD

Autopilot maintains constant pitch attitude
as determined by INS

Autopilot PITCH Switch

roll

21.

FLCC Data Word Dot LightOn (amber)

Indicates failure in FLCC22.

ECA Data Word Dot LightOn (amber)

Indicates failure in ECA

23.

Spare

-

-

24.

Rate Gyro Speed Detect Dot
Light

On (amber)

Indicates rate gyro failure

attitudeasFigure 1-50. (Sheet 4)

1-137

T.O IF-16A-l

AOASYSTEM
The system consists of two AOA transmitters locatedon each side of the nose radome, AOA porta on thefuselage-mounted air data probe. an ADe, an AOA
correction device in the CADC, an AOA indexer. anda vertical scale AOA indicator. In flight, the airflow
direction is sensed by the conical AOA probes and the
ADA ports of the fuselage-mounted air data probe.
The AOA signals from all three sources are sent to theADe for comparison and correction to true AOA
which is used for LEF scheduling. A true AOA signal
is sent from the ADC to the FLCe and is used forstabilization and AOAlg limiting. The CADC converts
an indicated AOA signal received from the ADe to 8
second true AOA signal for use by the AOA indexer,
the AOA indicator, and other avionic equipment.

AOA Indicator
Refer to figure 1-51. TheAOA indicator, located on the
instrument panel, displays actual AOA in degrees.
The indicator has a vertically moving tape display
indicating an operating range of -5 to approximately+32 degrees. The tape is color coded from 9-17 degrees
to coincide with the color-eoded symbols on the AOA
indexer.
AOA Indexer

Refer to figure 1·51. The AOA indexer, located on the
top left side of the glareahield, consists of threecolor-coded symbols arranged vertically. The indexer
provides a visual bead-up indication of aircraft AOA
by illuminating the symbols individually or in
combinations as shown. The indexer ligbts display
AOA correction (based on approximately 13 degrees
AOA). This correction may be used during landing
approaches as visual direction toward optimum
landing AOA. The AOA indexer operates continuously with the LG handle up or down.
A dimming lever, located on the left side of theindexer, controls the intensity of the lighted symbols.

The indexer lights are tested by activation of the MAL& IND LTS switch on theTESTswitch panel. The test
should be performed with the dimming lever in thebright position.
HUD AOA Display

Refer to figure 1-51. The HUD AOA bracket andfligbtpath marker provide a visual head-up indication of aircraft AOA. The fligbtpath marker alignedwith the top oftbe bracket indicates 11 degrees AOA.
The fligbtpath marker centered on the bracket
indicates 13 degrees AOA. The flightpath marker
aligned with the bottom of the bracket indicates 15
degrees AOA. The HUD AOA display is only available
with the NLG lowered.
I

AOA Displays
INDICATORINDEXER•q.. ~il

~•

IS

•

~

~
~

13

Q

~

•

•

~•

~

II

HUD DISPLAY

L
[-6-

~'LOW

~~S----~';:. ,:>- ON .".'0

~O~

.0-

I;

~1>- FAST_.

~

ATTITUDE

Ia3sr~~~-

Figu11! 1-51.

1-138

T.O.1F-16A-1

AIR DATA SYSTEM

Static Pressure Ports

Refer to figure 1-52. The air data system uses probes
and sensors to obtain static and total air pressures,
AOA, sideslip, and air temperature inputs. These air
data parameters are processed and supplied tovarious systems.

Two flush-mounted static pressure ports used formeasuring sideslip are located on the fuselage left
and right sides aft of the forward equipment bay
doors. These two ports provide inputs to a differential
pressure sensor for angle-of-sideslip measurement.
The measurement is also used to compensate thethird AOA source error.
Probe Heat Monitor ~I

Proper AOA transmitter and fuselage air data probe
operation is essential for safe flight operation.
Interference from foreign objects (especially ice,internal or external) or improperly installed AOA
transmitters can result in erroneous AOA data at
weight off wheels. Reporting of false high AOA
concurrently from two sources can cause the FLCC tocommand full nose down pitch which is impossible forthe pilot to stop. Ground use of probe covers protects
the system from foreign objects and moisture
intrusion. Ice on/in the probes is eliminated by usingprobe heat prior to takeoff.
Ai r Data Probes

Two air data probes provide data inputs to the air
data system. One air data probe (pitot probe) ismounted on the nose and provides a dual source ofstatic and total pressure. The other air data probe is
mounted on the forward right side of the fuselage andprovides a source of AOA, sideslip, static pressure,and total pressure.

The probe heat monitor monitors current flow to thepitot, fuselage air data, and AOA probes (total
temperature probe current is not monitored). If thecurrent flow decreases below a certain value, themonitor illuminates the PROBE HEAT caution light.
The monitor operates anytime the aircraft isairborne, regardless of the PROBE HEAT switch
position.

•

PROBE HEAT Switch

The PROBE HEAT switch is located on the TEST
switch panel. The pitot, fuselage air data, AOA, andthe total temperature probe heaters are on anytime
the aircraft is airborne, regardless of the PROBEHEAT switch position.
Functions are:
• PROBE HEAT - On the ground, this positionenergizes the pitot, fuselage air data, AOA, and thetotal temperature probe heaters ~ and the probe •heat monitor.
• OFF - On the ground, circuits deenergized.

AOA Transmitters•

The AOA transmitters are mounted on each side of
the radome and each provides four signals to the ECA
proportional to local AOA. The probe of thetransmitter protrudes through the radome to alignwith the airstream.

• ~ TEST - On the ground and in flight, this position •performs a functional test of the probe heatmonitoring system. The PROBE HEAT caution
light flashes 3-5 times per second. If the caution
light does not illuminate or if it illuminates but does
not flash, the probe heat monitoring system is
inoperative. The test feature does not verify proper
operation of the probe heaters.

Total Temperature Probe

AIR DATA CONVERTER (ADC)

The total temperature probe provides the CADC withan analog signal which is required for true airspeed
and air density computation. The probe is located onthe ~ underside of the right forebody strake,
LESS ~ the left side of the engine nacelle air inlet.

The ADC is comprised of the PSA and ECA. The PSA
contains sensors which convert pneumatic inputs
from the pitot probe and the fuselage air data probe
pressure ports into electrical signals. The PSA
supplies static and impact pressure signals andsingle AOA signals to the ECA.

Change 51-139

10. 1F- 16A- 1

Air Data System Schematic (Typical)
FLUSH STATIC PORTrRlGHTSiDE I
I
0

I
I

0

0

0

I

0

L ___

AOA TRANSMITTER
LEFT/RIGHT SIDE

I
PSR

0

SIDESLIP DIFFERENTIALPRESSURE SENSORCAUTION LIGHT PANEL--JVERTICAL VELOCITY INDHEMISPHERICAL(Portion of fuselagemounted air data probe)

----l

ECA

L--l
Pa 2
CE~====:===tt$===Et;:z::j AI R DATACONVERTER
PT3
(ADC)

FUSELAGE~

I
I

MOUNTED
AIR DATA
PROBE

~~~~
Pa 1

L ___ _

PSA

I
I
I

l_

FLUSH STATIC PORT'LEFT S I D E - i

I
I

0

00

AIRSPEED/MACH IND

1
L _ _o_~FLCS
COMPUTERCAUTION LIGHT PANELAOA INDEXER

CENTRAL AIR DATACOMPUTER (CADC)LEGEND:I- - ELECTRICAL
PT = TOTAL PRESSURE
Ps = STATIC PRESSURE
Pa = AOA
P,e = PROBE STATIC PRESSURE
PSL = STATIC PRESSURE, LEFT
PSR = STATIC PRESSURE, RIGHTTOTALTEMPERATURE
PROBEUSING SYSTEMS:• ALTIMETER• FIRE CONTROL SYSTEM (HUD AND FCC)• AOA INDICATOR• IFF TRANSPONDER• PROPULSION• NAVIGATION• ECS• FLIGHT CONTROLS• AUTOPILOT (ECA)• SELF TEST (ECA)• LESS ~ FLIGHT LOADS RECORDER• [8Q] SFDR
1 F-16A-1-1086X@

Figure 1-52.
1-140

Change 12

T. O. 1F- 16A- 1

The ECA converts signals from the PSA into four
identical signals for each required parameter (AOA,
static pressure, and impact pressure). The ratio of
impact to static pressure is generated within the ECA
and is used along with AOA and static pressure for
LEF scheduling. This pressure ratio, AOA, andimpact pressure are supplied to the FLCe for gain
scheduling.
Monitoring is provided in the ECA to detect single
and dual malfunctions of the triplex sensor signals. A
signal malfunction in any of the sensor signals
illuminates the ADC caution light. A dual malfunction of static or impact pressure or pressure ratio
signals results in the following:• Illumination of the ADC, FLT CONT SYS, and LE
FLAPS caution lights and the STBY GAINS light(on the FLCP).
• Activation of FLCS standby gains. Refer to
STANDBY GAINS, this section.
• Loss of autopilot.
Pitot probe tip icing results in erroneously low
airspeed indications, illumination of the ADC caution
light, and flight control gains scheduled for low
airspeed flight conditions.
ADC Caution Light

The ADC caution light, located on the caution light
panel, illuminates whenever a single or dual failure
occurs in the sensing of static and/or impact pressures
or AOA. The light also illuminates due to an error in
impact/static pressure computation.
The ADC and LE FLAPS caution lights and DUAL
FC FAIL warning light may illuminate for high AOA(above 29 degrees) and/or sideslip maneuvers. After
recovery, reset lights and continue normal opera-

tions.
CENTRAL AIR DATA COMPUTER (CADC)

I•

Refer to figure 1-53 for a signal flow diagram
showing the systems interacting with the CADC.
The CADC receives total and static pressures, AOA,
and total temperature inputs, converts the inputs
into digital data, and then transmits the data to theusing systems. The CADC has continuous BIT andinitiated BIT features; initiated BIT is run during theFLCS self-test.

detected. If there is a mach signal failure fromthe CADC, the IPW200 I EEC, I PW2201 ENGINEFAULT caution light also illuminates.

WARNING, CAUTION, AND INDICATOR LIGHTS
Refer to figure 1-54. Warning, caution, and indicatorlights are used throughout the cockpit to callattention to a condition or to allow an item to beeasily read. Red warning lights and the amber
MASTER CAUTION light are all located on the edgeof the glareshield. All of the lights, except the
MASTER CAUTION light, are described under theirrespective systems.
The warning and caution lights (except MASTER
CAUTION) are not press-to-test or press-to-reset
lights. Pressing these lights releases them from theirmodules and deactivates them. To reengage areleased light, pull it out slightly and then press toreengage the module.
VOICE MESSAGE SYSTEM (VMS)

The VMS provides a warning message, a caution
message, or discrete messages. The fixed volume I
voice message does not blank other audio and,
therefore, may not be heard.
The warning message (WARNING-WARNING
pause WARNING-WARNING) is automatically
activated 1.5 seconds after illumination of any
warning light on the glareshield.
The caution message (CAUTION-CAUTION) is
automatically activated 7 seconds after the illumination of any light on the caution light panel. If the IMASTER CAUTION light is reset immediately
after its illumination, the voice caution message
does not occur.
The warning/caution messages are reset for subsequent activation by:• Resetting the HUD WARN RESET for VOIce
warning.
• Resetting the MASTER CAUTION for VOIce
caution.

CADC Caution Light• Eliminating the condition that originally activatedthe lights and messages.

The CADC caution light, located on the caution light
panel, illuminates whenever a malfunction is

Discrete voice messages are provided when certain I
conditions occur.

1- 141

T. O. 1F- 1 6A- 1

CADC and Interfacing Systems
STATIC PRESSURE
NOSE-MOUNTED
AIR DATA (PITOT)PROBE

CENTRAL
AIR DATA___ TOTAL TEMPERATURECOMPUTER(CADC)TOTAL PRESSURE

PRESSURE ALTITUDE (29.92)

DISCRETE BASED ON
CALIBRATED AIRSPEED,PRESSURE ALTITUDE, AND
PRESSURE ALTITUDE RATE

T.O./LAND CONFIGWARNING LIGHTTOTALTEMPERATURE
PROBE

p

....

ALTIMETER

BAROMETRIC SETTING

AIMS CODEPRESSURE ALT VALID.
.. IFF TRANSPONDERp....
AVIONICSYSTEMS
AND
IpW220\ EDUCADC STATUS
AIR DENSITY RATIOPRESSURE ALTITUDE

TRUE TEMPERATURE

MACH NUMBER.. ENVIRONMENTALCONTROL SYSTEM.. PRESSURE ALTITUDE (29.92)
TRUE AIRSPEED....

CALIBRATED AIRSPEED

PRESSURE ALTITUDE (29.92)

AOA

PRESSURE ALTITUDE (BARO)pp.. MACH NUMBER

CALIBRATED AIRSPEED.. PRESSURE ALTITUDE (BARO)TRUE TEMPERATURE.. STANDARD FLIGHT

..
p

AOA

COMMAND WORDDATA RECORDER(SFDR) ~po

MACH NUMBERp

CADC FAIL STATUS

I

BIT ENGAGE

PRESSURE ALTITUDE

ELECTRONIC RESETCALIBRATED AIRSPEEDpop

ELECTRONICCOMPONENTASSEMBLY(ECA)

po

FLIGHT LOADSRECORDER LESS ~AOA.. ~ PRESSURE ALTITUDE- PRESSURE ALTITUDE RATE~ALTITUDE ENGAGE

115 VAC, 1

0, 400 Hz

ELECTRICAL POWER...
DISCRETES.. AOA INDEXERpt
AOA
TRANSMITTERS

FLIGHT CONTROL
PANEL

MACH NUMBER

AOA

BIT ENGAGEpo

PROPULSIONCONTROL SYSTEM..

AOA INDICATORELECTRONIC RESETCADC FAIL STATUS

CAUTION LIGHT
PANEL

1 F-16A-1-1 087X@

Figure 1-53.

1-142

Change 1210.lF-16A-1• Descent is occurring after takeoff.

The VMS does not function with WOW. However, it
can be tested by pressing the MAL & IND LTS buttonon the TEST switch panel. During test, each word isheard one time. The VMS is poweredby ~ battery bus No.1, LESS ~ the battery bus.

• Barometric altitude is below the entered MSL
ALOWvalue.

VOICE MESSAGE Switch I!l ~Messages are:
• ALTITUDE-ALTITUDE - Advises that:• Radar altitude is below the entered radar ALOW
value.

I
I
I

Refer to T.O. IF-16A-34-1-1, ~ T.O. IF-16A-34-1-3
or~ T.O. IF-16A-34-1-4 for a detailed description.
• BINGO-BINGO - Advises that the bingo fuelwarning has been activated. Refer to T.O.
IF-16A-34-1-1, ~ T.O. IF-16A-34-1-3 or 1M T.O.
IF-16A-34-1-4 for a detailed description.
• IFF - Advises that the IFF code should bechanged. Refer to T.O. IF-16A-34-1-1, ~ T.O.
IF-16A-34-1-3 or 1M T.O. IF-16A-34-1-4 for a
detailed description.
• JAMMER - Not operable in flight. (Message isheard during ground test.)

I
I• LOCK-LOCK- Advises that radar has locked on totarget. Refer to T.O. IF-16A-34-1-1, ~ T.O.
IF-16A-34-1-3 or 1M!! T.O. IF-16A-34-1-4 for a
detailed description.
• PULLUP-PULLUP-PULLUP-PULLUP - Advisesthat an immediate pullup is required. Refer to T.O.
IF-16A-34-1-1, ~ T.O. IF-16A-34-1-3 or 1M T.O.
IF-16A-34-1-4 for a detailed description.
All voice messages have priority over the low speed
warning tone and LG warning horn. Voice messages
are also prioritized.

Refer to figure 1-3. The VOICE MESSAGE switch,
located aft of the stick, is a two-position switch.
Positions are marked VOICE MESSAGE and
INHIBIT. During normal operation, the switch issafety-wired in VOICE MESSAGE. Placing theswitch to INHIBIT disables all voice messages.
INHIBIT should only be used to clear a voice message
which repeats abnormally. Placing the swikh back to
VOICE MESSAGE enables normal operation.
MASTER CAUTION LIGHT

The MASTER CAUTION light, located on the left
upper edge of the glareshield, illuminates shortly
after an individual light on the caution light panel
illuminates to indicate a malfunction or specificcondition exists. It does not illuminate in conjunction with the warning lights. The MASTER
CAUTION light may be reset by depressing the faceof the light unless it is illuminated by the FLT CONTSYS or ELEC SYS caution lights. The light should bereset as soon as feasible so that other caution lights
can be monitored should additional malfunctions or
specific conditions occur. Unless it is reset, theMASTER CAUTION light remains illuminated aslong as the individual caution light is illuminated.
The light can be checked by depressing the MAL &IND LTS button on the TEST switch panel. ~ The
MASTER CAUTION light is a repeater and cannot
be reset individually.
CAUTION LIGHT PANEL

Priority sequence is:• PULLUP.
• ALTITUDE.
• WARNING.
• BINGO.
• CAUTION.
• JAMMER.
• LOCK.
• IFF.

Refer to figure 1-54. The caution light panel is locatedon the right auxiliary console. The following twolights must be reset at the respective system control
panels:• FLT CONT SYS - FLCP, FCS CAUTION RESET
button.
• ELEC SYS - ELEC control panel, ELEC CAUTIONRESET button. Exception: Certain aircraft battery
charging system or FLCS battery system failures
can result in a nonresettable ELEC SYS caution
light. The caution light may appear nonresettable Iin situations where the ELEC SYS caution light is
rapidly flashing or cycling on and off.

Change 13

1-143T.0. 1F-16A-1

Warning, Caution, and Indicator
Lights ~ IBF I (Typical) Block 10

ADA
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FIgure 1-54. (Sheet 1)

1- 144

U-UIA 11088 IX.

1 0 . lF-16A-1

AR STATUS/NW S
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ECM POD CONTROL PANELS

Figure } ·54. (S heet 2)

1-145

T.O.1F- 16A-l

Warning, Caution, and Indicator
Lights IBRI (Typical) Block JO

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IF-I6A,- 1-1089- I X.

Figure 1·54. (Sheet 3)

1-146

T.O. 1F-16A-1
AR STATU S/NWS
INDICATORI ROY,

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IF· l 6A· l 089·zxe

Figure 1·54. (Sheet 4)1-147

T.O. 1F-16A- l

Warning, Caution, and Indicator
Lights [8] IBFI(Typical) Block J5

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INDEXER

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Figure I -54. (Sheet 5)

1- 148

If · 16 ... · '-1090- 1X.

IO, IF-16A-l
AR STATUS/NWS
INDICATOR

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IPW220 I CAUTION LIGHT PANel• LESS~I~B3~ o m.~~~

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CHAFF/FLARECONTROL PANEL

11 16A- ' 1090-2,, 61

ECM POD CONTROL PANELS

Figure 1·54. (Sheet 6)
Change 12

1-149

T. O. 1F-16A-1

Warning, Caution, and Indicator
Lights JBRJ (Typical) Block 1.5

AOA
INDEXERTHREAT WARNINGSTICK rn9~:

OVERRIDE
LIGHT.

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Figure 1-54. (Sheet 7)

1-1 50to.1F.16A.1
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Figure 1-54. (Sheet 8)
Change 12

1·151

T.o. 1F-16A- 1• AVIONICS - FCNP, FALT ACK button.

are not otherwise obvious from cockpit indications.
The PFL displays failed subsystem, degree ofseverity, specific test, number of occurrences and timeof first fault.

• LE FLAPS, ADC, CADC - FLCP, SERVO ELECRESET switch (ELEC position).

There are two conditions which may display PFL
faults, a normal pilot report and a pilot fault recall .

The following caution lights may be reset at thesystem control panels:• 1PW2201 ENGINE FAULT - ENG & JET START

panel, AB RESET switch (AB RESET position).

The normal PFL report is structured to minimize
actions required to isolate avionics or PW220 engine
faults and their impacts and/or limitations on
current modes of operation. The MASTER CAUTION light illuminates, providing a visual indication of an aircraft system malfunction. The
AVIONICS or I PW2201 ENGINE FAULT caution
lights illuminate to isolate the fault to its respective
system. Further isolation to the particular subsystem is available by depressing the FALT ACK button I
on the FCNP. This action results in a fault code
display on the FCNP (5 seconds for each previously
unacknowledged fault) and resets the MASTER
CAUTION, AVIONICS, orl PW2201 ENGINE FAULT
caution lights. The FCNP display returns to normal
subsequent to the display of all previously
1

MAL & IND LTS Test Button

The MAL & IND LTS test button, located on the TEST
switch panel, operates relays which test theillumination of all warning, caution, and indicatorlights (except the FLCP SERVOS and dot lights), theLG warning horn, voice messages, and ECM pod
control panel lights.
PILOT FAULT LIST (PFL)

Refer to figure 1-54.1. The PFL is displayed on the
FCNP located on the ~ ~ left console. The PFL isdesigned to display faults of concern to the pilot which

I

Pilot Fault List (PFL)(Typical)2

6Q\.:::J~==============~

F
/

N
A
V

[
0

~

AI. T CAL

\

POS
I

,

OEST

CRUIS~WINO,.-:~ ~ ~ ~ If\

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TEST'

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T HOG

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NORM

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IMODElnlDATAln FALTSEL U OPT U ACK

1. Degree of Severity (0 thru 6)
2. Test Number Failed
3. Failed Subsystem
4. Number of Occurrences
5. Time of First Occurrence
6. FALT ACK Button

f:\ r-:'\ f:\~~ ~~8~~DEST\l...J ~ V

G==

1F-1SA-1-1092X@

Figure 1-54.1.

1-152

Change 3

T. O. 1F- 16A- 1
unacknowledged faults in the PFL. The PFL isdesigned to store acknowledgment of each fault.
Once a fault has been acknowledged, repetitive
illumination of the AVIONICS or I PW2201 ENGINEFAULT caution lights by an intermittent fault isblocked to prevent annoyance reports.
The PFL recall allows review of current faults. The
pilot can initiate a recall of current fault status bydepressing the FALT ACK button when there is no
unacknowledged pilot fault present. Depression of
the FALT ACK button clears all faults in the PFL(except radar self test faults 004 through 092),
initiates a survey of all current malfunctions, andrebuilds the PFL. From the rebuilt PFL, the current
faults are reported in sequence on the FCNP for
periods of approximately 5 seconds each until all
faults contained in the list have been displayed.

I

Refer to T.O. IF-16A-34-1-1, I@I T.O. IF-16A-34-1-3
or ~ T.O. IF-16A-34-1-4 for a detailed description.

[A] identification light are controlled from the EXT

LIGHTING control panel.
Anticollision Strobe Light

The anticollision strobe light is masked to minimize
projections in the cockpit.
Position/Formation Lights

Refer to figure 1-56.
Air Refueling Lights

The AR floodlight shares the housing of the top
fuselage formation light. The light is directed aft toflood the receptacle, fuselage, wing, and empennage.
The AR slipway contains embedded lights on eachside of the slipway. These lights are enabled when the
AIR REFUEL switch is in OPEN.
Vertical Tail-Mounted Floodlight ~LIGHTING SYSTEM
EXTERIOR LIGHTING

Refer to figure 1-55. All of the exterior lights except the landing and taxi lights and ~ and ~ ~A white light is mounted on the upper leading edgeof the vertical tail and is directed forward to flood the
AR receptacle and upper fuselage. The lightilluminates by the OPEN position of the AIR
REFUEL switch.

EXT LIGHTING' Control Panel [8JIBFI(Typical)
3 4

12
ANTI
~ L COLLISION-POSITION -FLASH

WING/TAIL-

I

1. ANTICOLLISION Switch
2. FLASH STEADY Switch
3. WING/TAIL Switch
4. FUSELAGE Switch
5. AERIAL REFUELING Knob
6. MASTER Switch
7. FORM Knob-OFF -E GOFF

STEADY--DIM--XH

TT

FUSELAGE-BRT--FORM(QDj~ G~T

MASTERAERIAL REFUELING

~~T W

NORM

OFF
1
~==I=~

1
7

6

5
IF-16A-l-0041A ®

Figure 1-55. (Sheet 1)

Change 13

1-152.1/(1-152.2 blank)

T.O. 1F-16A-1

EXT LIGHTING Control PanellAllIE] (Typical)
CONTROL
1. ANTICOLLISION Switch

POSITIONFUNCTIONANTICOLLISION

Turns on the white anticollision strobe(flash) light on top of vertical tailOFF

Turns off the anticollision strobe light

FLASH

Causes the lights controlled by WING!I'AIL
switch to flash when turned on

STEADY

Causes the lights controlled by WING!I'AIL
switch to light steady when turned on

BRT

Turns on the red and green wingtip andinlet lights and white light at trailing edgeof vertical tail brightOFF

Turns off the white light at trailing edge ofvertical tail and the inlet lights. Allows the
red and green wingtip lights to be controlled
by the FORM knob

DIM

Turns on the red and green wingtip lights
Block 10 bright, Block 15 dim. Turns on the
red and green inlet lights and the whitelight at the trailing edge of the vertical tail
dim

BRT

Turns on the white floodlights at the base of
the vertical tail brightOFF

Turns off the white floodlights

DIM

Thrns on the white floodlights dim

5. AERIAL REFUELING Knob

Variable from off to
BRT

Varies the AR slipway lights from off tobright if AIR REFUEL switch is in OPEN

6. MASTER Switch

NORM

Enables all exterior lights excoot landing
ID lights
and taxi and ~ and [QE] Ifm

OFFDisables all exterior lights except landing
and taxi and ~ and ~ Ifm IA1 ID lights

Variable from off to
BRT

Varies the white formation lights on top andbottom of fuselage and, when the WING/
TAIL switch is OFF, varies the red andgreen wingtip lights from off to bright

2. FLASH STEADY Switch

3. WING!I'AIL Switch

4. FUSELAGE Switch

7. FORMKnob

Figure 1-55. (Sheet 2)

Change 1

1-153

IT. O. 1F- 16A-1

Exterior Lighting (Typical)

ANTICOLLISIONSTROBE (WHITE)BLOCK 15 VERTICAL TAIL MOUNTEDFLOODLIGHT (WHITE)POSITION/FORMATION
LIGHT (GREEN)(TOP AND BOTTOM)(AFT SIDE)
AR (FLOOD)(WHITE)POSITION
LIGHT (WHITE)(BOTH SIDES)FORMATION
LIGHT (FUSELAGE LABEL ON
CONTROL PANEL)(FLOOD) (WHITE)(FORWARD SIDE)FORMATION LIGHT(WHITE)POSITION LIGHT(R-GREEN)(L-RED)POSITION/FORMATION
LIGHT (RED)(TOP AND BOTTOM)TAXI LIGHT (RMLGILANDING LIGHT (LMLG)

10 LIGHT (WHITE)
IADIANDI1m~CAJ

1F-16A-1-1093xe

Figure 1·56.

Landing and Taxi Lights• OFF - Disables ID light.

A white landing light is located on the left MLG strut.
The light is angled to illuminate the landing area. A
white taxi light is located on the right MLG strut. The
lights are turned on by the three-position LANDINGTAXI lights switch located on the LG control panel.
The switch has positions of LANDING, OFF, andTAXI. The light goes off during LG retraction if theswitch is left in either the LANDING or TAXI
position.

• TEST - Momentary position which allows afunctional test of the ID light on the ground or whenthe LG is extended.

Identification (ID) Light ~A white, flush-mounted target ID light is installed in
the left forward equipment bay door. The beam points
70 degrees left of forward and 10 degrees above theaircraft horizontal centerline plane. The ID light iscontrolled by the three-position ID LIGHT switch
located on the MISe panel.
Functions are:
• ID LIGHT - The ID light is turned on after LG isretracted.

1-154

Identification (ID) Light ~ 1m ~A white, flush-mounted target ID light is installed in
the left forward equipment bay door. The beam points
70 degrees left of forward and 10 degrees above theaircraft horizontal centerline plane.
The light is turned on by the TAXI position of the
LANDING TAXI lights switch after LG is retracted.

INTERIOR LIGHTING
Refer to figure 1-57. The interior LIGHTING control
panel contains the power and intensity controls forthe primary (console and instrument) and secondary(flood) lighting systems for the cockpit. The HIGH
INT position of the FLOOD CONSOLES knob
provides thunderstorm lighting.

T. O. 1F- 16A- 1
FREQ DISP Knob lim

PRIMARY CONSOLES Knob

The FREQ DISP knob has a cw arrow pointing toward
BRT. The knob controls the lighting of the ReF! from
off to bright.

The PRIMARY CONSOLES knob has a cw arrow
pointing toward BRT. Rotating the knob cw varies theintensity of the primary and auxiliary console
lighting from dim to bright.

MAL & IND LTS Switch

The MAL & IND LTS switch has positions of BRT
and DIM and a. spring-loaded unmarked center
position. If the PRIMARY INST PNL knob is on,
momentary activation of the switch to DIM places
the lighting system to the dim condition. The system
automatically returns to the BRT condition if the
FLOOD CONSOLES knob is turned past the detentto HIGH INT, if the PRIMARY INST PNL knob isturned off, or if nonessential dc power is lost. The
BRT condition can be manually selected anytime.
The switch controls the light intensity of all thewarning, caution, and indicator lights except thefollowing:

PRIMARY INST PNL Knob

The PRIMARY INST PNL knob has a cw arrow
pointing toward BRT. Rotating the knob cw varies theintensity of the forward instrument panel lightingfrom off to bright.

NAV/FREQ DISP Knob ~ [IllThe NAVIFREQ DISP knob has a cw arrow pointing
toward BRT. The knob controls the lighting of the
FCNP and RCFI from off to bright.

Cannot be dimmed:• FLCP fault lights

Interior LIGHTING Control Panel (Typical)
2

w[@
L
I
G
H
T
I

N

6

G

~

BRT

--FLOOD--MAL &INO LTS

INST PNLCONS1..ES

4

BRT
DIM

OFf HIGH INTW~

1. PRIMARY CONSOLES Knob
2. PRIMARY INST PNL Knob
3. 00 [jf] NAV/FREQ DISP Knob
FREQ DISP Knob

IIBI

BRT

\
5

rnBJ

~

BLOCK 10BLOCK 15

4. MAL & IND LTS Intensity Switch
5. FLOOD INST PNL Knob
6. FLOOD CONSOLES Knob
1F-16A-1-1094xe

Figure 1-57.

1-155

T. O. 1F- 16A- 1

HIGH INT, the MAL & IND LTS automatically go tobright and the alphanumeric displays controlled by
the NAVIFREQ DISP knob go to the highest intensity
level. CCW rotation past a certain point restores thealphanumeric displays to the intensity level set by
the NAVIFREQ DISP knob, but the MAL & IND LTS
switch must be manually reset to DIM, if dim isdesired.

Individually dimmed:• AOA indexer.
• ARlNWS indexer.
• ECM pod control panel.

.

• FCNP and RCFI display.

I

LESS~ IFF control panel REPLY and TEST lights.

• ~ AIFF control panel REPLY light.
• SCPo• TWS indicators.
FLOOD INST PNL Knob

The FLOOD INST PNL knob has a cw arrow pointing
toward BRT. Rotating the knob cw varies thefloodlights intensity from off to bright.
FLOOD CONSOLES Knob

The FLOOD CONSOLES knob rotates from OFF to
HIGH INT. Rotating the knob cw varies the console
floodlights intensity from off to bright. If rotated to

Utility Light

Refer to figure 1-58. The utility light, located on theright console, includes three controls: a pushbutton
switch to allow momentary operation at the highest
intensity level, an OFF DIM BRTrotary knob to allow
continuous operation at any desired intensity level,
and a lens housing which, when rotated, adjusts thebeam from flood to spot. To release the light from its
stowed position, lower the knurled collar at the baseof the light and it will pop free. The light can be lockedback into position by placing the body of the lightparallel to the sidewall fairing and pushing down
firmly on the light assembly. With the canopy closed,
the utility light may be attached to an adjustablesliding holder located on the right body positioning
handle (towel rack). The light is powered by ~ battery bus No.1, LESS ~ the battery bus.

Utility Light and Spotlights (Typical)
PUSHBunON SWITCHROTARY KNOB/QUICKDISCONNECT
PLUG

HOLDER ON
RIGHT BODY
POSITIONINGHANDLE(TOWEL RACK)UTILITY LIGHT

Figure 1-58.

1-156

Change 3

1F-16A-1-1 095X@

10.lF-16A-1
Cockpit Spotlights

Refer to figure 1-58. The cockpit spotlights are locatedunder the upper left and right glareshields. In thestowed position (horizontal, facing forward), thespotlight is off.
The spotlight is turned on by pulling the spotlightbarrel downward. Illumination intensity is controlled
by turning the knurled barrel (dimmer). To turn the
spotlight off, return it to the stowed position.
Overrotation of the knurled barrel may cause
breakage of the bulb or rheostat. The light is poweredby the ~ battery bus No.1, LESS I!Q] the batterybus.

ESCAPE SYSTEMCANOPYThe canopy is a two-piece, plastic, bubble-type,transparent enclosure. The forward part is asingle-piece windshield-canopy transparency which
is hinged at the aft end and is unlatched, opened, orclosedllatched by an electrically operated actuatorwith a manual backup. 00 A smaller fixedtransparency fairs to the fuselage aft of the seat. The
canopy may be jettisoned by internal controls forin-flight or ground escape and by external controls forground rescue. An inflatable pressurization seal on
the cockpit sill mates with the edge of the movable
canopy. A noninflatable rubber seal on the canopy
prevents the entry of water when the cockpit is notpressurized.
The canopy provides some bird strike protection. Bird
strikes on centerline at approximately eye level may
produce enough canopy deflection to shatter the HUD
combiner glass and cause rearward propagation of adeflection wave. Deflection of the canopy in the areaof the pilot's helmet has been observed to be 1 to 2
inches during bird strike tests that were consideredsuccessful. Successful completion of canopy bird
strike testing (4 pound bird at 350 or 550 knots,
depending on canopy) requires that the canopy not
deflect more than 2 114 inches in the area of the pilot's
helmet. This may be a consideration for adjustingseat height, especially while flying at lower altitudes
with helmet-mounted equipment. Impacts off center
may not shatter the HUD glass. High energy bird
strikes may cause canopy penetration or larger
deflection waves.
CANOPY CONTROLS AND INDICATORS

Refer to figure 1-59.

MANUAL CANOPY CONTROL Handcrank

00 rnf] An internal MANUAL CANOPY CONTROL
handcrank manually performs the same function as
the canopy switch. Due to the strength required toopen the canopy with the handcrank, the methodshould be considered a last resort.
An external flush-mounted CANOPY handcrank
receptacle just opposite the inside manual drive isused for ground crew manual operation of the canopy.
Canopy Handle 00 [M]The canopy handle, located on the canopy sill justforward of the throttle, hinges down to cover andprotect the internal canopy switch. The handle alsofunctions to inflate/deflate the canopy pressure seal,
to turn the CANOPY warning light off/on, and tomechanically prevent the canopy actuator fromunlatching. The canopy handle should be in the up(unlock) position prior to lowering the canopy.
The FLCS batteries are disconnected from the FLCS
inverters when the canopy is partially opened.
CANOPY JETTISON

Pulling the external canopy jettison D-handle,
located on either side of the fuselage, initiates the
canopy jettison sequence independent of seatejection.
Depressing the button, located on either side of theinternal CANOPY JETTISON T-handle, and pulling
the T-handle initiates the canopy jettison sequence
independent of seat ejection. 00 [M] There are twodifferent CANOPY JETTISON T-handle configurations which may require different handgrips to insure
proper canopy jettison. If the CANOPY JETTISON
T-handle is mounted so that the words CANOPYJETTISON engraved on the T-handle are upright,
then an underhand grip should be used to insure astraight pull and jettison. If the CANOPY JETTISON
T-handle is mounted so that the words CANOPYJETTISON are inverted, an overhand or underhand
grip may be used.
Pulling the ejection handle (PULL TO EJECT),
located on the front of the ejection seat, initiates the
canopy jettison sequence followed by the seat ejection
sequence.
EJECTION SEAT

Refer to figure 1-60. The ACES II ejection seat is afully automatic emergency escape system. One ofthree ejection modes is automatically selected.
Mode 1 is a low airspeed, low altitude mode during
which the recovery parachute assembly is deployedalmost immediately after the ejection seat departs
the aircraft.

Change 111-157

10. 1F-16A- 1

Canopy Controls and Indicators (Typical)
NOTE:
If the internal canopy switch is in the up position,
the canopy cannot be closed using the external
canopy switch.

2
1

LESS

a3
3 !CANOPY!

5

3!CANOPY!

6(Left side panel,
forward cockpit)

6

4

7 10

8

9

4

1. [AJ~ Canopy Handle (shown in unlocked position)
2. [8Jrn£l Canopy Switch (internal) (spring-loaded tocenter from down position)
3. CANOPY Warning Light
4. Ejection Handle (PULL TO EJECT)
5. [8Jrn£l MANUAL CANOPY CONTROL Handcrank
6. CANOPY JETTISON T-Handle
7. Canopy Jettison D-Handle (each side of fuselage)
8. CANOPY Handcrank Receptacle (external)
9. Canopy Switch (external) (spring-loaded to center)
10. Canopy Lock Access Plug (external)

1 F-16A-1-0045A@

Figure 1-59. (Sheet 1)

1-158

lO. 1F-16A-l

Canopy Controls and Indicators (Typical)
1.

2.

3.

FUNCTIONCONTROUINDICATORPOSITIONIINDICATION[!J rnfI Canopy Handle

Up

Unlocks canopy

Down

Locks canopy

Up

Opens canopy

CenterStops canopy motion

Down

Closes and latches canopy

Off

Canopy lockedOn

Canopy unlockedrm[!J
Canopy Switch(internal) (spring-loaded tocenter from down position)

Canopy Warning Light
(red)

4.

Ejection Handle (PULL TO
EJECT)

Pull

Jettisons canopy and ejects seat5.

[!J lIE1 MANUAL CANOPYCONTROL Handcrank

Rotate ccw

Opens canopy

Rotate cw

Closes and latches canopy

6.

CANOPY JETrISON
T-Handle

Pull (depress either button). [!J mEl There are two
CANOPY JETrISON
T-handle configurations
which may require different handgrips

Jettisons canopy independent of seat ejection

7.

Canopy Jettison D-Handle(each side offuselage)

Pull (approximately 6feet) (either handle)

Jettisons canopy independent of seat ejection

8.

CANOPY Handcrank
Receptacle (external)

Rotate cw

Opens canopy

Rotate ccw

Closes and latches canopy[!J UpOpens canopy

9.

Canopy Switch (external)(spring-loaded to center
position)[!] Aft

CenterStops canopy motion[!J Down

Closes and latches canopy

00 Fwd10.

Canopy Lock Access Plug(external)

Remove access plug

Access to unlock internal canopy handle
Refer to EMERGENCY ENTRANCE AND
CREW RESCUE, Section III

Figure 1-59. (Sheet 2)

1-159

10. 1F- 16A- 1

Ejection Seat Controls and Indicators(Typical) (FIXED PITOTS)
EJECTION MODE SELt-:~
INSPECTION HOLE -

__ ..

AFT,

11
SIDE VIEWUWARSINSTALLED

INERTIA REEL STRAPSRETAINING PIN(YELLOW)

10RECOVERY PARACHUTEASSEMBLY

11

RED PIN
PROTRUSION
INDICATES EXPENDEDSEQUENCERSEAWARSINSTALLED

SEAT/HOSE
JH.!l;~~- DISCONNECT

LAPBELT ANDSURVIVALKIT STRAPSSHOULDER HARNESS STRAPS/RECOVERY PARACHUTE RISERS

7 ------w~:Y_.

I

SEAWARS/UWARSNOT INSTALLEDI>

KIT DEPLOYMENTFWD AUTOMATIC
AFT
MANUAL(ROTATED 180°)

SURVIVAL KIT(UNDER SEATPAN LID)

1. Emergency Oxygen Green Ring
2. SHOULDER HARNESS Knob
3. Ejection Safety Lever
4. RADIO BEACON Switch
5. Survival KIT DEPLOYMENT Switch
6. Ejection Handle (PULL TO EJECT)
7. EMERGENCY MANUAL CHUTE Handle
8. Survival Kit Ripcord
9. Electronic Recovery Sequencer Battery Indicator
10. SEAT NOT ARMED Caution Light
1 1. [jBJ EJECTION MODE SEL Handle

RADIO BEACON

MAN~AUTO(ROTATED 180°)

1F-16A-1-1097B@

Figure 1-60. (Sheet 1)

1-160

Change 3

1.0. 1F- 16A-1

Ejection Seat Controls and Indicators(Typical)(FLIP-UP PITOTS)
INSPECTION HOLE -ENVIRONMENTAL
SENSOR PITOTS(EXTENDED DURING~____- - - - - - EJECTION)

t.:~

_ ..

.,..

EJECTION MODE SEL
AFT ......

...... !D.O

SIDE VIEW

11INERTIA REEL STRAPSRETAINING PIN(YELLOW)ft

RED PIN
PROTRUSION
INDICATES EXPENDEDSEQUENCER

ENVIRONMENTAL SENSOR~;«i:!~- PITOT ~STOWED)5EMERGENCY OXYGEN HOSEQUICK-DISCONNECTSEATIHOSE
L.a'~-- DISCONNECT

LAPBELT AND SURVIVALKIT STRAPSSHOULDER HARNESS STRAPS/RECOVERY PARACHUTE RISERS

8---~

>

I

7
T

KIT DEPLOYMENTRETAINING PIN(BOTH SIDES)
1 (YELLOW)

FWD AUTOMATIC
Ifl
MANUAL(ROTATED 180°1

SURVIVAL KIT(UNDER SEATPAN LlDI

4~3

1. Emergency Oxygen Green Ring
2. SHOULDER HARNESS Knob
3. Ejection Safety Lever
4. RADIO BEACON Switch
MAN~AUTO
6. SEAT NOT ARMED Caution Light
6. Survival KIT DEPLOYMENT Switch(ROTATED 180°)
7. Ejection Handle (PULL TO EJECT)
8. EMERGENCY MANUAL CHUTE Handle
9. Survival Kit Ripcord
10. Electronic Recovery Sequencer Battery Indicator
11.
EJECTION MODE SEL Handle

EMERGENCY OXYGENPRESSURE INDICATOR

RAD 10 BEACON

H

1F-16A-1-1098A@

Figure 1-60. (Sheet 2)

Change 1

1-161

T.O. 1 F-16A-l

Ejection Seat Controls and Indicators(Typical)
CONTROUINDICATORPOSITIONIINDICATIONFUNCTION

1.

Emergency Oxygen Green
Ring

Pull

Activates emergency oxygen

2.

SHOULDER HARNESSKnobUNLOCKED (aft)Unlocks inertia reel and allows free movement of shoulder harness

LOCKED (forward)

Locks inertia reel and prevents forward

movement of shoulder harness

3.

4.

5.

6.

I

Ejection Safety Lever

RADIO BEACON Switch

SEAT NOT ARMED
Caution Light (amber)
Survival KIT DEPLOYMENTSwitch

Up

Prevents pulling of ejection handle

Down

Allows ejection handle to be pulled

MAN (green dot visible)

Selects manual mode

AUTO (red dot visible)

Selects automatic modeOn

Ejection safety lever up

OffEjection safety lever down

A (forward)

Selects automatic mode

M (aft)Selects manual mode

7.

Ejection Handle (PULL TO
EJECT)

Pull

Jettisons canopy and ejects seat

8.

EMERGENCY MANUALCHUTE Handle

Pull

After ejection - ballistically deploys therecovery parachute assembly and releases
lapbelt and inertia reel straps and unlatches
the seat pan lid

9.

Survival Kit Ripcord

Pull

Deploys survival kit

10.

Electronic Recovery
Sequencer Battery Indicator

White

Sequencer thermal batteries not activated - operationalRed

Sequencer thermal batteries activated - notoperational

AFTINORMISOLO

Selects ejection mode

I

11.

~ EJECTION MODE SEL
Handle (LIFT TURN)

Refer to F -16B AIRCRAFr, this section

Figure 1-60. (Sheet 3)1-162

T. O. 1F- 16A- 1
Mode 2 is an intermediate airspeed, low altitude
mode during which a drogue chute is first deployed
to slow the ejection seat followed by the deploymentof the recovery parachute assembly. Mode 3 is a highairspeed/high altitude mode in which the sequenceof events is the same as mode 2, except that
automatic pilot/seat separation and deployment of
the recovery parachute assembly are delayed until
safe airspeed and altitude are reached. Controls areprovided to adjust seat height and lock shoulder
harness.
Eiection Seat Controls and Indicators

Refer to figure 1-60.

safety lever is in the up (vertical) position. [ID The
caution lights illuminate independently of each other.
SEAT ADJ Switch

Refer to figure 1-3. The SEAT ADJ switch is locatedon the right cockpit sidewall outboard of the stick.
Center position is spring-loaded off. The up position
raises the seat, while the down position lowers theseat. The seat adjustment motor is protected by athermal relay which interrupts electrical power
when overheated. After a I-minute cooling period,
the motor should operate normally. The motor is
powered by nonessential ac bus No.2.

Eiection Handle

Shoulder Harness Straps/Parachute Risers

The ejection handle (PULL TO EJECT) is sized for
one-handed or two-handed operation and requires apull of 40-50 pounds to activate. The handle remains
attached to the seat by a wire cable after activation.

The upper torso restraints consist of shoulder
harness straps which also act as parachute risers.
The inertia reel straps are attached to the parachuterisers.

SHOULDER HARNESS Knob

Inertia Reel Straps

The SHOULDER HARNESS knob unlocks theinertia reel when in the aft position and locks it whenin the forward position. If high longitudinal
deceleration force or high shoulder harness strap
playout velocities are encountered, the inertia reelautomatically locks.
EMERGENCY MANUAL CHUTE Handle

The EMERGENCY MANUAL CHUTE handle islocked while the ejection seat is in the aircraft. Afterejection, the handle provides a backup to theautomatic pilot/seat separation and recovery parachute deployment system. The handle must be pulled
approximately 6 inches. The first 2 inches of pullballistically deploys the recovery parachute assembly; however, the seat will still be attached by thelapbelt and the inertia reel straps, and the survival
kit under the latched seat pan lid. Continued pull
releases the lapbelt and inertia reel straps andunlatches the seat pan lid.

The inertia reel straps may be manually released
after ejection by pulling the EMERGENCYMANUAL CHUTE handle approximately 6 inches.
Lapbelt

The lower torso restraint is the lap belt. The lapbelt
may be released after ejection by pulling the
EMERGENCY MANUAL CHUTE handle approximately 6 inches.
Survival Kit

The survival kit is stowed under the seat pan lid. The
KIT DEPLOYMENT switch has a manual (aft) or
automatic (forward) position which selects the modeof postejection survival kit deployment. Pulling the
kit ripcord handle deploys the kit which remains
attached by a 25-foot lanyard.

M

Eiection Safety Lever

RADIO BEACON Switch ~ [M!l ~The ejection safety lever mechanically safeties (in the
up/vertical position) or arms (in the downlhorizontal
position) the seat ejection handle.

The RADIO BEACON switch allows the pilot to select
AUTO or MAN. In AUTO (red dot visible), the beacon
activates after pilot/seat separation. In MAN (green
dot visible), the beacon does not activate. The beacon
may be activated when on the ground if the RADIOBEACON switch is placed to AUTO or it can beremoved and manually operated as desired. The
beacon transmits on 243.0 MHz.

SEAT NOT ARMED Caution Light

The SEAT NOT ARMED caution light, located on thecaution light panel, illuminates when the ejection

Change 12

~

1-163

•

T. O. 1F- 16A- 1
Emergency Oxygen

Emergency oxygen supply is automatically activatedduring ejection or may be manually activated bypulling the emergency oxygen green ring.

survival KIT DEPLOYMENT switch is in AUTO. If
the parachute is equipped with SEAWARSfUWARS, •the parachute risers are automatically released
approximately 2 seconds after entering salt water.
•~

EJECTION SEAT OPERATION

Seat ejection is initiated by pulling the ejection
handle (PULL TO EJECT). This action retracts theshoulder harness straps and locks the inertia reel,fires the initiators for canopy jettison, and ignites two
canopy removal rockets. As the canopy leaves theaircraft, lanyards fire two seat ejection initiators.
A rocket catapult propels the seat from the cockpit
exposing the seat environmental sensor pitots andactivating the emergency oxygen. The recovery
sequencer selects the correct ejection mode, ignites
the stabilization package (STAPAC) rocket and the
trajectory divergence rocket, and (if in mode 2 or 3)
initiates the drogue gun.
If the automatic pilot/seat separation and recovery
parachute deployment system fails, pulling the
EMERGENCY MANUAL CHUTE handle approximately 6 inches ballistically deploys the recovery
parachute assembly and releases the lapbelt andinertia reel straps and unlatches the seat pan lid.

Seat ejection also automatically selects AIFF
EMER and performs an escape zeroize operation bypurging coded electronic information associated withthe following equipment:• DTC.
• AIFF mode 4.
EJECTION MODE ENVELOPES

Refer to figure 1-61.
EJECTION SEQUENCE TIMES

Refer to figure 1-62.
CANOPY JETTISON/SEAT EJECTION.

Refer to figure 1-63.
EJECTION SEAT PERFORMANCE

The life raft, survival kit, and radio beacon antenna
are deployed following pilotlseat separation when the

1-164

Change 3

Refer to figure 1-64.

10. 1F- 16A- 1

Ejection Mode Envelopes

I- 16

w

w

r-~

LI.

-h--

o
o

r--r--

,

~

\

w

c

i'

~ 4

" r--."-

\

i: 8

5
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..............

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""-

1\

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MODE 3 - - I -

" t\.

\

t/)
t/)

w

r\.

-'

g: 0
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100

200

300

400

500

"

600

700

AIRSPEED - KIAS

1F-16X-1-0011X®

Figure 1-61.

Ejection Sequence Times
NOTE:

TIME (SECONDS)
EVENT

•

In mode 3, events after drogue deployment
are delayed until within mode 2 envelope.
Recovery parachute deploys 1 second after
entering mode 2 envelope.

• rID Times in the aftlforward sequenceincrease to include a O.33-second delay forthe rear seat and a 0.73-second delay forthe forward seat. In SOLO, the forward seatis delayed 0.33 second.

•

Canopy jettison time varies from 0.75
second at 0 KIAS to 0.13 second at 600 KIAS.
Ejection begins when canopy jettison initiates seat lanyards.

Mode 1Mode 2

1.

Catapult Initiation

0.0

0.0

2.

Drogue Gun Fired

NA

0.17

3.

Drogue Chute Inflated

NA

0.38

4.

Parachute Fired

0.20

1.17

5.

Seat/Drogue Separation

NA

1.32

6.

Pilot/Seat Separation

0.45

1.42

7.

Recovery Parachute Inflated

1.8

2.8

8.

Survival Kit Deployed

5.5

6.3

•
•
•

Figure 1-62.

Change 11

1-165

.....

:-I

I

0-0--

9

Canopy Jettison/Seat Ejection

II
I

0--

()

:::r
o

MODE 1»

MODE 2

I

:J

•

CO

m

(J) T = 2.8(J) T = 1.8~@T=1.42~~@ T = 1.32

'>

----

~
~

=:aE)

~

~

@T~1.17FCD@T ~ O.3Sgf cS)~~

®T=O.17~

~-

~_o

~_u

~-

~-

•
•
•
1 F-16X-1-0012X ®

Figure 1-63.

T. O. 1 F- 1 6A- 1

Ejection Seat Performance ~MINIMUM EJECTION ALTITUDE vs DIVE ANGLE AND SPEEDNOTES:• ZERO PILOT REACTION TIME
• WINGS LEVEL• SEA LEVEL• 95th PERCENTILE PILOT

1600

I I

!

~\~~

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./

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200,..,~v

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....

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....

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. - --

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w
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MODE 2

I
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30

45

60

75

DIVE ANGLE -90

DEGREES
1F-16X-1-1 002X ®

Figure 1-64. (Sheet 1)

1-167

10. 1F-16A-l

Ejection Seat Performance [8]MINIMUM EJECTION ALTITUDE vs SINI{ RATE AND SPEEDNOTES:• ZERO PILOT REACTION TIME
• WINGS LEVEL• SEA LEVEL• 95th PERCENTILE PILOT

600~~-r~~rT~~~-r~~~~.-rTr~-r~-r~~~~-r~~rT'-~~,'-~,~
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Figure 1-64. (Sheet 2)

1-168

T. O. 1 F- 1 6A- 1

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Figure 1-64. (Sheet 4)1-170

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Figure 1-64. (Sheet 5)1-171

T. O. 1F- 16A- 1

Ejection Seat Performance [mMINIMUlVl EJECTION ALTITUDE vs SPEED, DIVE ANGLE,
AND BANI{ ANGLENOTES:• ZERO PILOT REACTION T'ME
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1-172

75

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1F-16X-1-1007X@

T. O. 1F- 16A- 1

OXYGEN SYSTEM

Oxygen System Controls and Indicators

The oxygen system consists of a liquid system and an
emergency gaseous system.

Refer to figure 1-65 for description of the oxygen
system controls and indicators.
Oxygen Duration

LIQUID OXYGEN SYSTEM

Refer to figure 1-66.

A 5-liter liquid oxygen system provides breathingoxygen to a diluter demand oxygen regulator. The
regulator provides for selection of normal diluted
oxygen and 100 percent oxygen m and selection of
PBG. Quick-disconnects are used to expedite egressing the aircraft on the ground. Oxygen duration
varies depending upon altitude, regulator settings,
and usage.
Pressure Breathing for G (PBG)m

The PBG position of the oxygen regulator provides
pressure breathing above 4g's to enhance g tolerance
and to reduce pilot fatigue. ECS air used to inflate theg-suit is also used by the oxygen regulator to control
the amount of pressurization supplied to the oxygenmask, helmet, and vest.

EMERGENCY OXYGEN SYSTEM

The emergency oxygen system consists of ahigh-pressure bottle with a pressure regulator
mounted on the left side of the ejection seat. The hose
is routed to the right side of the seat.
The system is activated:• Automatically upon ejection .
• Manually by pulling the emergency oxygen green
ring located on the left aft side of the seat.
OXYGEN SYSTEM SCHEMATIC

Refer to figure 1-67.

1-173

T.0 .1 F·16A·l

Oxygen System Controls and Indicators(Typical)
2

1

6

1

OXYGEN REGULATORPRESSURE OEWAHO

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now

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1 OXYGEN REGULATOR FLOW Indalo..
2 OXYGEN SUPPlY Indicato..
3 Mode leYer {psG lever lock!
4 OiNter lever
S. EMER lever
6. OXYGEN SUPPlY PRESSURE IndielitOf
7. SUPPlY lever
8. EMeRGENCY leverg, 00 ~ UQUIO OXYGEN Ou.ntiry IndatOf
10, [!] [!f] OXY QTY Indicato.. Test SWitch, 1. OXY lOW Caution Light
12. [!] 1M] CABIN PRESS AlT
13. Cockpit OXY FLOW Indicator

Figure 1·65. (Sheet 1)

1-174

OXY
QTY

10

T.O. 1 F-16A-1

Oxygen System Controls and Indicators(Typical)
1.

2.

CONTROUINDICATORPOSITIONIINDICATIONOXYGEN REGULATORFLOW Indicator

White

Indicates oxygen flow

Black

Indicates no oxygen flow

Oxygen pressure (psi)Indicates gaseous oxygen pressure at regulator in psi

PBG (lever lock)

Provides oxygen supply to mask, helmet
bladder, and vest. Pressure breathing as afunction of g is available

ON

Provides oxygen supply to mask, helmet
bladder, and vest. Pressure breathing as afunction of g is not available

OFF

Turns off oxygen supply

NORM

Provides regulated mixture of cockpit air andoxygen as determined by cockpit pressure
altitude

100%02

Provides regulated 100 percent oxygen

NORM

Provides nonnal operation. Positive pressure
is provided if cabin pressure altitude exceeds
28,000-32,000 feet

EMER

Provides 100 percent oxygen under positive
pressure. This position is used by the pilot totest for leaks

TEST MASK

Provides 100 percent oxygen under positive
pressure. This position is used for testing bylife support maintenancem

Oxygen pressure (psi)Indicates gaseous oxygen pressure at regulator in psi

m SUPPLY LeverON'furns on oxygen supplyOFF'furns off oxygen supply

m OXYGEN SUPPLY
Indicator

3.

4.

5.

m Mode Leverm Diluter Lever

m EMERLever

6.

LESS
OXYGEN
SUPPLY PRESSURE
Indicator

7.

LESS

FUNCTION

Figure 1-65. (Sheet 2)

1-175

T.O. 1F-16A-1

Oxygen System Controls and Indicators(Typical)
CONTROIJINDICATOR
LESS
Lever

8.

m EMERGENCY[!]rnE LIQUID OXYGEN

9.

POSITIONIINDICATIONFUNCTIONNORMAL

Provides normal operation. Positive pressure
is provided if cabin pressure altitude exceeds
28,000-32,000 feet

EMERGENCY

Provides 100 percent oxygen under positive
pressure to the mask. This position is used by
the pilot to test for leaks

TEST MASK

Provides 100 percent oxygen under positive
pressure to the mask. This position is used fortesting by life support maintenance

0-5 liters

Indicates quantity of liquid oxygen

OXYQTY

Drives pointer ccw toward O. As pointerpasses 0.5 liter, OXY LOW caution lightcomes on. Light goes off after the switch isreleased and the pointer passes 0.5 liter

Quantity Indicator

I!JrnE OXY QTY Indicator

10.

Test Switch

11.

OXY LOW Caution Light(amber)On

Indicates oxygen quantity is less than 0.5
liter or system pressure below 42 psi

12.

~rm CABIN PRESS ALT

0-50,000 feet

Indicates cockpit pressure altitude

White

Indicates oxygen flow

Black

Indicates no oxygen flow

OXY FLOW
I 13. Cockpit
Indicator

Figure 1-65. (Sheet 3)

1-176

Change 2

T.O. 1 F-16A-1

Oxygen DurationOXYGEN DURATION (HOURS)COCKPITPRESSURE
ALTITUDE(FEET)

35,000

DILUTERLEVER(POSITION)

5

4

3

2

1

100%

30.94

24.75

18.56

12.37

6.19

NORM

30.94

24.75

18.56

12.37

6.19

100%

22.63

18.11

13.58

9.05

4.53

NORM

23.00

18.40

13.80

9.20

4.60

100%

17.48

13.98

10.49

6.99

3.50

NORM

21.72

17.37

13.03

8.69

4.34

100%

13.19

10.55

7.91

5.28

2.64

NORM

24.43

19.55

14.66

9.77

4.89

100%

10.62

8.49

6.37

4.25

2.12

NORM

29.86

23.89

17.92

11.94

5.97

100%

8.53

6.83

5.12

3.41

1.71

NORM

29.86

23.89

17.92

11.94

5.97

INDICATED QUANTITY (LITERS)AND UP

30,00025,00020,00015,00010,000NOTES:
1.

Oxygen duration increases as cockpit pressure altitude increases because there is less ambient pressure
acting upon the lungs. Therefore, a smaller quantity of oxygen at altitude expands the lungs to the same
size that they were at sea level.

2. With two people, decrease NORM 100% oxygen duration by 50 percent.
3. mUse ofPBG may reduce the listed available oxygen times.

1-177

10, 1 F- 16A-l

Oxygen System Schematicm (Typical)

CAUTIOf'\LIGHT

OXYGEN MASKHELMET
BLADDERCONNECTORPRESSURE SIGNALFROM ANTI-G VALVEQUANTITY
INDICATOR

VEST
CONNECTOR[KJlliEJ

PRESSURE RELIEF VALVE(FUNCTIONAL WHEN VEST
IS NOT CONNECTED)ON VALVE WITH
EMERGENCY OXYGENGREEN RING
EMERGENCY OXYGENPRESSURE INDICATORRELIEF VALVE

V':'UH2Z02TI2=E-C]2illillillillZZ2ZII(OVERBOARD
'"
~
VENT

r

SUPPLY

::.\

I
I t::
I l
I ::f:
I
I':'::':"':':~'

FILL/BUILDUP/
VENT VALVEPRESSURE
CLOSINGVALVE

LEGEND:o::::::.:':'.:.::

:}
):;

::\

ELECTRICAL
FILL CONNECTOR

B

QUICK DISCONNECT~CHECK VALVE

L"j~o.·."·j~

LIQUID OXYGEN

f::.:.::.:.::.:.:.::.]GASEOUS OXYGENLOX CONVERTER (REMOVABLE)
1F-16A-1-1102X®

Figure 1-67. (Sheet 1)

1-178

T. O. 1F- 16A- 1

Oxygen System Schematic LESSm (Typical)

CAUTIONLIGHT

OXYGEN MASKPRESSURE
SWITCHQUANTITY
INDICATOR[I][@
HARNESSMOUNTEDCONNECTOROXY FLOW
INDICATORPILOT TO SEATON VALVE WITH
EMERGENCY OXYGENGREEN RING

EMERGENCY OXYGENPRESSURE INDICATOR~-~~~~~~~~~~ OVERBOARD~~~~~~~/~~

VENT

VENT

---1

,

I
L

. ·.Jl . .·. '.....:.
o· ••

BUILDUP COIL

I
I
I
I
I
I
I
I
I

I
I

LEGEND:
ELECTRICAL

loU

FILL CONNECTOR

B

QUICK DISCONNECTElCHECK VALVE

0
'00

,° '1

! .....: . j

LIQUID OXYGENGASEOUS OXYGEN--------------~LOX CONVERTER (REMOVABLE)
IF-16A-l-ll03X®

Figure 1-67. (Sheet 2)

1-179

T.O. 1F- 16A- 1

COMMUNICATIONS SYSTEMI
I•

The aircraft communications system consists of a
LESS ~ HAVE QUICK UHF radio, ~ UHF, a VHF
radio, ~ [8J an HF radio, a radio channel/frequency
indicator (RCF!), a COMM control panel, LESS lBNl
and provisions for a secure voice system.

When the HQ radio is off, the WOD data is not lost,
but is stored in nonvolatile preset memory. Whenthe radio is initially turned on, the WOD must betransferred to the volatile memory. To transfer WOD
data to the volatile memory, select preset mode andrepeat the steps described above.

ANTENNA LOCATIONSSingle WOD transfer is done automatically at power up if the single WOD method was used
for loading.

Refer to figure 1-68.

Multiple Word of Day (MWOO)HAVE QUICK (HQ) SYSTEM LESS ~The HQ system provides normal and antijamming,
air-to-air, and air-to-ground UHF communication
capability. The usual operating mode for an HQ UHF
radio is in the normal mode where the radio uses 1 of
7000 channels. The antijamming (AJ) mode uses afrequency hopping scheme to change the channel orfrequency many times per second. Because theparticular frequency used at any instant depends on
the precise time of day (TOD), both participating HQ
UHF radios must have clocks which are synchronized. In addition, the HQ UHF radio uses word ofday (WOD) and net number in the AJ mode.
Single Word of Day (WOO)

The WOD entry is normally entered before flight, but
it is possible to enter it in flight. WOD is entered byusing the six preset channels 15-20. The WOD
elements are entered with the manual frequency
knobs and the PRESET button. For a new WOD
entry, start at preset channel 15 and enter an element
using the same method as in entering preset
frequencies in the normal mode. Progressively selectthe next higher preset channel (16-20) and continue
to enter WOD elements into preset memory. A shorttone is heard when the PRESET button is depressed
for channel 20. This tone indicates that the WOD
element for channel 20 was accepted and transferredto the radio. The WOD is stored in the nonvolatile
preset channel memory until a WOD transfer occurs.
This procedure overwrites preset channels 15-20.
Single WOO Transfer

The WOD must be transferred from the nonvolatile
preset channel memory to the volatile WOD memoryfor HQ operation. Progressively select the next lower
channel (19-15) until a double beep is heard atchannel 15. A double beep indicates that all WOD
elements have been transferred to the volatilememory.
1-180

Change 12

I HQ 11.2 I

MWOD capability allows for up to six WOD's anddates to be stored at one time into nonvolatile MWOD
memory. If the radio is turned off or power is lost
after entry ofMWOD and date codes, the data is not
lost; therefore, the information remains intact until
manually changed. The six most recent WOD's are
retained. If a WOD with a duplicate date is entered,
the new entry takes precedence. Mter the MWOD
information is entered, the operational date must beentered so the radio can select the proper WOD. If
power is lost, the operational date is lost and must
be reentered. However, if a TOD is received, theoperational date is automatically entered as part of
the TOD message. To access MWOD load, selectpreset mode, enter 220.025 in channel 20, anddepressI HQ I I I PRESET, I HQ I I . 2 I LOAD
(M-LOAD is displayed) button. Next, select manualmode and load word of day elements in channels
20-15
by I HQ II I depressing the TONE button I HQ I I . 2 I positioning the T-TONE switch to
TONE. After each entry is made, a tone is heard. Afterthe first WOD is entered, the corresponding date codeis entered by selecting channel 14. The date code isrepresented by the format 3AB.OOO, where AB equals
the day of the month; e.g., 315.000 indicates the 15thof the month. Press the I HQ II I TONE button,
I HQ 11.2 I T-TONE switch to complete the dateentry. A double beep indicates the date code has beenentered. When using MWOD, the current date must
be set into channel 1 prior to receiving TOD or thecurrent date must be part of the received TOD
message. Without date information, the radio cannot
select the current WOD, and a steady tone results
when entering AJ mode. Mter entering WOD(s)and/or setting operational date, select 220.000, presetchannel 20, and depress PRESET button to enter theverify/operate mode. The UHF radio is inoperative
when in any mode other than verify/operate. MWOD
capability can be determined by selecting 220.000,
channel20
presetmode,
andpressing
the I HQ II I PRESET, I HQ II. 2 I LOAD (M-LOAD
is displayed) button. Select channel 19; if one or two
beeps are heard, the radio is not equipped with
MWOD capability. If no beep is heard, the radio has
MWOD capability. Verify a WOD for a particular date

T.O. 1F- 16A- 1

Antenna Locations (Typical)

TACAN UPPER

IFF, UHF UPPER

IAOIMISSILE TUNINGANTENNA (RIGHTAND LEFT SIDE)RADAR

GLIDE SLOPE/LOCALIZERRADAR ALTIMETERBLOCK 10-..../'~~~~---;.=----- THREAT WARNINGBLOCK 10 IFF, [AQj(4)BLOCK 10 UHF LOWERBLOCK 15
THREAT WARNINGBLOCK 15
IFF, UHF LOWER

TACAN
LOWER
1 F-16A-l-0052A @Figure 1-68.

1- 181

T. O. 1F- 16A- 1

is loaded by selecting channel 20, manual mode, andby selecting the date to be verified with the manualfrequency knobs. Select CHAN 19, then CHAN 20. A
single beep signifies that a WOD for that date ispresent. The single WOD entry procedure remains
valid for radios that are equipped for MWOD
operation. Preset channel 20 is reserved for MWOD
mode selection.
Automatic MWOD Loading I HQ II . 2 I

The HAVE QUICK II PHASE II radio allows
automatic MWOD loading with a KYK-13/TSEC filldevice. Insure radio power is on. Lift access door toreveal FILL connector and insure mode switch on filldevice is in OFF position. Install fill device in FILL
connector and place mode switch on fill device to ON.
Frequency/status display displays FILL. Selectdesired channel on fill device and depress LOAD
button. A series of beeps is heard and thefrequency/status display displays WOD OK ifWOD
is valid and successfully received. IfWOD is invalid
or unsuccessfully received, the frequency/status
display displays BAD. If BAD is displayed, the filldevice must be disconnected and reloaded and theloading procedure must be repeated. If WOD OK isdisplayed, repeat the loading procedure for subsequent channels. Insure WOD OK is displayed after
each WOD is loaded. The CHAN display displays
CHAN 1. Operational date may be loaded bypositioning the frequency switches to OP DATE andsetting the T-TONE switch to TONE. After all
WOD's are loaded, place mode switch on fill device to
OFF and disconnect FILL connector. Close accessdoor. Radio returns to previous mode and thefrequency/status display displays previous settings.
Time-ot-Day (YOD) Transmission

The TOD entry is normally entered before flight, but
it is possible to enter it in flight. It is possible totransmit timing information in both normal and AJ
modes by momentarily pressing the I HQ II I TONE
button, IHQ II. 2 I T-TONE switch. In the normalmode, a complete TOD message is transmitted,
while in the AJ mode, only an abbreviated time
update is transmitted. A time transmission allows atime update if one radio has drifted out ofsynchronization.
Time-of-Day (TOO) Reception

Reception is possible in both normal and AJ modes.
The radio automatically accepts the first TOD
message received after the radio is turned on.
Subsequent messages are ignored unless the T
position is selected with the I HQ I I I A-3-2-T

1-182

Change 12knob, IHQ I I . 2 I T-TONE switch. The radio thenaccepts the next TOD update in either normal or AJ
mode, provided the TOD update arrives within 1
minute of the time the T position has been selected.
To receive time in the normal mode, position
the IHQ II I A-3-2-T
knob, I HQ 11.2 IT-TONE
switch to T and return to a normal channel in either
manual or preset mode. I HQ II I To receive a time
update in AJ mode, rotate the A-3-2-T knob to the T
position and then back to the A position. IHQ I I . 2 I To receive a time update in AJ
mode, rotate A-3-2 knob to A and place T-TONE
switch to T momentarily. A TOD update (time tick)can now be received on the selected AJ net for 1
minute.
Net Number

After TOD and WOD are entered, any valid AJ net
number can be selected by using the manualfrequency knobs.
Antijamming (AJ) Mode Operation

An intermittent tone is heard in the headset if aninvalid AJ net is selected. A steady tone is heard if
TOD was not initially received or if WOD was notentered.

In the AJ mode, the radio can receive and process two
simultaneous transmissions on the same net.
Conferencing is controlled by the WOD loaded inchannel 19. If the WOD in channel 19 ends with 00 or
50, conferencing is possible. If the WOD in channel 19
ends with 25 or 75, conferencing is disabled.
If the function knob is set to BOTH and guard channel
jamming is encountered, switching the function knob
to MAIN negates the jamming.
HAVE QUICK (HQ) UHF RADIO LESS ~The UHF radio provides line-of-sight communications. UHF transmissions are made by holding the
UHF VHF transmit switch on the throttle to the UHF
position. Frequencies range from 225.000-399.975
MHz. The guard receiver monitors the guard
frequency of 243.0 MHz. The control panel allows
selection of 14 preset channels (19 if MWOD loading
procedure is used). The UHF radio is poweredby [@] battery bus No.2, LESS ~ the battery bus.
UHF Radio Control Panel

Refer to figure 1-69. The UHF radio control panel islocated on the the left console.

•

T. O. 1F- 16A- 1

I

UHF Radio Control Panel (Typical) LESS IAN\
18

m
m

FR:Q

7
8
9

14
15

10
11
12

17
18

5
6

13

10

1

3
4

[illS
ill .®

2JT~ ~ ~ ~

v

15

7-=

SQUELCH

3 4 5

J

=

CHAN

GOSQ

~~STO e I~~-

I

-

~

6

J

"

14~~~iDF~"AlN 8 SQUElC:~LPRESETGR~~~ 7
BOTH WF

~

9

10

13

f=TONE\OFF~ON~---~=-----\--11

10

9

8

HAVE QUICK 11.2

HAVE QUICK II
1. Preset Channel Card and Access Door
2. FILL Connector
3. CHAN Display
4. Frequency/Status Display
5. CHAN Knob
6. STATUS Button-I~(§ ~;(f5;~ O~~I

ffi[SE~1

VOL

13 12

ZER

\9]7

/I/1~ffilffiI~'::F~ '''~:®F UV),2 OF~
"IN

17

LOAD

19

ill

®

CHAN~

16

2

m
16~lmll,.~ ~1 ;/;1=)]1:;]~~J ~L 03xl'~;1

/
CH
1

1

--~-~-.

7. Manual Frequency Knobs
8. Mode Knob
9. SQUELCH Switch
10. VOL Knob
11. T -TONE Switch
12. TONE Button

13. Function Knob
14. A-3-2 Knob
15. A-3-2-T Knob
16. TEST DISPLAY Button
17. LOAD Button
18. ZERO Switch
1 F-16A-1-1 105X@

Figure 1~69.

Function Knob

Mode Knob I HQ I I . 2 I

Functions are:Functions are:• OFF - Power off.

• MNL - UHF frequency is selected by manuallysetting the five frequency knobs.

• MAIN - Power on, UHF operating on selected
frequency.
• BOTH - Normal operation, plus receiving on guard
frequency.
• ADF - Not operational.
Mode Knob I HQ I I

Functions are:
• MANUAL - UHF frequency is selected by manuallysetting the five frequency knobs.
• PRESET - UHF frequency is determined by the
CHAN knob.
• GUARD - The main receiver and transmitter areautomatically tuned to the guard frequency and theguard receiver is disabled.

• PRESET - UHF frequency is determined by the
CHAN knob.
• GRD - The main receiver and transmitter areautomatically tuned to the guard frequency and theguard receiver is disabled.
CHAN Knob

The CHAN knob permits the selection of 1 of 19(MWOD) or 14 (single WOD) preset frequencies(channels 1-19) with the mode knob at PRESET and
the I HQ II I A-3-2-T, I HQ 11.2 I A-3-2 knob
ineither 2 or 3. Preset channels used for WOD storagecannot be used as preset channels for normal radio
operation. Frequencies set for each channel can be
manually written on a channel frequency card locatedon the access door. Preset channel frequencies are set(or changed) as follows:

Change 121-183

10. 1F- 16A-1• Function knob - MAIN or BOTH.
• Mode knob - PRESET.
• Manual frequency knobs - Set to desired frequency.

SQUELCH Switch

Functions are:
• ON - Enables squelch circuit which helps toeliminate background noise in normal reception.

• CHAN knob - Set to desired channel.

• OFF - Disables squelch circuit to permit receptionof a weak signal.

• Lift access door.

TONE Button I HQ I I• Depress I HQ I I I PRESET, I HQ I I . 2 I LOAD
button under access door.

Depressing the TONE button in normal or AJ modeinterrupts reception and transmits a tone signal and
TOD for HQ on the selected frequency. Simultaneously pressing the TONE button in conjunction
with the A-3-2-T knob in T position starts the
emergency startup of the TOD clock. Pressing the
TONE button with channels 1, 14, 15-20 selected inmanual mode enters MWOD data when in load mode.
Pressing the TONE button in erase mode erases all
MWODdata.

Manual Frequency Knobs

The five manual frequency knobs allow manual
selection of frequencies in steps of 0.025 MHz from
225.000-399.975 MHz.
A-3-2-T Knob I HQ I I I

CHAN Display I HQ I I . 2 I

Functions are:

The CHAN display displays selected channel number
when mode knob is positioned to PRESET. If in
M-LOAD or FMT.CNG operating mode, the selected
memory location is displayed. Display is blank whenthe mode knob is positioned to MANUAL or GUARD.

• A - Selects AJ mode.
• 3 - Allows manual selection of frequencies.

Frequency/Status Display I HQ II. 2 I• 2 - Allows manual selection of frequencies.
• T - Momentary position which enables the radio toaccept a new TOD for up to 1 minute after selection.
Also used in conjunction with the TONE button for
emergency startup of the TOD clock when TOD is
not available from external sources.
A-3-2 Knob I HQ I I . 2 I

Functions are:
• A - Selects AJ mode.
• 3 - Allows manual selection of frequencies.
• 2 - Allows manual selection of frequencies.

The frequency/status display displays several conditions of radio operation. Four programming modes
control AJ operation. These modes are VE RIO P,
M-LOAD, ERASE, and FMT.CHG. These modes areaccessed by placing the CHAN knob to 20, rotating
the mode knob to PRESET, and keying in theappropriate frequency for the applicable mode.
Frequencies are 220.000 for VERJOP, 220.025 for MLOAD, 220.050 for ERASE, and 220.075 for
FMT.CHG. VER/OP is displayed when the STATUS
button is depressed. M-LOAD is displayed when
MWOD's are loaded manually. ERASE is displayedwhen WOD's are erased by depressing the TONE
button or by selecting O. FMT.CHG is displayed whenfrequency management training net frequencies arechanged by selecting a channel with the mode knob
and changing frequencies with the manual frequency
knobs.
Other displays are:

VOL Knob

The VOL knob controls the volume of the audio
signal.

1-184• FILL - A fill device is connected to the FILL
connector to load MWOD's.
• BAD - A WOD was unsuccessfully received.

T. O. 1F- 16A- 1• WOD OK - A correct WOD has been successfullyreceived.

STATUS Buffon I HQ 11.2 I
When the STATUS button is depressed, selected
preset channels and manual frequencies may beverified on proper displays.

UHF Radio Control Panel

Refer to figure 1-69.1. The UHF radio control panel islocated on the left console.
Function Knob

Functions are:

T-TONE Switch I HQ II. 2 /• OFF - Power off.
Functions are:• T - Momentary position which enables the radio
to accept a new TOD for up to 1 minute after
selection. Also used in conjunction with the TESTDISPLAY button for emergency startup of the
TOD clock when TOD is not available fromexternal sources.
• TONE - In normal or AJ mode, the TONE positioninterrupts reception and transmits a tone signal
and TOD for HQ on the selected frequency.
Selecting TONE with channels 1, 14, 15-20
selected in manual mode enters MWOD data whenin LOAD mode. Selecting TONE when in ERASE
mode erases all MWOD data.

TEST DISPLAY Buffon I HQ I I . 2 I
Depressing the DISPLAY TEST button illuminates
all segments of the frequency/status display and theCHAN display for a functional test.

ZERO Switch IHQ 11.2 /Depressing the ZERO switch erases all MWOD data
regardless of radio switchlknob positions or radio
operating mode.
FILL Connector I HQ 11.2 I

The FILL connector connects fill device to the radio.
The FILL connector allows automatic MWOD
loading.
UHF RADIO~

The
UHF
radio
providesline-of-sightcommunications. UHF transmissions are made byholding the UHF VHF transmit switch on the throttle
to the UHF position. Frequencies range from
225.000-399.975 MHz. The guard receiver monitors
the guard frequency of 243.0 MHz. The control panel
allows selection of 20 preset channels. The UHF radiois powered by the battery bus.

• MAIN - Power on, UHF operating on selected
frequency.
• BOTH - Normal operation, plus receiving on guard
frequency.
• ADF - Not operational.
Mode Knob

Functions are:
• MANUAL - UHF frequency is selected by manuallysetting the five frequency knobs.
• PRESET - UHF frequency is determined by thepreset channel knob.
• GUARD - The main receiver and transmitter areautomatically tuned to the guard frequency and theguard receiver is disabled.

CHAN Knob
The CHAN knob permits the selection of 1 of 20 preset
frequencies with the mode knob at PRESET.
Frequencies set for each channel can be manually
written on a channel frequency card located on theaccess door. Preset channel frequencies are set (orchanged) as follows:• Function knob - MAIN or BOTH.
• Mode knob - PRESET.
• Manual frequency knobs - Set to desired frequency.
• CHAN knob - Set to desired channel.
• Lift access door.
• Depress PRESET button under access door.

Change 121-184.1

T.O. 1F- 16A-l

UHF Radio Control Panel (Typical) IANI

1. Preset Channel Card andAccess Door
2. CHAN Knob
3. Manual Frequency Knobs
4. Mode Knob
5. SQUELCH Switch
6. VOL Knob (inactive)
7. TONE Button
8. Function Knob
9. PRESET Button

9

8

7

6

5

4
1F-16AN-1-0074X ®

Figure 1-69.1.

1- 184.2

Change 1210. 1 F- 16A- 1
Manual Frequency Knobs

The five manual frequency knobs allow manual
selection of frequencies in steps of 0.025 MHz from
225.000-399.975 MHz.

ANT SEL Panel[8J I B FI (Typical)

VOL Knob

The VOL knob controls the volume of the audio
signal.

SQUELCH Switch

Functions are:
• ON - Enables squelch circuit which helps toeliminate background noise in normal reception.
• OFF - Disables squelch circuit
unhampered use of a weak signal.

topermit
1. IFF ANT SEL Switch
2. UHF ANT SEL Switch
IF-16A-l-ll06X®

TONE Button
Figure 1-70.

Depressing the TONE button interrupts reception
and transmits a tone signal on the selected frequency.

UHF ANT SEL SWITCH ~ ~Refer to figure 1-70. The three-position UHF ANTSEL switch is located on the ANT SEL panel.
Functions are:
• UPPER - Upper antenna is used to receive andtransmit signals.
• NORM - The antennas cycle between upper andlower to provide omnidirectional antenna pattern.
• LOWER - Lower antenna is used to receive andtransmit signals.

VHF RADIOThe VHF radio provides line-of-sight voice communication. VHF transmissions are made by holding the
UHF VHF transmit switch on the throttle to the VHF
position. Transmission and reception are available
for AM from 116.000-15l.975 MHz and for FM from30.000-87.975 MHz. Only reception is available from108.000-115.975 MHz. Twenty channels may bepreset. Operation may be either on narrow band or
wide band. Narrow band is used for all normal
operations LESS ~ and wide band is automatically •selected for secure voice.
VHF Radio Control Panel

Refer to figure 1-7l. The VHF control panel is locatedon the right console.

Change 121-185

10. 1F- 16A- 1

VHF Radio Control Panel (Typical)[]] mE] BLOCK 15[]] mE] BLOCK 10
aD
aD

®

B]

0

OJ

@]

~.

~

1

0 0 O·

2

®

3
4

®

~

@.

cI

I

2QJ0rau11

o 0 O·~I I 2PR£SE~313'" I,~, II" ::~12
I

DD

5

14

rnffi BLOCK 15

r-ru

f1l

L2J

1. Manual Frequency Indicator
2. Manual Frequency Knobs
3. Preset Channel Indicator
4. Preset CHAN Knob
5. Preset Channel Card and Snap-On Cover
6. FM SQUELCH Control (Preset)
7. Band LOCK OUT Switch (Preset)

4

8. WB MEM LOAD Switch
9. AM SQUELCH Control (Preset)
10. VOL Knob
11. SQ DIS TONE Switch
12. Function Knob
13. Mode Knob
14. LOAD Button
1 F-16A-1-11 07X@

Figure 1-71.

Function Knob

• Preset CHAN knob - set to desired channel.

Functions are:• ~ [B] ~ LOAD button - depress momentarily.

• OFF - Power off.

• LESS ~ [B] ~ and ~ ~ Remove snap-on
cover and position bandwidth!M:EM LOAD switch
to MEM LOAD and release.

• TR - Power on, VHF operating on selected
frequency.

Mode Knob• DF - Not operational.

Functions are:Preset CHAN Knob

The preset CHAN knob permits selection of presetchannels with the mode knob at PRE. Frequencies set
for each channel can be manually written on a presetchannel card located on a snap-on cover.
Preset channel frequencies are set (or changed) as
follows:• Function knob - TR.
• Mode knob - MAN.
• Frequency knobs - set to desired frequency.
1-186• EMER AM - Provides reception and transmission
on a preset guard frequency between
119.000-124.000 MHz. To select this function,
first set the function knob to TR and then set themode knob to EMER AM.
• EMER FM - Provides reception and transmission
on a preadjusted guard frequency between
38.0-43.0 MHz. To select this function, set thefunction knob to TR and the mode knob to EMERFM.
• MAN - Frequency selected by the manualfrequency knobs.
• PRE - Frequency selected by the preset channelknob.

10, IF-16A~1
Manual Frequency Knobs

WB MEM LOAD Switch rn:Bl

The four manual frequency knobs allow manual
selection of frequencies with the mode knob in MAN.

Functions are:
• WB - Wideband is automatically selected for secure
voice.

VOL Knob

The VOL knob controls the volume of the audio
signal.

• NB - This position is used for all normal operations.
• MEM LOAD - Switch may be momentarily held inthis position to set a manual frequency into a presetchannel. When the switch is released, it automatically returns to the NB position.

SQ DIS TONE Switch

Functions are:
• SQ DIS - Disables the squelch. Squelch remains
disabled until the switch is returned to the center
position.

HF RADIO I8Q] [BJ• TONE - Switch may be momentarily held in thisposition to transmit a 1000 Hz tone. When theswitch is released, it automatically returns to thecenter position.

The HF radio provides over-the-horizon communication using voice operation. HF transmissions are
made by holding the VHF HF UHF transmit switch
on the throttle to the HF position. Frequencies range
from 2-29.9999 MHz. The control panel allows
selection of 20 preset channels. The HF radio is
powered by the essential dc bus No.2.

LOAD Button B3 [A] [M]

HF Radio Control Panel ~ [A]The LOAD button may be depressed momentarily to
set a manual frequency into a preset channel.

Refer to figure 1-72. The HF control panel is locatedon the right console.

• Center Position - Normal squelch operation.

HF Radio Control Panel fADl ~
3

4

6

5

~~~

H

F

USB

LOAD

12

11

10

OFF

~ OF~~
i

i

i
9

8

I

7

1. Preset Channel Select Knob
2. CHAN Display
3. PRE Tune Light
4. XMIT Light
5. MAN Tune Light
6. FREQ Display
7. Frequency Select Switches
8. OFF VOL Control Knob
9. OFF SOL Control Knob
10. LOAD TEST Switch
11. USB AM LSB Select Switch
12. UNTUNED Light
1F-16A··1-1108X®

Figure 1-72.

1-187

I

I

T.O. 1F- 16A- 1
OFF VOL Control Knob

Frequency Select Switches

Moving the OFF VOL control knob cw out of the OFF
detent turns the HF radio on and controls the volumeof the audio signal.

Five three-position spring-loaded frequency selectswitches select frequencies from 2-29.9999 MHz.
Upward switch movement increases the numericalvalue and downward movement decreases the value.
Holding the switch up or down for more than 1 second
slews the number.

OFF SQL Control Knob

Functions are:• OFF - Disables squelch to allow all received signals
to be heard in the headset.
• SQL - The seven detents cw from the OFF position
enable increasing levels of squelch. Each progressive position requires a stronger received signal inorder to break the squelch and be heard in theheadset.

FREQ Display

The FREQ display has a dual function. The six
illuminated indicators show the selected numericfrequency in MHz during normal operation. The fouralphanumeric characters to the right of the frequency
decimal point also show system fault codes.
LOAD TEST Switch

USB AM LSB Select Switch

Functions are:• USB - Selects upper sideband of carrier frequency
produced by modulation.
• AM - Selects amplitude modulation of radio carrier
frequency.
• LSB - Selects lower sideband of the carrier
frequency produced by modulation.
Preset Channel Select Knob

The rotary preset channel select knob is acontinuously turning knob which individually selects
up to 20 preset frequencies, increasing in number as
the knob is rotated cwo To set a preset channel:

The three-position LOAD TEST switch presets
selected frequencies when placed to the LOAD
position, clears PBIT-detected fault codes from the
FREQ display when momentarily placed to TEST,
and starts IBIT when placed to the TEST position.
The switch must be returned to center at thecompletion of IBIT to clear the display.
PRE Tune Light

The PRE tune light illuminates to show that thefrequency selected in the FREQ display is a presetchannel frequency.
XMIT Light

The XMIT light illuminates to show that the radio istransmi tting.

• Turn OFF VOL control knob on.
• Turn the preset channel select knob to the channel
number to be set.
• Move the frequency select switches to select thedesired frequency.
• Momentarily move the LOAD TEST switch to the
LOAD position, then release.
CHAN Display

1\vo illuminated indicators on the CHAN display
show the selected channel from 1-26.

1- 188

MAN Tune Light

The MAN tune light illuminates to show that thefrequency selected in the FREQ display is manually
selected using the frequency select switches.
UNTUNED Light

The UNTUNED light illuminates to show that the
selected active frequency does not have tuning datastored. The sequence to store tuning data is initiatedby keying the radio for a few seconds. The UNTUNED
light goes off when tuning data is stored.

T. o. 1F- 16A- 1
Built-In Test (BIT) Function

The HF radio has a periodic or PBIT and initiated or
IBIT function. Placing the LOAD TEST switch to
TEST starts the initiated built-in test (lBIT) function.
All indicators illuminate briefly for a 4-second visualdisplay check, while the control panel conducts
various internal checks. If no control panel faults
exist, the word TEST appears on the FREQ display as
the receiver/transmitter (RT) signals the start of the
22-second IBIT sequence. If no digits, invalid digits,
or fault code C appears on the FREQ display, a control
panel fault exists and all further IBIT's stop. IBIT
failure displays consist of a two-letter fault code for
either one of four line replacement units (LRU's), anda single numeric fault code to indicate, whenapplicable, failure of shop repairable units (SRU's),
which are subassemblies of LRU's.

Radio Channel/Frequency Indicator(Typical)

The associated fault codes are:
1. FREQ Pushbutton Switch
2. UHF Radio Transmission Indicator Light (White)
3. VHF Radio Transmission Indicator Light (White)
4. Radio Channel/Frequency Indicator
5.
UHF VHF Select Pushbutton (Unmarked)• RT - Receiver transmitter.
• CD - Coupler.

m

1F-16A-1-0057X@• RD - RF feedline (interface).
Figure 1-73.

• C - Control panel.
The numeric SRU fault code appears when any of the
RT or CU SRU's fail, in conjunction with the fault
code for the applicable LRU. There are no SRU fault
codes for the RF feedline or control panel. If no faults
exist at the completion of IBIT, the FREQ display will
read PASS until the LOAD TEST switch is returnedto the off position.
The PBIT is a less extensive, continuous self-test that
occurs during normal radio operation. The LRU fault
code is the only code displayed if the PBIT detects afault. Moving the LOAD TEST switch from TEST to
off clears the fault code from the FREQ display,allowing the previously selected frequency toreappear. It also assures that this same fault code
does not reappear unless the radio is turned off, thenback to on.

RADIO CHANNEL/FREQUENCY INDICATOR (RCFI)

Refer to figure 1-73. The RCFI, located on the
instrument panel, displays the selected channel orthe manually selected frequency. When presetchannel is displayed, the corresponding frequency
may be verified by depressing the FREQ pushbutton.

~

With UHF radio in operation, depressing the UHF
VHF pushbutton (unmarked) displays either UHF or
VHF channel/frequency as indicated by the U or V
indicator light.
LESS ~ Two small indicator lights on the right sideof the indicator display a U or V during each radiotransmission; however, the RCFI does not display a
VHF channel/frequency. The UHF radio must be on
for the indicator lights to be displayed.

COMM CONTROL PANEL [AJ rnfl

Refer to figure 1-74. The COMM control panel islocated on the left console.
HOT MIC CIPHER Switch

Functions are:• HOT MIC - Provides an intercom for pilot to groundcrew or for AR. Activation of UHF VHF transmit
switch on the throttle overrides this function.
• OFF - Disables HOT MIC CIPHER.
• CIPHER - Momentary position which limits
reception to secure voice only. This function is
operable only when radio is operating in secure
voice mode.
1-189

T. O. 1F- 16A- 1

COMM Control
Panel [8J I BF I (Typical)

MSL Tone Knob

The MSL tone knob has a cw arrow pointing to INC.
Rotating the knob cw increases the volume of the tone
from the AIM -9 missile being monitored.
MASTER ZEROIZE SWITCH [IDmERefer to figure 1-3. The MASTER ZEROIZE switch
is a guarded two-position toggle switch. When
positioned to MASTER ZEROIZE, coded electronic
information associated with the following equipment
is purged:

2

3

4

• LESS ~ Secure voice.
• LESS M IFF mode 4.

I• DTC (if installed).
1. HOT MIC CIPHER Switch
2. INRCOM/SV Knob
3. THREAT Tone Knob
4. MSL Tone Knob

• ~ AIFF mode 4.
• ~ Seat ejection also automatically performs anescape zeroize operation; refer to EJECTION SEATOPERATION, this section.

1 F- 1 6A-l-0058X @Figure 1-74.

INRCOM/SV Knob

The INRCOM/SV knob has a cw arrow pointing to
INC. CW rotation increases the volume of both the• intercom LESS ~ and secure voice systems.
Functions are:
• Monitoring and volume control of voice communication between pilot and ground crew or tanker boom
operator, AIM-9 missile tone, and TWS composite
audio tone.
• Monitoring of systems individually volume controlled from the respective UHF, VHF, TACAN, ILS,
and ~ [!] HF control panels.

I• Monitoring of fixed volume warning tones (LG andlow speed warning tone, TWS missile launch tone,
LESS ~ and IFF mode 4 audio monitor) and voicemessages.

SECURE VOICE SYSTEM LESS IANI
The secure voice system is used in conjunction with
the UHF and VHF radios to provide secure voicecommunications. Operation requires the selected
UHF or VHF radio to be operating and the desiredchannel selected. The system enciphers voicemessages before they are transmitted and deciphers
voice messages which are received. Normal radio
procedures (turn-on, warmup, and channel selection)
are not otherwise affected by the secure voice system.
SECURE VOICE CONTROL PANEL [!] ~Refer to figure 1-75. The secure voice panel is locatedon the right console.
POWER Switch

The POWER switch is a two-position lever lockswitch. The ON position enables the secure voicesystem.
ZEROIZE Switch

THREAT Tone Knob

The THREAT tone knob has a cw arrow pointing to
INC. CW rotation increases the volume of the TWS
composite tone.

1-190

Change 12

Refer to figure 1-3. The ZEROIZE switch is a guarded
lever locked switch. When activated, the switch
purges all variables of the secure voice system,
disabling secure voice operation,IT.O.1F-16A-1

Secure Voice Control
Panel ~ IBF I (Typical)

PLAIN Cipher Switch

The PLAIN cipher switch is a three-position rotaryswitch.
Functions are:
• CIRAD 1 - Provides secure voice operation withUHF.

1

• PLAIN - Allows normal radio communications.

m1
KY
58
R

• CIRAD 2 - Provides secure voice operation with
VHF.

C
U

During CIRAD 1 or CIRAD 2 operation, plain text
messages can still be received when the HOT MIC
CIPHER switch (located on the COMM control panel)
is not in the CIPHER position.
1. PLAIN Cipher Switch
2. MODE Switch
3. POWER Switch (Lever Lock)
4. FILL Switch
5. DELAY Switch (Lever Lock)
6. ZEROIZE Switch (Guarded)
1F-16A-1-0059X®

Figure 1-75.

DELAY Switch

The DELAY switch is a two-position lever lock switch.
The delay function is not used.

When operating in secure voice (CIRAD 1 or CIRAD
2), the volume of the received messages is controlled
by the COMM control panel. The radio set volume
control is disabled.
When in CIRAD lor CIRAD 2, do not select GUARD
position on the UHF radio or EMER AMlFM on the
VHF radio. This action may result in permanent lossof power to the radio selected.
For normal radio operation, the PLAIN cipher switch
must be in PLAIN even though the POWER switch isin OFF. If the PLAIN cipher switch is in the CIRAD
1 position, there is no UHF or VHF side tone. If the
PLAIN cipher switch is in the CIRAD 2 position, theVHF radio is inoperative.

FILL Switch

INERTIAL NAVIGATION SET (INS)The FILL switch is a six-position rotary switch. Eachposition selects the memory location that will be used
for the selected operating mode.

The INS is a prime sensor for aircraft velocity,
attitude, and heading and is a source of navigation
information. The INS consists of the INU, the FCNP,
and the INU battery.

MODE Switch

The MODE switch is a three-position rotary switch.

The INS in conjunction with the CADC and FCC I
provides:Functions are:

• Present position with update and storage capability.
• Current winds.

• OP (operational) - Normal operation for transmitting and receiving plain and cipher messages.

• Groundspeed and drift angle.

• LD (load) - Allows the secure voice system to accept
a variable from a fill device or another KY-58.

• Great circle course computation with steering
provided to 32 points.
•• RV (receive variable) - This position is used in• Instantaneous and maximum g data for display in
the HUD.

remote keying operations.

1- 191

T.O. 1 F-16A-1

I

Refer to T.O. IF-16A-34-1-1, ~ T.O. IF-16A-34-l-3
or 1M!] T.O. IF-16A-34-l-4 for a detailed description.

TACAN Function KnobTACTICAL AIR NAVIGATION (TACAN)• OFF - Power off.

SYSTEM• REC - Receive mode. The system receives signals
which result in a bearing and course deviation
display on the HSI and audio in the headset.
TACAN range window (MILES) on HSI isshuttered.

The TACAN system provides continuous bearing anddistance information from any selected TACAN
station within a line-of-sight distance up to
approximately 390 miles, depending upon terrain
and aircraft altitude. Only distance information ispresented when a DME navigational aid is selected.
There are 252 channels available for selection. Twoantennas, one on top and one on bottom of thefuselage, provide omnidirectional coverage regardless of aircraft attitude.
The TACAN works in conjunction with the INSTR
MODE select panel and the HSI and through the
COMM control panel for audio output. TACAN
information is presented on the HSI.
TACAN CONTROL PANEL

Refer to figure 1-76. The TACAN control panel islocated on the left console.

TACAN Control
Panel (Typical)

Functions are:

• TIR - Transmit/receive mode. Same as REC and, in
addition, interrogates the TACAN ground station
for DME information; distance (nm) is displayed in
the HSI range window (MILES).
• AlA REC - Air-to-air receive mode is not used.
• AlA TIR - Air-to-air transmit-receive mode.
TACAN system interrogates and receives signals
from aircraft having air-to-air capability, providing slant range (nm) distance between aircraft
operating 63 TACAN channels apart. (KC-lOA also
provides bearing information.) Up to five aircraft
can determine distance from a sixth lead aircraft.
Lead aircraft can only determine distance fromone aircraft. Audio identification is not provided.
VOL Knob

The VOL knob controls the volume of the audio
signal.
X/Y Channel Ring

The Xry channel ring is the outer portion of theunits knob. The ring provides for the selection of X
or Y channels as viewed in the channel display.
Each mode has 126 channels available.
Units Channel Knob

The units channel knob is used to select the units digitof the channel (0-9) as viewed in the channel display.
Hundreds-Tens Channel Knob

The hundreds-tens channel knob is used to selectthe tens and hundreds digits (00-12) as viewed in
the channel display.
CHANNEL Display
1. CHANNEL Display
2. VOL Knob
3. Function Knob
4. X/Y Channel Ring
5. Units Channel Knob

6. TEST Pushbutton
7. Test Light (Red)
8. Hundreds-Tens
Channel Knob

Figure 1-76.

1-192

Change 13

1 F-16A-1-1112X@

The CHANNEL display displays the digital readoutof the selected TACAN channel.
TEST Pushbutton

The TEST pushbutton initiates system self-test when
depressed momentarily. The self-test checks theentire system except the antennas.

T. O. 1F- 16A- 1
TEST Light

The TEST light illuminates when a malfunction
occurs during manual or automatic system self-test.
If the system fails self-test in the TIR mode but not in
the REC mode, the TACAN may be used for bearing
information.

INSTRUMENT LANDING SYSTEM (ILS)The ILS provides precision approaches to runways
equipped with localizer, glide slope, and marker
beacon equipment. Localizer identification signals
are supplied to the headset for station identification.
The glide slope and localizer receivers supply glide
slope and localizer deviation data to the deviationbars on the ADI and HUD; the HSI also displays
course or localizer deviation data. Two warning flags,
designated LOC and GS, appear on the ADI when
deviation data is invalid. A course deviation warning
flag appears on the HSI if localizer deviation data isinvalid. HUD symbology consists of localizer andglide slope deviation bars. Dashed deviation bars
indicate invalid data. Deviation bars are rollstabilized with tic marks positioned at the one-dot
and two-dot deflections.
The symbology automatically displayed on the HUD
with ILS selected is the same as LG down with thefollowing exceptions:• The great circle steering symbol is replaced by the
ILS deviation bars.
• The lower HUD windows, except distance-to-destination, are not blanked unless the NLG is lowered
and inertial velocity exceeds 80 knots.
• The HUD altitude scale does not change from
100-foot increments to 20-foot increments until
NLG is lowered.
• AOA bracket is not displayed until NLG is lowered.
The flight director displays command steering dataon the HUD when selected on the FCNP. Commandsteering symbology consists of a circle, a tic markpositioned at the top of the circle, and a referencemark/caret positioned at the heading/ground trackscale. The flight director circle is referenced· to the
FPM and appears when localizer data is valid. Properuse of the flight director requires that the localizer beintercepted from a heading no more than 45 degrees
from the localizer course using bank angles of 30
degrees or less. When the aircraft is within two dots
deflection of the localizer deviation bar, the flight
director commands a turn to roll out on the localizer

course. The tic mark appears on the flight director
circle when glide slope deviation nears center,indicating that pitch steering data is valid. The glide
slope should be intercepted from a position that iswings level and on the localizer course. The reference
caret indicates the heading required to maintain thecourse selected (magnetic heading scale displayed) or
ground track error relative to the course selected(magnetic ground track scale displayed). Refer to ILS
PROCEDURES, Section II.
The ILS flight director is designed to intercept theglide slope from below while in approximately level
flight. If the aircraft approaches the glide slope fromabove, there is no pitch steering data. The flight
director symbol remains on the horizon, displaying
bank steering data, and the symbol X appears over
the tic mark, indicating that pitch steering data isinvalid. Valid pitch steering is provided after the
glidepath is intercepted.
The marker beacon receiver operates on a fixed
frequency of 75 MHz. Refer to MRK BCN LIGHT, thissection.
ILS CONTROL PANEL

Refer to figure 1-77. The ILS control panel is locatedon the ~ ~ right console, ~ left console.
ILS Frequency Knobs

The frequency knobs allow individual selection of 40
localizer frequencies ranging from 108.10-111.95
MHz in 0.05 MHz increments using odd frequencies(.1, .3, etc.). Each localizer frequency selected ispaired with a glide slope frequency (329.15-335.00
MHz). The localizer frequency selected is displayedon the control panel.
ILS PWR Switch

The PWR switch is an outer ring on the frequency
knob. In the OFF position, power is removed fromlocalizer, marker beacon, and glide slope receivers.When the switch is placed to PWR, power isapplied to the localizer, marker beacon, and glide
slope receivers.
ILS VOL Ring

The VOL ring is an outer ring on the frequency knob
and adjusts the volume of the localizer station
identification signal. CW rotation increases volume.
MRK BCN Light

Refer to figure 1-3. The MRK BCN light is located on
the instrument panel. When the aircraft is over amarker beacon facility, the light illuminates green
and blinks according to the code of the marker beacon.
Change 12

1-193

T. O. 1F- 16A- 1

ILS Control Panel(Typical)ILS

rTOr'O]
3
1. ILS PWR Switch
2. Frequency (Units) Knob
3. Frequency Display
4. Frequency (Tenths) Knob
5. ILS VOL Ring

1F-16A-1-0061A@

IFF ADVISORY FUNCTION

The FCC provides an IFF advisory to change IFF
codes within a determined period of time. Two times(in minutes) are displayed: the FCNP LMD displays
time remaining to initiate advisory (countdown time);
the FCNP RMD displays the set time between
advisories. The countdown time is displayedwhenever FCC power is on; the countdown time
restarts when FCC power is recycled or when the
FCNP ENT switch is depressed after a countdown
time is selected. When the countdown time reaches 0,
IFF flashes in the HUD, the VMS issues an aural IFF
advisory, the countdown time is reset to the time
between advisory value, and the countdown begins
for the next IFF code change reminder. The HUD IFF
advisory is reset by positioning the DRIFT C/O switch
to TEST; however, fuel advisories must be acknowledged and cleared prior to display of IFF advisory in
the HUD. At FCC initialization, the last value
entered for time between advisories is retained. Timebetween advisories is loaded by the DTU.
LOSE and BORE messages cannot be reset and havepriority over the IFF advisory. Refer to T'O.I
IF-16A-34-1-1, ~ T.O. 1F-16A-34-1-3 or ~ T.O.
IF-16A-34-1-4 for a detailed description.
IFF CONTROL PANEL

Figure 1-77.

IFF SYSTEM
The air-to-surface IFF system provides selective
identification feature (SIF), automatic altitude
reporting, LESS 1M!] and (mode 4) encrypted IFF.
Normal operation is possible in any of five modes:

Refer to figure 1-78. The IFF control panel is locatedon the left auxiliary console.
MODE 4 CODE Knob 1M!]

The mode 4 CODE knob is inoperative.
MODE 4 CODE Knob LESS 1M!]

• Mode I - Security identity.

The MODE 4 CODE knob must be pulled out before
it can be moved to the ZERO position and isspring-loaded from HOLD to the A position.

• Mode 2 - Personal identity.

Functions are:

• Mode 3/A - Traffic identity.
• LESS ~ Mode 4 - Encrypted identity.

• Mode C - Altitude reporting.
The equipment does not perform interrogation but
only transmits coded replies to correctly codedinterrogations. Modes 2 LESS ~ and 4 code settings
are set into the receiver-transmitter on the groundand thus are fixed for anyone flight. Modes 1 and 3/A
codes are set at the control panel. All modes can beturned on or off at the control panel. Mode C provides
altitude information from the CADC to the ground in
100-foot increments.
1-194

Change 13• HOLD - Both code settings can be retained after
flight by positioning the MODE 4 CODE knob to the
HOLD position momentarily prior to placing the
MASTER knob to OFF or removal of power.
• A and B - The positions select the preset code for A
orB.
• ZERO - Both code settings zeroize if the MASTER
knob is in any position except OFF. Both codes areautomatically zeroized when the IFF is turned off
after landing if the HOLD mode is not used.
REPLY Light ~The REPLY light is inoperative.

T. O. 1F- 16A- 1• LOW - Only local (strong) interrogations arerecognized and answered.

IFF Control Panel[8J IBF (Typical)• NORM - Full range recognition and reply occur.
Transmitted power from the IFF system is the samefor both the LOW and NORM positions.

I• EMER - The knob must be pulled outward toposition to EMER. When so positioned, an
emergency-indicating pulse group is transmitted
each time a mode 1, 2, or 3/A interrogation isrecognized.
RAD TEST MON Switch

The RAD TEST MON switch is spring-loaded fromthe TEST position to the OUT position.
Functions are:8

9
1. MODE 4 CODE Knob
2. REPLY Light (Greenl
3. TEST Light (Green)
4. MASTER Knob
5. RAD TEST MON Switch

7

6. Mode TEST Switches (4)7. IFF IDENT Switch
8. Code Wheels
9. MODE 4 Switch
10. MODE 4 Monitor Switch• OUT - The radiation test and monitor circuits areinoperative.
• TEST - Used for preflight check.
• MON - Used with the built-in test capability.
IFF IDENT Switch

1 F- 1 6A-l-l 1 14X@

Functions are:Figure 1-78.

•

REPLY Light LESS ~The REPLY light illuminates green to indicate mode
4 replies. The light is operative only when the mode
4 monitor switch is in AUDIO or LIGHT.
TEST Light

When a mode TEST switch is held in TEST, the TEST
light illuminates green to indicate the mode isoperable. When any of the mode test switches areplaced to ON and the RAD TEST MON switch is inMON, the light may light because of externalinterrogations.
IFF MASTER Knob

Functions are:

I · OFF - Removes power from the equipment LESS~
and also zeroizes mode 4 code settings unless the
HOLD function is used. The knob must be pulled
outward to rotate from STBY to OFF.
• STBY - The equipment is turned on and warmed up
but does not transmit.

• IDENT - Spring-loaded out of IDENT. When so
positioned, the lIP timer is energized for 15-30
seconds. If a mode 1, 2, or 3/A interrogation isrecognized within this 15-30 second period, lIP
replies are made.
• MIC - When UHF VHF transmit switch on throttle
is placed to UHF, lIP timer is energized for 15-30
seconds. If a mode 1, 2, or 3A interrogation ispresent, lIP replies are made.
• OUT - Transmission of lIP replies is prevented.
Mode TEST Switches

Four mode TEST switches, located on the IFF control
panel, are labeled M-1, M-2, M-3/A, and M-C from left
to right to correspond to mode 1, mode 2, mode 3/A,
and mode C. Each switch has positions of TEST, ON,
and OUT. The OUT position disables the transmitterreceiver for the mode selected. The ON position
enables the transmitter-receiver to reply tointerrogations for the mode selected. If more than one
switch is placed to ON, the transmitter-receiver
replies to interrogations for all modes selected. The
switches are spring-loaded to the ON position fromthe TEST position. Modes 1, 2, 3/A, and C can betested by holding the associated mode TEST switch in
TEST. If the TEST light comes on while the mode
TEST switch is held in TEST, the related mode isfunctioning properly.
Change 12

1-195

10. 1F- 16A- 1
Code Wheels

Functions are:Two sets of thumb wheels are labeled MODE 1 and
MODE 3/A. MODE 1 thumb wheel allows selection of
32 different codes. MODE 3/A thumb wheel provides
the capability of setting 4096 codes. Code digits oneach thumb wheel are read in windows recessed in
the face of the panel.

• UPPER - Upper antenna is used to receive andreply to interrogation signals.
• NORM - The system selects the antenna which isreceiving the best signal.
• LOWER - Lower antenna is used to receive andreply to interrogation signals.

MODE 4 Switch ~IFF IDENT Button

The MODE 4 switch is inoperative.
MODE 4 Switch LESS ~Functions are:
• ON - Mode 4 operation is enabled. Lever-locked
position.
• OUT - Mode 4 operation is disabled.
MODE 4 Mon itor Switch ~The MODE 4 monitor switch is inoperative.

The IFF IDENT button, located on the MISC panel,
provides an alternate method of initiating the lIP
function of the IFF system. Pushing the button
momentarily causes the lIP timer to energize for
15-30 seconds. If a mode 1, 2, or 3/A interrogation isrecognized within this 15- to 30-second period, lIP
replies are made. [N§ [NQ] ~ mIl The alternate IIP
function is not installed.
I FF Caution Light LESS !8Nl

The IFF caution light, located on the caution light
panel, illuminates whenever an inoperative mode 4capability is detected, provided the mode 4 computeris installed in the aircraft and the MASTER knob is
not in OFF. Specific discrepancies monitored by the
IFF caution light are as follows:

MODE 4 Monitor Switch LESS~• Mode 4 codes zeroized.
Functions are:
• AUDIO - Monitoring of mode 4 interrogations isprovided by an audio tone on the intercom andmonitoring of mode 4 replies is provided by the
REPLY light on the IFF control panel. An audio toneof 300 to 3000 Hz is generated when valid mode 4interrogations are being received. The frequency of
the tone depends on the number of interrogations
being received; i.e., the higher the number ofinterrogations, the higher the frequency. The
REPLY light illuminates green when valid mode 4
replies are being transmitted.
• LIGHT - Switches out the audio tone and provides
monitoring only by the REPLY light.
• OUT - Both the audio tone and the REPLY light are
inoperative.
IFF ANT SEL Switch ~ mIl• Failure of the system to reply to proper mode 4interrogation.
• Automatic self-test function of the mode 4 computer
reveals a faulty computer.

AIFF SYSTEM IAD/
TRANSPONDER ~The transponder provides selective identification
feature (SIF), automatic altitude reporting, and(mode 4) encrypted IFF. Operation is possible in any
of five modes:• Mode 1 - Security identity.
• Mode 2 - Personal identity.
• Mode 3/A - Traffic identity.
• Mode 4 - Encrypted identity.

The three-position IFF ANT SEL switch is located on
the ANT SEL panel.

1-196

Change 12• Mode C - Altitude reporting.

•T.O. 1 F-16A-1

The transponder does not perform interrogation but
only transmits coded replies to correctly codedinterrogations. Mode 4 code settings are set into thereceiver-transmitter on the ground and thus are fixed
for anyone flight. Mode C provides altitude
information from the CADC to the ground in 100-footincrements.
Modes 1,2, and 3/A code settings are controlled fromthe SCPO Refer to T.O. IF-16A-34-1-3 for a detailed
description of switch functions.
INTERROGATOR ~The interrogator provides interrogation of other IFF
systems. Interrogation of other IFF systems ispossible when in anyone of four modes:• Mode 1 - Security identity.
• Mode 2 - Personal identity.
• Mode 3/A - Traffic identity.
• Mode 4 - Encrypted identity.
The AIFF provides interrogation of IFF systems
within range recognition. Location and classificationof a particular aircraft as a friend or unknown ispossible by comparison SIF codes; During mode 1, 2,
or 3/A interrogations, the interrogator is commanded
to the same code as the transponder. During mode 4interrogation, the interrogator is commanded to the
M4A or M4B code as selected on the SCP and isindependent of the mode 4 (A or B) transponder code
as selected on the AIFF control panel. Selection of adifferent mode or deselection of mode 4 on the AIFF
control panel commands the interrogator to the samemode 4 transponder code.

time is selected. When the countdown time reaches 0,
IFF flashes in the HUD, the VMS issues an aural IFF
advisory, the countdown time is reset to the time
between advisory value, and the countdown begins
for the next IFF code change reminder. The HUD IFF
advisory is reset by positioning the DRIFT C/O switch
to TEST; however, fuel advisories must be acknowledged and cleared prior to display of IFF advisory in
the HUD. At FCC initialization, the last value
entered for time between advisories is retained. Timebetween advisories is loaded by the DTU.
LOSE and BORE messages cannot be reset and havepriority over the IFF advisory. Refer to T.O.
IF-16A-34-1-3 for a detailed description.
AIFF CONTROL PANEL ~Refer to figure 1-79. The AIFF control panel, locatedon the left auxiliary console, enables transponder
operation only.

AIFF Control PanellADI~IBFIThe FCC provides an IFF advisory to change IFF
codes within a determined period of time. Two times(in minutes) are displayed: the FCNP LMD displays
time remaining to initiate advisory (countdown time);
the FCNP RMD displays the set time between
advisories. The countdown time is displayedwhenever FCC power is on; the countdown time
restarts when FCC power is recycled or when the
FCNP ENT switch is depressed after a countdown

3

2

9

8
IFF ADVISORY FUNCTION ~(Typical)

5

7 6

4

1. Mode 4 CODE Knob
2. M-4 REPLY Light (Green)
3. AIFF MASTER Knob
4. M-C Switch
5. M-3/A Switch
6. M-2 Switch
7. M-l Switch
8. M-4 Switch
9. Mode 4 Monitor Switch
1F-16A-1-1115X@

Figure 1-79.

Change 13

1-197

I

T.O. 1F- 16A- 1
Mode 4 CODE Knob ~M-4 Switch [@]Functions are:
The mode 4 CODE knob must be pulled out before it
can be moved to the ZERO position and IS
spring-loaded from HOLD to the A position.

• M-4 - Mode 4 operation is enabled. Lever-locked
position• OUT - Mode 4 operation is disabled.

Functions are:Mode 4 Monitor Switch l@l• HOLD - Both code settings can be retained after
flight by positioning the mode 4 CODE knob to the
HOLD position momentarily prior to placing the
MASTER knob to OFF or removal of power.
• A and B - The positions select the preset code for A
or B.
• ZERO - Both code settings zeroize if the MASTER
knob is in any position except OFF. Both codes areautomatically zeroized when the AIFF is turned off
after landing if the HOLD mode is not used.

M-4 REPLY Light ~I

The M-4 REPLY light illuminates green when valid
mode 4 replies are being transmitted. The light isenabled only when the mode 4 monitor switch is in
AUDIO or LIGHT.

AIFF MASTER Knob ~Functions are:
• AUDIO - Monitoring of mode 4 interrogations isprovided by an audio tone on the intercom andmonitoring of mode 4 replies is provided by the M-4
REPLY light on the AIFF control panel. An audio
tone of 0 to 1500 Hz is generated when thetransponder is not replying to valid mode 4interrogations. The frequency of the tone depends
on the number of interrogations received; i.e., thehigher the number of interrogations, the higher thefrequency. The M-4 REPLY light illuminates green
when valid mode 4 replies are being transmitted.
• LIGHT - Switches out the audio tone and provides
monitoring only by the M-4 REPLY light.
• OUT - Both the audio tone and the M-4 REPLY
light are disabled.
M-l Switch ~Functions are:• M-l- Mode 1 operation is enabled.
• OUT - Mode 1 operation is disabled.

Functions are:• OFF - Removes power from the equipment and also
zeroizes mode 4 code settings unless the HOLD
function is used. The knob must be pulled outward
to rotate from STBY to OFF.
• STBY - The equipment is turned on and warmed up
but does not transmit. Interrogator functions areavailable.

M-2 Switch ~Functions are:• M-2 - Mode 2 operation is enabled.
• OUT - Mode 2 operation is disabled.
M-3/A Switch ~Functions are:• M-3/A - Mode 3/A operation is enabled.

• NORM - Full range recognition and reply occur.
• EMER - The knob must be pulled outward toposition to EMER. When so positioned, an
emergency-indicating pulse group is transmitted
each time a mode 1, 2, or 3/A interrogation isrecognized. The mode 3/A code is automaticallychanged to 7700.
1 -198Change 4• OUT - Mode 3/A operation is disabled.
M-C Switch I8Q]

Functions are:• M-C - Mode C operation is enabled.
• OUT - Mode C operation is disabled.

T. O. 1 F- 16A- 1
IFF ANT SEL Switch [8] [MI ~The three-position IFF ANT SEL switch is located on
the ANT SEL panel.
Functions are:
• UPPER - Upper antenna is used to receive andreply to interrogation signals.
• NORM - The system selects the antenna which isreceiving the best signal.

causes the lIP timer to energize for 15-30 seconds. If
a mode 1,2, or 3/A interrogation is recognized within
this 15-30 second period, lIP replies are made.
IFF Caution Light ~The IFF caution light, located on the caution light
panel, illuminates whenever an inoperative mode 4capability is detected, provided the mode 4 computeris installed in the aircraft and the MASTER knob is
not in OFF. Specific discrepancies monitored by the
IFF cau:ion light are as follows:• LOWER - Lower antenna is used to receive andreply to interrogation signals.

• Mode 4 codes zeroized.

IFF IDENT Button ~• Failure of the system to reply to proper mode 4
in terrogation.

Refer to figure 1-3. The IFF IDENT button is locatedon the MISe panel. Pushing the button momentarily• Automatic self-test function of the mode 4 computer
reveals a faulty computer.

1-199

T.O. 1 F-16A-lmagnetic heading in the event of an INS failure(indicated by the AUX warning flag on the ADD.

INSTR MODE SELECT PANEL
Refer to figure 1-80. The INSTR MODE select panel,
located on the instrument panel, provides for theselection of the displays on the HSI and the AD!.
When an ILS mode is selected by the INSTR MODE
knob, the localizer and glide slope deviation bars arealso displayed on the HUD. ILS information isprovided to the HUD from the ILS regardless of thefire control/navigation mode selected.

INSTR MODE Knob

Refer to figure 1-80 for functions and details of the
INSTR MODE knob located on the INSTR MODE
select panel. Refer to figure 1-82 for details of the ADI
and HSI in the instrument modes.
Navigation Aids and Display

INSTR HOG Knob !AI mEl

Refer to figure 1-81.

Refer to figure 1-80. The INSTR HDG knob, locatedon the INSTR MODE select panel, has arrows
poin ting in cw and ccw directions. The knob is pushed
and turned to set the INS heading to a known

Instrument Modes

Refer to figure 1-82.

INSTR MODE Select Panel (Typical)
BLOCK 10BLOCK 15BLOCK 15BLOCK 10

*
~
ILSI

MODE
TeN NAVILS/

r~NiVII

S ~1

1. INSTR MODE Knob
2. INSTR HDG Knob

T
R

1F-16A-1-0064A @Figure 1-80.

1-200

T. O. 1F- 16A- 1

Navigation Aids and Display (Typical)
INSTR
MODE
SELECTPANEL(IMSP)

ADI

OO~

Hue
VOL

ILS

®

ILS
CONTROLPANEL

CHANNELTACANCONTROL
PANEL(( I 0 0
T

C

N

<aD

/:VOL~

05:9~
©

IMSP
INSTRUMENT
MODE
SELECTEDRANGE
INDICATOR

RANGE TOTACANSTATIONOR DMENAV AIDCOURSE
ARROW &COURSE
SELECTEDMANUALLY
SELECTEDLOCALIZERCOURSEMANUALLY
SELECTEDCOURSENAV

ILS/NAV

~REC
AlA

orr

T/R

®
BeN

MRKBCNID
HSI
HSI

ILS{TCN

TCN

~

0

Xl]

RANGE
TO INSDESTINATION* BLOCK 10

COURSEDEVIATION
INDICATOR

TO- FROM
INDICATORLOCALIZERDEVIATIONOUT OF
VIEW

BEARING
POINTERATTITUDESPHERELOCALIZERDEVIATION
BAR

GLIDE
SLOPE
DEVIATIONBAR

LOCALIZERDEVIATION

GLIDE
SLOPE
DEVIATIONOUT OFVIEW

OUT OF
VIEW

LOCALIZERDEVIATION

GLIDE
SLOPE
DEVIATION

BEARINGTO TACANSTATION

DEVIATION
FROM
SELECTEDCOURSE

INS
ROLL
AND
PITCH
ATTITUDE

IN VIEW* INACTIVE** DEVIATIONFROMSELECTEDCOURSEMANUALLY
SELECTEDLOCALIZERCOURSEADIOUT OF
VIEW

LOCALIZERDEVIATION** BLOCK 15

BEARINGTO INSDESTINATION

1F-16A-1-1117X8

Figure 1-81.

1-201

T.o. 1F- 16A- 1

Instrument Modes (Typical)
TeN
NOTE:
The TCN mode has no effecton HUD symbology.

ADI
INS MAGNETIC HEADINGHEADING REFERENCE MARKER
i

MODE
ILSI

TeN NAY

ILSI

r~NAV
T
R

TACAN COURSE SIGNAL --;;o-----.;:-;,-'--;-~~UNRELIABLE

INSTR MODE
SELECT PANEL(IMSP)

HSI

NAV
NOTE:
The NAV mode has no effecton HUD symbology.

ADI
INS MAGNETIC HEADINGHEADING REFERENCE MARKER

RANGE TO INSDESTINATION---l-

INS MAGNETIC HEADING --r"t---;:-;:.~_~
SIGNAL UNRELIABLE

INSTR MODE
SELECT PANEL(IMSP)

DEVIATION FROMSELECTED COURSE

HSI
Figure 1-82. (Sheet 1)

1-202

1F-16A-1-1118X@

T. O. 1F- 16A- 1

Instrument Modes (Typical) IANIILSJTCN~

HORIZON LINE~~ PITCH SCALE (CLIMB)

.~

~

"D~'J

U

CALIBRATEDAIRSPEED

10'-I

-

-

I
06

~/O 20~ c 5 ' - -~~~ ~~~EK::DI/11,01 15-= IIIAI I
5c_

LOCALIZER

1I~

DEVIATION

VI//
10L-_~~!.. I \,

SELECTED COURSEU\\(LIMITED)'---' \\

nl:-02,2

I(j\.\\\-~02,l~L ~~t5_0"1 \\\\\)\"" L::JLOCALIZERDEVIATIONIU

HUD
PITCH SCALE (DIVE)

DEVIATION
BAROMETRIC ALTITUDE
RADAR ALTITUDE

150>03~11\\'
ALTITUDE LOW\ __ -"10o~ rr
FLiGHTPATH
MARKER (FPM)

GLIDE SLOPE~01'O-~_ __ "5.\U

INS MAGNETIC HEADING.----,=

07----'5j;A~~\ (±CROSSWIND COMPONENT)/ROLL SCALE

SETTINGSTEERPOINT NUMBER

COMMAND STEERING(CMD STRG)(©
NOTES:
1. Glide slope and localizer
bars indicate deviation
only, not steering.

2. Command steering isreferenced to the
localizer course
displayed on the FCNP.

ADI

ILSI

TeN NAV

ILSI

r~AV
T
R

INSTR MODE
SELECT PANEL(lMSP)

LOCALIZERDEVIATION

HSI

1F-16AN-1-0085X @Figure 1-82. (Sheet 2)
Change 12

1-202.1

T. O. 1 F- 1 6A- 1

Instrument Modes (Typical) IANII LS/NAV-e:====HORIZON LINE.'i.':
ID/";\1U~

CALIBRATEDAIRSPEEDAOA BRACKET(NLG LOWERED)~~ PITCH SCALE (CLIMB)

I

-

~~~

-

!..

I
06

'/O 20~ c 5'---01
IIMI' I\,-----'5

DEVIATIONVI

lOL __n~\ \"\ __ -"10LOCALIZERDEVIATION(LIMITED)nl~_ 02,1~\\\\"I'NU
ROLL SCALE~I--~oi II I

ROLL ATTITUDE

L:J

SETTINGSTEERPOINT NUMBER

COMMAND STEERING(CMD STRG)~ LOCALIZER SIGNALUNRELIABLE

INS PITCH AND

GLIDE SLOPEDEVIATION

150>03~1\\\'
ALTITUDE LOW

GLIDE SLOPEDEVIATION____- + - - - - l , C

INS MAGNETIC HEADING~02,0__ \ II~~\\\
BAROMETRIC ALTITUDE~L ~~6~~"'1 \I\\\)RADAR ALTITUDE

HUD

PITCH SCALE (DIVE)

U\\'----' \\=-02,2___ -"5//

SELECTED COURSE

r---'I=-1I~

FLiGHTPATH
MARKER (FPM)I

07

15-= -

LOCALIZER

o~ [

~

10'--

j;~

~~\ (±CROSSWIND COMPONENT)/NOTES:
1. Glide slope and localizer
bars indicate deviation
only, not steering.

2. Command steering isreferenced to the
localizer course
displayed on the FCNP.

~©

~------------------~

ADI

ILS/

TeN NAV

ILS/

~wNAV
S

ILS LOCALIZER SIGNAL---7--7---~
UNRELIABLE

T
R
INSTR MODE
SELECT PANEL(lMSP)

LOCALIZERDEVIATION

HSI
Figure 1-82. (Sheet 3)

1-202.2

Change 12

IF- 16AN-l-0086X @T. O. 1F- 16A- 1

Instrument Modes (Typical) IANIILS HUD DISPLAYS

.{?>
l2'
-- - - - - - - - - - - - - - - - - - - -l- - - - - - - - --------41.....--------------·-1
INBOUND LOCALIZER COURSE= 090 ()

\

CD

(j)

(0

RUNWAY

CD®

CD
I

I

\

I

1

01
45 0 INTERCEPT HEADING(090° - 45° = 045 0
AIRCRAFT HEADING)

J-

1

02

--1-----1-;• CMD STRG SYMBOL APPEARS (NO PITCHSTEERING PROVIDED)• LOCALIZER DEVIATION BAR IS SOLID (VALID)• A RIGHT TURN IS COMMANDED ICMD STRGSYMBOL IS TO RIGHT OF FPMj

,

CD

1

I

\
I

1

02

1

03

INITIAL AIRCRAFTHEADING
015 0=

30° BANK ACHIEVED

CD

CD

I

1

01

I
I

\
I

1

04

02

1

05

--1-----1--1---+------101-----

1

T

-'-

NO CMD STRG SYMBOL• LOCALIZER AND GLIDE SLOPE DEVIATION BARSARE DASHED (INVALID)• COURSE CARET IS LIMITED• HORIZON LINE IS DISPLAYED FOR REFERENCE•

AIRCRAFT IS ON 45

0

INTERCEPT HEADING

1 F-16AN-1-0087X @Figure 1-82. (Sheet 4)
Change 12

1-202.3

10.lF-16A-1

Instrument Modes (Typical) IANIILS HUD DISPLAYS

I

I

I

04

05

09

______~r\~------__• A RIGHT TURN IS COMMANDED (CMD STRG SYMBOL IS TO
RIGHT OF FPM)
• LOCALIZER DEVIATION BAR INDICATES APPROXIMATELY 2
DOTS LEFT

CD

\
1

I

I

1

06

• AFTER TIC MARK APPEARS ON TOP OF CMD STRG SYMBOL,PITCH STEERING IS PROVIDED• AIRCRAFT IS ON LOCALIZER COURSE AND ABOVE GLIDE
SLOPE

I

I

1

09

08

07

o

• 30° BANK ACHIEVED• LOCALIZER DEVIATION BAR INDICATES APPROXIMATELY 1
DOT LEFT

(j)
1

I

08

t

1

AIRCRAFT IS ON LOCALIZER COURSE AND GLIDE SLOPE(ALL SYMBOLS ARE ALIGNED)

@

I

I

09

07

• AIRCRAFT IS ON LOCALIZER COURSE (CMD STRG SYMBOL,COURSE CARET, AND LOCALIZER DEVIATION BAR ARE ALL
ALIGNED)• COURSE CARET INDICATES 083° VS 090° HEADING DUE TO
CROSSWIND

1

Iv
I

08

I

1

09

f---I----I--,

AIRCRAFT IS LEFT OF LOCALIZER COURSE AND GLIDE SLOPEIS INVALID (CMD STRG SYMBOL HAS X ON IT. NO PITCHSTEERING IS PROVIDED)

1F-16AN-1-0088X@

Figure 1-82. (Sheet 5)

1-202.4

Change 12

lO.lF-16A-l

Instrument Modes (Typical) LESS IANII

ILS/TCNPITCH SCALE (CLIMB)INSTANTANEOUS G'S

\

I \

I

AOA BRACKET(NLG LOWERED)

5,--

~
07

06

CALIBRATEDAIRSPEED ________

I

I

rl

\

I

5

I

_"

LOCALIZER::DEVIATION

L__

L'·''';/~~~Il6lnll
KI_

DEVIATIONSCALE

BAROMETRIC-

009D09

ALTITU DE

5L--HUD--05~

FLiGHTPATH
MARKER (FPM)( ± CROSSWIND COMPONENT)(LIMITED)GLIDE SLOPE

~

20-=

SELECTED COURSE

Il \~02~ROLLI= 1.1 , - - I ~I1L J
HORIZON LlNE-----iir::~C\

INS MAGNETIC HEADINGCOMMANDSTEERING(CMD STRG)PITCH SCALE (DIVE)
LOCALIZERDEVIATIONGLIDE SLOPEDEVIATION-+-+-----ll~

1--+-f----,.4__

NOTES:

LOCALIZER SIGNALUNRELIABLE

1. Glide slope and localizer
bars indicate deviation
only, not steering.
2. Command steering isreferenced to the
localizer course
displayed on the FCNP.

INS PITCH AND
ROLL ATTITUDE

ADI

ILS/

TeN NAVRANGE TO TACANSTATION OR DME
NAV AID - - - - - + - , . . .
ILS/

r~AV

ILS LOCALIZER SIGNAL - - ; - 7 - - - - , . ' ' ' - ' - UNRELIABLE

T
R
LOCALIZER DEVIATIONINSTR MODE
SELECT PANEL(IMSP)

HSI
1 F-16A-1-1119A @Figure 1-82. (Sheet 6)

Change 121-203

10. 1F- 16A- 1

I Instrument Modes (Typical) LESS IANII LS/NAVPITCH SCALE (CLIMB)INSTANTANEOUS G'S\I

AOA BRACKET(NLG LOWERED)

06
I

Irl

CALIBRATED= 1.1
AIRSPEED ~20~

II~I_ 15-;C
II I
10,

HORIZON
LINE

I

I

I

I

;- 02,2

LOCALIZERDEVIATIONGLIDE SLOPE,OR
BAROMETRIC
ALTITUDE

5~--HUD--~5~~

FLiGHTPATH
MARKER (FPM)

SELECTED COURSE± CROSSWIND COMPONENT)(LIMITED)

C 02

"L

009009II I(,,,,,,,~--

LOCALIZERDEVIATION

INS MAGNETIC HEADING

Il \ ~:~~ATION
I,--- I -----'.

\

I

---,5
07

5,------\COMMANDSTEERING(CMD STRG)PITCH SCALE (DIVE)(©

GLIDE SLOPEDEVIATION--l---+----Ii~

I-+f--::,.j__

NOTES:

LOCALIZER SIGNALUNRELIABLE

1. Glide slope and localizer
bars indicate deviation
only, not steering.

2. Command steering isreferenced to the
localizer course
displayed on the FCNP.

ADIILS LOCALIZER SIGNAL -:---;.----,.!~UNRELIABLE

INSTR MODE
SELECT PANEL(lMSP)

LOCALIZERDEVIATION

HSI
Figure 1-82. (Sheet 7)

1-204

Change 12

1 F-16A-1-1120A @10. 1F- 16A- 1

Instrument Modes (Typical) LESS IANII

ILS HUD DISPLAYSINBOUND LOCALIZER COURSE = 90~0---------------------\------------------------------........---.~ I. RUNWAY

CD

45 c

0

CD

CD

®

CD
01

45" INTERCEPT HEADING(90
45 () = 045 n
AIRCRAFT HEADING)
0

CD
CD
CD

_

02

--I~\.!.I~/j

-1-

~

/

~

• CMD STRG SYMBOL APPEARS (NOPITCH STEERING PROVIDED)• LOCALIZER DEVIATION BAR IS SOLID (VALID)• A RIGHT TURN IS COMMANDED (CMD
STRG SYMBOL IS TO RIGHT OF FPM)
0203

INITIAL AIRCRAFTHEADING = 015

030 BANK IS ACHIEVED
0

01

0)

02

04

05

1

~I_ -I~l,.dl.J1-1- _,----~:-~~c)~-/----------------1

1

NO CMD STRG SYMBOLLOCALIZER AND GLIDE SLOPE DEVIATIONBARS ARE DASHED (INVALID)• COURSE CARET IS LIMITED• HORIZON LINE IS DISPLAYED FOR
REFERENCE__ ,-\l.-d I.J).
~

1_ -

,

/

----------~~()~----------------••AIRCRAFT IS ON 45

0

INTERCEPT HEADING

1 F-16A-1-1121 X@

Figure 1-82. (Sheet 8)

Change 12

1-205

T.O. 1F- 16A- 1

I Instrument Modes (Typical) LESS JANJILS HUD DISPLAYS

08

CD

05

04

I

09

I

1

I

1

if'
\\1

"'-

11/

./

(" '\

---

...... . /6• AFTER TIC MARK APPEARS ON TOP OF CMD STRGSYMBOL, PITCH STEERING IS PROVIDED• AIRCRAFT IS ON LOCALIZER COURSE AND ABOVE GLIDESLOPE• A RIGHT TURN IS COMMANDED (CMD STRG SYMBOL IS TO
RIGHT OF FPM)
• LOCALIZER DEVIATION BAR INDICATES APPROXIMATELY 2
DOTS LEFT

CD

07

06

08

09

I

1

I

1

/

• 30 BANK ACHIEVED• LOCALIZER DEVIATION BAR
INDICATES APPROXIMATELY
1 DOT LEFT
0

(j)

08

09

I

I

AIRCRAFT IS ON LOCALIZER COURSE AND GLIDE SLOPE(ALL SYMBOLS ARE ALIGNED)

®

07

08

I

1/\

if'

• AIRCRAFT IS ON LOCALIZER COURSE (CMD STRG SYMBOL,COURSE CARET, AND LOCALIZER DEVIATION BAR ARE ALL
ALIGNED)• COURSE CARET INDICATES 083 VS 090 HEADING DUE TO
CROSSWIND
0

0

~

-

09

I

I

-I -

-

-

-

I

1

r- -

1

AIRCRAFT IS LEFT OF LOCALIZER COURSE AND GLIDESLOPE IS INVALID (CMD STRG SYMBOL HAS X ON IT.
NO PITCH STEERING IS PROVIDED)

1F-16A-1-1122X@

Figure 1-82. (Sheet 9)

1-206

Change 12

T. O. 1F- 16A- 1

FLIGHT INSTRUMENTS
Refer to figure 1-83. The flight instruments arelocated on the instrument panel. The instruments
listed below are common and are not illustrated indetail.
• Airspeed/Mach Indicator.
• Clock.
• Magnetic Compass.

displayed by a moving airspeed-mach pointer against
a fixed airspeed scale. Mach number, which is readagai~st the airspeed-mach pointer, is displayed by arotating mach scale. The range of the indicator is from80-850 knots and from 0.5-2.2 mach.
The maximum allowable airspeed pointer indicates
800 knots at sea level. Higher airspeeds are indicatedas altitude is increased. This indication is not a valid
maximum allowable airspeed cue: it should bedisregarded. Refer to MAXIMUM: AIRSPEEDOPERATING LIMITATIONS, Section V.
The SET INDEX knob is used to set the airspeed
reference index.

• Servo-Pneumatic Altimeter.
• Standby Attitude Indicator.

STANDBY AnlTUDE INDICATOR (SAl)• Vertical Velocity Indicator.

ALTIMETERI

The servo-pneumatic altimeter is a dual modepressure altitude indicator with a range of -1000 to+80,000 feet. The operating mode is manually
selected by the lever located at the lower right cornerof the instrument. In the ELECT (primary) operating
mode, the altimeter is electrically driven by the
CADC. In the PNEU (secondary) operating mode thealtimeter is pneumatically operated by staticpressure supplied by the pitot-static system. Should
the CADC or altimeter servo malfunction (or LESS (i])
the. right A0J:. probe heater circuit breaker open
wh,lle the altimeter is operating in ELECT), thealtimeter automatically reverts to the pneumatic
mode and the PNEU flag appears on the face of thealtimeter. The PNEU flag may also appear when
accelerating or decelerating through the transonicregion or while performing high g maneuvers.
The barometric setting knob, located at the lower left
corner of the instrument, is used to set the desired
altime~er sett~ng. [B] .rnB The barometric setting of
the altimeter IS electncally transmitted to the CADC
as a manual input correction for the pressure altitudedisplay on the HUD. ~ ~ l!ffiI ~ The barometric setting is shown in millibars. ~ The barometric setting is shown in inches of mercury.

AIRSPEED/MACH INDICATOR
The

airspeed/mach indicator ispneumaticallyo~erated ?y total and static pressure supplied by thepltot-statIc system, The indicator displays indicatedairspeed and mach number. Indicated airspeed is

The . SAl is a self-contained, electrically poweredver:tIcal gyroscope that mechanically positions thea~tItude sphere of the indicator to display aircraft
pItch and roll attitudes. Manual caging of thegyroscope is accomplished by pulling the PULL TO
CAGE knob at the lower right corner of the indicator.
~e knob is held out until the sphere is caged to zero
pItch and roll indication and then released.
Adjustment of the miniature aircraft reference
symbol is accomplished by rotation of the PULL TO
CAGE knob,
Since the SAl is mounted in the instrument panel atan angle, it indicates a pitch angle of 4 degrees lessthan the ADI when pitch trim knobs on bothindic,ators are s7t at the pitch trim index. If caging is
reqUIred, t~e aIrcraft should be flown wings level,
constant altitude, and at an AOA of approximately +4
degrees. When caged on the ground, allow 2 minutes
prior to taxi.
An OFF warning flag appears whenever electrical
power is lost or whenever the PULL TO CAGE knob
is pulled. Mter power loss, the indicator continues toprovide usable attitude information for approximately 9 minutes. The gyroscope of the indicator isunrestricted in roll but is limited to approximately± 85 degrees in pitch.
The indicator can develop errors during aerobatic
maneuvering, primarily when pitch is near 90
degrees. If these errors exceed 7 degrees after
returning to level flight, erection is cut off. If thisoccurs, the gyro does not automatically erect andmust be manually caged to eliminate the error. The
indicator is powered by ~ battery bus No, 1,
LESS ~ the battery bus.

Change 51-207

T.0.1F-16A-l

Flight InstrulTlents (Typical)

8

1. AOA Indexer

BLOCK 10

2. Standby Attitude Indicator
3. Vertical Velocity Indicator

BLOCK 15

4. Altimeter
5. Attitude Director Indicator
6. Horizontal Situation Indicator
7. Magnetic Compass
8. Clock.
9. AOA Indicator
10. Airspeed/Mach Indicator

10

8~

I-U

BLOCK 10BLOCK 15
1F-1SA-1-1127X@

Figure 1-83.

1-208

T.O. 1F-16A-1
VERTICAL VELOCITY INDICATOR (VVI)

The VVI provides rate of climb/descent informationand operates from static pressure supplied by theaircraft pitot-static system. Range of the indicator is
6000 fpm climb or dive.
MAGNETIC COMPASS

The magnetic compass is a self-contained indicatorwhich shows the heading of the aircraft with respect
to magnetic north. Adjustable compensating magnets
in the compass provide the means for cancellingmagnetic disturbances originating within the aircraft. A deviation correction card for the compass islocated immediately below and aft of the compass.

on a compass rose. The face of the RSI is a compasscard driven by the INS so that aircraft magnetic
heading is always read at the upper lubber line.
The HDG set knob provides the means for rotating
the heading reference marker to the desired heading.
Once set, the heading reference marker rotates withthe compass card. The heading reference marker
provides a reference to the heading select mode of the
autopilot.
The CRS set knob provides the means for selecting
anyone of 360 courses. To select a desired course,
rotate the head of the course arrow to the desired
course on the compass card and check the course
selector window for the precise setting. Once set, thecourse arrow rotates with the compass card.

ATTITUDE DIRECTOR INDICATOR (ADI)

Refer to figure 1-84. The ADI displays pitch and roll
attitude infonnation supplied by the INS. The ADI is
not limited in pitch or roll and displays any aircraft
attitude accurately. The instrument displays turnrate which is presented in standard turn needle
format. The tum rate needle is driven by the rate
gyroscope transmitter which senses the aircraft turnrate and displaces one needle width in response to a
1-1I2-degree/second tum rate. The slip indicator(ball) is a self-contained item. The pitch trim knob isused to adjust the attitude sphere to the desired pitch
attitude in reference to the miniature aircraft. In
certain modes of operation, the indicator displays ILS
glide slope and localizer deviation information. Refer
to NAVIGATION AIDS AND DISPLAY, this section.
The OFF warning flag may indicate failure of either
the INS or the AD!. A momentary OFF and/or AUX
flag, even with proper attitude being displayed, may
indicate impending failure of the ADI or INS data to
the AD!. The GS warning flag indicates that the glide
slope deviation bar is unreliable. The LOC warning
flag indicates that the localizer signal is unreliable.
The AUX warning flag signifies that the INS has
failed or is operating in a less precise attitudecondition and that HSI heading must be set to a
known heading by the ~ ~ INSTR HDG knob on
the INSTR MODE select panel.
HORIZONTAL SITUATION INDICATOR (HSI)

I Refer to figure 1-85. The HSI displays a horizontal orplan view of the aircraft with respect to thenavigation situation.
The miniature aircraft symbol in the center of the HSI
is fixed and comparable to an aircraft superimposed

The bearing pointer provides bearing information to
TACAN station or INS destination. Refer to
NAVIGATION AIDS AND DISPLAY, this section.
The range indicator provides a readout of distance in
nm to a TACAN station, DME navigational aid, orINS destination. Loss of TACAN or DME signal or an
unreliable signal causes a warning flag to cover therange indication window when either ILSITCN or
TCN is selected. When NAV or ILSINAV is selected,
an INS failure causes the warning flag to cover therange indication window. Loss of power to the RSI
may cause the OFF warning flag to come into view.
CLOCK

The clock, located on the right auxiliary console, is an
8-day, manually wound clock with provisions for anelapsed time indication up to 60 minutes.

STANDARD FUG.HT DATA RECORDER(SFDR) ~The aircraft contains an SFDR which stores engine
usage data, aircraft service life data, and mishap
investigation data. Engine and service life data are
stored for approximately the last seven flights.
Mishap investigation data storage is limited toapproximately the last 15 minutes of flight time;
however, there are automatic special event provisions
in flight which protect an approximate 30 seconds
worth of data from overwrite until the next flight. A
total of five special events (automatic and manual)
may be saved. Special event data save also generates
an FDR 024 MFL.

Change 2

1-209

IT. O. 1F- 16A- 1

Special events are:

• Canopy unlocked in flight.

• Acceleration (Nz ) greater than 9.5g or greaterthan -3g.

• Engine overs peed.

• AOA greater than 29 degrees or greater than orequal to -5 degrees.

1-210• Engine subidle .
• Throttle in OFF in flight.

T.O. 1F- 16A-l

Attitude Director Indicator (ADI) (Typical)
2

5

4

3

19

rr:-t--t-- 6

7
8

17

r-r---916
15

14 13 1211

10

1. GS Warning Flag - Glide Slope Unreliable
2. Attitude Sphere
3. Pitch Scale
4. Upper Bank Index Pointer5. Localizer Deviation Bar
6. LOC Warning Flag - Localizer Signal Unreliable
7. AUX Warning Flag - INS Navigation andHeading Unavailable
8. Pitch Trim Index
9. Pitch Trim Knob
10. Rate-of-Turn Scale
11. Rate-of-Turn Needle
12. Slip Indicator (Ball)
13. Lower Bank Index Pointer14. Lower Bank Scale
15. Miniature Aircraft Symbol
16. Glide Slope Deviation Bar and Pointer17. OFF Warning Flag - Attitude Sphere/INS Unreliable
18. Glide Slope Deviation Scale
19. Horizon LinelE-1BX-l-0013X®

Figure 1-84.

Horizontal Situation Indicator (HSI)(Typical)
1

2

3

5

19
18 --++----77.~~~17 _--++-~-r--"I....I-I
16

9

15
10

13

12

1. Range Indicator
2. Warning Flag - Range Indicator
3. Heading Reference Marker
4. Upper Lubber Line
5. Course Selector Window
6. Bearing Pointer7. OFF Warning Flag - HSI Power
8. TO-FROM Indicator
9. Miniature Aircraft Symbol
10. CRS Set Knob
11. Course Arrow - Tail
12. Bearing Pointer - Tail
13. Lower Lubber line
14. HOG Set Knob
15. Course Deviation Indicator
16. Course Deviation Scale
17. Warning Flag - Course Deviation
18. Course Arrow
19. Compass Card11
1F-16X-l-0014X@

Figure 1-85.

Change 7

1-211

IT. O. 1F- 16A-1

F- 168 AIRCRAFT
Only those items which are different, significant, or
unique to the forward or rear cockpit are discussed in
the following paragraphs.

ENGINE CONTROLS IPW2001
Only a throttle, BUC switch, and FUEL MASTER
switch are provided.

ENG CONT Switch IPW2201 mm
Refer to figure 1-13. The ENG CONT switch, locatedon the left console, is a two-position switch.
Functions are:
• SEC - Manually selects SEC (SEC caution lightilluminates) regardless of forward cockpit ENG
CONT switch position.

Throttle I PW200 I (g• NORM - (Guarded position) Returns control toforward cockpit switch position.

The throttle is limited in certain functions:

FUEL SYSTEM• Cannot be advanced from OFF to IDLE.
• Cannot be advanced from MIL to AB.

Fuel Control Panel IIBl

BUC Switch I PW200lilm

Refer to figure 1-22. The fuel control panel, located on
the left console, contains a guarded FUEL MASTER
switch. The FUEL MASTER switches in both cockpits
must be in the MASTER (on) position to permit fuelflow to the engine. Either switch when positioned to
OFF shuts off all fuel flow to the engine.

Refer to figure 1-7. The BUC switch, located on theleft console, is a two-position switch.

SPEEDBRAKE SYSTEM

Functions are:

SPD BRK Switch ~• BUC - 'furns EEC off (EEC caution lightilluminates) and transfers to BUC when throttle isin OFF or above BUC IDLE regardless of forward
cockpit EEC BUC switch position. The throttle may
have to be advanced past BUC IDLE to havetransfer occur.

Refer to figures 1-9 and 1-15. The SPD BRK switch,
located on the throttle, is spring-loaded to the off(center) position and must be held in either the open(aft) position or close (forward) position.

• Cannot be retarded from IDLE to OFF.
• Does not have a throttle friction control.

I• OFF - (Guarded position) Returns control toforward cockpit switch position.

ENGINE CONTROLS I PW2201
Only a throttle, ENG CONT switch, and FUELMASTER switch are provided.

Throttle I PW2201

m

The throttle is limited in certain functions:• Cannot be advanced from OFF to IDLE.
• Cannot be advanced from MIL to AB.
• Cannot be retarded from IDLE to OFF.
• Does not have a throttle friction control.

1-212

Change 1

DRAG CHUTE SYSTEM IBEI ~DRAG CHUTE Switch [HE! ~ ~Refer to figure 1-40. The DRAG CHUTE switch,
located on the LG control panel, is a three-position
guarded switch used to deploy and release the dragchute. Manually holding the DRAG CHUTE switch in
REL prevents drag chute deployment when the [IE]switch is placed to DEPLOY.
Manually holding the DRAG CHUTE switch in REL
after[lE] drag chute deployment (switch remaining in
DEPLOY) ~ andallows the deployed drag chuteto release if airspeed is above 60 knots, LESS m ~allows the deployed drag chute to release.

m

Manually holding the DRAG CHUTE switch in
DEPLOY prevents drag chute release when the [IE]switch is moved from DEPLOY to NORMIREL.

10. 1F- 16A-1
LANDING GEAR (LG) SYSTEM~ All LG functions are duplicated except that the

BRAKES channel, ANTI-SKID, and LANDING TAXILIGHTS switches are not present. In addition, the
DN LOCK REL button function is different.
DN LOCK REL Button ~Refer to figure 1-36. The DN LOCK REL button
located on the LG control panel, when depressed:
unlocks the LG handle electrically to allow movement
to the UP position in case the left MLG WOW switch
has failed to the ground position. The DN LOCK REL
button has no mechanical unlock function.
FLIGHT CONTROL SYSTEM

Simultaneous inputs to the forward and aft sticks (orrudder pedals) are added together to position theflight control surfaces accordingly. Unintentionalright roll inputs can occur as a result of leg/knee
interference with the stick.

I!ml The flight controls are abbreviated and contain
only the stick, rudder pedals, and warning, caution,and status lights. The FLCP and MANUAL TRIM
panels are not available.
STICK CONTROL OPERATIONWith the STICK CONTROL switch in FWD, the stick
indicator displays the word FWD. Depressing theforward cockpit paddle switch locks out the rear
cockpit stick, rudder, and MPO commands. The
OVRD light in both cockpits illuminates, indicating
override is activated. Depressing the rear cockpit
paddle switch has no effect on the forward cockpit
controls.
With the STICK CONTROL switch in AFT, the stick
indicator displays the word AFT. Depressing the rear
cockpit paddle switch locks out the forward cockpit
stick, rudder, and MPO commands and transfers
control of NWS to the rear cockpit. The OVRD light inboth cockpits illuminate, indicating override isactivated. Depressing the forward cockpit paddle
switch has no effect on the rear cockpit controls.
STICK CONTROL Switch mEl

Refer to figure 1-17. The STICK CONTROL switch islocated on the TEST switch panel.
Functions are:
• FWD - The aft cockpit flight control functions arelocked out as long as the paddle switch is held
depressed.

• AFT - The forward cockpit flight control functions
are locked out, and NWS control is transferred to
the rear cockpit as long as the paddle switch is held
depressed.
Stick Indicator ~Refer to figure 1-86. A three-position stick indicator,
located on the instrument panel, indicates the
position of the STICK CONTROL switch bydisplaying the word AFT or FWD as applicable.
Diagonal stripes appear when power is removed orappear momentarily during switching.
Stick OVRD Lights [IDRefer to figure 1-86. The stick OVRD lights, locatedon the forward and rear cockpit instrument panels,
illuminate when the paddle switch is used to takecontrol.
ESCAPE SYSTEM [ID
EJECTION MODE SEL Handle ~Refer to figure 1-60. The EJECTION MODE SEL
handle is located on the right auxiliary console.
Functions are:
• NORM• Activation of ejection system from rear cockpit
results in canopy jettison, then a .33-second delay
followed by only rear seat being ejected.
Activation of ejection system from forward cockpit
following rear seat ejection results in forward seatejection.
• Activation of ejection system from forward cockpit
results in canopy jettison, a .33-second delay, aftseat ejection, and a .4-second delay followed byforward seat ejection.
• AFT - Activation of ejection system from either
cockpit results in canopy jettison, a .33-second
delay, aft seat ejection, and a .4-second delay
followed by forward seat ejection.
• SOLO - Activation of ejection system from forward
cockpit results in canopy jettison, then a .33-second
delay followed by only forward seat being ejected.
Activation of ejection system from aft cockpit
results in canopy jettison, then a .33-second delay
followed by only aft seat being ejected. Simultaneous activation from forward and aft cockpits
results in unsequenced ejections.
Change 7

1-213

I10. 1 F- 16A-l

Stick Indicator and
OVRD Light 00(Typical)l UJ

FLIGHT INSTRUMENTS ~An accelerometer is installed and the altimeterbarometric pressure set knob is not connected to the
CADC.
Accelerometer ~Refer to figure 1-89. The accelerometer is self-contained and mechanically indicates acceleration
acting along the vertical axis of the aircraft. The
accelerometer is graduated in g units and has three
indicating pointers. The main pointer displays
instantaneous changes in acceleration. The positive
and negative auxiliary pointers indicate the maximum positive and negative acceleration experienced.

~
2~

The auxiliary pointers retain their highest readings
until the PUSH TO SET knob is depressed.

18

RADAR/ELECTRO-OPTICAL (REO) DISPLAY ~1. OVRD Light (Amber)
2. Stick Indicator
1F-16A-1-1130A®

Figure 1-86.

COMMUNICATIONS SYSTEM rID

Refer to figures 1-87 and 1-88. Control panels for theUHF, TACAN, ILS, and ~ VHF and control
indicators/pushbuttons are located in both cockpits.

Refer to figure 1-3. The HUD CTVS video signal canbe displayed on the REO by placing the VIDEO SEL
switch to HUD.
STORES MANAGEMENT SUBSYSTEM (SMS) ~Refer to figure 1-3. The SCP is a display only. The edge
pushbuttons are not functional.
FIRE CONTROL RADAR (FeR) ~Refer to figure 1-90 and T.O. IF-16A-34-1-1, ~ T.O.
IF-16A-34-1-3, or ~ T.O. IF-16A-34-1-4. Limited
radar operation is afforded.
THROTTLE AND STICK SWITCHES/CONTROLS ~When power is first applied to the aircraft, theforward cockpit control indicators/pushbuttons illuminate to indicate that the UHF, TACAN, ILS,
and ~ VHF are controlled by the forward cockpit
control panels. Control can be taken from, but notgiven to, the other cockpit. When an indicator/pushbutton is not illuminated and control is desired,depress the indicator/pushbutton.
INSTR MODE SELECT PANEL lID

Refer to figure 1-80. If the INSTR MODE knob is in
the same position in both cockpits, the rear cockpit
HSI course slaves to that selected in the front
cockpit. I!ml The INSTR MODE select panel does nothave an INSTR HDG knob.

1-214

Change 13

Refer to figure 1-90 for the switches/controls locatedon the throttle and stick.
ARMT CONSENT SWITCH ~Refer to figure 1-3. The ARMT CONSENT switch,
located on the instrument panel, is guarded in the
ARMT CONSENT position. The switch is in series
with the MASTER ARM switch in the forward cockpit
and must be in ARMT CONSENT to enable thenormal release of any store.
SERVICING DIAGRAM

Refer to figure 1-91 for servicing/specifications
information.

IT.o. 1F- 16A-1

TAKE CONTR Panel IBFI (Typical)

1. VHF Button/Indicator
2. UHF Button/Indicator
3. TACAN Button/Indicator
4. ILS Button/Indicator

1F-16A-1-1131X®

CONTROLl.

~VHF

POSITION/INDICATIONFUNCTION

Depressed (green)

VHF radio control in forward cockpit

Unlighted

VHF radio control in rear cockpit

Depressed (green)

UHF radio control in forward cockpit

Unlighted

UHF radio control in rear cockpit

Depressed (green)

TACAN control in forward cockpit

Unlighted

TACAN control in rear cockpit

Depressed (green)

ILS control in forward cockpit

Unlighted

ILS control in rear cockpit

Button/Indicator

2. UHF
ButtonlIndicator
3. TACAN
Bu ttonlIndicator
4. ILS ButtonlIndicator

Figure 1-87.

1-215

T.O. 1F- 16A- 1

RADIO SEL Panel IBRI (Typical)
4

3

2

1. UHF Button/Indicator
2. VHF Button/Indicator
3. TACAN Button/Indicator
4. ILS Button/Indicator
5. INTERCOM Knob
6. THREAT Tone Knob
7. MSL Tone Knob
8. HOT MIC Switch

R
AS

oE
I L

HOT MICMSL

THREAT

INTERCOM~ ~======- ===~c====
8

(®

7

6

5
1 F-16A-1-0078A@

CONTROL
1. UHF
ButtonlIndicator

2.

m VHF

FUNCTIONPOSITIONIINDICATION
Depressed (green)

UHF radio control in rear cockpit

Unlighted

UHF radio control in forward cockpit

Depressed (green)

VHF radio control in rear cockpit

Unlighted

VHF radio control in forward cockpit

Depressed (green)

TACAN control in rear cockpit

Unlighted

TACAN control in forward cockpit

Depressed (green)

ILS control in rear cockpit

Unlighted

ILS control in forward cockpit

ButtonlIndicator
3. TACAN
ButtonlIndicator
4. ILS ButtonlIndicator

5. INTERCOM Knob

Controls rear cockpit intercom volume

6. THREAT Tone Knob

Controls volume of rear cockpit TWS composite threat tone

7. MSL Tone Knob

Controls volume of rear cockpit tone from the AIM-9 missile being monitored

8. HOT MIC SwitchOFF
HOTMICIIntercom not activated
Provides a rear cockpit intercom for cockpit,
ground crew, and AR. Activation of UHF VHF
transmit switch overrides the HOT MIC function

Figure 1-88.

1-216

T. O. 1F- 16A- 1

Accelerometer I BR I (Typical)
2

BLOCK 10
1. Main Pointer
2. Auxiliary Positive Pointer
3. Auxiliary Negative Pointer4. PUSH TO SET Knob

BLOCK 15

3

4

1F-16A-1-0079B@

Figure 1-89.

Throft/e and Stick Switches/Controls ~BLOCK 10BLOCK 15STATUSSWITCH/CONTROLSTATUSSWITCH/CONTROLTHROTILE

THROTILE
Dead/Dead

DOG FIGHT/Missile Override

Dead/Dead

SPD BRK

ActiveSPD BRK

ActiveRDR CURSOR/ENABLE

Active/Active

DOG FIGHT/Missile Override* ANT ELEVMAN RNG/UNCAGE
UHF VHF* RDR CURSOR/ENABLE* ANT ELEV
* MAN RNG/UNCAGE

ActiveDead/Dead

UHF VHF

Active/Active

CAMERA/GUN
WPN REL

Active** Acfivel** Active
Active/Active

STICK

STICK
TRIM

Active/Active

ActiveTRIM

ActiveDead/Dead

CAMERA/GUN

Active/Dead

ActiveWPN REL

ActiveActive/Active* DESIG/RET SRCH

Active/Active* DSG/RS

NWS/A/R DISC

Active/Active

DO

Dead

MSL STEP

Dead

DS

PaddleActiveActive (only to gain
control of *functions)

NWS A/R DISC

Active/Active* Control gained by activation of RDR CURSOR/ENABLE switch.

MSL STEP

Dead

PaddleActive* Control gained by activation of DS switch.
** Inactive in EEGS mode.

•

Figure 1-90.

1-217

10. 1F- 16A-1

Servicing Diagram (Typical)
DOOR 3302--------------~
LESS[8N] SECURE VOICEPROCESSORI

COCKPIT (EJECTION SEAT(S))
EMERGENCY GASEOUS
OXYGEN BOTTLE(S)

DOOR 2318 -----------,.
BLOCK 15EXTERNAL POWER CABLE
RECEPTACLE

NLG WHEEL WELL
LG/HOOK EMERGENCY
PNEUMATIC BOTTLE(NITROGEN)

DOOR 2306 -------------.,BLOCK 10EXTERNAL POWER CABLE
RECEPTACLE

I
I

r------

DOOR 3306 --------..
LESS[8N] IFF MODE 4

DOOR [K] 2105,[[j 2107
GROUND COOLING
RECEPTACLE
DOOR 2317~ ENGINE AND EMS

GO - NO - GO INDICATORSPANEL 3408 (TOP)-----..
EPU (MONOPROPELLANT)DOOR 3308
LIQUID OXYGENLESS ~ [iijiV22ij] ENG I N E
SYSTEM FAULT FLAGDOOR 3206
AMMUNITION LOADINGPANEL 3402 (TOP)LEF POWER DRIVE UNIT

DOOR 3412 (TOP)EPU OIL SERVICE

DOOR 3216
HYDRAULIC SYSTEM AGROUND TEST ANDSERVICE CONNECTIONS~------- DOOR

3115
HYDRAULIC SYSTEM B
GROUND TEST ANDSERVICE CONNECTIONS~DOOR 3101
HYDRAULIC SYSTEM B
ACCUMULATOR

DOOR 3208
EPU (NITROGEN)HYDRAZINE LEAK
DETECTOR

LMLG WHEEL WELL
HALON RESERVOIRDOOR 3202
HYDRAULIC SYSTEM A
ACCUMULATOR

DOOR 3103
GROUND REFUELING
RECEPTACLE

DOOR 3204
DEFUELING RECEPTACLE

DOOR 3316 - - - - - - > 'ENGINE OIL SERVICE
JOAP SAMPLE

DOOR 4113
FLCS ACCUMULATORS (2)DOOR 4304 --------./BRAKE/JFS ACCUMULATORS (2)DOOR 4214[N!l [NQ] DRAG CHUTE
ACCUMULATOR
1F-16A-1-1134B@

Figure 1-91. (Sheet 1)

1-218

Change 12

T. O. 1F- 16A- 1

Servicing Diagram (Typical)
SPECIFICATIONSSERVICEABLE ITEM
USAF

FUel

ENGINE/JFS

NATO

MIL-T-5624, JP-4

F-40

MIL-T-5624, JP-5

F-43 OR F-44

MIL-T-83133, JP-8

F-34

JET A, B (COMMERCIAL)NONEJET A- 1 (COMMERCIAL)

F-35

ENGINE *MIL-L-7BOBJ OR LATER

ADG/CSD/EPU

MIL-L-7BOB

HYDRAULIC SYSTEMS A AND B

MIL-H-5606
MIL-H-B3282

GASEOUSMIL-O-27210, TYPE I

LIQUID

MIL-O-27210, TYPE II

EXTERNAL ELECTRICAL
POWER

115 (±15) YAC, 400 (±30) HZ

AlM32A-60A

NONE

NITROGENGASEOUSBB-N-441 A, TYPE I, GRADE BNONEFUEL TANK INERTING
AGENT (OPTIONAL)

LIQUID

HALON 1301

NONE

MONOPROPELLANT (EPU)

LIQUIDHYDRAZINE (70% N2H4, 30%
H2 O )NONEOIL

HYDRAULIC FLUID

0-148

OXYGEN

H-515
H-537
NONE

·IF NECESSARY, ENGINE LUBRICATING OILS MIL-L-7S0S (NATO CODE 0-14S) AND MIL-L-23699 (NATO CODE 0-156) MAY BE MIXED. AT
THE FIRST OPPORTUNITY THEREAFTER, THE OIL SHALL BE DRAINED AND FLUSHED, AND THE ENGINE SERVICED WITH THE PROPER
LUBRICATING OIL AS SPECIFIED IN THE APPLICABLE ENGINE TECHNICAL ORDERS.

Figure 1-91. (Sheet 2)
1-219/(1-220 blank)IT. O. 1F- 16A- 1

SECTION II
NORMAL PROCEDURESTABLE OF CONTENTS

Page

Introduction ..
2-1
Flight Planning ....
2-1
2-1
Takeoff and Landing Data Card .
Cockpit Designation Code.
2-2
2-2
Weight and Balance ...........
Preflight Check
2-2
2-2
Exterior Inspection
Cockpit Access .
2-2
2-2
Before Entering Cockpit ........
2-3
Cockpit Interior Check .........
Left Console .
2-4
2-4
Left Auxiliary Console ..........
Instrument Panel ..
2-5
Right Auxiliary Console .....
2-5
Right Console .
2-5
After Cockpit Check Is
Complete - Verify ..
2-6
2-7
Before Starting Engine.
Starting Engine IPW200 I .
2-8
Starting Engine I PW2201 .
2-9
2-11After Engine Start.
Before Taxi
2-16
2-17
Taxi .....
2-17
Before Takeoff.
2-20
Takeoff
2-20
Normal Takeoff ...
Takeoff With Asymmetric Stores ......
2-20.1
Operating Airspeeds below 10,000
Feet MSL
2-21
2-21
Climb
Climb/ln-FlightiOperational Checks ..
2-22
2-22
Air Refueling Procedures ...
2-22
DescentlBefore Landing ....
Landing o.
2-23Normal Landing ..
2-23
Short Field Landing (Dry Runway) ...... 2-26
2-26
Crosswind Landing ......
2-26
Touch-and-Go Landing ..
After Landing
2-26
Prior to Engine Shutdown .....
2-27
Engine Shutdown ...............
2-28
2-28.1
Scramble .......
Preflight .....
2-28.1
Aircraft On Quick Response Status ...... 2-29
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Expanded FLCS Self-Test Procedures ....
ILSProcedures ....
Exterior Inspection ....
Danger Areas ......
Turning Radius and Ground
Clearance .. o....
G-Suit Hose Routing ..
Strange Field Procedures ........
Fuel Management LESS
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2-29
2-29
2-29
2-30
2-30
2-30
2-30
2-30
2-30
2-30
2-30
2-31
2-31
2-31
2-31
2-31
2-31
2-32
2-33
2-34
2-34
2-34

2-34
2-34
2-34

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Hot Refueling ..
Hot Refueling Precautions .......
Prior to Hot Pit Entry ...........
Prior to Hot Refueling ........
During Hot Refueling ....
Hot Refueling Complete .....
Aircraft De-icing with Engine Operating
Prior to Application of De-icing Fluid , ..
During De-icing.
After De-icing is Complete
Prior to Takeoff .
Quick Turnaround ....
Prior to Engine Shutdown ...........
Special Procedures
EPU Hydrazine Support at Non-F-16
Bases ...
Supplemental Procedures ..
BUC Ground Start Procedures

•

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INTRODUCTION
This section provides the actions required for normal
operation of the aircraft. Amplification is included
only when special considerations or techniquesshould be observed. A complete knowledge of Section
III, EMERGENCY PROCEDURES, and Section V,
OPERATING LIMITATIONS, is required prior toflight.
FLIGHT PLANNING

Refer to ToO. 1F-16A-1-1.
TAKEOFF AND LANDING DATA CARD

A takeoff and landing data card is provided in the
Flight Crew Checklist for recording computed data.

Change 11

2- 1

T.O.1F-16A-1

I WARNING I

COCKPIT DESIGNATION CODE

An asterisk (*) preceding steps is used to highlightprocedures for [ID aircraft which apply to both cockpits.

If winds exceed 30 knots, open the
canopy only as far as needed to enter/exit
the cockpit. Decreasing the canopy angle
reduces the possibility that the canopy
can be blown past full open.

WEIGHT AND BALANCE

For maximum GW limitations, refer to Section V,
OPERATING LIMITATIONS. For weight and balance
information, refer to the individual aircraft Form F(DD Form 365-4) and the Weight and Balance
Handbook.

2.

PREFLIGHT CHECK

Ladder - Position on cockpit sill.

~

A failure of the canopy actuator
could allow the canopy to fall duringtransit. Do not position the ladderuntil the canopy motion has stopped.

Check AFTO Form 781 for aircraft release and storesstatus.

NOTE

EXTERIOR INSPECTION[ID Cockpit access is gained by the useof two entrance ladders or moving one
ladder as required.

Refer to figure 2-5 for normal preflight inspection.
COCKPIT ACCESSBEFORE ENTERING COCKPIT

Refer to figure 2-1.
1.

Canopy - Open by positioning external canopy switch to the up position.

*1.

Ejection seat - Check.
• Ejection safety lever - Safe (up).
• Safety pins (2) - Removed (ejection safety
lever and EMERGENCY MANUAL CHUTE
handle).

Cockpit Access(Typical)• Survival KIT DEPLOYMENT switch - Asdesired.
EXTERNAL CANOPY SWITCH(LOWER SURFACE OF LEFTSTRAKEIIWARNING

1

A Parachute Landing Fall (PLF) withthe survival kit not deployed may resultin injury.
• ~ [QE] IN;] ~ RADIO BEACON switch - Asdesired.
• Emergency oxygen bottle - 1800 psi minimum.
• Quick-disconnect (left side) - Connected.
• Recovery parachute - Free from damage andgrease.

1 F-16X-1-0045A @Figure 2-1.

2-2

Change 11• Inertia reel straps retaining pin (yellow) - Visible in circular inspection hole.
• Parachute risers and SEAWARSIUWARS (ifinstalled) - Free from damage and grease.

T.O.1F-16A-1

NOTE

Rapidly pulling on the shoulder
harness to check the locking mechanism may cause damage to the locking
mechanism.

UWARS BIT is a maintenance function.
Do not perform UWARS BIT since
excessive BIT checks will degrade
battery life.
• Lanyards from canopy to seat - Check.

Change 52-2.1/(2-2.2 blank)

10. 1 F- 16A-l• Emergency oxygen
nect - Connected.

hose

quick-discon-• Environmental sensor pitot tubes - Clear of
obstructions and if flip-up pi tot tubes are
installed, safety pins (2) removed.
• Quick-disconnect (right side) - Connected.
• Electronic recovery sequencer battery indicator - White indication.

18. G'fj DSPL POWER switch - OFF.
19.

OXYGEN REGULATOR - OFF and 100%.

20.

Utility light - OFF and secured.

COCKPIT INTERIOR CHECKThe checklist steps are not arranged in a mandatoryorder.
*1.

Loose or foreign objects - Check.
A thorough cockpit interior preflight checkshall be accomplished prior to each flightwith emphasis on loose or foreign objects
that might cause injury to personnel ordamage to the aircraft.

*2.

Harness and personal equipment - Fasten.
Attach the parachute risers to the harness.
Attach and adjust survival kit straps.
Secure and adjust the lapbelt. Mter
fastening lap belt, hold right side of lapbelt
buckle stationary and shake and push left
side. Also insure that latching mechanism
has reset back to its original position and isflush with buckle to insure a positive lock.
Connect oxygen hose, g-suit hose, m vesthose, and communication leads. Check
operation of the shoulder harness locking
mechanism.

• CANOPY JETTISON T-handle - Secure,safety pin removed.
• Lapbelt retaining pins (yellow) - Protruding, both sides.
• Lapbelts - Secure. (Pull up and forward,
both sides.)• Survival kit straps - Secure. (Pull up andaft, both sides.)2.

MAIN PWR switch - OFF.

3.

Chaff/flare programmer - Check lAW T.O.
1F-16A-34-1-1CL-1, ~ T.O. 1F-16A-34-13CL-1 or~ T.O. 1F-16A-34-1-4CL-1.

I~ For solo flight:

4.

Ejection seat - Safe, straps secure, pins
removed.

5.

CANOPY JETTISON
safety pin removed.

T-handle - Secure,

SPD BRK switch - Center.

7.

FUEL MASTER switch - MASTER (guarddown).

8. I PW200 I BUC switch - OFF (guard down).
9. I PW2201 ENG CONT switch - NORM (guarddown).
10. ALT GEAR handle - In.
11. ALT FLAPS switch - NORM.
GND JETT ENABLE switch - OFF.

13. ~!NQI DRAG CHUTE switch - NORM.
14.

1• A partially locked lapbelt may open
during maneuvering flight or ejection
sequence.

6.

12.

IWARNING

HOOK switch - UP.

• Unobserved g-suit hose m or vest hose
disconnects can contribute to gray/blackout or loss of consciousness.
• Failure to adjust survival kit straps toachieve a snug fit between the pilot
and seat pan lid may result in injury
during ejection.
• Failure to adequately secure andtighten lapbelt may result in inability
to reach and operate the MPO switch
during out-of-control situations.

15. ARMT CONSENT switch - ARMT CONSENT (guard down).
16.

EJECTION MODE SEL handle - SOLO.

17.

Interior LIGHTING
knobs off.

control

panel - All

Do not adjust seat height with survival
kit straps or lapbelt disconnected asdamage to the ejection seat or stick
may occur.
Change 13

2-3

T.O. 1 F-16A-1

NOTE

13.

ENG FEED knob - NORM.

• Refer to figure 2-8. The recommended
routing of the g-suit hose is directly
under the torso harness and aft of thesurvival kit strap to reduce thepossibility of a g-suit hose disconnect.
G-suit hose routing must provide
sufficient slack to allow for maximum
mobility since in-flight reconnect isextremely difficult.

14.

AIR REFUEL switch - CLOSE.

*15.

COMM control panel- As desired.

*16.

TACAN - As desired.

17.

TRIM/AP DISC switch - NORM.

18.

ROLL, YAW, and PITCH TRIM - Center.

19.

EPU switch - NORM.

20.

MAIN PWR switch - OFF.

21.

FCC PWR and INS PWR buttons - Off.

22.

DIR AIM, OAP 1, OAP 2 - Depress (asdesired).

23.

FCNP FUNCTION knob - OFF.

24.

IPW200 I ~ ~ STARTING FUEL switch-• Excess g-suit hose must be properly
routed to prevent MANUAL TRIM
panel interference during flight.
• ~ Incorrectly routed or unsecured
g-suit hose may result in throttleinterference or disconnected g-suit
hose.

•

• ~ ~ A PLD and an inflatable lumbar

support pad may be worn in any
combination with or without the lower
backrest pad installed.

AUTO LEAN.
25. IPW200 I ~ rn£l EEC BUC switch - EEC.

*3.

Rudder pedals - Adjust.
Adjust rudder pedals so that legs are flat on
the seat cushion to prevent the right leg
from hitting stick or injury during ejection.

Left Console

26. IPW200 II!!Rl BUC switch - OFF (guard down).
27. I PW2201 ~ ~ AB RESET switch - NORM.
28. j PW2201 ~ ~ ENG CONT switch - PRJ.

1.

PROBE HEAT switch - OFF.

29. I PW2201 I!!Rl ENG(guard down).

2.

~ STICK CONTROL switch - As briefed
when ~ occupied; FWD for solo flight.

30.

JFS switch - OFF.

3.

FLCS PWR TEST switch - NORM.

*31.

UHF function knob - BOTH.

4.

DEFOG lever - As desired.

*32.

UHF SQUELCH switch - ON.

5.

SERVOS ARM switches (3) - Center.

33.

Radar - OFF.

34.

Throttle - Verify freedom of motion, then OFF.

6.

SELF TEST switch - OFF.

*35.

SPD BRK switch - Forward.

7.

LE FLAPS switch - AUTO.

*36.

DOG FIGHT switch - Center.

*8.

ALT FLAPS switch - NORM.

9.

ECM power - Off.

10.

VIDEO SEL switch - As desired.

*11.

FUEL MASTER switch - MASTER (guard
down ~ rn£l and safety-wired).

*3.

ALT GEAR handle - In.

TANK INERTING switch - OFF.

4.

STORES CONFIG switch - As required.

12.

2-4

CONT

switch - NORM

Left Auxiliary Console

1.

IFF MASTER knob - STBY.

2. l@l AIFF MASTER knob - STBY.

Change 12

T.O. 1F- 16A- 1

5.

LANDING TAXI LIGHTS switch - OFF.

*6.

GND JETT ENABLE switch - OFF.

7.

BRAKES channel switch - CHAN 1.

8.

ANTI-SKID switch - ANTI-SKID.

*9.

EMER STORES JETTISON button - Coverintact.

*10.

HOOK switch - UP.

*11.

LG handle - DN.

Instrument Panel

4. ~ EJECTION MODE SEL handle - ffi1
NORM or AFT (as briefed), LESS ffi1 AFT.

i WARNING I• SOLO position should not be selectedwith the rear seat occupied.
• ~ Ejection initiated with the EJECTION MODE SEL handle in NORM
may incapacitate the ~ pilot atairspeeds greater than 180 knots. Useof the NORM position is therefore notrecommended.

1.

AUTOPILOT switch - OFF.

Right Console

2.

ROLL switch - As desired.

*1. ILS - As desired.

3.

PITCH switch - As desired.

2.

VHF function knob - TR.

4.

MASTER ARM switch - OFF.

3.

VHF mode knob - As desired.

5.

rn.Bl ARMT CONSENT switch - ARMT CON-

4. [@l [8] HF OFF VOL control knob - OFF.

SENT (guard down).
5.

LESS [8Nl NUCLEAR CONSENT switch - OFF I(guard down).

7. ~ ID LIGHT switch - OFF.

6.

[8] ~ RDR ALT switch (if installed) - OFF.

8. ~ LESS ~ LASER ARM switch - OFF.

7.

~

*9. ~ INQ] DRAG CHUTE switch - NORM.

*8.

Wristrest and armrest - As desired.

10.

*9.

Interior LIGHTING control panel - As desired.

6.

LESS ~ SMS PWR switch - As required.

HUD control panel - Set.

*11. Altimeter - Set.
*12.

REO knobs - As required.

*13.

INSTR MODE knob - As desired.

MASTER ZEROIZE switch - OFF.

10. ~ LESS ~ SNSR PWR switches (3) - OFF.
11.

EXT LIGHTING control panel- As required.

*1.

Clock - Set.

12. [8] ~ VOICE MESSAGE switch - VOICEMESSAGE.
If safety wire is broken, notify maintenance.

2.

FUEL QTY SEL knob - NORM.

13.

3.

EXT FUEL TRANS switch - NORM.

14. TEMP knob - AUTO.

Right Auxiliary Console

NOTE

If a wing transfer valve fails and WING
FIRST is selected, fuel flows onto theground from the overboard vent line
under the left wing (outboard of thefuel pylon). Selecting NORM shouldstop the fuel spillage.

15.

MASTER light switch - NORM.

AIR SOURCE knob - NORM.

If AIR SOURCE knob is in OFF or
RAM, electronic equipment may bedamaged.

Change 12

2-5

TO. 1F- 16A- 1

I

16.

LESS ~ Secure VOIce POWER switch (ifinstalled) - OFF.

LESS ill Perform the following:• Pressure - Check 50-120 psi.
a.

PLAIN cipher switch - PLAIN.
• SUPPLY lever - ON.

17.

C0 AVIONICS POWER switches (5) - As
required.

18.

CHAFF/FLARE switches (4) - OFF.

19.

ANTI ICE switch - AUTO/ON.

IWARNING

I WARNING I
Certain oxygen regulators provide norestrictions to breathing when the
SUPPLY lever is in OFF. Cockpit air
only is supplied, and hypoxia occurs as
cockpit altitudes which require supplemental oxygen are reached.

1

If engine anti-ice is on or has been on
during ground operation, heat in theinlet strut may be sufficient to cause
injury if touched.

• Diluter lever - NORMAL OXYGEN.
• EMERGENCY lever - NORMAL.
• FLOW indicator - Check.
• EMERGENCY lever - EMERGENCY
Check for positive oxygen pressure andmask and hose/connector leakage.

If there is visible moisture andambient temperature is 45°F (7°C) or
less, place the ANTI ICE switch to ON.

• EMERGENCY lever - NORMAL.

20.

IFF ANT SEL switch - NORM.

AFTER COCKPIT CHECK IS COMPLETE - VERIFY21.

UHF ANT SEL switch - NORM.

22.

C0 ~ DSPL POWER switch - OFF.

The following items are those important switches
that, if not correctly positioned, could cause a safety
hazard and/or improperly operated systems during
engine start.

*23.

Oxygen system - Check.

*1.

FUEL MASTER switch - MASTER (guard
down [ru mIl and safety-wired).

2.

ENG FEED knob - NORM.

3.

EPU switch - NORM.

ill Perform the following:• Pressure - Check 50-120 psi.
• Mode lever - PBG/ON (as required).

4. I PW200 I ~ rn£l STARTING FUEL switchAUTO LEAN.

• Diluter lever - NORM.

5. I PW200 I ~ rnfl EEC BUC switch - EEC.

• EMER lever - NORM.
• FLOW indicator - Check.
• EMER lever - EMER.
Check for positive oxygen pressure andmask and hose/connector leakage.
• EMER lever - NORM.
2-6

Change 12

NOTE
I PW200 I ~ rn£l The EEC BUC switch

can be in OFF with the guard down.
6. IPW200 I ~ BUC switch - OFF (guard down).
7. I PW2201 [BJ rnfl ENG CONT switch - PRI.
8. I PW2201 ~ ENG(guard down).

CONT

switch - NORM

T.O. 1F-16A-1*9.

Throttle - OFF.

*10.

HOOK switch - UP.

I

WARN INGNOTE• A JFS RUN light flashing once persecond indicates a non-critical failurein the JFS system. If the JFS operates
normally for engine start, the mission
may be continued. Notify maintenanceafter flight .

I

lID If either HOOK switch is in DN, the• A JFS RUN light flashing twice persecond indicates a critical failure in
the JFS system. The JFS will not
operate. Place the MAIN PWR switch
to OFF and notify maintenance.

hook extends when aircraft battery
power is applied.
*11.

LG handle - DN.

I

WARNINGI• Verify lights on:• FLCS PWR (4).

Insure that the LG handle is confirmed
fully down. The LG handle can be in anintermediate position allowing LG
extension and/or safe indications;however, the LG handle is not locked
and LG retraction could occur during
subsequent ground operations.
12.

• ACFT BATT TO FLCS.
2.

MASTER ARM switch - OFF.

• Verify lights on:

13. AIR SOURCE knob - NORM.
*14.

FLCS PWR TEST switch - TEST and hold.
Activation of the FLCS PWR TEST switch
with the MAIN PWR switch in BATT checks
that the FLCS batteries are discharging andthat inverter output is good.

• FLCS BATT (4).

Loose or foreign objects - Check.
• FLCS PWR (4).

I

WARNINGI

Loose or foreign objects in the cockpit
can cause entanglement or obstruction
with critical controls (e.g., ejection
handle, stick, rudder pedals, throttle,
canopy handle, etc.).

BEFORE STARTING ENGINE
1.

MAIN PWR switch - BATT.

• Verify lights off:
• FLCS PMG.
• ACFT BATT TO FLCS.
3.

FLCS PWR TEST switch - Release.

4.

MAIN PWR switch - MAIN PWR.
• Verify warning/caution lights on:• FLT CONT SYS.
• ELEC SYS.
• HYD/OIL PRESS.

To prevent possible depletion of battery power, do not allow MAIN PWR
switch to remain in BATT or MAIN
PWR for more than 5 minutes without
engine running or external power
applied.

• IpW2001

BUC.

• I PW2201

SEC.

• ENGINE.
Change 11

2-7

T.O. 1 F-16A-1

locked or unlocked. If this occurs, holdthe switch in the down (close) position
as soon as possible to relieve thejammed condition and prevent canopy
actuator motor burnout. If required,manually crank canopy toward theclosed position to relieve the jammed
condition. If possible, place MAINPWR switch to OFF and have theaircraft battery disconnected by theground crew. This action removes
power from the canopy actuator.

Do not apply external electrical power
for more than 30 minutes without
cooling air; longer operation may
damage electronic components whichcannot be turned off. Prior to applying
external power without cooling air,insure that all avionic equipment is off;
UHF radio operation is permissible.

NOTE
8.
Brake pedal deflection of 1116 inch
activates the brakes and bleeds thebrake/JFS accumulators. To avoid
brake activation and loss ofbrake/JFS
accumulator pressure, do not rest feeton brake pedals.
5.

EPU GEN and EPU PMG lights - Confirm off.

I WARNING I
If either light is illuminated, turn the
MAIN PWR switch to OFF. Insure that
the EPU safety pin remains installed
and notify maintenance. (The. EPU
will activate using hydrazine if the
EPU safety pin is removed.)

STARTING ENGINE IpW200i
Refer to figure 2-6 for danger areas. During a normal
start, the JFS RUN light comes on within 30 seconds
after moving the JFS switch to START 2. Mter the
JFS RUN light illuminates, engine rpm starts toincrease smoothly. At approximately 15 percent rpm,
the hydraulic and fuel pressures increase and theBUC caution light goes off. The throttle should be
advanced to IDLE after rpm reaches 20 percent.
Engine light-off occurs within 20 seconds after
throttle advance and is indicated by an ignition
vibration followed by an increase in rpm followed byan increase in FTIT. Without external power
connected, only the RPM and FTIT indicators
function until the generator is on line.

I WARNING I

At approximately 45 percent rpm, the main generator
comes on line. At approximately 50 percent rpm, the
JFS automatically shuts down. At approximately 55
percent rpm, the ENGINE warning light goes off. The
EEC trims to idle rpm within approximately 1minute.

[!J Failure of the canopy actuator could

FTIT increases smoothly at a moderate rate andnormally peaks at less than 575°C.

6.

Communications - Established.

7.

Canopy - As desired.

allow the canopy to fall during transit.
Keep hands and arms out of the path of
canopy travel during opening or
closing.

• To prevent possible engine damage,stow or secure loose cockpit items prior
to engine start with the canopy not
closed.
• ~ [Iffi When lowering or raising the

canopy handle, insure that the canopy
switch is not bumped to the up (open)
position before handle is completely
2-8

Chocks in place, fireguard posted, and intake
and other danger areas clear (ground crew).

Change 13l.

JFS switch - START 2.

2.

BUC caution light - Check off.

3.

Throttle - Advance to IDLE at 20 percent
rpm minimum .

If the FUEL MASTER switch IS
discovered in OFF during engine start(i.e., after the throttle is moved toIDLE), discontinue the start. Positioning the FUEL MASTER switch to
MASTER with the throttle out of OFF
may cause a hot start or tailpipe fire.

T. O. 1F- 16A- 1

If the engine has been run more than
30 minutes ago and less than 3-112
hours ago, do not advance the throttle
to IDLE until 40 seconds after the JFS
RUN light comes on. The time the JFS
RUN light was on during previous
start attempts is cumulative towardthe 40 seconds.

NOTE• If the HYD/OIL PRESS warning lightgoes off and stays off prior toapproximately 25 percent rpm, suspect failure of the engine oil pressure
sensor and abort the aircraft. This
check is not valid if less than 30
seconds elapse from movement of theMAIN PWR switch out of OFF until
both hydraulic pressures are above
1000 psi (i.e., low oil pressure timedelay has not expired; in which case,the HYD/OIL PRESS warning lightgoes off when hydraulic pressuresexceed 1000 psi, but may reilluminate
if the 30-second time delay expires
while oil pressure is still low ).
• If the main generator fails to come online automatically during engine
start, do not attempt reset. Although
the generator may reset, a problem
may still exist. Notify maintenance.

8.

OIL pressure - 15 psi (minimum).

9.

NOZ POS - 70-95 percent.

10.

RP1tI - 60-73 percent.

11.

FTIT - 575°C or less.

12.

HYD PRESS A & B - 2850-3250 psi.

13.

Six fuel pump lights (ground crew) - On.

NOTE
FFP light may blink at idle rpm .
If the main fuel shutoff valve is wired open, do not
perform step 14:

14.

Main fuel shutoff valve (ground crew) - Check.

I WARNING I
Do not make stick inputs while groundcrew is in proximity of control surface.
15.

JFS doors (ground crew) - Verify closed.

16.

Throttle cutoff release - Check.
Verify that the cutoff release does notremain in the actuated position by attempting to retard the throttle to OFF without
depressing the cutoff release. If throttle
moves to OFF, remain in OFF and notify
maintenance.

NOTE
4.

ENGINE warning light - Off (approximately
55 percent rpm).

*Engine at idle and check:5.

JFS switch - Confirm OFF.

Failure to perform this check afterengine start can result in an undetected stuck throttle cutoff release,which may lead to an unintentional
engine shutdown.

STARTING ENGINE IpW2201
If JFS switch does not automatically
return to OFF, turn JFS off. (Notify
maintenance after flight.)

6.

HYD/OIL PRESS warning light - Off.
The HYDIOIL PRESS warning light may
not go off until rpm is increased to 70
percent. If the warning light comes on again
at idle, notify maintenance.

7.

FUEL FLOW - 500-1500 pph.

Refer to figure 2-6 for danger areas. During a normal
start, the JFS RUN light comes on within 30 seconds
after moving the JFS switch to START 2. Mter the
JFS RUN light illuminates, engine rpm starts toincrease smoothly. At approximately 15 percent rpm,
the hydraulic and fuel pressures increase. The SEC
caution light goes off by approximately 18 percent
rpm. The throttle should be advanced to IDLE afterengine rpm reaches 20 percent. Immediately afteradvancing the throttle to IDLE, the SEC caution lightilluminates for about 3 seconds to indicate automatic

Change 11

2-9

T. O. 1 F- 1 6A- 1

DEEC/SEC ground self-test. Engine light-off occurs
within 20 seconds after throttle advance and isindicated by an ignition vibration followed by anincrease in rpm followed by an increase in FTIT.
Without external power connected, only the RPM and
FTIT indicators function until the main generator ison line.
At approximately 45 percent rpm, the main generator
comes on line. At approximately 50 percent rpm, the
JFS automatically shuts down. At approximately 55
percent rpm, the ENGINE warning light goes off.

• If the main generator fails to come online automatically during engine
start, do not attempt reset. Although
the generator may reset, a problem
may still exist. Notify maintenance.
• The SEC caution light illuminates
immediately after throttle advance forabout 3 seconds indicating automatic
DEEC/SEC ground self-test.
4.

•

FTIT increases smoothly at a moderate rate andnormally peaks at less than 575°C.
1.

JFS switch - START 2.

2.

SEC caution light - Check off.

3.

Throttle - Advance to IDLE at 20 percent
rpm minimum.

ENGINE warning light - Off (approximately
55 percent rpm).
•*Engine at idle and check:5.

JFS switch - Confirm OFF.

If JFS switch does not automatically• If the FUEL MASTER switch is
discovered in OFF during start (i.e.,
after the throttle is moved to IDLE),
discontinue the start. Positioning the
FUEL MASTER switch to MASTER
with the throttle out of OFF may cause
a hot start or tailpipe fire.
• If the engine has been run less than 4hours ago, do not advance the throttle
to IDLE until one minute after the JFS
RUN light comes on.

NOTE• If the HYD/OIL PRESS warning lightgoes off and stays off prior toapproximately 25 percent rpm, suspect failure of the engine oil pressure
sensor and abort the aircraft. This
check is not valid if less than 30
seconds elapse from movement of theMAIN PWR switch out of OFF until
both hydraulic pressures are above
1000 psi (i.e., low oil pressure timedelay has not expired; in which case,the HYD/OIL PRESS warning lightgoes off when hydraulic pressuresexceed 1000 psi, but may reilluminate
if the 30-second time delay expires
while oil pressure is still low ).

2-10Change 14

return to OFF, turn JFS off. (Notify
maintenance after flight.)
6.

HYD/OIL PRESS warning light - Off.
The HYD/OIL PRESS warning light may
not go off until rpm is increased to 70
percent. If the warning light comes on again
at idle, notify maintenance.

7.

FUEL FLOW - 500-1500 pph.

8.

OIL pressure - 15 psi (minimum).

9.

NOZ POS - 70-95 percent.

10.

RPM - 60-73 percent.

11.

FTIT - 575°C or less.

12.

HYD PRESS A & B - 2850-3250 psi.

13.

Six fuel pump lights (ground crew) - On.

NOTE
FFP light may blink at idle rpm.

T.O. 1F-16A- 1
If the main fuel shutoff valve is wired open, do not
perform step 14:

14.

Main
fuel
crew) - Check.

I

shutoff

WARNING

valve

NOTE

Proper VMS operation is verified by
the presence of each word in priority
sequence (i.e., PULLUP, ALTITUDE,WARNING, etc.). A brief LG warning
horn is also heard prior to the
WARNING and CAUTION words.

(groundI
2.

INS - Align:

Do not make stick inputs while groundcrew is in proximity of control surface.

• FCC PWR button - Off.

15.

JFS doors (ground crew) - Verify closed.

• FCNP FUNCTION knob - NORM or STOR
HDG.

16.

Throttle cutoff release - Check.
Verify that the cutoff release does notremain in the actuated position by attempting to retard the throttle to OFF without
depressing the cutoff release. If throttle
moves to OFF, remain in OFF and notify
maintenance.

• FCNP DATA knob - POS.
• INS PWR button - On.
• Alignment coordinates - Enter for NORM
align.
• FCNP DATA OPT button - Verify alignstatus on I S display.

NOTE

Failure to perform this check afterengine start can result in an undetected stuck throttle cutoff release,which may lead to an unintentional
engine shutdown.

AFTER ENGINE START• FCC PWR button - On.
NOTE• Entry of alignment coordinates isrequired even if internal coordinates
are exactly equal to parking spot
location .
•1.

TEST switch panel - Check:a.

~ PROBE HEAT switch - PROBE HEAT.

PROBE HEAT caution light off. If tLe
caution light illuminates, one or moreprobe heaters are inoperative or amonitoring system failure has occurred.
b.

Failure to enter alignment coordinates flags the alignment as degraded(NAV RDY light does not flash).

@~ PROBE HEAT switch - TEST.

PROBE HEAT caution light flashes 3-5times per second. If the caution light
does not illuminate or if it illuminatt~s
but does not flash, the probe heatmonitoring system is inoperative.

3.

l\IFL - Clear.

4. 1EW200 I UFC/BUC - Check.
Allow the enE,rine to operate approximately
2 minutes prior to starting the BUC check.
May be delayed until the BEFORE TAKEOFF check.
NOTE

C.

ff) PROBE HEAT switch - OFF.

d.

FIRE & OHEAT DETECT button - Test.

LESS ~ To preclude degrading the INSalignment, complete the UFC/BUC
check within 4.5 minutes after starting the alignment or defer check untilalignment is complete.

e.

OXY QTY test switch - Test and ~heck.

• Throttle - IDLE.

f.

MAL & IND LTS button - Te~t.

• EEC BUC switch - OFF.

Change 14

2-11

T.o. 1 F-16A-l• Nozzle - Closes smoothly to 35-50 percent.
If movement is erratic or hesitant, contact
maintenance.
• FTIT - Increases approximately 30°C.
• RPM - Decreases to UFC idle.

lation. If transfer does not occur, movethe throttle to OFF and inform
maintenance. If the engine autoaccelerates during transfer to or when
operating in BUe, shut down theengine and have maintenance investigate before further operation.
NOTE• EEe caution light - On.
• Throttle - Increase up to 70 percent rpm asconditions permit.

This throttle technique differs fromthat required during a BUe start.
• BUe caution light - On.

NOTE• Unless local circumstances dictate alower thrust setting, approximately 70
percent rpm should be used to reduce
the possibility of a stall if transfer
occurs immediately.
• In extremely cold conditions, 70
percent rpm may result in the throttle
being at or forward ofBUC IDLE.
• EEC BUC switch - BUC.

• Throttle - Verify engine response
throttle movement, then BUe IDLE.
• Throttle - Advance to 75 percent rpm.
• EEe BUC switch - OFF.
• Throttle - Retard to IDLE at first indicationof rpm increase.
• BUC caution light - Off.

• BUC IDLE detent drops in place (audible
click).
• If the engine fails to transfer back to

If an immediate transfer to BUC

occurs when BUC is selected and thethrottle is below BUe IDLE, rpm
decreases rapidly and the engine may
stagnate if the throttle is advanced.
Should immediate transfer occur
move the throttle to OFF and inforr~
maintenance.
• Throttle - Rotate outboard; advance quickly
past the BUe IDLE detent (to approximately
midrange); then rotate inboard and retard to
BUe IDLE, all in one smooth continuous
motion.

UFC and the throttle is retarded to
IDLE, the engine will probably stalland may stagnate. If this occurs, shut
down the engine. If no other problems
are indicated but FTIT remains above
500°C, JFS START 2 can be used to
motor the engine until FTIT is below
200°C.
• If the throttle is not retarded immedi-

ately after the first indication of rpm
increase, the rpm could accelerate tomore than 80 percent very quickly.
• EEC BUC switch - EEC.
• EEC caution light - Off.
• Nozzle - Opens steadily (70-95 percent).

Transfer (illuminated BUe caution
light) usually occurs within 1-2 seconds after the above throttle manipu-

2-12• FTIT - Decreases.
• RPM - Verify EEC idle.

10. 1F-16A- 1
5.

I PW2201 SEC - Check after the engine has
run at idle for at least 30 seconds. May bedelayed until the BEFORE TAKEOFF check.
• Throttle - IDLE.
• ENG CONT switch - SEC.
• NOZ POS - Less than 5 percent.

• D light - Steps 23-25.
The FLCS BATT lights and the ELEC SYS caution
light do not illuminate during these steps.
a.

LE FLAPS switch - AUTO.

b.

PITCH, ROLL, and YAW TRIM - Center.

c.

FCS CAUTION button and SERVOELEC RESET switch - Activate SERVO
and FCS CAUTION simultaneously andthen ELEC momentarily.

• SEC caution light - On.
• Throttle - Verify engine response
throttle movement, then IDLE.

to0). FLT CONT SYS caution light - Off.
d.

• ENG CONT switch - PRI.

CONFIG

warning(2). Test No. indicates 00.

• SEC caution light - Off.
Flight controls - Cycle.

SELF TEST switch - TEST.
(1). T.O.ILAND
light - On.

• NOZ POS - 70-95 percent.

6.

•

e.

ADV SLEW switch - ADV momentarily.
(1). Test No. indicates 00.

NOTE(2). MAL light is on.
To assist in warming the hydraulicfluid and removing air from thehydraulic system, maximum stick andrudder pedal inputs should be madeprior to running FLCS self-test.
7.

f.

(1). STBY GAINS light on in test No. 1.
(2). Self-test sequences through test No.
42.

FLCS self-test - Initiate and monitor.
NOTE• Entering system altitude during FLCS
self-test results in an incorrect system
altitude.
• During FLCS self-test, the CADC
provides a mach signal to the engine
that results in an rpm increase of
approximately 2 percent.

ADV SLEW switch - ADV momentarily
twice; hesitate between advances.

(3). Program stops.
(4). Test No. indicates 43.
g.

ADV SLEW switch - ADV momentarily.
(1). Tests sequence through test No. 47.
(2). Program stops.
(3). Test No. indicates 48.

• The FLCS PWR lights illuminate
individually during FLCS self-test as
follows:

h. ADV SLEW switch - ADV momentarily.
(1). Test No. advances to 51.

• A light - Steps 1-4.
• Blight - Steps 11-13.

(2). FLT CONT SYS, P, R, Y, and all five
SERVOS lights - On.

• C light - Steps 17-19.

(3). Program stops.

Change 1

2-13

T.O. 1 F-16A-l
1.

FCS CAUTION button and SERVOELEC RESET switch - Activate SERVO
and FCS CAUTION simultaneously andthen ELEC momentarily.

15.

• TEST - FR, AL pointers indicate 2000(± 100) pounds and totalizer indicates 6000
( ± 100) pounds. FWD and AFT FUEL LOW
caution lights illuminate.

(1). All FLCS lights - Off.
J.

ADV SLEW switch - ADV momentarily.

FUEL QTY SEL knob - Check:

The following values are based on JP-4 or JP-5/8:(1). Test No. advances to 54.

• NORM - AL pointer indicates approximately
2800/2940 pounds. FR pointer indicatesapproximately 3100/3250 ( 1m 1800/1890)
pounds.

(2). FLT CONT SYS, P, R, Y, and all five
SERVOS lights - On.
(3). Program stops.

• RSVR - Each reservoir indicates approximately 460/480 pounds.

For expanded FLCS self-test, refer to SUPPLEMENTAL PROCEDURES, this section.

• INT WING - Each wing indicates approximately 5251550 pounds.

k. FCS CAUTION button and SERVOELEC RESET switch - Activate SERVO
and FCS CAUTION simultaneously andthen ELEC momentarily.

• EXT WING - Each external wing tankindicates approximately 2300/2420 pounds(for full tanks).

(1). All FLCS lights - Off.

1.

SELF TEST switch - OFF. End of normal
FLCS self-test.

• EXT CTR - FR pointer indicates approximately 1800/1890 pounds (for full tank). AL
pointer drops to zero.

IWARNING I

NOTEThe sum of the individual fuel tanks
and the totalizer should agree wi thin
± 100 pounds with only internal fuel or± 300 pounds with external fuel.

If required to shut down and restart

the engine, the FLCS battery test must
be reaccomplished prior to start.
NOTE• FUEL QTY SEL knob - As desired.
While the engine is shut down, theaircraft battery continues to power the
FLCS with the MAIN PWR switch in
BATT or MAIN PWR. Place MAINPWR switch to OFF to conserve
aircraft battery power.

16.

EPU FUEL quantity - 95-102 percent.

After FLCS self-test completed:17.

Trim - Check:

8.

ECM knobs - As required.

NOTE

9.

SPD BRK switch - Cycle.

10.

TWS - As desired.

*11.

WHEELS down lights - Three green.

If the FLCS self-test is turned on, theTEF's retract, allowing observation oftrim inputs from the cockpit. Insure
FLCS self-test is turned off after trim
is checked.

12.

HUD - As desired.

*13.

REO - As desired.

*14.

SAl - Set.

2-14

Change 4• TRIMIAP DISC switch - DISC.
• Stick TRIM button - Activate in roll andpitch.

I

T.O. 1F-16A- 1

NOTE• No control surface motion.

If both the RDY and DISC lights are on,then one or more of the AR/FLCS
relays are failed. This condition must
be corrected prior to AR.

• No TRIM wheel or indicator motion.
• TRIWAP DISC switch - NORM.
• Stick TRIM button - Check and center.

• [ID AIR DISC button - Depressed. DISC lighton, RDY light off; 3 seconds later, RDY lighton, DISC light off.

• Control surface motion.
• TRIM wheel and indicator motion.

• AIR REFUEL switch - CLOSE, RDY lightoff.

• Rudder trim check.
• YAW TRIM knob - Check and center.

*22.

*18. rn:J FLCS override - Check:• One brake pedal- Depress.

• mEl STICK CONTROL - Selected cockpit.
• ~ Stick indicator - As selected.

• Confirm (ground crew) - Brake activates.

• Selected cockpit paddle switch - Depress.

• Opposite brake pedal- Depress.

• OVRD lights - On (both cockpits).

• Confirm (ground crew) - Initial brake does
not activate.

• Selected cockpit stick - Operative.

• Repeat above steps for opposite brake.

• Other cockpit stick - Inoperative.
*19.

MPO - Check.

23. I PW2201 Anti-ice - Check:• Stick - Full forward and hold; note horizontal tail deflection.

• Insure that the engine has been stabilized atidle rpm for 1 minute with ANTI ICE switch
ON.

• MPO switch - OVRD and hold.

• ANTI ICE switch - OFF.

• Confirm that horizontal tail trailing edges
move farther down.

• Monitor FTIT when the ANTI ICE switch
is moved to OFF. IfFTIT does not decrease
by at least 10°C within 15 seconds, abort
the aircraft.

• Stick and MPO switch - Release.
• Confirm that the horizontal tail returns to
its original position.
*20.

Operate controls - All surfaces respondnormally; no FLCS lights on.

*21.

AR system (if required) - Check.

• ANTI ICE switch - AUTO or ON as required.
24.

• AIR REFUEL switch - OPEN, RDY lighton, DISC light off.

Brakes - Check both channels; then returnto CHAN 1.

EPU GEN and EPU PMG lights - Confirm
off.
If either light is illuminated, cycle the EPU
switch to OFF and then back to NORM. If
either light stays on, abort the aircraft.

2-15

10. 1F-16A- 1
25.

EPU switch - OFF.

• TACAN function knob - TIR. (Allow 90
seconds for warmup.)

NOTE• RSI CRS knob - Set COURSE 180.

To prevent inadvertently selecting
ON, lift only the aft section of the split
guard.

• INSTR MODE knob - TCN.

26.

Ground safety pins (ground crew) - Remove.

27.

EPU switch - NORM.

28.

Intercom (ground crew) - Disconnect.

• TACAN TEST light - Flashes momentarily.

29.

Avionics - Program as required and verify(manual or data transfer cartridge).

• RSI range and course warning flags - In

·SMS.

• RSI bearing pointer may temporarily slew
to 270 degrees.

• TACAN TEST button - Depress momentarily (less than 1 second).

VIeW.

• RSI range and course warning flags - Outof view.
When AIM -9's are loaded, the SMS
switch must be maintained in the SMS
position while taxiing or airborne. If
recycling of the SMS switch is required, power should not be off longer
than 8 seconds at anyone time.

• RSI range indicator - Indicates 000 (± 0.5)
nm.
• RSI bearing pointer - Indicates 180 (± 3)degrees.

• Steerpoints.

• RSI CDI - Center within ± 1/2 dot.

• Profile data (including bingo fuel).

• RSI TO-FROM indicator - TO.

• DVAL - As required.

33. ~ [!] RF radio - As required.

NOTE

IWARNINGI
System altitude error results if D-value altitude calibration is accomplishedprior to completion of the first four
steps of the FLCS self-test.
30.

Refer to figure 2-6. Ground operationof RF radio is hazardous to personnel
and poses an ignition source forflammable fluids and electroexplosive
devices. Minimum safe distances must
be observed.

Radar - As desired.

*34.
In environments of high humidity,
delay turning the FCR on for approximately 4 minutes. This action precludes damage to FCR caused bymoisture in the ECS immediately afterengine start.
31.

Avionic BIT's - As desired.

32.

TACAN preflight check - Complete.

2-16

Change 12

Seat - Adjust to design eye.

BEFORE TAXI
1.

Canopy - Close and lock.
The canopy may be partially opened fortaxiing if required for increased visibility.

2.

LESS[M!] HAVE QUICK radio - Set and check(if required).

I

IO.lF-16A-1
brake use may cause residual heatdamage to brakes and may increase
the probability of a subsequent brakefire.

• Function knob - MAIN or BOTH.
• Mode knob - PRESET.
• For single WOD, enter the WOD in presetchannels 15-20 and, starting with channel
20, rotate the preset channel knob ccw until
a double beep is heard.
• After the MWOD and current date areentered, select channel 20, manually set
220.000 MHz, and depress the PRESET
button. Failure to complete this step renders
the radio inoperative.
• Enter the TOD by requesting a TOD
transmission. To receive TOD update for thenormal mode (UHF-TOD), position the
I HQ II I A-3-2-T knob, I HQ II . 2 IT-TONE
switch to T and then return to a normal
channel in either manual frequency or presetchannel mode. To receive TOD update for the
AJ mode, I HQ II I rotate the A-3-2-T knob to
T and then back to A, I HQ I I . 2 I place
T-TONE switch to T. A TOD abbreviated time
update can be received on the selected AJ net
on which the TOD is being transmitted. In
either mode, the transmission must be
received within 1 minute of selecting T.

TAXI
Refer to figure 2-7 for turning radius and groundclearance.
*1.

• To minimize heat buildup in the brakeassemblies, do not ride the brakes tocontrol taxi speed. Use one firm
application of the toe brakes to slow
the aircraft.
• Hydraulic fluid ignites at relatively
low temperatures. If advised of ahydraulic leak around the brakes,
request firefighting equipment and do
not taxi.
NOTE

Constant rudder pedal pressure
should not be required to maintain astraight ground track during taxi
(based on level surface and no asymmetries or crosswind) .

• Select an AJ net number with the manualfrequency knobs.
• , HQ II*3.

Brakes and NWS - Check.

I A-3-2-T,' HQ 11.2 I A-3-2 knob - A.

Altimeter and altitude indications - Set andcheck.
Check that the altimeter readings in
ELECT and PNEU are ± 75 feet of a known
elevation and are x ± 75 feet of one another.
Also check that the ELECT mode altimeterreading is x ± 75 feet of the altitude displayed in the HUD.

4.

Exterior lights - As required.

5.

FCNP FUNCTION knob - NAY.

6.

Chocks (ground crew) - Remove.

I WARNING I
If the aircraft has flown in the past 2
hours, do not use the parking brake
except for an emergency. Parking*2.

Heading - Check.

*3.

Flight instruments - Check for proper operation.

BEFORE TAKEOFF* 1.

ALT FLAPS switch - NORM.

2.

Trim - Check pitch and yaw trim centeredand roll trim as required.

IWARNING

1

Failure to center yaw trim may result
in severe control difficulties during
takeoff.
3. IPW200 Ilm rn£] STARTING
switch - AUTO LEAN.

Change 1

FUEL

2-17

T.0.1F-16A-1

4.

I PW200 I [8J

rntJ EEC BUC switch - EEC.

• ill OXYGEN mode lever - As required.

5. IPW200 l!!ml BUC switch - OFF (guard down).
PW2201 [8J rntJ ENG CONT switch -

6.

I

7.

I PW2201 !!ml ENG CONT

switch - NORl\1(guard down).

8.

Speedbrakes - Closed.

9.

Canopy - Close, lock, light off.

10.

IFF - Set and check.

I

PRI.

13.

External tanks (if installed) - Verify feeding.
If three tanks are installed, verify that thecenterline tank is feeding. This action
checks that pressurization is available toall tanks.
FUEL QTY SEL knob - NORM.
NOTEThe FUEL QTY SEL knob must be in
NORM for operation of the automatic
forward fuel transfer system, LESS ~trapped fuel warning, and for the
BINGO fuel warning computation to
be based on fuselage fuel.

I

14.

STORES CONFIG switch - As required.

*15.

GND JETT ENABLE switch - As required.

*16.

Harness, leads, and anti-g system - Check.

I• ill Failure to select PBG when PEG
operation is desired can contribute tograylblackout or loss of consciousness.

• ill Placing the OXYGEN mode lever to
PEG without wearing the CRU -94/P
harness-mounted connector can expose
the pilot to injury producing levels ofoxygen pressure should a failure occur
within the oxygen regulator.

11. [till] AIFF - Set and check.
12.

WARNING17.

EPU - Check:
NOTE• Any failure of the INU to maintain fullnavigation capability during the EPU
check indicates a possible weak ordepleted INU battery. This condition
could result in loss of INS data after
any inflight power interruption.
•@ Verify normal navigation capability

after any aircraft power interruption.
The LN-93 RLG may shut down andautorestart to only the attitude modewith the FCNP FUNCTION knob in
NAY.
• ECM control panel - Power off.

• mOXYGEN mode lever - PBG.
• Anti-g system - Test.
Depress TEST button to check that g-suit
inflates, comfort zippers are closed andremain closed, and hose remains connected
and that pressure is supplied tomask, helmet bladder, and vest.

m

IWARNINGI
Delayed or slow inflation of the g-suit
could indicate impending anti-g system failure. Complete loss of g-protection could occur with no warning.

2-18

Change 12

NOTE• EPU may be checked anytime afterengine start.
• If power is applied to the ECM pod

during the EPU check, voltage transients may cause nonrecoverable software faults within the pod.
• EPU safety pin (ground crew) - Checkremoved.
• OXYGEN - 100tJr .
• Engine rpm - Increase 5 percent above
normal idle.

T.O. 1 F-16A-l

If electrical bus cycling is experi-

enced, abort the aircraft and do not
taxi.
Attempting EPU/GEN TEST at idle
rpm results in abnormally low EPU
run speed. Low EPU speed may cause
electrical bus cycling and possible
damage to electrical equipment. Indications may include blinking EPU
GEN light or audible clicking ofelectrical con tactors as the bus cycles.

• EPU/GEN TEST switch - EPU/GEN andhold. Check lights:• EPU AIR light - On.
• EPU GEN and EPU PMG lights - Off (maycome on momentarily at start of test).

Change 8

2-18.1/(2-18.2 blank)

T. O. 1 F- 1 6A- 1• FLCS PWR lights - On.

IWARNING• EPU run light - On for a minimum of 5
seconds.

If the RDR ALT switch is not in RDR

ALT, the ALTITUDE voice message is
not available for descent warning after
takeoff or AGL ALOW.

NOTE
If EPU run light does not come on

19. ~ LESS I@I SNSR PWR panel:

within 10 seconds after EPU/GEN is
selected, release EPU/GEN TEST
switch, advance the throttle to idle
rpm plus 10 percent, and reselect EPUI
GEN. If the EPU run light does notilluminate within 10 seconds or if itcycles on and off, abort the aircraft.

a. TF POD switch - OFF.
b. EO POD switch - OFF.
c.

• EPU/GEN TEST switch - OFF.
20.

• Voltage transients may cause fault lightindications. Reset FLT CONT SYS, CADC,AVIONICS, I PW200 I EEC, and I PW2201
ENGINE FAULT caution lights as necessary.

I WARNING I• OXYGEN - NORMAL.

I WARNING I• Airflow is indicative of a failed highpressure bleed valve which could
result in an EPU overspeed in flight. If
airflow is detected, abort the aircraft.
• If required to shut down and restart
the engine after completing the EPU
check or if the EPU safety pin isreinstalled for any reason, the EPU
check must be reaccomplished prior toflight to insure proper EPU operationand closure of the high-pressure bleed
valve.
• ECM control panel- Power as required.
18.

IAJ rnEJ RDR ALT switch (if installed) - RDR
ALT.
Nuisance ALOW warnings may occur at LG
retraction or during intermediate level offs
until above the entered ALOW value.

RDRALTswitch - RDRALT.
Nuisance ALOW warnings may occur at
LG retraction or during intermediatelevel oft's until above the entered ALOW
value.

PROBE HEAT switch - PROBE HEAT.
Turn probe heat on at least 2 minutes prior
to takeoff anytime icing of probes is possible.
When probe icing is not suspected, delay
selection of probe heat as long as possible
prior to takeoff.

• Throttle - IDLE.

• EPU exhaust (ground crew) - Check for no
airflow.

I• Probe internal icing must be suspectedanytime the aircraft has been exposed
to near or below freezing conditions on
the ground. Internal icing may bedifficult to see and may remain present
even when current conditions do notappear conducive to ice formation .
• AOA probes may become hot enough toshut off probe heaters and cause
illumination of the PROBE HEAT
caution light. Refer to CAUTION
LIGHT ANALYSIS, Section III, todetermine if the shutdown was due tooverheat or system failure.
• If the probe heaters are on or have beenon, they may be hot enough to cause
serious injury if touched.
*21.

Ejection safety lever - Arm (down).

*22.

Flight controls - Cycle.

*23.

OIL pressure - Check psi.

*24. All warning and caution lights - Check.
25. Adjustable sliding holder (when utility lightis not in use) - 00 [§£I Full forward, rotated cw,
and secured.
Change 2

2-19

T.O.1F-16A-l

I WARNING I
Failure to properly secure utility lightand adjustable sliding holder can
result in stick interference.

TAKEOFFNORMAL TAKEOFF

I WARNING I
I PW200 I During AB operation, if thethrottle is retarded far enough to shut
off the AB (nozzle position less than 20
percent), the throttle should be positioned to MIL for a minimum of 1.5
seconds before reselecting AB. Failure
to do this may result in mislight andcould cause a compressor stall or
stagnation.
NOTE

Refer to figure 2-2. When ready for takeoff:• Advance throttle to approximately 80 percent rpm.
• Verify parking brake disengaged.
• Check engine instruments for normal indications.
• Release brakes prior to exceeding 80 percent rpm.
• Advance throttle to desired thrust.
NOTE

An engine runup check is not required
if conditions require a rolling takeoff.
Maximum FTIT and rpm vary with temperature andpressure altitude but stabilize in 5-15 seconds.

ml During an MIL takeoff, the nozzle closes, thenopens slightly, before stabilizing at 5 to 20 percent.
During hot weather operation expect the nozzle tostabilize at 25 percent or less.
Nozzle position during an MIL takeoff should neverexceed 30 percent after 5 seconds at MIL.

Spacing of less than 15 seconds
between aircraft when AB is used bypreceding aircraft increases the probability of an AB blowout or no light due
to hot gas ingestion.

m When airspeed is approximately 10 knots below
computed takeoff speed for non-AB or 15 knots for
AB, initiate rotation to establish takeoff attitude(8-12 degrees). Do not apply aft stick at airspeed
lower than 10-15 knots below computed takeoff
speed. Early rotation can lead to overrotation,
skipping, wallowing due to early lift-off, andincreased takeoff distance.
LESS m When airspeed is approximately 15 knots
below computed takeof( speed, establish takeoff
attitude (8-12 degrees).

As aircraft lifts off, LEF's extend downward. Retract

LG when safely airborne. Insure LG is up and locked
before exceeding 300 knots. TEF's retract when the
LG handle is raised.

• Since LG and TEF retraction occurs
simultaneously, LG retraction should
not be rushed after takeoff. The
reduction in lift may cause the aircraft
to descend and contact the runway.

Normal engine operation during an AB takeoff isindicated by:• Due to low aft stick forces required for
takeoff, use caution to avoid earlyrotation.

• The nozzle beginning to open within 5 seconds after
selecting AB.

• If any FLCS lights (LE FLAPS, P, R, Y,
ADC, or CADC) illuminate during
takeoff, a WOW switch problem may be
indicated. If the light(s) resets, themission may be continued. Write up
the occurrence after the flight.

., PW200 I Nozzle position of95 percent or less after 5

seconds at MAX AB.
2-20

Change 5

T. O. 1F- 16A- 1
NOTE

Insure that the LG handle is placed
fully up. The handle can stop in anintermediate position which retracts
the LG; however, the handle is notlocked and may lower under high g
conditions.

It is possible to exceed the roll trim
authority of the aircraft for an on
speed takeoff with a net asymmetric(rolling) moment less than aircraft
takeoff limits. Refer to ASYMMETRICSTORES LOADING, Section V.

TAKEOFF WITH ASYMMETRIC STORES

Roll trim should be set prior to takeoff withasymmetric stores to prevent wing drop. The amountof roll trim required for various asymmetric store
weights is shown in T.O. IF-16A-l-l, Part 2.

When ARI activates after takeoff, roll trim forasymmetric stores causes a rudder input that cancause aircraft yaw away from the wing with theasymmetric store. This yaw is easily controllable bypilot rudder inputs.

Change 11

2-20.1/(2-20.2 blank)

T. O. 1 F- 16A- 1

Normal Takeoff (Typical)~.
-., ,--,.

--.:~ ~ ------~===~-=---==---"'=---=,-_--==~

-

--_00

TAKE OFF AT
PREDETERMINEDTAKEOFF SPEED

USE NWS UNTIL RUDDER IS
EFFECTIVE (APPROXIMATELY
70 KNOTS)ROTATE AT
PREDETERMINEDROTATION SPEED

1F-16A-1-1136X@

Figure 2-2.

OPERATING AIRSPEEDS BELOW 10,000
FEET MSL
Below 10,000 feet MSL, airspeeds for areas outside
special use airspace are as follows:
300-350

Point-to-point navigation and formation
rejoins that do not occur on departure.
Descents into an MTR.
Non-lAP descents into the terminal
area.
Tactical initial.
If a route abort or unplanned climb
causes the aircraft to exit the MTR.

350-400

Formation rejoins on departure.

400-450

Initial entry airspeed for g-awarenessexercises.

CLIMB
The climb schedules are defined by airspeed/machnumber or mach number only. When airspeed/machnumber is shown, climb at the scheduled airspeed tothe scheduled mach number, then maintain the machnumber to the desired altitude. When starting a climb
at an altitude above the airspeed/mach transition
point, climb at the scheduled mach number.
The recommended MIL climb schedule to optimum
cruise altitude is tabulated versus drag index in T.O.
1F-16A-1-1, Part 3. For MIL, the schedule provides
optimum range performance_ The recommended MIL
climb schedules to altitudes other than optimum
cruise altitude are shown in T.O. 1F-16A-1-1 and T_O_
1F-16A-1CL-1.
The recommended MAX AB climb schedule istabulated versus drag index in T.O. 1F-16A-1-1, Part
3. The MAX AB climb schedule provides minimum
time to climb performance.

Change 11

2-21

T. O. 1F- 16A- 1

NOTE

CLiMB/IN-FLIGHT/OPERATIONAlCHECKSAt frequent intervals, check the aircraft systems,
engine instruments, cockpit pressure, and oxygen
flow indicator and system operation. Monitor fuel ineach internal and external tank to verify that fuel istransferring properly by rotating the FUEL QTY SEL
knob and checking that the sum of the pointers andtotalizer agree and that fuel distribution is correct.

I

WARNING• I PW200 I Higher fuel flows (refer to T. O.
IF-16A-l-l) are required to achievethrust when the nozzle is scheduled
partially open. The greatest impact onfuel flow occurs when flying at low
altitude with the throttle below midrange. As much as 500-600 pph
additional fuel flow may be required.
When flying at high altitude or withthe throttle above midrange, impact onfuel flow is negligible.
1.

Fuel - Check quantity/transfer/balance.

2.

FUEL QTY SEL knob - NORM.

Change 13

3.

Oxygen system - Check.

4.

Cockpit pressurization - Check.

I• Maximum fuel transfer rate is 18,000
pph from the 300-gallon fuel tank or
30,000 pph from the 370-gallon fueltanks. Maintaining fuel flow above
these values while the external tank(s)is feeding results in a decrease ofinternal fuel. Prolonged operation
under these conditions may result in
the rapid depletion offuselage fuel andrender fuel transfer by siphoning
action inoperative. Without siphoning
action, fuel transfer to the fuselage
tanks is provided by the wing turbine
pumps at a maximum rate of6000 pph.
A fuel flow rate greater than 6000 pph
continues to deplete fuselage fuel.
Under these conditions, the externalfuel tank(s) may appear slow to feed
and a fuel imbalance may result.
Prolonged AB operation in a three tankconfiguration may result in engine
flameout prior to depletion of external
fuel.

2-22

The FUEL QTY SEL knob must be in
NORM for operation of the automatic
forward fuel transfer system, LESS ~trapped fuel warning, and for the
BINGO fuel warning computation to
be based on fuselage fuel.

I

WARNINGIThe CABIN PRESS caution light does
not illuminate until cockpit pressure
altitude is above 00 22,500, LESS 00
27,000 feet. At ambient altitudes
where hypoxia is possible, do notremove the oxygen mask without firstchecking cockpit pressure altitude.
5.

I

Engine instruments - Check.

AIR REFUELING PROCEDURES
Refer to T.O. l-lC-1 for general AR procedures, andto AIR REFUELING PROCEDURES, Section VIII
for specific AR procedures.

DESCENT/BEFORE LANDING
1.

Fuel- Check quantity/transfer/balance.

2.

Final approach airspeed - Compute .

3.

DEFOG lever/cockpit heat - As required.

4.

Landing light - On.

*5.

Altimeter and altitude indications - Check
altimeter setting, ELECT versus PNEU
mode altimeter readings, and ELECT mode
altitude versus altitude displayed in HUD.
For subsonic flight below 20,000 feet MSL
with vertical velocity less than 500 fpm, thedifference between ELECT and PNEU
mode altitudes should not exceed 270 feetand the difference between the ELECT
mode altitude and the altitude displayed in
the HUD should not exceed 75 feet.

IT. o. 1F- 16A-1

I

WARNINGI

e An erroneous ELECT mode altitude can

e Failure to depress the LG handle down

be displayed without a CADC caution
light or a transfer to PNEU mode.

permission button prior to attempting
to lower the LG may result in damage
to the electrical solenoid.

e An erroneous altitude can be displayede Failure to reduce sink rate, particular-

in the HUD without a CADC caution
light.
*6.

Attitude references - Check ADIIHUD/SAI.

7.

ANTI ICE switch - As required.

LANDINGly at heavier GW's, may cause a firm
landing and structural damage orfail ure of the LG.

e lID Use of the paddle switch may cause
pitch and/or roll transients as control
is switched from one cockpit to theother.

e Avoid landing directly on approachNORMAL LANDING

Refer to figure 2-3. Fly initial at 300 knots. At the
break, retard throttle and open speedbrakes as
required. On downwind leg, when airspeed is below
300 knots, lower the LG. During base turn, recheckthe LG down and slow to computed final approach airspeed to arrive on final at 11 or 13 degrees AOA.
Check speedbrakes open and maintain computed
final approach airspeed/AOA on final. Rate of descentdecreases slightly when entering ground effect.
Reduce thrust gradually to continue the descent
while applying back stick to reduce sink rate to theminimum practical. Thrust can be reduced sooner
during an II-degree approach than during a 13degree approach. In either case, maintain amaximum of 13 degrees AOA while reducing sink rate
to the minimum practical.

IWARNING

1e Physically confirm that the LG handle
is fully down. The LG handle may
visually appear to be down when in anintermediate position. An intermediate position may allow LG extension
and/or safe indications; however, theLG handle is not locked and LG
retraction could occur during subsequent in-flight or ground operations.

e Failure of the ANTI-SKID switch canallow it to be bumped/placed towards
PARKING BRAKE while airborne. A
very small movement out of ANTISKID is sufficient to engage theparking brake. Landing with theparking brake engaged will result in
main tire failures upon touchdown.

end arresting cable to prevent possible
cable strike damage to nozzle,
speedbrakes, and ventral fins.

e Horizontal tail contact with the runway is possible if a large roll input ismade at or near touchdown.

NOTE

e Ifrn ~ Deploying the drag chute
above 170 knots may result in loss ofthe chute canopy.

elfrn [NQI Drag chute deployment below
90 knots may result in improper
deployment and damage to the chute.

e The LG warning horn and the T.O'!
LAND CONFIG warning light are
inhibited at approach airspeed above
190 (± 4) knots.

e rn Aft CG approaches may be characterized by increased pitch sensitivity
which will be most noticeable uponentering ground effect.

t:rrn ~ Deploy the drag chute (if desired) immediately
after touchdown. The nose may pitch up or downwhen the chute is deployed, but the motion is easily
controlled.

NOTE

e Ifrn !r!QJ To deploy the drag chute after
touchdown, lift the guard and switch inone single motion with the side of theindex finger.

e ~ !r!QJ When deploying the drag chutein a two-point aerodynamic braking
attitude, the drag chute may contact
the runway.
Change 13

2-23

10. 1 F-16A-l

Normal Landing Pattern (Typical)
LEVEL BREAK• THROTILE - AS REQUIRED• SPEEDBRAKES - AS REQUIREDAPPROACHING BREAKAIRSPEED - 300 KNOTS• AOA - 13 DEGREES(MAX)• LG- DN
• SPEEDBRAKES - AS
REQUIREDNWSSELECTED• ~INOIDRAG
CHUTE - AS
REQUIRED• SPEEDBRAKES FULLY OPEN
(AFTER NLG WOW)
• BRAKES -AS
REQUIREDTOUCHDOWNAOA - 13 DEGREES (MAX)FINAL

BASE LEG• AOA - 13 DEGREES (MAX)• LG - CHECK DN• SPEEDBRAKES - AS
REQUIRED
NOTE:•

FINAL APPROACH AIRSPEED/13 DEGREES AOA CROSS-CHECK .

• cru 125 KNOTS + 4 KNOTS PER 1000 POUNDS OF FUEL/STORE WEIGHTS.
ADD 8 KNOTS FOR 11 DEGREES AOA APPROACH .

• 00 129 KNOTS + 4 KNOTS PER 1000 POUNDS OF FUEL/STORE WEIGHTS.
ADD 8 KNOTS FOR 11 DEGREES AOA APPROACH.

I•

THE PRECEDING BASELINE AIRSPEEDS ARE BASED ON THE BASIC OPERATING WEIGHT FROM
T.O. 1F-16A-1-1 PLUS FULL AMMO. ACTUAL APPROACH AIRSPEED AT 11/13 DEGREESAOA MAY DIFFER BY +/-5 KNOTS DUE TO VARIATIONS IN AIRCRAFT CG.
IF-16A-l-0083A @Figure 2-3.

2-24

Change 2

T.O.1F-16A-l
Use two-point aerodynamic braking until approximately 80 knots; then fly the nosewheel to runway.
Maximum effective two-point aerodynamic braking isachieved at 13 degrees AOA. An AOA less than 11
degrees results in significantly reduced two-point
aerodynamic braking. Two-point aerodynamic braking below 80 knots is ineffective and increases groundroll.

• Do not touch down with brake pedals
depressed. A failure in either the
touchdown protection circuitry or an
MLG WOW switch can result in locked
wheels and blown MLG tires.
• Use a maximum of 13 degrees AOA for
two-point aerodynamic braking. Nozzle,
speedbrakes, and ventral fins may
contact runway if 15-degree pitch
angle is exceeded.
• During two-point aerodynamic braking, the speedbrakes (43 degrees orgreater open) may contact the cable.
• At WOW, AOA indications go toapproximately 13.5 degrees and are nolonger a valid indication of aircraft
attitude.
• If any FLCS lights (DUAL FC FAIL,
FLT CONT SYS, LE FLAPS, P, R, Y,
ADC, or CADC) illuminate at WOW, a
WOW circuit problem is probably at
fault. Use two-point aerodynamic
braking with caution since a WOW
circuit problem which affects two
FLCS branches (DUAL FC FAIL) may
also cause a pitchup. Write up theoccurrence after the flight.

• During the landing phase, large/rapidroll control inputs in reaction toturbulence or wake vortices will cause
temporary retraction of one andsometimes both flaperons. This retraction will decrease lift and may induce
a sink rate beyond the structural limitof the landing gear. During rapid
reversal of roll inputs, both flaperons
might move up to a position that will
illuminate the T.O.lLAND CONFIG
warning light. Illumination of multiple FLCS servo lights is also possible.
Be prepared to initiate a go-around if
wake turbulence is encountered.

NOTE

The LE FLAPS caution light may comeon shortly after touchdown if landing
conditions cause asymmetric wheel
spinup. Illumination of the LE FLAPS
caution light for this reason is normal.

•

Mter the nosewheel is on the runway, open the
speedbrakes fully and maintain full aft stick for
maximum three-point aerodynamic braking andwheel braking effectiveness.

• Crossing an arresting cable in athree-point attitude above 90 knots
ground-speed with a centerline store
may cause cable strike.
• Do not move SPD BRK switch to open
until the nosewheel is on runway as
speedbrakes may contact runway.
• Until WOW, forward stick pressure inexcess of approximately 2 pounds
results in full trailing edge down
deflection of the horizontal tails. This
horizontal tail deflection reduces
wheel braking effectiveness. At highspeeds in the three-point attitude,
forward stick results in excessive loads
on the NLG which can lead to nose tire
failure and possibly cause structural
failure of the NLG.
Smoothly apply moderate to heavy braking todecelerate to taxi speed. Using less than moderate
braking increases the likelihood of a hot brake(s).
NWS should not be engaged above taxi speed unlessrequired to prevent departure from prepared runway
surface.

IWARNING

1

NWS malfunctions at any speed may
cause an abrupt turn, tire skidding or
blowout, aircraft tipping, and/or
departure from the prepared surface.

Change 12

2-25

T.O. 1F- 16A- 1
SHORT FIELD LANDING (DRY RUNWAY)
NOTE

The following procedures should be
used any time stopping distance iscritical, whether due to a long, fast,heavy weight, or short field landing.
When stopping distance is critical, a normal approachshould be made. Select IDLE at or slightly before
touchdown. Touch down as near as possible to the endof the runway at 13 degrees AOA. ~ INQ] Deploy thedrag chute immediately after touchdown. Two-point
aerodynamic and wheel braking should be used with
the nose held up at 13 degrees AOA until the nose
falls. Pitch must be held at 13 degrees AOA if
two-point aerodynamic braking is to be effective.
Maximum effort braking is achieved by using thewheel brakes in conjunction with two-point aerodynamic braking. When the wheel brakes become
effective, the nose automatically lowers. This occurs
soon after brakes are applied.
After the nosewheel is on the runway, maintain fullaft stick, open the speedbrakes fully, and use
maximum wheel braking (antiskid on).
For landing on icy/wet runways, refer to LANDING
IN ICY OR WET CONDITIONS, Section VII.
CROSSWIND LANDING

The recommended technique for landing in acrosswind is to use a wing level crab throughtouchdown. At touchdown, the ARI switches out.
Undesirable yaw transients may occur if roll control
is being applied at this time. After touchdown,
perform two-point aerodynamic braking using therudder to maintain aircraft track down the runway
and flaperon to prevent wing rise. In crosswinds, theaircraft may drift downwind due to side loads
imposed by the crosswinds or travel upwind due toinsufficient directional control inputs/availability. As
the airspeed decreases, increasing amounts of rudder
are required to maintain track. Maintain two-point
aerodynamic braking until approximately 80 knots or
until roll or directional control becomes a problem. As
the pitch attitude decreases, the nose tends to alignitself with the ground track.
Aft stick and fully opened speedbrakes reduce
stopping distance. Apply brakes after nosewheel is on
the runway; however, if stopping distance is a factor,
refer to SHORT FIELD LANDING, this section. With
all LG on the runway, maintain directional control
with rudder, differential braking, and NWS if
required.
2-26

Change 11

During landing rollout, the main concerns are wingrise (roll contro!), weathervaning (directional control), and downwind drift. Wing rise is controlled byflaperon into the crosswind. Excessive flaperon
deflection degrades directional control. Use rudder
and differential braking to control ground track,
especially on wet or icy runways. Engage NWS if
required to maintain directional control and toprevent departure from the runway. Excessive
differential braking may result in a hot brakecondition. High rudder pedal force may result in a
yaw transient when NWS is engaged. NLG strut
compression is required to engage NWS but
sustained forward stick may result in full horizontal
tail deflection which decreases weight on the MLG
and thus reduces wheel braking effectiveness. !!rn [fill]
NWS engagement may be required with the dragchute deployed to control increased weathervaning
tendencies. However, the nose up pull of the dragchute may prevent early NWS engagement.

IWARNING

1!!rn ~ Be prepared to release the dragchute during the landing rollout if
directional control or downwind drifting becomes a problem.
NOTE[N§ INQ] Deploying the drag chuteduring two-point aerodynamic braking with a crosswind may complicate
aircraft control.
TOUCH-AND-GO LANDING

Perform a normal approach and landing. Aftertouchdown, maintain landing attitude, advance thethrottle, close the speedbrakes, and perform a normal
takeoff.

AFTER LANDING
IWARNINGI
Do not use parking brake. Use only
chocks, if available, or minimum
possible toe brake pressure to hold theaircraft stationary. Parking brake use
may cause residual heat damage tobrakes and may increase probability ofsubsequent brake fire. [N§ ~ Parking
brake may be used if drag chute was
deployed on landing.

10. 1 F-16A-l
NOTE~ NWS disengagements are possible

when taxiing with CG near thein-flight aft limit.

PRIOR TO ENGINE SHUTDOWN
1. EPU safety pin (ground crew) - In.
NOTE

1. ~ [fjQI DRAG CHUTE switch - Release as
required.

Installation of the EPU safety pinshould be delayed un til after engine
shutdown under the following conditions:• The ground crew recovering theaircraft is not familiar with F-16
danger areas.

~ ~ Turn aircraft into the windprior to releasing drag chute.
2.

PROBE HEAT switch - OFF.

• The aircraft is being recovered byemergency
response
personnel(landing with activated EPU, hotbrakes, etc.)

I WARNING I

Place the EPU switch to OFF prior toengine shutdown if the EPU safety pin
is not installed.

If the probe heaters are left on, they
may be hot enough to cause serious
injury if touched.
2.

FCNP - Record data.
• MFL (FCC on) - Record.

Prolonged ground operation of probeheat may cause failure of AOA probe
heaters.

•3.

ECM power - Off.

4.

Speedbrakes - Close.

• INS miscellaneous data (FCC off) - Record.
• @) Locations 19 (RER), 20 (CEP), 62 (align*5.

Ejection safety lever - Safe (up).

6.

IFF MASTER knob - STBY.

7.

LESS ~ IFF mode 4 - HOLD .

8. ~ AIFF MASTER knob - STBY.
9.

~ AIFF mode 4 -10.

LANDING TAXI lights - As required.

11.

~ MASTER ZEROIZE switch - As required.

12.

Canopy handle - Up.

events), 64 (NAV events), 66 (VX), and 67(VY).

• LESS@) Locations 128 (RER), 150 (VN), and151 (VE).
NOTE• @ID The INS is out of tolerance if the

HOLD.

NOTE

Unlock the canopy to insure that the
canopy seal is deflated before thecanopy is opened.
13.

• Steerpoint of current location - Select.

Armament switches - Off, safe, or normal.

RER is greater than 3 nmlhour oreither velocity is greater than 5 fps.
• LESS @ID The INS is out of tolerance if

the RER is greater than 3 nmlhour oreither velocity is greater than 3.9
knots.
3.

HUD CAMERA switch - OFF.
NOTE

The camera must be turned off at least
15 seconds prior to engine shutdown toallow the tape to unthread.

Change 12

2-27

T.O. 1 F-16A-1

4.

TWS, INS, and FCNP - OFF:

4.

eTWS- Off.

JFS RUN light - Check.
Notify maintenance if the JFS RUN light isflashing after the throttle is placed to OFF.

Mter main generator drops off line:e INS PWR button - Off.

5.

EPU GEN and EPU PMG lights - Confirm off.

I

e FCNP FUNCTION knob - OFF.
NOTEWARNINGI

If either light is illuminated, turn the
MAIN PWR switch to OFF. Insure that
the EPU safety pin remains installed
and notify maintenance.

The INS requires aircraft power for 10
seconds after the INS is turned off toinsure flight data is stored in memory.
5.

Radar - OFF.

6.

MAIN PWR switch - OFF.

6.

LESS ~ PROBE HEAT switch - PROBEHEAT and then OFF to allow check of heaters
after shutdown.

7.

Oxygen hose, survival kit straps, lapbelt, g-suit
hose, m and vest hose - Disconnect, stow.
e Stow oxygen connector in bracket on rightsidewall. Insure oxygen hose does not
protrude beyond console edge.

I WARNING I• Stow lapbelt and survival kit straps on seatcushion.

If the probe heaters are on or have beenon, they may be hot enough to cause
serious injury if touched.

Prolonged ground operation of probeheat may cause failure of AOA probe
heaters.

ENGINE SHUTDOWN
When ready to shut down the engine, oil scavenge
should be performed, conditions permitting.
1.

2.

Throttle - Advance to 75-78 percent rpm(stabilize for 5-10 seconds).

• Use both hands to disconnect g-suit hose to
avoid excessive force on the hose-to-console
connection.

• One-handed or brute force disconnectsof the g-suit connection will causeinternal damage to the hose at thehose-to-console connection.
• Failure to properly stow lapbelt,
survival kit straps, oxygen connector,
g-suit hose, and oxygen hose may
cause damage to consoles and to theejection seat during seat adjustment.
8.

OXYGEN REGULATOR - OFF and 100%.

Throttle - Retard to IDLE for 1-2 seconds.
NOTE

Do not wait for engine to stabilize inidle prior to shutdown. Timing begins
when the throttle reaches IDLE.
Waiting longer than 1-2 seconds after
the throttle reaches IDLE negates theeffects of the scavenge shutdown.
3.

Throttle - OFF.

2-28

Change 13• Failure to position the oxygen regulator to OFF and 100% may result inparticulate contamination of the regulator and subsequent damage.

• m To avoid damage to the oxygen
regulator, do not pull the knob on theend of the mode lever when moving themode lever from ON to OFF.

10.lF-16A-l
9.

Avionics - As required.

10.

Canopy - Open.

IWARNINGSCRAMBLEI• If winds exceed 30 knots, open the
canopy only as far as needed to enter/exit
the cockpit. Decreasing the canopy angle
reduces the possibility that the canopy
can be blown past full open.

• ~ A failure of the canopy actuator could
allow the canopy to fall during transit.
Keep hands and arms out of the path of
canopy travel during opening or closing.
NOTE

If the internal canopy switch is left inthe up position, the canopy automatically opens if closed from the outside.

PREFLIGHT

Perform the following preflight inspections prior toplacing the aircraft on quick response status:
1.

EXTERIOR INSPECTION.

2.

BEFORE ENTERING COCKPIT.

3.

COCKPIT INTERIOR CHECK .

4.

BEFORE STARTING ENGINE.

5.

STARTING ENGINE.

6. AFTER ENGINE START (include EPU checkbut do not remove MLG ground safety pins).
7.

Aircraft cocked for scramble - Per local
policies and directives.

Change'3

2-28. , /(2-28.2 blank)IT. O. 1F- 16A- 1
AIRCRAFT ON QUICK RESPONSE STATUS

12.

If the above actions were not completed prior toscramble, normal preflight procedures should beused.

I WARNING I

1. FLCS power - Check.
With MAIN PWR switch in BATT, verify
that FLCS inverter output is good (FLCS
PWR lights) with the aircraft battery andwith the FLCS batteries.

•

2.

MAIN PWR switch - MAIN PWR.

3.

Engine - Start.

4.

Canopy - Close and lock.

5.

Instruments - Check.

6.

INS - Align.

If either light is illuminated, the EPU
activates using hydrazine when the
EPU safety pin is removed.
13.

• The FCNP displays FCC FAIL for
approximately 20 seconds after themain generator comes on line. FCNP
data entry is disabled during this time
period.

EPU - Check (if EPU safety pin was
installed since last EPU check).

I WARNING I
If the EPU safety pin is reinstalled forany reason, the EPU check must be
reaccomplished prior to flight to insure
proper EPU operation.

NOTE• Prior to aligning the INS using a stored
heading alignment, accomplish a gyrocompass alignment with NAV RDY
flashing, m and do not select the NAV
mode, LESS @IDselect the NAV mode forapproximately 10 seconds, turn the
FCNP FUNCTION knob to OFF for atleast 10 seconds prior to power being
removed, and do not move the aircraft.

EPU GEN and EPU PMG lights - Confirm
off.
If either light is illuminated, cycle the EPU
switch to OFF, then back to NORM. If either
light stays on, abort the aircraft.

14.

Chocks and safety pins (ground crew)Remove.

*15.

Brakes and NWS - Check.

*16.

Ejection safety lever - Armed (down).

*17.

Flight control surfaces - Cycle.

18.

IFF - As required.

19. ~ AIFF - As required.

HOT REFUELING

7.

FLCS self-test - Accomplish to test No. 43.

HOT REFUELING PRECAUTIONS

8.

Radar - As desired.

9.

SMS - As desired.

10.

HUD - As required.

11.

FCNP FUNCTION knob - NAV.

Perform de arming prior to entry to the hot refuelingpit. If suspected hot brakes or other unsafe condition
exists, do not enter refueling area. Follow groundcrew directions into the refueling area and establish
communications with the ground crew. If amalfunction is suspected, stop refueling.

NOTE• When positioning the FCNP FUNCTION knob from STOR HDG to NAV, do
not hesitate in NORM or alignment
will be lost.
• If time permits, the FCNP FUNCTION
knob can be moved to NORM at
anytime to initiate a full gyrocompassalignment.

Hot refueling is prohibited with an activated EPU,
hung ordnance, hot brakes, or fuel leakage in vicinityof AR receptacle during AR. Safety pins for stores andgun must be installed. In the refueling area, use
minimum thrust for taxiing, ~ lru and HF radiotransmissions are prohibited except in emergency.
PRIOR TO HOT PIT ENTRY

1. AFTER LANDING checks - Complete.
2.

AIR REFUEL switch - OPEN; RDY light on.
Change 11

2-29

T.O. 1F-16A-1
NOTE

3.

EPU switch - OFF.

• [ID With the AR switch in OPEN, NWS
can be engaged or disengaged fromeither cockpit regardless of the position of the STICK CONTROL switch
and without using the paddle switch .

4.

EPU safety pin (ground crew) - Removed.

5.

EPU switch - NORM.

6.

Intercom (refueling supervisor) - Disconnect.

7.

Taxi clear of refueling area and configure
aircraft as required.

• With the AR switch in OPEN, the
ARINWS light remains off whenever
NWS is engaged.
*3.

TACAN function knob - OFF.

*4.

GND JETT ENABLE switch - OFF.

5. f@lAIRCRAFT DE-ICING WITH ENGINEOPERATING

rru HF OFF VOL control knob - OFF.

PRIOR TO HOT REFUELING

Aircraft de-icing with the engine operating shall beaccomplished IAW guidance in T.O. 42C-1-2, Chapter
6, and local procedures.

Perform the following actions prior to refueling:
1.

EPU safety pin (ground crew) - Installed.

*2.

Personal equipment leads (except oxygenand communication) - As desired.

3.

Canopy - As desired.

Insure all cockpit items are secure
prior to opening the canopy. With the
canopy open, the engine is susceptible
to FOD from loose cockpit items.
4.

Brake and tire inspection (ground crew)Complete.

5.

Intercom with refueling supervisor - Established.

PRIOR TO APPLICATION OF DE-ICING FLUID

1.

Aircraft - Parked with nose into wind.

2.

Throttle - IDLE.

3.

Non-essential avionics - OFF.

4.

AIR SOURCE knob - OFF.

DURING DE-ICING

1.

Monitor de-icing operation.
Fluid spray should not be directed forwardof the wing leading edge.

The canopy transparency must be
cleaned prior to flight if de-icing fluid
has contacted it.

DURING HOT REFUELING*1. Be alert for visual or voice signals fromrefueling supervisor.
*2.

Terminate refueling if intercom contact islost - Visual signal.

*3.

Ground control radio frequency - Monitor.

*4.

Insure hands are visible to ground crew.

AFTER DE-ICING IS COMPLETE

1.

AIR SOURCE knob - NORM.

2.

Avionics - As required.

3.

Taxi interval- Maintain 300 foot spacing to
avoid degradation offluid anti-ice capability
on trailing aircraft.

HOT REFUELING COMPLETE

1.

AIR REFUEL switch - CLOSE.

2.

EPU GEN and EPU PMG lights - Confirm
off.

2 -30

Change 11

PRIOR TO TAKEOFF

1.

Aircraft surfaces check - Complete (groundcrew).

••

T. O. 1F- 16A- 1

QUICK TURNAROUNDPRIOR TO ENGINE SHUTDOWN

1. AFTER LANDING checks - Complete.
2. PRIOR TO ENGINE SHUTDOWN checks Complete.
3. Communication with ground crew - Establish (if required).
4. ENGINE SHUTDOWN checks - Complete.
5. Aircraft setup - lAW local procedures.

SPECIAL PROCEDURES• If the engine has been run more than
30 minutes ago and less than 3-1/2
hours ago, do not advance the throttle
to IDLE until 40 seconds after the JFS
RUN light comes on. The time the JFS
RUN light was on during previous
start attempts is cumulative towardthe 40 seconds.
After light-off:Light-off is indicated by ignition vibration andincreasing rpm. Iflight-off does not occur within 15
seconds, very slowly advance the throttle until it
does occur and then stop all throttle movement.

EPU HYDRAZINE SUPPORT AT NON-F-16 BASES

4. Allow rpm to increase and begin to stabilize(approximately 10 seconds),

At non-F-16 bases, the pilot is responsible for theaircraft. Response actions shall be limited to
identification of a hydrazine or EPU problem,
isolation, containment, and minimal dilution withwater (1 part water to 1 part hydrazine). No major
neutralization, maintenance, or hydrazine servicing
capability is planned for transient bases. If ahydrazine leak or EPU incident occurs on a basewhere no disaster response force or bioenvironmental
support is available, the pilot must insure that theaircraft is isolated and the leak contained. Refer to
ACTIVATED EPUIHYDRAZINE LEAK, Section III.

If the throttle is advanced earlier, theengine may stall, precluding a successful start since rpm will not accelerate
past approximately 45 percent.

SUPPLEMENTAL PROCEDURESBUC GROUND START PROCEDURES I PW200i

The following steps may be substituted for steps 1,2,
and 3 under STARTING ENGINE, this section.
•• Do not mistake BUC IDLE for IDLE. If
the throttle is forward of the IDLE
position, a hot start may occur.

1. JFS switch - START 2.

2. EEC BUC switch - BUC.
• EEC caution light - Verify on.
3. Throttle-Advance to IDLE at approximately
25 percent rpm.

As rpm begins to stabilize:

5. Throttle - Advance slowly to produce a
steady rpm rise.
• Advance the throttle slowly and smoothly tothe backside of the BUC IDLE detent toproduce a steady rpm increase similar to anormal UFC start.
• Monitor FTIT during the start; FTIT should
not exceed 600°C. If the throttle is rapidly
advanced to obtain this FTIT, a hot start may
result. If FTIT reaches this value, stop
throttle advance, wait for FTIT to stop
increasing, and then continue the throttle
advance.
When at the backside of the BUe IDLE detent:
6. Throttle - Pause, rotate outboard, andsmoothly advance past the detent2-31

T. O. 1F- 16A- 1• Pause (2-3 seconds minimum) at thebackside of the BUC IDLE detent to allow
FTIT and rpm to stabilize. Then rotate thethrottle outboard and advance slowly into
BUC IDLE.

• If the engine fails to transfer to UFC

and the throttle is retarded to IDLE,
the engine will probably flame out. If
this occurs, shut down the engine.
10. EEC BUC switch - EEC.
• EEC caution light - Verify off.

·1btal time to advance the throttle from IDLE
to BUC IDLE will be a minimum of 30
seconds.

I• Nozzle - Opens steadily (70-95 percent).
• FTIT - Decreases.

• After stabilized in BUC IDLE, allow 2
minutes minimum at BVC IDLE.

• RPM - Verify EEC idle.
11. EPU switch - Cycle to OFF, then NORM.
NOTE

Momentary cycling of main generator
power during engine start may generate an EPU start command and cause
the EPV GEN and EPU PMG lights toilluminate. This is not an abnormal
occurrence during a BVC start. If thelights go off when the EPU switch is
cycled and no other abnormal indications are present, flight may becontinued .

• To prevent possible engine damage, do
not allow FTIT to exceed 680°C.

•• The nozzle may remain nearly fullopen after a BUC start due toinsufficient air load. Throttle should
not be advanced beyond 75 percent
rpm with nozzle above 30 percent.
7. Throttle - Advance to 75 percent rpm.
8.

EEC BUC switch - OFF.

EXPANDED FLCS SELF·TEST PROCEDURES• BUC caution light - Verify off.

After program stops at test No. 54:
1. SERVO ELEC RESET switch - ELEC.

To avoid a possible engine stall, allow
the engine to idle at least 2 minutes at
BUC IDLE before placing EEC BUC
switch to OFF.

•

9. Throttle - Retard to IDLE at first indicationof rpm increase.

2.

SERVOS ARM switches:a. HORIZ - R ARM.
b.

FLAPERON - R ARM.

c.

RUDDER - ARM.

3. ADV SLEW switch - ADV momentarily.
a. Test No. advances to 57.
• If the throttle is not retarded immedi-

ately after first indication of rpm
increase, the engine could quickly
accelerate to more than 80 percent
rpm.

2-32

Change 2b. FLT CONT SYS, P, R, Y, and all five
SERVOS lights - On.
c.

DUAL FC FAIL warning light - On.

d. Program stops.

T. O. 1F- 16A- 1

4. Operate controls - Right horizontal tail,right flaperon, and rudder may respondinitially and then lockout; the surfaces do notrespond to further control inputs. Lefthorizontal tail and left flaperon respond.
NOTE

Deflection of those control surfaces
which are not locked out may be lessthan normal.
5. SERVOS ARM switches:a. HORIZ - L ARM.
b. FLAPERON - L ARM.
c.

RUDDER - Disarm.

6. ADV SLEW switch - ADV momentarily.
a. Test No. advances to 58.
b. FLT CONT SYS, P, R, Y, and all five
SERVOS lights - On.

c.

Program stops.

12. SERVOS ARM switches:a. HORIZ - R ARM.
b. FLAPERON - R ARM.
c.

RUDDER - ARM.

13. ADV SLEW switch - ADV momentarily.
a. Test No. advances to 62.
b. FLT CONT SYS, P, R, Y, and all five
SERVOS lights - On.
c.

Program stops.

14. SERVOS ARM switches - All three disarm.
15. FCS CAUTION button and SERVO ELECRESET switch - Activate SERVO and FCS
CAUTION simultaneously and then ELEC
momen tarily.
16. ADV SLEW switch - ADV momentarily.

c.

DUAL FC FAIL warning light - On.
a. Test No. advances to 63.

d. Program stops.
b. Program stops.
7.

Operate controls - Left horizontal tail andleft flaperon may respond initially and thenlock out; the surfaces do not respond tofurther control inputs. Right horizontal tail,right flaperon, and rudder respond.

c.

MAL light - On.

d. ECA data word dot light - On.
e. FLT CONT SYS, P, R, and Y lights - On.

8. SERVOS ARM switches:a. HORIZ - Disarm.
h. FLAPERON - Disarm.
9. FCS CAUTION button and SERVO ELECRESET switch - Activate SERVO and FCS
CAUTION simultaneously and then ELEC
momentarily.
10. SERVOS ARM switches:a. HORIZ - L ARM.

17. FCS CAUTION button and SERVO ELECRESET switch - Activate SERVO and FCS
CAUTION simultaneously and then ELEC
momen tarily.
a. MAL light - On.
b. ECA data word dot light - On.
18. SELF TEST switch - OFF.

*19. Operate controls - All surfaces respondnormally; no FLCS lights on.

b. FLAPERON - L ARM.
ILS PROCEDURES

11. ADV SLEW switch - ADV momentarily.
a. Test No. advances to 61.

1. ILS PWR switch - PWR. (Allow 1 minute forwarmup.)b. FLT CONT SYS, P, R, Y, and all five
SERVOS lights - On.

2. Localizer frequency - Set. (Verify reliable
reception. )2-33

T.O. 1 F-16A-l

3.

INSTR MODE knob - ILS/TCN or ILSINAV.

4.

HSI - Set inbound localizer course.

5. Verify proper indication/symbology on ADI,
HSI, and HUD.

STRANGE FIELD PROCEDURES
Refer to Air Force/Command guidance.

FUEL MANAGEMENT LESSm

For flight director steering:
Refer to T.O. IF-16A-5-2 for in-flight CG limits.
6.

FCNP:
a. FUNCTION knob - NAV.

Refer to figure 2-4 for totalizer readings and fueldifferentials.

b. DATA knob - MISC.

Refer to figure 5-9 for AOA and rolling limitations.

c.

Fuel management procedures may be required toprevent operational CG's aft of the limits established
for departure resistance. Fuel management (otherthan monitoring) is not required when carrying
category II loadings with the STORES CONFIG
switch in CAT III.

DATA OPT button - Depress once.

d. Alpha display - Shows LOC.
e. Verify/enter inbound localizer course.
f.

MODE SEL button - Depress.

g. MODE SEL indicator light - On.
h. HUD - Verify flight director symbology
present.

I WARNING IDANGER AREAS

Selection of AFT ENG FEED prior toreaching the specified totalizer reading or failure to select NORM ENG
FEED when the required differential
is reached may result in filling theforward and right wing tanks, causing
an undesirable lateral asymmetry andreducing departure resistance.

Refer to figure 2-6.

NOTE

EXTERIOR INSPECTION

Refer to figure 2-5 for normal preflight inspection.

TURNING RADIUS AND GROUND CLEARANCE

Refer to figure 2-7.
G-SUIT HOSE ROUTING

Refer to figure 2-8 for the recommended g-suit hoserouting.

2-34• The following procedures may result inforward fuel distributions which will
uncover the red portion on the AL
pointer.
• Select AFT ENG FEED when thespecified fuel remaining is reached, if
necessary, to establish the specified
forward/aft fuel system differential.

T.O. 1F-16A-l

Fuel Management LESS EDNOTES:(1)

Loading categories I, II, and III are the AOA androlling limitations shown in Sedion V.

(2)

Totalizer reading denotes fuel quantity to begin
fuel balancing. Pointers denote fuel spread for~ minimum forward tank heavy or (I] maximum
aft tank heavy.

(3)

BEFORE BALANCING limits - The limits apply
anytime prior to achieving the specified pointerspread as long as fuel transfer starts at the
specified totalizer reading.

(4)

AFTER BALANCING limits - These apply
whenever the specified pointer spread exists.

(5)

IF BALANCING DELAYED limits - The.. apply
when fuel transfer is delayed below the specified
totalizer value until the specified pointer spreadis achieved or when the pointer spread can nolonger be maintained because of low fuelconsiderations.

Figure 2·4.

(6)

CONFIGURATION ADJUSTMENT - For changes
in baseline loading, add to or subtrad from thetotalizer and pointer spread adjustments shown on
the fuel quantity indicator. Individual adjustments
are additive.

(7)

Do not exceed 15 degrees AOA or perform
maximum command rolling maneuvers.

(8)

No adiustments required to baseline fuelmanagement if configuration has no ammunition or cases and 0 or 1 AIM-9. [I] Two pilots are
required.

(9)[I] No adiustments required to baseline fuelmanagement if configuration has full ammunition or cases, 0 AIM-9's, and 1 pilot.

(10) [I] A maximum aft heavy distribution of 1100
pounds is only fuel balancing adiustment
required if configuration has full ammunition orcases,l AIM-9, and 1 pilot.

(Sheet 1)

2-35

~
I

W
0.

fuel Management LESS

m

:-I

--

P
"'T1
I

0.

(")
::T

o

~

F-16A

F-J6A

~

cc
CD
.....,LOADING CATEGORY II

LOADING CATEGORY I
LIMITS (1)BASELINE LOADINGCONFIGURATIONFUEL BALANCINGREQUIRED (2)BEFORE AFTER IF BALBALANCE BALANCE DELAYED
(5)
(3)
(4)

BASELINE LOADINGCONFIGURATIONTOTALIZERFULL AMMO OR
CASES & 0, 1,
OR 2 AIM-9's

06000
POINTER SPREAD

CAT I

CAT I

CAT I

300 MIN
FWD HEAVYCONFIGURATIONADJUSTMENTS(6) (8)eND AMMO OR
CASES

ADJUSTTOTALIZER-100

ADJUST
POINTERSPREADCAT I
e3 OR 4AIM-9's

-200

CAT I

CAT III

+300

(3)

(4)

(5)

CAT IICAT IICAT IIITOTALIZERFULL AMMO OR
CASES & 0, 1,
OR 2 AIM-9's
PlUS EMPTYPYLONS AT STA
3&7

POINTER SPREAD

CONFIGURATIONADJUSTMENTS

AD.lJST
TOTAl-

I

05800
500 MIN
FWD HEAVY

IZER

ADJUST
POINTERSPREADDOWNLOADPYLONS AT STA
3 & 7 AND UPLOAD 370--GALLON TANKS AT
STA 4 & 6

-200

+300

CAT IICAT IICAT IIIeNO AMMO OR
CASES

-100

+100

CAT IIICAT IICAT IIIe3, 4, 5, OR 6
AlM-9's

-200

+300

(6) (8)+100

FUEL BALANCINGREQUIRED (2)LIMITS (1)BEFORE AFTER IF BALBALANCE BALANCE DELAYEDeNO ADJUSTMENTS REQUIRED FOR NN CENTERLINE STORE, FOR
EMPTY PYLONS AT STA 4 & 6, OR FOR ACMI PODS CARRIED IN
LIEU OF AIM-9's

e NO ADJUSTMENTS REQUIRED FOR ANY CENTERLINE STORE OR ACMI
POD CARRIED IN LIEU OF AN AlM-9.

Figure 2-4.

(Sheet 2)

fuel Management LESS m
f-J6A

f-168

LOADING CATEGORY III

BASELINE LOADINGCONFIGURATIONFUEL BALANCINGREQUIRED (2)FULL AMMO OR
CASES & 0, I,
OR 2 AIM-9's
AND MY CAT
III LOADINGTOTALIZERCONFIGURATIONADJUSTMENTS

LOADING CATEGORY I

LIMITS (1)BEFORE AFTER IF BALBALANCE BAlANCE DELAYED
(3)

(4)

(5)

CAT IIICAT III

300 MIN

FULL AMMO OR
CASES & 0, I,
OR 2 AlM-9's
WITH 2 PILOTS

04800

CAT III(7)(6)

ADJUST
POINTERSPREAD-NO AMMO OR
CASES

-100

+200

CONFIGURATIONADJUSTMENTS

CAT IIICAT IIICAT III
(7)

-100

(3)

(4)

(5)

POINTER SPREAD

CAT I

CAT I

CAT I

CAT IIICAT I

CAT III

I

1200 MAX
AFT HEAVY

FWD HEAVY
ADJUSTTOTALIZER- LAUNCHER AT
STA 2 & 8

FUEL BALANCINGREQUIRED (2)TOTALIZER

06000
POINTER SPREAD

BASELINE LOADINGCONFIGURATION

LIMITS (1)BEFORE AFTER IF BALBALANCE BALANCE DELAYED(6) (8) (9) (10)ADJUSTTOTALIZER

ADJUST
POINTERSPREAD-NO AMMO OR
CASES

-100

-100

-3 OR 4
AlM-9's

-200

-300

-ONE PILOT

-200

-300

+100

- NO ADJUSTMENTS REQUIRED FOR ANY CENTERLINE STORE OR ACMI
POD CARRIED IN LIEU OF AN AlM-9.

- NO ADJUSTMENTS REQUIRED FOR ANY CENTERLINE STORE OR ACMI
POD CARRIED IN LIEU OF AN AlM-9.

n
:::T

a

::l

co

CD
'-I

:-I

9

"T1

.....
I

~

0-

I

W

'-I

Figure 2-4.

(Sheet 3)~~fuel Management LESS m:-I

9....0-,....>,

."

n
::T
Q

f-'68

f-'68

:l

CQ

LOADING CATEGORY II

CD......

BASELINE LOADINGCONFIGURATIONFULL AMMO OR
CASES & 0, I,
OR 2 AIM-9's
PLUS EMP1YPYLONS AT STA
3 & 7, 2 PILOTSCONFIGURATIONADJUSTMENTS(6) (8) (9)FUEL BAlANCINGREQUIRED (2)LOADING CATEGORY III

LIMITS (1)BEFORE AFTER IF BALBALANCE BALANCE DELAYED
(3)

(4)

(5)

CAT IICAT IICAT IIBASELINE LOADINGCONFIGURATIONFUEL BALANCINGREQUIRED (2)FULL AMMO OR
CASES & 0, I,
OR 2 AIM-9's
WITH 2 PILOTS

04800

LIMITS (1)BEFORE AFTER IF BAl.
BALANCE BALANCE DELAYED
(3)

(4)

(5)

CAT IIICAT IIICAT IIITOTALIZER

04800
POINTER SPREADTOTALIZER

1000 MAX
AFT HEAVY

ADJUSTTOTALIZER-200-NO AMMO OR
CASES

-100

-100

-3, 4, 5, OR 6
AIM-9's

-200

-300

-ONE PILOT

-200

-300

-300

1300 MAX(7)

AFT HEAVY

ADJUST
POINTERSPREADDOWNLOADPYLONS AT STA
3 & 7 ANDUPLOAD370-<lAlLONTANKS AT STA
4&6

POINTER SPREAD

CONFIGURATIONADJUSTMENTS

CAT IIICAT IIICAT III

ADJUST
POINTER(6)

ADJUSTTOTALIZER-NO AMMO OR
CASES

-100

-200

SPREADCAT IIICAT IIICAT III(7)

CAT IIICAT IIICAT III-ONE PILOT

-200

-300

- LAUNCHER AT
STA 2 & 8

-100

-100

-NO ADJUSTMENTS REQUIRED FOR NN CENTERLlN: STORE OR ACMI
POD CARRIED IN LIEU OF AN AlM-9.

- NO ADJUSTMENTS REQUIRED FOR Am CENTERLINE STORE, FOR
EMPTY PYLONS AT STA 4 & 6, OR FOR ACMI PODS CARRIED IN
LIEU OF AlM-g's.

Figure 2-4.

(Sheet 4)I10.lF-16A-1

Exterior Inspection (Typical)
NOTE:

Check aircraft for loose doors andfasteners, cracks, dents, leaks, andother discrepancies.

J.
K.

LEF - CONDITION.
STORES AND PYLONS - SECURE (PREFLIGHTlAW T.O. 1 F-16A-34-1-1, ~ T.O.
1 F-16A-34-1-3 or
T.O. 1 F-16A-34-1-4) .

L.

NAV AND FORM LIGHTS - CONDITION.

M.

FLAPERON - CONDITION.

IMJI

•

AFT FUSELAGE - C
1. TAIL:
A. ADG -CHECK.
B.
C.
D.
E.
1 F-16A-l-Q084X®

NOSE-A
1. FORWARD FUSELAGE:
A. EXTERNAL CANOPY JETTISON D-HANDLES(2) - ACCESS DOORS CLOSED.
B. PITOT-STATIC PROBES (2) - COVERSREMOVED.
C. AOA PROBES (2) - COVERS REMOVED; SLOTSCLEAR; FREEDOM OF MOVEMENT CHECKED;ALIGNMENT CHECKED (ROTATE PROBES FULLYTOWARD FRONT OF AIRCRAFT (CCW ON THE
LEFT; CW ON THE RIGHT) AND VERIFY
BOTTOM SLOTS SLIGHTLY AFT OF 6
O'CLOCK AND TOP SLOTS FORWARD); SET IN
NEUTRAL POSITION (BOTTOM SLOT AT 4
O'CLOCK ON THE RIGHT SIDE AND 8
O'CLOCK ON THE LEFT SIDE).
D. STATIC PORTS (2) - CONDITION.
E. ~ AND~ 10 LIGHT - CONDITION.
F.
RADOME - SECURE.
G. ENGINE INLET DUCT - CLEAR.
H. EPU FIRED INDICATOR - CHECK.
I.
ECS RAM INLET DUCTS - CLEAR.

!1m Ifrn

F.
G.
H.

CSD OIL LEVEL - CHECK.
BRAKE/JFS ACCUMULATORS - CHARGED(REFER TO FIGURE 2-5.1).
HOOK - CONDITION AND PIN FREE TO MOVE.
~DRAG CHUTE ACCUMULATORCHARGED (REFER TO FIGURE 2-5.3).
VENTRAL FINS, SPEEDBRAKES, HORIZONTALTAILS, AND RUDDER - CONDITION.
~ M DRAG CHUTE HOUSING - CONDITION.
ENGIN1: EXHAUST AREA - CONDITION.

MI.

NAV AND FORM LIGHTS - CONDITION.

J.
K.

~ VERTICAL TAIL LIGHT - CONDITION.

L.

FLCS ACCUMULATORS - CHARGED (REFER TO
FIGURE 2-5.3).
JFS DOORS - CLOSED.

LEFT WING & CENTER FUSELAGE - 0
1. LEFT WING:
A. FLAPERON - CONDITION.
B. NAV AND FORM LIGHTS - CONDITION.
C. STORES AND PYLONS - SECURE (PREFLIGHTlAW T.O. 1F-16A-34-1-1, ~ T.O.
1F-16A-34-1-3 or ~ T.O. 1F-16A-34-1-4l.

CENTER FUSELAGE & RIGHT WING - B
1. RIGHT MLG:
A. TIRE, WHEEL, AND STRUT - CONDITION.
B. UPLOCK ROLLER - CHECK.
C. DOOR AND LINKAGE - SECURE.
D. LG SAFETY PIN -INSTALLED.
E. TAXI LIGHT - CONDITIDN.
2. RIGHT WING:
A. HYDRAZINE LEAK DETECTOR - CHECK.
B. EPU NITROGEN BOTTLE - CHARGED (REFER TO
FIGURE 2-5.2).
C. EPU OIL LEVEL - CHECK.
D. HYD SYS A QTY AND ACCUMULATOR - CHECK.
E.
GUN-RNDS COUNTER AND RNDS LIMIT - SET.
F.
SECURE VOICE PROCESSOR - CHECK.
G. EPU EXHAUST PORT - CONDITION.
H. I PW200 I DOOR 3308, ENGINE SYSTEM FAULTFLAG - CHECK.
I.
I PW2201 BLOCK 10 DOOR 2338, BLOCK 15
LESS M DOOR 2306 ENGINE AND EMS
GO-NO GO INDICATORS - CHECK.

D.
E.
F.

2.

3.

4.

LEF - CONDITION.
FUEL VENT OUTLET - CLEAR.
HYD SYS B QTY AND ACCUMULATOR-CHECK.
LEFT MLG:
A. TIRE, WHEEL, AND STRUT - CONDITION.
B. UPLOCK ROLLER - CHECK.
C. DOOR AND LINKAGE - SECURE.
D. LG SAFETY PIN -INSTALLED.
E.
LG PIN CONTAINER - CHECK CONDITION.
F.
LANDING UGHT - CONDITION.
fUSELAGE:
A. GUN PORT - CONDITION.
B. IFF - CHECK.
C. AVTR - CHECK.
D. ~ DOOR 2317 ENGINE SYSTEM FAULTFLAG - CHECK.
UNDERSIDE:A. NLG TIRE, WHEEL, AND STRUT - CONDITION.
B.

NLG TORQUE ARMS - CONNECTED, PIN
SECURE AND SAFETIED.

C.

NLG DOOR AND LINKAGE - SECURE.

D.

LG/HOOK EMERGENCY PNEUMATIC BOTTLE
PRESSURE - WITHIN PLACARD LIMITS (REFER
TO FIGURE 2-5.2).

Figure 2-5.

Change 13

2-38.1/(2-38.2 blank)•

T.o. 1 F-16A-l

BrakelJFS
Accumulators
Pneumatic Servicing
TEMPERATURE

OFI-44 to -36-35 to -27-26 to -18-17 to-9-8 to-1o to 89 to 17
18 to 26
27 to 35
36 to 44
45 to 53
54 to 62
63 to 71
72 to 80
81 to 89
90 to 98
99 to 107
108 to 116
117 to 125
126 to 135

PRESSURE
PSIG
1475-1625
1525-1675
1575-1725
1625-1775
1675-1825
1725-1875
1775-1925
1825-1975
1875-2025
1925-2075
1975-2125
2025-2175
2075-2225
2125-2275
2175-2325
2225-2375
2275-2425
2325-2475
2375-2525
2425-2575

Figure 2-5.1.

EPU Nitrogen &Alternate LG/Hook
Bottles Pneumatic
Servicing
TEMPERATURE

OFPRESSURE
PSIG

100 and higher
50 to 100
10 to 50-60 to +103250-3500
2850-3250
2500-2850
2000-2500

Figure 2-5.2.

Drag ChutelFLCS
Accumulators
Pneumatic Servicing
TEMPERATURE
ofPRESSURE
PSIG

100 and higher
50 to 100
10 to 50-60 to +10

1300-1400
1200-1300
1100-1200
950-1100

Figure 2-5.3.

Change 13

2-39

T.O. 1F-16A-l

Danger Areas
ENGINEENGINE Fl00-PW-20012 20~~~~~~~~M~A~X~A~B~T]H~R~UES~Tt:~~~aiZiliil 30l;;~~~_il 20 ~ENGINE INLETDANGER AREA

PLANE OFROTATION JFS

o

20

40

60

c!:l..L,-J 20
14 0

160

DISTANCE FROM EXHAUST NOZZLE -

FEET

80

100

120

180

200

PLANE OF ROTATIONENGINE TURBINE

4

"c2
~

o w

"~

4

l;;
~

TIRES

o

10

WITH HOT BRAKES, AVOID INFLATED

20

30

40

50

60

70
DI STANCE FROM EXHAUST NOZZLE -

80
FEET

90

8
100

MLG TiRe Sloe AREA WITHIN 300 fEET
FOR 45 MINIJTES AFTER AIRCRAFT HAS

STOPPEO. IF REOUIA[O, APPnOACH
FROM THE FRONT OR REAlI
FJREFIGHTING PURPOSES---j----j2+-f---j 1 ~•~~-r-+-+-4----j0

is
~

a:
w

-l--l---J- ; --1--1--11 "~l;;
2
I NGlN[ INLl T

DANGCR ARr A

o

10

30

60

·/0

80

D I STAN CF FROM EXHAUST NOZZLE -

FEET

20

40

50

'0

~

100

H IliA 1 1139C ..

Figure 2·6. (Skeel 1)

2-40

Chang e 13

T.O. 1F-16A-l

Danger Areas

ENGINE Fl00-PW-200/220

RADIATION AND TEMPERATURE

UPPER TACANANTENNA

UPPER UHF/I FF

ANTENNANOTES:• Dista nce from radar disc to
forward tip of redoma _ 5 feet

• ECM pods radiate in a

fan panern fora and aft.

!AQ] [A] HF ANTENNAECM PODVHf ANTENNALOWER lACAN ANTENNAIADI UPPER IFF
ANTENNAS 14 )CENe AIR-MOTOR

EXHAUST 900 OF
BLOCK 15 LOWER UHF/IFF ANTENNA

F~'~::;~~RRADARAANTENNAJFS EXHAUST 1'75 0 F
EPU EXHAUST 1600 of

ecs EXHAUST 325 - F

BLOCK 10 IFF ANTENNA. fAQ] 14) IFF ANTENNABLOCK 10 LOWER UHf ANTENNAMINIMUM SAFE DISTANCEFROM ANTENNAS IN FEETOPERATING
TRANSMITTERS

VOLATILE FLUIDSPERSONNEL

EEDUPPER AND LOWER UHF/IFF

-

1

-

UPPER AND LOWER lACAN

-

1

VHF

1

RADAR ALTIMETER

1

IADIIAJ HF

•1

•1

·260

fiRE CONTROL RADAR AND ~ CWl

30

120

120

6

AN/ AlO- 131

-

,.

6

AN/ AlQ- 176

-

6

"

31

6

6

6

6

6

AN/ AlO- l'9

AN/ ALQ- 184

AN/ ALD- leB-

ORC- SO-O'

I

6• Minimum safe distance is d istance from aircraft surface.

H- HIA-l-00869:t

Figure 2-6. (Sheet 2)

Change 8

2-41

T. O. 1F- 16A- 1

Danger Areas
HAZARDOUS NOISE LEVEL AREAS

ENGINE F100-PW-200/220

MIL THRUSTMAX AB THRUST
FEET
NOTE:
Noise levels are
symmetrical andmay be alteredby surrounding
obstacles.

LIMIT 15MINUTESPER DAY
WITH EAR
PROTECTION

LIMIT 5MINUTESPER DAY
WITH EAR
PROTECTION

IDLE THRUSTNOISE EXPOSURE LIMITS
130
125

FEET
120
115

< 110

III
"0

I 105
....J

0..
VJ

100
95
90

1\\

""

~ ~ I"-

1\

\

'"

85
80

LIMITS WITH

~L/MITSW~-~.!It!..E~M.Yi~

-

~~SW

...f!!i.OUT fAR---"':::~CT/ON-

012345678
DURATION - HOURS EXPOSURE LIMIT PER DAYNOTES:• Attenuation factors and exposure limits are for typicalearmuffs and earplugs.
• SPL - Sound pressure level.
• dBA - Adjusted (human ear response) decibels.
1F-16A-1-1141X®

Figure 2-6. (Sheet 3)2-42

T.O. 1F-16A- 1

Turning Radius and Ground Clearance

NOTES:• Turn radius can be
reduced to 38 feet byusing inside brake.
• Maximum NWS deflection.
• All distances areapproximate.

* • Avoid arresting cable
support donuts whentaxiing with centerline
stores installed.
FEET

PIVOT POINT~~B7' L MISSILER MISSILEWARNING

I

With large asymmetries, reduce taxi speed,
avoid sharp turns, and, when possible, turntoward heavy wing to prevent possible tipping.

TI
3 FEET 6 INCHES

5 FEET

5 FEET

2 FEET

4 FEET 6 INCHES

16 FEET 8.5 INCHES-'1 F-16A-1-0088A ®

Figure 2-7,

2-43

T.O. 1 F-16A-l

G-Suit Hose Routing (Typical)

LAPBELTTORSO

1i1Jj(!J.e-~~~Lfiff!-;:"'--I- HARNESSSURVIVAL KITSTRAPG-SUIT HOSE

1F-16X-1-001SX@Figure2-B.
2-44

~"""""""""""

T. o. 1F- 16A- 1

SECTION III

EMERGENCY PROCEDURESTABLE OF CONTENTS

Page

INTRODUCTION ....................... .

3-3

WARNING AND CAUTION LIGHT ~ f1)
AND ENGINE PILOT FAULT LIST
ANALySIS ........................... .

3-4

FORMAT ............................... .

3-4

GROUND EMERGENCIES .............. .
Fire/OverheatlFuel Leak (Ground) ...... .
Hot Start (Ground) .................... .
Hung StartIN 0 Start .................. .
Engine Autoacceleration (Ground) ...... .
Antiskid Malfunction (Ground) ......... .
Brake Failure ......................... .
Hot Brakes ........... '" ............. .
Main Generator Failure (Ground) ...... .
Emergency Entrance and Crew
Rescue. '" ......................... .
Emergency Ground Jettison ............ .
Ground Egress ........................ .
Hot Refueling Emergency .............. .
Activated EPUIHydrazine Leak ........ .
NWS Failure/Hardover ................ .

3-28
3-28
3-28
3-28
3-28
3-28
3-29
3-29
3-30

TAKEOFF EMERGENCIES .............. .
Delayed Rotation ...................... .
Abort ................................ .
Engine Malfunction on Takeoff ......... .
Engine Failure on Takeoff ............. .
AB Malfunction on Takeoff ............. .
Low Thrust on Takeoff or at Low
Altitude (Non-AB) I PW200' ........... .
Low Thrust on Takeoff or at Low
Altitude (Non-AB) I PW2201 ........... .
Engine Fire on Takeoff ................ .
LG Fails To Retract ................... .
LG Handle Will Not Raise ............. .
Blown Tire on Takeoff ................. .

3-34
3-34
3-35
3-36
3-36
3-36

IN-FLIGHT EMERGENCIES ............ .
CANOPY Warning Light On ........... .
Canopy LosslPenetration in Flight ...... .
Drag Chute Deployed in Flight IffE] Ifrn .. .
Cockpit Pressurerremperature
Malfunction ........................ .

3-31
3-31
3-31
3-34
3-34
3-34

3-36

3-37
3-37
3-38
3-38
3-38
3-39
3-39
3-39
3-39
3-40

EQUIP HOT Caution Light ............. 3-40
Ejection ........................ " ..... 3-40.1
Ejection (Immediate) . . . . . . . . . . . . . . . . . 3-43
Ejection (Time Permitting) ........... 3-43
Failure of Canopy To Separate ........ 3-43
Ejection Seat Failure. . . . . . . . . . . . . . . . . 3-43
Ditching.. ......... ... .... ... .. . ....... 3-43
Electrical System Failures. . . . . . . . . . . . . . 3-44
FLCS PMG Failure .................. 3-44
Main Generator Failure (in Flight) . . . . 3-44
Main and EPU Generator Failure ..... 3-45
EPU Malfunctions ................... 3-48
Uncommanded EPU Operation..... 3-48
Abnormal EPU Operation . . . . . . . . . . 3-48
FLCS Battery Discharge ............. 3-49
Partial Electrical Power Loss ......... 3-50
Aircraft Battery Failure . . . . . . . . . . . . . . 3-50
Emergency Power Distribution ........ 3-50.1
Engine Malfunctions IPW200 I . . . . . . . . . . . . 3-58
Engine Fire IPW200 I . . . . . . . . . . . . . . . . . . 3-58
OVERHEAT Caution Light IPW200 I ... 3-59
Engine Vibrations IPW200 I ............ 3-59
Oil System Malfunction I PW200 I . . . . . . . 3-60
EEC Caution Light IPW200 I . . . . . . . . . . . 3-60
FTIT Indicator Failure I PW200 I ....... 3-61
Zero RPMlErroneous RPM
Indication IPW200 I . . . . . . . . . . . . . . . . . 3-61
Abnormal Engine Response IPW200 I ... 3-62
AB BlowoutlFailure To Light
IPW200 I . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-65
Engine Stalls I PW200 I . . . . . . . . . . . . . . . . . . . 3-65
AB-Associated Engine Stalls
I PW200 I . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-65
Non -AB Engine Stalls IPW200 I ........ 3-65
Engine Stagnation [PW200 I ........... 3-66
Engine Stall Recovery I PW200 I ........ 3-66
Inlet Buzz I PW200 I ..................... 3-67
Engine Failure or Flameout I PW200 I ..... 3-67
Tower Shaft Failure IPW200 I .......... 3-67
Low Altitude Engine Failure or
Flameout IPW200 I ................. 33~6770'
Airstarts IPW200' ...................... .
High Altitude Airstart Considerations

~

~

~'III

~~~

~'III

~

~

...~

~

~

~

~

~

~~~

~
~
...~

~

~

~~....
~
~

..~
~...
~
~

~
~'III

LO~~:~d~ ·Ai~s~~rt ·C~~~id~~~ti~~~ .. 3-78 ~IPW200 I ...........................

3-78 ~~

Change 133~~

~111111'"1111111111111
~

~

~
~

~
~
~
~

~
~

~

~~
~
~

T. o. 1F- 16A-1
Airstart - UFC I PW200 I . . . . . . . . . . . . . . .
Airstart - Hot (UFC) I PW200 I .........
Airstart - BUC IPW200 I ..............
Engine Malfunctions I PW2201 . . . . . . . . . . . .
Engine Fire I PW2201 . . . . . . . . . . . . . . . . . .
OVERHEAT Caution Light I PW2201 .. ,Engine Vibrations I PW2201 ............
Oil System Malfunction I PW2201 . . . . . . .
ENGINE FAULT Caution Light ~ W

3-78
3-80
3-80
3-84
3-84
3-85
3-86
3-86

E~~~~~AULT C~~ti~~ Li~h; ....... 3-87

LESS ~, LESS W I PW2201 .......... 3-87
SEC Caution Light I PW2201 ........... 3-88
FTIT Indicator Failure I PW2201 ....... 3-88
Zero RPMlErroneous RPM
Indication I PW2201 .. . . . . . . . . . . . . . . . 3-88
Abnormal Engine Response I PW220 I •. , 3-89
Nozzle Failure I PW2201 ............... 3-90.1
AB BlowoutlFailure To LightI PW2201 .......................... , 3-90.2
Engine Stalls I PW2201 .................. , 3-91
AB-Associated Engine Stalls
I PW2201 . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-91
Non-AB Engine Stalls I PW2201 ........ 3-91
Engine Stagnation I PW2201 ........... 3-91
Engine Stall Recovery I PW220 I ........ 3-92
Inlet Buzz I PW220 I ..................... 3-92
Engine Failure or Flameout I PW2201 ..... 3-92
Tower Shaft Failure I PW2201 .......... 3-93
Low Altitude Engine Failure or
FlameoutlPW2201 ................. 3-93
Airstarts I PW2201 . . . . . . . . . . . . . . . . . . . . . . . 3-96
High Altitude Airstart Considerations~~ ~§~i::~~:::~p::::;deratiO~S:: ;~;;;~~~

Flameout Landing .....................
Overhead Approach ............... , ..
Straight-In Approach ................
IMC Penetration .....................
Landing Phase ......................
After Touchdown ....................
Flameout Landing Procedures ........
Jettison ...............................

3-105
3-106
3-107
3-107
3-112
3-112
3-112
3-114

~~~ FL;;~~:~;[i:~~:~~~~.:::::: .::.:::::::: ~:tt~

AD ClAir Data Malfunctions .......... 3-115
CADC Malfunction IPW200 I or
I PW2201 LESS 1Mi), LESS f11l .......... 3-116

~

~

3-2

Change 12

CADC Malfunction: PW2201 ~ W ..... 3-116
P, R, and/or Y Malfunctions .......... . 3-117
Servo Malfunction .................. . 3-118
LEF Malfunction (Symmetric) ....... . 3-119
LEF Malfunction (Asymmetric) ...... . 3-120
TEF Malfunctions .................. . 3-121
Trim Malfunction .................... 3-121
Stick Interference. . . . . . . . . . . . . . . . . . .. 3-121
Fuel Malfunctions . . . . . . . . . . . . . . . . . . . . .. 3-122
Fuel Management System PFL ....... 3-122
Fuel Leak . . . . . . . . . . . . . . . . . . . . . . . . . .. 3-122
Fuel Low ............................ 3-123
Hot Fuel or Gravity Feed . . . . . . . . . . . .. 3-124
Fuel Imbalance ...................... 3-124
Trapped External Fuel ............... 3-125
Hydraulic Malfunctions . . . . . . . . . . . . . . . .. 3-127
Single Hydraulic Failure ............. 3-127
System A Failure. . . . . . . . . . . . . . . . .. 3-127
System B Failure. . . . . . . . . . . . . . . . .. 3-127
Dual Hydraulic Failure .............. 3-128
System B and Main Generator
Failure (PTO Shaft) ............... 3-129
INS Failures. . . . . . . . . . . . . . . . . . . . . . . . . .. 3-130
INS Computer Failure LESS @) ....... 3-131
Total INS Failure. . . . . . . . . . . . . . . . . . .. 3-131
Out-of-Control Recovery ................ 3-131
Oxygen Malfunction . . . . . . . . . . . . . . . . . . .. 3-133
PBG Malfunction m . . . . . . . . . . . . . . . . . . .. 3-133
Smoke or Fumes ....................... 3-133
LANDING EMERGENCIES. . . . . . . . . . . . . .. 3-134
Straight-In Landing. . . . . . . . . . . . . . . . . . .. 3-134
Simulated Flameout (SFO) Landing ..... 3-134
Controllability Check. . . . . . . . . . . . . . . . . .. 3-135
Cable Arrestment ...................... 3-136
Net Arrestment ....................... . 3-136
Landing With a Blown Tire ............ . 3-137
Landing With a Blown Main Gear
Tire ............................. . 3-137
Landing With a Blown Nose Gear
Tire ............................. . 3-138
LG Extension Malfunctions ............ . 3-138
LG Handle Will Not Lower .......... . 3-139
LG Fails To Extend .................. 3-140
Alternate LG Extension .......... .. 3-140.1
Landing With LG UnsafelUp ....... 3-140.2
Antiskid Malfunction (Landing) ....... 3-140.2
Asymmetric Stores (Landing) ........... 3-141
NLG WOW Switch Failure . . . . . . . . . . . . .. 3-145
Drag Chute Failure ruJ M . . . . . . . . . . . . .. 3-146

, I I I I I I'~Irlll~~~~INTRODUCTION

Th\, j()llowing information provides general landing
guidance:
Land As Soon As Possible

This section covers the operation of the aircraft
during emergency/abnormal conditions. It includes a
discussion of problem indications and corrective
actions as well as procedural steps when applicable.
Adherence to these guidelines insures maximum
safety for the pilot and/or aircraft. The situations
covered are representative of the most probablemalfunctions. However, multiple emergencies,
adverse weather, or other factors may require
modification of the recommended procedures. Only
those steps required to correct or manage the problem
should be accomplished. When dealing with emergency/abnormal conditions, it is essential to determine
the most correct course of action by using soundjudgment and a full understanding of the applicablesystem(s). When practical, other concerned agencies(i.e., flight lead, tower, etc.) should be advised of the
problem and intended course of action.
When a voice WARNING or CAUTION message isheard, check cockpit indications; then refer to theappropriate emergency procedure for corrective
action.
When structural damage or any other failure that
may adversely affect aircraft handling characteristics is known or suspected, a controllability checkshould be performed.
Certain steps (e.g., MASTER CAUTION reset, ELECCAUTION reset, FLCS CAUTION reset) areintentionally omitted from the numbered procedures.
Pilots are expected to perform these actions without
prompting, when warranted.
Three basic rules, which apply to all emergencies, areestablished:
1.

MAINTAIN AIRCRAFT CONTROL.

2.

ANALYZE THE SITUATION AND TAKE
PROPER ACTION.

3.

LAND AS THE SITUATION DICTATES.

An emergency will be declared. A landing should beaccomplished at the nearest suitable airfield
considering the severity of the emergency, weather
conditions, airfield facilities, lighting, aircraft GW,
and command guidance.

Land As Soon As Practical
Emergency conditions are less urgent and, although
the flight is to be terminated, the degree of theemergency is such that an immediate landing at thenearest suitable airfield may not be necessary.

IWARNING

1• The canopy should remain closed
during all emergencies that could
result in a crash or fire such as crashlandings, aborted takeoffs, and arrestments. The protection the canopy
affords far outweighs the isolated riskof entrapment due to a canopy
malfunction or overturn.
• Ejection is preferable to sliding into anarrestment cable with the NLG collapsed. The cable may slide up over thenose with unpredictable and potentially dangerous consequences to anyone in the cockpit(s).
• If it appears that the aircraft willdepart a prepared surface above
normal taxi speed during an aborted
takeoff or a landing and go-around is
not possible, eject since breakup of
cockpit structure may occur. Retracting the LG to prevent departure from aprepared surface is not recommended
since the MLG will probably not
retract symmetrically.
• If remaining with an aircraft that willdepart a prepared surface, shut down
the engine, if feasible. This action
reduces the potential for fire, reduces
engine damage, and permits EPU
turnoff if an MLG WOW signal is lost.

~

".

~

T. O. 1F- 16A- 1

~~

~

~

"

I WARNING AND CAUTION LIGHT [AN]fl3 AND ENGINE PILOT FAULT LISTANALYSIS••

3~ 3~2,

3~3

Refer to figures 1,
WlJ I1'l andfor analysis
and amplification of warning and caution light andlBNl fI3 engine pilot fault list procedures. Fault trees
show interrelationships with examples of problem
events and corrective action.

FORMAT
The format of Emergency Procedures differs slightly
between the Checklist and the Flight Manual.
Procedures in the Checklist have been grouped bym;dt'unctior I categ' 11'\ (en~! 11. ele' "j t,~al, ", l: I to
prm'ide m:tximum in-flight utility, !11 the Flight
Manual, proL'pdure:-; :Ire listl'd by th(, phase of' flightin which the emergpncy may occur. 111 t he Checklist,
some procedures are split into two independentside-by-sidc· series of steps and are ~('parated by astraight line: in the Flight Manual, thp~p side-by-side
steps appear in a continuous column and can be
identified by repeat numbering of steps following
conditional statements beginning with the word if.
Amplification following procedural steps in the Flight
Manual is repeated in the Checklist under theheadings Inoperative Equipment, Other Indications,or Other Considerations.
A thorough review of the layout and content of the
Checklist and Flight Manual is recommended prior to
in-flight use.

~I I I I I~~~~~~~T1!,~_
Warning Light Analysis
WARNING LIGHT
ENG FIRE---

ORf}~'~!~(CORRECTIVE ACTION/REMARKS

CAUSE

I

ENG FIRE

I

Refer to ENGINE FIRE

Engine compartment fire

PROBABLE
MALFUNCTION~------j IN FIRE DETECTSYSTEM

ENG FIRE- - OR

LAND AS SOON AS POSSIBLENOTES:
Refer to:
• GROUND EMERGENCIES.
• GROUND EGRESS.
• ENGINE FIRE.
• EJECTION.

THROTTLE TOMINIMUM PRACTICAL THRUSTENGINE FIREPERSISTSWARNING LIGHT

TO/LOGORCONFIGlG Handle - UP
(No Malfunction)

I

TO/LOGOR

AIR

CONFIGI

CORRECTIVE ACTION/REMARKSCAUSE

lO.jLANDCONFIGI

EJECT

Aircraft configuration wrong for
takeoff or landing

A. For in-flight LG problems, refer to

Any of the following:
1. lG not down and locked
2. TEF's not fully down

B. If TEF's do not extend, execute land-

BELOW 10,000 FEET
MSL, <190 KNOTS
AND AT LEAST 250
FPM DESCENT

LG/TEF'S NOTDOWN **LANDING EMERGENCIES

ing using normal AOA (approximately
20 knots faster than normal)

INCREASE AIRSPEED, DECREASESINK RATE, OR PLACE LG
HANDLE- ON** NOTE: LG warning horn sounds.
* GND

TEF'S NOT FULLY'-----T"---------~ DOWN

LG Handle - ON
(Malfunction)

ABORT AIRCRAFT* NOTE: Confirm self-test OFF.
ONE OR
MORE
GREEN
LIGHTSNOT ONAIR

ANY/ALL LG NOTDOWN AND
LOCKED ***.UI

BELOW 10,000 FEET
MSL, <190 KNOTS
AND AT LEAST 250
FPM DESCENTSEE ANOTES:• LG warning
horn sounds.
• Red light in
LG handle on.

3 GREEN
LIGHTS

TEF'S NOT FULLY
DOWN

SEE B

1F-16A-1-1144X<tFigure 3-1. (Sheet 1)_

"T.~1~ M' M' M' M' M' M' M' M' M' M' .6T ~~ Warning Light Analysis
WARNING LIGHT

I

I

~

DUAL
Fe FAILI

CORRECTIVE ACTION/REMARKS

CAUSE

I

Dual AOA failure in the air data system(ADC caution light illuminates after first
failure)A. Refer to ADC/AIR DATAMALFUNCTIONS

Dual electronic failure in one axis of
the FLCS (P, R, or Y light illuminatedafter the first failurel

B. Refer to P, R, AND/OR Y

Illuminates to indicate a second servo
failure in a single branch after an ISA
is armed

C. Refer to SERVO MALFUNCTION

MALFUNCTIONSDUAL AOA SOURCEFAILURE (CADC AND..------~ ADC MFL MAY ALSO 1 - - - -..... SEE A
COME ON) (LEF'S MAY
STILL PROGRAMNORMALLY)

YES""------I

POWER FAILURE IN AT
LEAST TWO FLCS
I----~ SEE B
BRANCHESALWAYS GIVESFLT CO NT SYSCAUTION LIGHT~

NO

P, R, Y L1G HTSIN ANYCOMBINATIONELECTRONICFAILURE IN AT
LEAST TWOBRANCHES OF AT
LEAST ONE FLCSAXIS (MAY ALSOHAVE SERVO
LIGHT)SEE B

SERVOS LIGHTAND
PREVIOUSLY
ARMED

DUAL SERVO
MALFUNCTIONSEE CCHECK FLCP~~~

1F-16A-1-1145A@

Figure 3-1. (Sheet 2)

3-6

Change 2

Warning Light Analysis
CAUSE

CORRECTIVE ACTION/REMARKS

Low pressure in one or both hydraulicsystems, low engine oil pressure, or oilpressure switch failure

Check pressure indicators to determine system and condition

WARNING LIGHT

I PRESSHYD/OILI, Refer to:I• OIL SYSTEM MALFUNCTION• HYDRAULIC MALFUNCTIONSI PRESS I
HYD/Oll

CHECK OIL AND
HYD PRESS
INDICATORSOILOIL PRESSURENORMAL

L..---4

OIL PRESSURE
BELOW 15 PSIMONITOR_OIL PRESSURE
SWITCH FAILURE

f---

HYD

r--------------I

ENGINE OIL
SYSTEM MALFUNCTION, MINIMIZE THROTTLE
MOVEMENT.
LAND AS SOON
AS POSSIBLE- - + OIL PRESSURE
INDICATOR~REFER TO OIL
SYSTEM
MALFUNCTIONSYSTEM A
REFER TOSYSTEM A
HYD PRESS~ SYSTEM A
INDICATOR ONLY f - - - HYDRAULICFAILURE
L...-.
_ _ _ _ _- - - 'BELOW 1000 PSI
FAILURESYSTEM B
HYD PRESS
INDICATOR ONLY
BELOW 1000 PSINO

Is EPUrun light
on?

SYSTEM B
HYDRAULICFAILURE~REFER TOSYSTEM B
FAILUREYESIs ELEC
SYS caution
light on?

YES

NO

DUAL HYDRAULICREFER TO DUAL
FAILURE--+ HYDRAULICFAILUREPTO SHAFT' - - - - - - - - i FAILURE

BOTH HYD PRESS
INDICATORS BELOW 1000 PSI
WITH EPU SWITCHIN NORM OR ON-

TOTAL HYDRAULIC FAILURE(EPU MAY STILL RUN AND
POWER THE EPU GENERATOR,BUT EPU RUN LIGHT DOESNOT ILLUMINATE UNLESSSYSTEM A IS RESTORED)REFER TO SYSTEM B--+ AND MAIN GENERATORFAILURE (PTO SHAFT)

4

EJECT

I
1F-16A-1-1146X®

Figure 3-1. (Sheet 3)

WARNING LIGHTI

CANOPYCAUSE

CORRECTIVE ACTION/REMARKS

Canopy hooks or lock not secure

A . • Push handle down farther into locked(outboard) position or if on the
ground, inform maintenance• Refer to CANOPY LOSS/PENETRATION IN FLIGHT

I

!

j

POSITION HANDLE
DOWN
t----..,~ SEE ACANOPYNOTSECURECANOPY HOOKSNOT FULLYENGAGEDWARNING LIGHT

,

CAUSE

Red light in endof handle

LG or LG door not in positioncommanded by LG handlet----..,~SEE A

CORRECTIVE ACTION/REMARKS
A. Refer to LG FAILS TO RETRACT
B. Refer to LG FAILS TO EXTEND

ANY/All lG OR
lG DOORS NOTFULLY UP AND
LOCKEDt---------------------.... SEE ACHECKWHEELSDOWN
LIGHTS

NONlllUMINATED LIGHT
INDICATES ASSOCIATED
lG IS NOT DOWN AND
LOCKED (MAY ALSO HAVE 1----. SEE B
lG WARNING HORN
AND T.O./lAND CONFIGWARNING LIGHT)

IN FLIGHT

1F-16A-1-1147X®

Figure 3-1. (Sheet 4)

""""""~~~~
Warning Light Analysis
WARNING LIGHT

CAUSEFlashing WARN symbol in HUD~i~1

Check for specific illuminated warning light-:::::::;~
1=

One or more red glareshield warning
lights illuminated

CORRECTIVE ACTION/REMARKS

...

WARN

Reset by toggling WARN RESET switch on HUD
control panel

5 [&

--

l I

--

iii

\~

J

LESS~ Flashing TRAP FUEL
warning symbol in HUD~~

With the FUEL QTY SEL knob in NORM,
TRAP FUEL appears when the fuselage
fuel is 500 pounds below the fuselage capacity and the total fuel is 500 pounds
greater than fuselage fuel

TRAP FUEL

Reset manually by placing DRIFT C/O switch toTEST~ -0-[1}l I

I J

IAlT lOW I

I RDR IOR

ALT

Refer to TRAPPED EXTERNAL FUEL

Radar altitude is lower than AGL AlOW
value (with LG up)

Climb above AGl ALOW value or establish aminimum 1200 fpm climb

Ground clobber warning is activatedPullup immediately using maximum allowable g(no less than 4) to avoid ground impact

LOW

Figure 3-1. (Sheet 5)'"~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ha;2~3~

T. O. 1F- 16A- 1

Warning Light Analysis
I

WARNING LIGHT

I

!

CAUSE

CORRECTIVE ACTION/REMARI<SI

A. Place throttle to OFF and initiate airstart.
Refer to AIRSTARTS! Engine stagnation or flameout

I

I

ENGINEI

Check rpm and FTIT indications

i

~

I

--------------Engine alternator failure/malfunction

I

I

----~

I

I

B. Refer to ZERO RPM/ERRONEOUS RPM
INDICATIONI
Engine overtemperature!
!

C. Land as soon as possible

Engine warning system failure, RPM/ 'D. Land as soon as practical
FTIT indicator failure, or engine
history recorder failure

CHECK RPM AND
FTIT INDICATIONSYES~ENGINESTAGNATIONSEE A

ENGINE
FLAMEOUTSEE A

ENGINE
ALTERNATOR
FAILURE/
MALFUNCTION

SEE B

NO

YES

L - -_ _ _ _ _ _-I

YESENGINE OVERTEMPERATURE) - - - - - - - - - . . SEE CNOENGINE WARNING SYSTEMFAILURE, RPM/FTIT INDICATORFAILURE, OR ENGINE HISTORY
RECORDER FAILURE

1--------.. SEE D
1F-16A-1-1148A@

Figure 3-1. (Sheet 6)~, ' , Ie, ~ ~ ~ ~ ~ ~ ~~ ~ ~ ~ ~ ~. ~'r'T" ~I'nr~' ~'F' ~ ~ ~ ~6' ".
T.O. IF-16A 1

I

Warning Light Analysis 'PW22ij]
WARNING LIGHTI

ENGINEI

CORRECTIVACTION/REMARKSCAUSEEngine stagnation or flameout

A. Place throttle to OFF and initiate airstart.
Refer to AIRSTARTS

Engine alternator failure/malfunction

B. Refer to ZERO RPM/ERRONEOUS RPM
INDICATION

Check rpm and FTIT indications

Engine overtemperature

C. Land as soon as possible

Engine warning system failure or
RPM/FTIT indicator failure

D. Land as soon as practical

CHECK RPM AND
FTIT INDICATIONSENGINESTAGNATION

SEE A

I
I
ENGINE
FLAMEOUT

SEE A

II

YES
ENGINE'-----------1 ALTERNATOR

FAILURE/
MALFUNCTION

YES" ' > - - - - - - - 1 ENGINE OVERTEMPERATURE

SEE B. - - - - - - -..... SEE CNOENGINE WARNING SYSTEM

O""----------i FAILURE OR RPM/FTIT

1---------. SEE D

INDICATOR FAILURE
IF-1I5A-1-1149A ®

Figure 3-1. (Sheet 7)

I
I

"

~r I " ~ ~ ~ ~,T.O. 'I F-16A-l-

nr ~,p~' ~rr ~lfJIr ~ ~ ~·r.v ~Caution Light Analysis ~?:QQJ
1~~1

CORRECTIVE ACTION/REMARKSCAUSECAUTION LIGHT

Check for specific caution light on caution lightponel

One or more caution lights on

Reset MASTER CAUTION light

NOTE
The MASTER CAUTION light does not reset when
the ELEC SYS or FLT CO NT SYS caution light isilluminated. The ELEC CAUTION RESET buttonmust be depressed or the electrical malfunction
cleared to extinguish ELEC SYS and MASTER
CAUTION lights. For FLT CONT SYS caution
light, the FLCS SERVO ELEC RESET switch must
be placed to ELEC, FCS CAUTION RESET button
depressed, or the malfunction cleared to extinguish FLT CONT SYS and MASTER CAUTION
lights.

AVIONICS

ANTI
SKID
HOOK

EEC

ATF NOTENGAGED
RADAR
ALT
EQUIP
HOT

BUC

IFF

NWS
FAIL
CABIN
PRESS
OXV
LOW

FUEL
HOT
SEAT NOT
ARMED

ECM

NUCLEARSTORESCONFIGPROBE HEATANTI-SKID

FLT CONTSVS

FLT CONTSVS

FWD
FUEL LOW
AFT
FUEL LOW

HOOK

ADC

FWD
FUEL LOW
AFT
FUEL LOWCADCFUEL HOTNWS FAILLE
FLAPSOVERHEATLE FLAPSCABIN PRESS

IFF

CADC

ELEC SVSOVERHEATOXV LOWELECTSVSAVIONICSEQUIP HOTNUCLEARSEAT NOT
ARMED

RADAR
ALT
STORESCONFIG

EEC

ADC

BUC

PROBE HEATBLOCK 10BLOCK 15

Figure 3-2. (Sheet 1)

CORRECTIVE ACTION/REMARKSCAUSECAUTION LIGHT

Check for specific caution light on caution lightpanel

One or more caution lights on
I

MASTER I~~~~I~s~Reset MASTER CAUTION light

NOTE
The MASTER CAUTION light does not reset when
the ELEC SYS or FLT CONT SYS caution light isilluminated. The ELEC CAUTION RESET buttonmust be depressed or the electrical malfunction
cleared to extinguish ELEC SYS and MASTER
CAUTION lights. For FLT CONT SYS caution
light, the FLCS SERVO ELEC RESET switch must
be placed to ELEC, FCS CAUTION RESET button
depressed, or the malfunction cleared to extinguish FLT CONT SYS and MASTER CAUTION
lights.

ADC

FWD
FUEL LOW

ANTI-SKID

FLT CONTSYS

FWD
FUEL LOWAVIONICS

ANTI
SKID

FLT CONT
SYS

AFT
FUEL LOW

HOOK

ADC

AFT
FUEL LOWATF NOTENGAGED

HOOK

CADC

FUEL HOTNWS FAILLE
FLAPSOVERHEAT

RADAR
ALT

NWS
FAIL

LE FLAPSCABIN PRESS

IFF

CADC

ENGINE
FAULT

EQUIP
HOT

CABIN
PRESS

ELEC SYS

OVERHEATOXY LOW

ELECT
SYS

SEC

*IFF

OXY
LOW

AVIONICSEQUIP HOTNUCLEAR

-----

FUEL
HOT

ECM

* NUCLEARSEAT NOT
ARMED

RADAR
ALT

ENGINE
FAULT

-----

SEAT NOTARMEDSTORESCONFIGPROBE HEATSECSTORESCONFIGPROBE HEAT

-----

-----

-----

-----

BLOCK 10BLOCK 15

~IO.lF-16A-l

Caution Light Analysis~---------CAUTION LIGHT

CAUSE

i

CORRECTIVE ACTION/REMARI<S

I Single or dual failure in either pres- A. Refer to ADC/AIR DATA M/\LFUNCTIONS
ADC

I

I

I

ADC

I

I

I

I

FLT CONT
SYS

I

I

DUAL
FC FAIL

I

I
I

ADC

FLT CONT
SYS

I

I

ADC

I

sure sensing/computation system or
AOA system
:

I

CADC

LE
FLAPS

I

I

1

STBY
GAINS

1
I

I

LE
FLAPS

CADC

I

I

I

FAILURE IN PRESSURESENSING/COMPUTATIONSYSTEM OR AOA SYSTEM

LESS

m

icing conditions
YES
AIR DATAPROBE HEAT SWITCH _present with erroneously ">-------1 PROBE ICING t - - - - 1 VERIFY PROBE HEAT
low airspeed?

SEE A

NO

CONT

1 F-16A-1-1150B®

Figure 3~2. (Sheet 3)~,l,;e,~~~~~~~~~~~

T. O. 1F- 16A- 1

Caution Light Analysis
CONT

YES

ABNORMAL
FLIGHT
CONDITIONS

SEE A

NOTE:
May also have DUAL FC
FAIL warning light and
FLT CONT SYS, CADC
(with ADC MFl), and LE
FLAPS caution lights on.

NO

YES

DUAL MALFUNCTION OF
AOA

YES

DUAL MALFUNCTION IN

SEE A

NO

> - - - - - - - - f PRESSURE SENSINGISEE ACOMPUTATION SYSTEM
NOTE:
Lose autopilot.
NO

SINGLE FAILURE IN
PRESSURE SENSINGI>-------~ COMPUTATION SYSTEMOR PNEUMATIC ADASOURCENO

SEE A
SINGLE FAILURE OF AOA'---------------1 SOURCE (LEFT OR RIGHT .--------~AOA TRANSMITTER)NOTE:Will also have ADC MFL
and OFF warning flag on
ADA indicator.
1F-1 6A-1-1 1 S1AtSl

Figure 3-2. (Sheet 4)
Change 5

3-15

T.0.1F-16A-l

Caution Light Analysis
CAUTION LIGHTSI

FlT CONT

SYS

CAUSE

CORRECTIVE ACTION/REMARKS
NOTE:
Depress FCS CAUTION RESET
button to clear FLT CONT SYS
and MASTER CAUTION lights.

I

P. R, or Y light located on FLCP

Single failure in one axis of the FLCS

A. Refer to P, R, AND/OR Y
MALFUNCTIONS

P, R, and Y lights located on FLCP
(three lights)

Single branch power failure in FLCS

B. Refer to P, R, AND/OR Y

SERVOS light(s) (any) on FLCP

Single servo valve failure

C. Refer to SERVO MALFUNCTIONSERVOS lights (all simultaneously)
on FLCP

Failure of either hydraulic system A,
system 8, or momentary drop inpressure

D. Refer to SERVO MALFUNCTION(single light)

MALFUNCTIONSYESSEE DUAL FC FAIL>-----.t WARNING LIGHTANALYSISNO

P, R, OR Y
LIGHT ONSINGLE FAILURE
IN ONE AXISOF FlCS

SEE A

P, R, AND Y
LIGHTS ONSINGLE BRANCHPOWER FAILURE
IN FLCSSEE B

ANY SERVOS
LIGHT

SINGLE FAILURE
IN ISA SERVOSEE CCHECK FLCP FOR
LIGHTSFAILURE OF EITHERHYDRAULIC SYSTEM A,
ALL SERVOS
1 - - - - 4 SYSTEM 8, OR
LIGHTS ON
MOMENTARY DROP IN
SIMULTANEOUSLYPRESSUREYESRESET SERVOSAND CONTINUE
FLIGHTNO
SEE D

1 F-16A-1-1152X@

Figure 3-2. (Sheet 5)

3-16

'" ~ ~ ~ ~ ....l'r ~',c1'r ~llfl~r' ~"T1W' ....'~~"l~'."':~~~i'~1f'

~

~

T.O.

..,

F- 16A- 1

Caution Light Analysis-.

CORRECTIVE ACTiON/REMARKSCAUSECAUTION LIGHT

I

CAOC

I

IPW20011

CADC

II

EEC

IPW22011

CADC

II

ENGINE

Malfunction of CADC or input(sl to it

FAULT

A. Refer to CADC MALFUNCTION

I Loss of valid mach data to engine

B. Refer to EEC CAUTION LIGHT

Loss of valid mach data to engine

B. Refer to CADC MALFUNCTION

MALFUNCTION OF,----------1 CADC OR INPUTIS) I----------~ SEE A
TO IT

L....-_ _ _ _ _ _ _ _~LOSS OF VALID
MACH NUMBER. . . . - - - - - - - - -... SEE BDATA TO ENGINECAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKS

A. LE FLAPS switch - AUTO (no mal-

LE FLAPS switch in LOCK

function)

I

LE

FLAPS

LEF malfunction or asymmetry detected

B. Refer to LEF MALFUNCTION (SYM-

I

METRIC)

C. Refer to ADC/AIR DATA MAL-AOA malfunction

FUNCTIONSStandby gains

SYSTEM MANUALLY
r---------~ LOCKED IN PRESENTPOSITION

1----------..... SEE A

LEF MALFUNCTION
SEE B
r-----.. . OR
ASYMMETRY
t-------......

'-------4

FLCS
MALFUNCTION

I---~SEE

C

1F-16A-1-1153B8

Figure 3-2. (Sheet 6)

I

,

,

,
,
,
,

-

.-

-

,"
, DU
,
,.
,

~~,,. ~ ~ ~ ~10. I F-16A-'..

~.., ~ ~ ~ ~ ~---~-~

Caution Light Analysis
CAUTION LIGHT

CAUSEFailure(s) of electrical system--* Check ElEC control panel for illuminated light(s)FlCS PMG light (in flight)

Failure of FLCS PMG

A. Refer to FLCS PMG FAILURE

MAIN GEN, EPU run, EPU AIR, and(possibly) EPU HYDRAZN lights

Failure of main generator

B. Refer to MAIN GENERATOR FAILURE (IN
FLIGHT)NOTE:
If the FFP fails or system A hydraulic pressure
is lost, refer to HOT FUEL or GRAVITY FEED.

MAIN GEN, EPU GEN, EPU run, EPU
AIR, and (possibly) EPU HYDRAZN
lightsFailure of main and EPU generators

C. Refer to MAIN AND EPU GENERATOR FAILURE

MAIN GEN, EPU GEN, and EPU
PMG lights on. EPU run light off

Failure of main and EPU generators

D. Refer to MAIN AND EPU GENERATOR FAILURE

MAIN GEN, EPU GEN, EPU PMG,
FLCS PMG, ACFT BAn TO FLCS
and/or FLCS BAn lights

Failure of all generators

E. Refer to MAIN AND EPU GENERATOR FAILURE

Failure of aircraft battery or battery
charging system

F. Refer to AIRCRAFT BAnERY FAILURE

One or more of the four FlCS batteries
supplying power to inverter{s) as indicated by individual branch lights

G. Refer to FLCS BAnERY DISCHARGE* FLCS BAn lights A, B, C, and/or D

One or more FLCS batteries failed toconnect to the flight control system
when main generator came on line (onground only)

H. Notify maintenance*NOTE:

Certain aircraft battery charging system or FLCS battery system failures can result in a nonresettable ElEC SYS caution light.

-

_

J

,
I

* ACFT BAn FAIL light

No ELEC control panel lights,

and depress ElEC. CAUTION RESET button

SYS

,I

CORRECTIVE ACTION/REMARKS

,

,

,
J

,
,

Figure 3-2. (Sheet 7)

3-18

Change 3"lllllllllllh

I
I
I
I
I
I

""""""~~~~
Caution Light Analysis
I

I

SYS

I

~ SEE A

MAIN GEN, EPU RUN,
AND AIR (AND POSS- IBLY HYDRAZN) LIGHTS• MAIN GENERATOR FAILURE• EPU GENERATOR OPERATING• FLCS PMG AND EPU PMG
OPERATING~ SEE BI--

MAIN GEN, EPU GEN,EPU RUN, AND EPU
AIR (AND POSSIBLY
HYDRAZN) LIGHTS ON-• MAIN AND EPU GENERATORSFAILURE• EPU PMG IS OPERATING AT
PROPER SPEED• FLCS PMG IS OPERATING~ SEE C-

MAIN GEN, EPU GEN,AND EPU PMG LIGHTS
ON. EPU RUN LIGHT
OFF-• MAIN GEN FAILURE• EPU GEN, EPU PMG FAILED OR
UNDERSPEED• FLCS PMG OPERATING• AIRCRAFT BATTERY POWERINGBATTERY BUS(ES)--. SEE D-• MAIN GEN FAILURE• EPU GEN, EPU PMG FAILEDOR UNDERSPEED• FLCS PMG FAILURE• AIRCRAFT BATTERY AND/OR
FLCS BATTERIES POWERING
FLCS• AIRCRAFT BATTERY POWERINGBATTERY BUS(ES)~ SEE E

AIRCRAFT BATTERY SYSTEMFAILURE~ SEE F-. SEE G-. SEE H

r-----I* CHECK ELEC CONTROLPANEL FOR ILLUMf--f--INATED LlGHT(S) AND
DEPRESS ELEC CAUTIONRESET BUTTON*NOTE:
Certain aircraft battery charging
system or FLCS battery system
failures can result in anonresettable ELEC SYS
caution light.

~

NONE OF THE FLCS BRANCHESARE RECEIVING POWER FROM
THE FLCS PMG

ELEC

f---

FLCS PMG LIGHT(IN FLIGHT)

MAIN GEN, EPU GEN,EPU PMG, FLCS PMG,
AND ACFT BATT TO
FLCS AND/OR FLCS
BATT LIGHTS ON. EPU
RUN LIGHT OFF

H·

1-

---1

* FLCS BAn LlGHT{SI~ONE OR MORE FLCS BATTERIESARE SUPPLYING POWER TO
FLCS INVERTER(S) AS INDICATEDNO ELEC CONTROLPANEL LIGHTS-ONE OR MORE FLCS BATTERIESDID NOT CONNECT TO THE
FLIGHT CONTROL SYSTEM (CANONLY OCCUR ON GROUND)ACFT BAn FAIL

LIGHT

-

'F-'6A-'-1154X@~

Figure 3-2. (Sheet 8)~~~~~~~~~~~~~I-~Caution Light AnalysislM
CAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKSAVIONICS

Several possible causes. Check display on
FCNP

Depress FALT ACK button on FCNP and note
fault list display. AVIONICS caution light resets
automatically

Ejection safety lever up (system safe)

When desired, rotate ejection safety lever down
(armed)Hook not up and locked

Normal landing, touchdown beyond approachend arresting gear

NWS system failure or loss of NWS electrical power

Normal landing, NWS is not available. Refer 10
NWS FAILURE/HARDOVER

I

SEAT NOT
ARMED

I

HOOK

I

NWS
FAIL

I

I

I
I

NOTES:
1. Does not illuminate for NWS failure dueIto hydraulic system B failure or NLG
strut overextension.
2. Possible indication that AR door cannot
be opened or closed.

I PW20011

BUe

I

FUEL
HOT

I

BUC, selected and engine operating in
SUC or main fuel pump pressure is low

Refer to ENGINE MALFUNCTIONS

Temperature of fuel to engine excessive

Refer to HOT FUEL or GRAVITY FEEDI
Forward reservoir contains:FWD
I FUEL LOW

I[B] Less than 400 pounds of fuel

Refer to FUEL LOW[ID Less than 250 pounds of fuelAft reservoir contains:

~

I

AFT
FUEL LOWI[8J Less than 250 pounds of fuel

Refer to FU EL LOW

lID Less than 400 pounds of fuel

I

LESS~

LESS~

I

IFF

I

NUCLEARI

ECM

I

STORESCONFIGI

Mode 4 sw is OFF; not coded; correct code
not selected (A or B); code does not match
code interrogation; or mode 4 is inoperative

Advisory

Malfunction in nuclear control circuitry

Advisory

Not used

None

STORES CONFIG switch is in incorrect
position or loading category in SMS software disagrees with actual GP/STORE/
LINE loading category

Verify STORES CON FIG switch is in proper position for aircraft loading category. Refer 10 T.O.
1 F- 16A- 1 -3 (Block 10) or 10. 1F- 16A- 1 -4 (Block
15 and~)

I

I

I

rn

LESS
STORES CONFIG switch in
CAT I with CAT II stores loading

If aircraft is loaded CAT II, the switch may beplaced either in CAT I or CAT III. However, if
CAT I is selected, the CAT" maximum bankangle change limits must be observed. Refer to
Section V

Figure 3-2. (Sheet 9)~,2,;e,~~~~~~~~

~"""""""""""
rO.1F-16A-l

Caution Light Analysis
CAUTION LIGHT

I

CABIN
PRESS

I

CORRECTIVE ACTION/REMARKS

CAUSE
Cockpit pressure altitude above 00
22,500, LESS 00 27,000 feet

BELOW 00
22,500, LESS ~27,000 FEET

A. Refer to COCKPIT PRESSURE/TEMPERATURE MALFUNCTION

MALFUNCTION
IN INDICATIONSYSTEM

CONTINUE
FLIGHT

COCKPIT PRESSURE
ALTITUDE EXCEEDS
00 22,500,
LESS 00 27,000 FEET....-----1 BECAUSE OFCHECK COCKPITPRESSURE
ALTITUDE

PRESSURIZATIONFAILURESEE A

ABOVE 00
22,500, LESS 00
27,000 FEETAIR SOURCEKNOB- NORM

CAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKS

E~UIP

Avionic equipment cooling air temperature/pressure insufficient

A. Refer to EQUIP HOT CAUTION LIGHT

I

OT

I

r--------I

~
~
~
~~

~
~
~
~
~

AIR SOURCEKNOB- NORMILLUMINATES WHEN
AVIONIC EQUIPMENT
COOLING AIR TEMPERATURE/PRESSURE
IS INSUFFICIENT
DEGRADED EQUIPMENT PERFORMANCE
AND/OR DAMAGECAN RESULT~

t--------...... SEE A~~~~

Figure 3-2. (Sheet 10)

Change 13

3-2~ ~

'-'-""llllfl""""'~~T. O. 1F- 16A-1~ Caution Light Analysis

~
~
~
~
~
~
~
~
~
~
~
~~

~

CAUTION LIGHT

CAUSE

CORRECTIVE ACTION/REMARKSI OVERHEAT I

Overheat condition in engine com partment, ECS or EPU bays, or MLG wheel
wells

A. Refer to OVERHEAT CAUTION LIGHT

THROTTLE TOMINIMUM PRACTICAL THRUSTOVERHEAT>-__Y_ES_~ CONDITIONYESIN ENGINE
COMPARTMENT

SEE A

NO

NO

TURN EPU OFF
IF IT WAS MANUALLY ACTIVATED

YES

YES

OVERHEAT" > - - - - - - 1 CONDITIONSEE A

IN EPU BAY
NO

CONT

1F-16A-1-1156X®

Figure 3-2. (Sheet 11)

3-22

""""""~~~
Caution Light Analysis
CONT

AIR SOURCEKNOB - OFF

YES

OVERHEAT CON-> - - - - - - - 1 DITION IN ECS

SEE A

BAY
NO

NO

LG -DNOVERHEAT CONDITION I - - - - - - - - - t LAND AS SOON
AS POSSIBLE
IN MLG WHEEL WELLS

YES

NO

OVERHEAT CONDITION IN ENGINE COMECS OR EPU BAYS, OR MLG
WHEEL WELLS STILL PRESENT~---I PARTMENT,NO

OVERHEAT CONDITION IN ENGINE COM~RTMEN~ ECS OR EPU B~S, OR MLG ~----------------~---~ SEE A
WHEEL WELLS STILL PRESENT

1F-16A-1-1157X®

Figure 3-2. (Sheet 12)-

Caution Light Analysis
CAUTION LIGHT

I

ANTI
SKID

I

CORRECTIVE ACTION/REMARKSCAUSEANTI-SKID switch OFF or system
malfunction

A. Refer to ANTISKID MALFUNCTION(GROUND) or ANTISKID MALFUNCTION(LANDING), as appropriate

ANTISKID

YES

SEE A

MALFUNCTIONNOPOSITION TO
ANTI-SKID

CAUTION LIGHT

II
I

ATF NOTENGAGED

RADAR
ALT

OXY
LOW

I~
I
I

CAUSE

CORRECTIVE ACTION/REMARKS

Not used

None

Malfunction of radar altimeter

Move RDR ALT switch to OFF

Oxygen pressure is below 42 psi or
quantity is below 0.5 liter

A. Descend to 10,000 feet cockpit
pressure altitude. Use EMERGENCY
oxygen if required

SEE A

NO

Figure 3-2. (Sheet 13)

QUANTITY
INDICATING
MALFUNCTION

1 E-16A-1-1158X

'IIIIIIIIIII~~~~
Caution Light Analysis PW200I

CAUTION LIGHT

[

EEC

CAUSE

CORRECTIVE ACTION/REMARKS

Engine alternator failed

A. Refer to ZERO RPM

Loss of valid mach data from CADC
or EEC malfunction

B. Refer to EEC CAUTION LIGHT

[

SEE B

NO
LOSS OF VALID MACHDATA FROM CADC

YES

YES

1. THROTTLE - MIL
NO

2. SLOW TO SUBSONICSEE B

SPEED USING SPEEDBRAKES AS REQUIREDNO

THROTTLE - MIDRANGE(80-85% RPM, FTIT 700°-

750°C)YES

ENGINE ALTERNATOR
FAILED; EEC INOP; MAIN
AND AB IGNITION INOP

SEE A
1F-16A-1-1159X@

Figure 3-2. (Sheet 14)

3-25

~

~

IO.lF-16A-l

Caution Light Analysis
CAUTION LIGHT

I
I

I

ENGINEFAULT

I PW2201 I

ENGINE
FAULT

I

I

PW220!

PW2201

I

SEC

I

CAUSE

~fn

I LESS ~LESS fn

I

I PROBE HEAT I

CORRECTIVE ACTION/REMARKS

Engine fault(s} detected

Refer to ENGINE FAULT CAUTION LIGHT

Engine AB malfunction

Refer to ENGINE FAULT CAUTION LIGHT

Loss of valid mach data from CADC
Engine operating in SEC or main fuelpump pressure is low

Refer to SEC CAUTION LIGHT

Ground: Probe heater failure, monitoring system failure, or one/both AOA
probe heaters have shut off to prevent
overheat

Ground: Place PROBE HEAT switch to OFF for 1
minute (caution light goes off when OFF isselected); then reselect PROBE HEAT. If caution
light comes on simultaneously with reselection of
PROBE HEAT, a probe heater or monitoring system
failure has occurred. If caution light does notcome on when PROBE HEAT is reselected, one/both AOA probe heaters were shut off to prevent
overheat

In flight: Probe heater(s} or monitoring system failure

In flight: Place PROBE HEAT switch to PROBE
HEAT, if required, and avoid areas of known orsuspected icing conditions~

Figure 3-2. (Sheet 15)~~2~ ;;e~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~Pilot Fault List
FCNP FAULT CODE

- Engine fANI t13
CAUSE

CORRECTIVE ACTION/REMARKS

ENG 003

Communication lost between EDU andmux busIlluminates AVIONICS caution light. A subsequent
engine fault causes a nonresettable ENGINEFAULT caution light

ENG 084

Anti-ice valve failed open and/or bleed
air temperature greater than 850~F

Reduce throttle setting to midrange unlessrequired to sustain flight. Operating the engine
above midrange with anti-ice system failed on
may result in engine stall. Land as soon as practical

ENG 085

Engine anti-ice valve failed in closed
position

Avoid areas of known or suspected icing conditions

ENG 086

The CADC supplied mach number tothe DEEC is no longer available

Supersonic stall protection is inoperative. Do notretard throttle below MIL while supersonic. If
CADC caution light is also on, refer to CADC
MALFUNCTION, this section

ENG 087

AB system failure detected

AB RESET switch - AB RESET. Land as soon aspractical if fault does not clear. AB operation ispartially or fully inhibited

ENG 088

Loss of Redundant FTIT signals received
by DEECMIL rpm is reduced 7 percent by DEECDEEC has detected a failed open or
missing nozzle

If a failed open or missing nozzle is suspected,
refer to NOZZLE FAILURE, this section

Communication lost between DEEC
and EDU

Do not retard throttle below MIL while supersonic.
Only ENG 084 PFL can be displayed

ENG 089

Figure 3-3.

~'IIIIIIIIIIIIIII~"'111.

~ ~~~FUI~~l EMERGENCIES~ :::~:~:~,R:E,;;~F~r~'~o~~':~~~::Ub~:)etected by

~
~~

~

~

~~
~~
~~
~

flames, smoke. explosion, signal from ground crew, orradio call. FTIT may exceed 680 c C and, if ac power isavailable. ENG FIRE warning or OVERHEAT
caution light may illuminate.
Throttle - OFF.

2.

.JFS switch - OFF.

~~
~
~~

~
~
~
~
~
~

Throttle - OFF,2.

FUEL MASTER switch - OFF.

ANTISKID MALFUNCTION (GROUND)

3.

FUEL MASTER switch - OFF.

4.

ENG FEED knob - OFF (if external power
applied ).

Monitordetected while the aircraft is moving above 5 knots,
the ANTI SKID caution light illuminates. In mostcases this represents the loss of a wheel speed sensor
signal, and the system switches to an alternate
braking mode. In this mode, if differential braking isapplied (15 percent or greater difference between
pedals), both brakes oscillate between pressure asmetered and no pressure. Braking effectiveness is
reduced by 50 percent or greater. If brake pedals arewithin 15 percent, the system uses the informationfrom the remaining good wheel speed sensor andstopping distance is increased by approximately 25
percent on both wet and dry runways. An ANTISKID caution light which only illuminates below 5
knots indicates a malfunction that does not affectbraking performance. Normal braking and antiskid
are available; however, system redundancy may
have been lost.

IfFTIT remains above 500°C:&ED Below 20 knots ground speed the alternate

If fire continues:fi.

Abandon aircraft.

~~ ~: ;~:~t~~cd~C~::~~l.

&ED If a failure affecting braking performance is1.

~~ ::Ts~:=~TF~~~~e~:~ooc. During engine start, if~~

vVlwll tht, aircraft i:-- full~ ;:-,topped, 1 ,lye chl)ck . .
installed or engage parking hrake Lem'\ the h:ltter:,on until chocks are in:"talled.

the FTIT increases at an abnormally rapid rate
through 575°C, a hot start can be anticipated.

13.

JFS switch - START 2.
Motor engine with JFS until FTIT reaches
200°C or for four minutes (JFS ground
operating limit), whichever occurs first.

HUNG START/NO START

Hung start - RPM has stopped increasing below
IDLE and FTIT is stabilized at less than 680°C.
No stcut - Light-off does not occur within 20 seconds.
1.

Throttle - OFF. Notify maintenance.

ENGINE AUTOACCELERATION (GROUND)

If the engine autoaccelerates on the ground, primaryconsideration should be given to shutting the engine
down as quickly as possible. With engine shut down,
only the brake/JFS accumulators are available tosupply hydraulic pressure for braking. Stop theaircraft by making one steady brake application.

3-28

Change 13

braking mode is less effective. Place BRAKES
channel switch to CHAN 2 and ANTI-SKID switch toOFF. Braking will be manual.
LESS ffiD In case of an antiskid failure, the ANTISKID caution light illuminates and the brakesystem automatically switches to pulsating pressure. In this mode, braking effectiveness is reduced
approximately 50 percent; however, in most cases,
braking effectiveness is as good as can be obtainedwith ANTI-SKID switch in OFF while avoiding
wheel lockup and its associated risk of control
difficulty. Short field landing distances are
increased approximately 60 percent for dry runway
and 25 percent for wet runway from those normallycomputed.

If the ANTI SKID caution light illuminates (with the
ANTI-SKID switch in ANTI-SKID):
1.

~ [NQ] DRAG CHUTE switch - DEPLOY (ifrequired).

IIIIIIIIIIII~~~
2.

Brakes - Apply as needed.

NOTE.00 Use of maximum symmetric pedal

Ipressure provides the best stopping
performance. Differential brake only
when essential for directional control. If
the ANTI SKID caution light illuminated above 5 knots groundspeed, theaircraft may oscillate due to pulsatingbrake pressure (if 15 percent or greater
differential pedal pressure is applied).
Changing brake channels may restore
normal braking.
• LESS iFE Maximum pedal pressure isrequired to obtain approximately 50
percent of normal braking force. Iflessthan maxim urn pedal pressure is used,braking is extremely degraded. The
aircraft will oscillate due to pulsatingbrake pressure. Changing brake channels will not restore normal braking
since the same antiskid signal is usedin both brake channels.

3.

NWS - Engage (if required).

If manual braking is desired or after aircraft is
stopped:
4.

00 BRAKES channel switch - CHAN 2.

5.

ANTI-SKID switch - OFF.

No antiskid protection is available
with the ANTI-SKID switch in OFF 00
and BRAKES channel switch in CHAN
2. Brakes should be applied withcaution to avoid wheel lockup andblown tires.

NOTE
LESS 00 Below normal taxi speed,
pulsating braking is only marginallyeffective. Stopping distance may be
shortened with antiskid off.

~

,.

-

~~~.~~. ~ ~-~T~~
T. O. 1F- 16A- 1
1.

:'-.1alfuncl II\nS in ,.\,;tems .. ,jeh ah·:·t norm;;; braking
urocedut,l' which
addressl';-' each ::-:p('cific ".\:-;tem, Une of the brake
failure Ir10des is the loss o( one brakv circuit. With thisfailure, hoth br~kt.'s an- .-:till 8\'ailable; howevlcr,
significantly m(II'!' ped:l] force than normal isrequired to achit'\,(~ a spl'cdk braking effectivenes::i.

BRAKES channel switch - Change channels.

arp descrihed iii the elllt'lg'enc.,

Another failure mode is lil:'s of brakes on one or both~lLG. Changing brake channels may return thesystem to normal operation. 'I\lrning the ANTI-SKID
switch to OFF irID and confirming BRAKES channel
switch in CHAN 2 may also restore braking; however,the system reverts to man nal con trol and antiskid
protection is lost. (Status of the power source for toe
brake transducers can be determined by testing theFLCS PWR lights on the TEST switch panel.) Releasebrake pedal pressure before changing channels or
turning off the ANTI-SKID switch to avoid
immediate brake lockup if braking returns. When
moving the ANTI-SKID switch, be very careful not toselect the PARKING BRAKE unless that is what is
intended. If directional control is a problem (such aswith one brake inoperative on landing roll), do nothesitate to use NWS. If conditions permit, consider ago-around if the brakes are found to be inoperative on
landing. Lower hook if a cable is available. If normal
brakes cannot be restored, do not hesitate to use theparking brake if a cable is not available. The lower the
groundspeed, the less chance there is for aircraft
damage when using the parking brake. If the aircraft
is accelerating, use the parking brake early. It may bepossible to cycle the parking brake on and off and stop
the aircraft; however, regardless of technique, use of
the parking brake may result in blown tires.
Another failure mode is a hydraulic leak in the brakeitself, which might not be apparent until after
two-point aerodynamic braking. In this case, if a cable
is not available, the aircraft should be stopped using
the good brake and N\VS for directional control. Oncethe aircraft is stopped, do not engage the parking
brake; use continuous pedal pressure on the goodbrake only. Failure to do so could deplete thehydraulic system and result in total brake failure
prior to chock installation.

Release brakes prior to changing
brake channels or turning antiskid off.
2.

irE BRAKES channel switch - CHAN 2.

3.

ANTI-SKID switch - OFF.

Release brakes prior to changing
brake channels or turning antiskid off.
4.

NWS - Engage (if required).

5.

HOOK switch - DN.

If arresting cable is not available or if at low
groundspeed:
6.

ANTI-SKID switch - Intermittent PARKINGBRAKE, then ANTI-SKID.

~
~

I~

~
~
~

~

,

~

C!!!~!D~If in a congested area, use the parking
brake immediately to stop.

,

HOT BRAKES

The pilot has the responsibility to determine when a
hot brake condition exists. The pilot evaluates thesituation by analyzing the variables that influence
brake temperature: GW, pressure altitude, OAT,
speed at brake application, etc. Refer to T.O.
IF-16A-l-1, PART 2, BRAKE ENERGY LIMITSMAXIMUM EFFORT BRAKING. Observations byground crewmembers should also be used as certain
malfunctions that result in overheated brakes, suchas dragging brakes, may not be readily apparent to
the pilot. Perform hot brake procedures anytime hot
brakes are suspected.

NOTE

It is impossible for the ground crew to avoid the hot
brake and engine intake danger areas while pinning
the EPU or chocking the aircraft. Therefore, if
conditions permit, the aircraft should be shut down
without pinning the EPU or chocking the wheels.

If conditions permit, consider a goaround if the brakes are found to be
inoperative on landing. An approachend cable arrestment is recommended.

Release brake pressure as soon as possible tominimize heat transfer between the brake surfaces
and the wheel. This action also relieves hydraulicpressure to the brakes, which if leaking, could feed ahydraulic fire.

Accomplish as many steps as required:~-~~~IIIIIIIIIIII-

Perform the following after any event that may resultin hot brakes:

-

,,
,
,
,
,

Use only minimum possible toe brakepressure to hold aircraft stationary
until engine is shut down and nose
wheel is chocked.

3.

EPU switch - OFF.

4.

Throttle - OFF.

• If a hot brake condition is a result of a
dragging brake, taxiing the aircraft
worsens the condition.

5.

Nose wheel- Chocked.

6.

MAIN PWR switch - OFF.

• Any leaking hydraulic fluid may be
ignited by hot wheel and brakesurfaces.

7.

Exit toward the front of the aircraft.

IWARNINGI-

Request firefighting equipment and proceed
directly to the designated hot brake area ornearest area clear of other aircraft andpersonnel.

l.

• Wheel fusible plugs may relieve tirepressure at anytime during the 15
minutes after brake application.
• With hot brakes, avoid inflated MLG
tire side area within 300 feet for 45
minutes after aircraft has stopped. If
required, approach from the front orrear for firefighting purposes only.
When in the hot brake area:
Align aircraft with nose into wind if possible.

2.

IWARNING I

If a brake fire occurs:
8.

Go to GROUND EGRESS, this section.

MAIN GENERATOR FAILURE (GROUND)

If the main generator fails on the ground, the FLCS
PMG and aircraft battery provide power for fullnormal braking (both channels). The EPU should
activate and provide power for NWS.
Stop and engage the parking brake prior toattempting to reset the generator.
If main generator resets and further taxiing isrequired, brakes should be checked carefully. Allow
the aircraft to begin rolling slowly and check fornormal braking. If normal braking is inoperative,
immediately engage the parking brake.

• Do not use the parking brake.

If main generator failure is indicated:• If battery power is not available, toe
brakes will be inoperative after engine
shutdown.

1.

Stop the aircraft.
'fum EPU on, if required, to obtain NWS.

2.

ANTI-SKID switch - PARKING BRAKE.

3.

OXYGEN - 100%.

4.

EPU switch - OFF.

• Do not turn MAIN PWR switch to OFF
until the nosewheel is chocked.
• Attempt to park in a level area tominimize risk of aircraft rolling if thebrakes should fail after shutdown.
• Delay engine shutdown until arrival of
firefighting equipment because hotwheels and brakes may ignite fueldrained overboard during engine shutdown.

3-30

If chocks are not installed, be preparedto immediately engage the parking
brake ifit disengages when the EPU isshut off.

Change 7,~~~~~~~~~~~~~~

T. O. 1F- 16A- 1
If further taxiing is required:

5.

4.

I WARNING I

MAIN PWR switch - BATT, then MAIN PWR.
Toe brakes and parking brake are available
with or without the EPU as long as the MAINPWR switch is not moved to OFF.

Exit over the left side (conditions
permitting) to avoid EPU exhaustgases.
5.

If main generator cannot be reset,

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

.IID Consider canopy jettison so rearseat occupant can egress more rapidly.

EMERGENCY ENTRANCE AND CREW RESCUE• Opening the canopy with the
MANUAL CANOPY CONTROL handcrank is extremely difficult. If immediate egress is required, the canopy
should be jettisoned rather thanopened with the hand crank.

Refer to figure 3-4 for emergency entrance and crew
rescue procedures.
EMERGENCY GROUND JETTISON
Ground jettison of the 300-gallonl370-gallon fueltanks results in the tank(s) striking the ground before
the pylon aft pivots release. The tank(s) will probably
rotate horizontally and may strike the LG. Use
EMER STORES JETTISON on the ground only as alast resort. Refer to EMERGENCY JETTISON, thissection.

If canopy does not raise:

6.

1.

Throttle - OFF.

2.

Ejection safety lever - Safe (up).

3.

Harness and personal equipment - Release.

OXYGEN - 100%.

I WARNING I• If jettison is unsuccessful, heat, blast,
and toxic gas from the rockets may
enter the cockpit.

GROUND EGRESSThe order of accomplishment of ground egress steps
depends on the nature of the emergency. For quickest
ground egress (without jettisoning the canopy), place
the canopy switch up and then prepare for exit while
the canopy is opening. However, if fire or danger of
explosion exists, accomplish steps necessary foregress prior to opening canopy to retain maximum
protection until ready for exit. Disconnect parachuterisers, lap belt, survival kit, and g-suit. Oxygen andcommunication leads are quick-disconnect. If required, the canopy can be jettisoned even after it has• been partially or fully opened. If the canopy is
restrained by debris or jammed by crash damage,attempted jettison may result in a portion of the
canopy rocket exhaust entering the cockpit. Thisexhaust may present a heat and blast hazard in the
cockpit; toxic gases are present and 100 percent
oxygen should be used.

Canopy - Open.

I WAR N I N G I

NWS is inoperative unless the EPU is
activated.
6.

EPU switch - OFF (time permitting).

• To prevent the flow of oxygen into the
cockpit after the oxygen hose is
disconnected, do not select
EMER,
LESS m EMERGENCY.

m

7.

Canopy - Jettison.

I WARNING I
Pulling the CANOPY JETTISON Thandle other than straight out may
cause the handle to jam. If the
CANOPY JETTISON T-handle ismounted so that the words CANOPYJETTISON engraved on the T-handle
are upright, an underhand grip should
be used. If the CANOPY JETTISON
T-handle is mounted so that the words
CANOPY JETTISON are inverted, an
overhand or underhand grip may be
used.
Change 53-31

10. 1 F- 1 6A-l

Emergency Entrance and Crew Rescue(Typical)
NORMAL

1. If time and conditions permit:

CANOPY SWITCH

a. Insert a 1/4-inch drive socket wrench/speed handle
into canopy handle lock access plug and rotate ccw
to remove plug.
b. Insert an 8-inch or longer piece of number 25 drillrod (or 1/8-inch rod) into opening and push inboard
to unlock canopy handle.

I.c.

c. Position external CANOPY switch to UP.

Positioning the external CANOPY switch to UP
prior to unlocking the canopy will overheat
the canopy actuator motor or pop the circuit
breaker.
d. If canopy is still not open, insert 1/4-inch drive
socket wrench/speed handle into the external
canopy handcrank receptacle and rotate cw(approximately [AI 52 or rnJ 87 revolutions
required to fully open canopy).

IEMERGENCyl

2.

I.d.

Open the canopy emergency release door and extend
the canopy jettison D-handle to full length of cable(approximately 6 feet). When the cable tightens, pullhandle hard to jettison the canopy.

WARNING

o

I

o

o0

0

@e.!Q()00

C-J

The canopy jettisons upward and back toward
the vertical tail with great force. Stand to theside and slightly aft of canopy to full lengthof the cable to avoid canopy rocket blast.

1.a.

ACCESS PLUG

\\
\

,,

\

.,' \

I.b.

,\ \

;,

"'''''"(' "(Locked)(Unlocked)[]] IBF ICANOPY HANDLE

RIGHT OR LEFT
SIDE OF AIRCRAFT
1 F-16X-1-0016X®

Figure 3-4. (Sheet 1)

3-32

"""""'I~~~­
Emergency Entrance and Crew Rescue(Typical)FLIP-UP PITOTS(STOWED)UWARSINSTALLEDSEAWARSINSTALLEDSEAWARS/UWARSNOT INSTALLEDEJECTION SAFETY
LEVER (SAFE)-EMERGENCY RESCUE

3.

Rotate ejection safety lever located onleft of seat to full up (vertical) position.

I WARNING I
To prevent possible seat ejection
during rescue, rotate ejection safety
lever located on left of seat to fullup (vertical) position.

4.

1-

SAFE
POSITION

Disconnect crewmember from lapbelt,
g-suit hose, survival kit straps, andparachute risers.

1F-16A-1-1161B@,-_

Figure 3-4. (Sheet 2)~~~~~~~~IIIIII~

~~~IIIIIIIIIIII

~
~
~
~

~
~

~

~ •

~

HOT REFUELING EMERGENCY

In the event of a fire or fuel leak/spill while refuelingin hot pit area, refer to FIRE/OVERHEATIFUELLEAK (GROUND), this section. In the event of fire in
the area of refueling operation (other than in the hot
pit area), have the refueling operation discontinued
and taxi clear.
ACTIVATED EPU/HYDRAZINE LEAK

If landing with an activated EPU or a hydrazine leakis detected while the engine is running:
Inform landing base of hydrazine leak or EPU
operation and request bioenvironmental services
support.

~.
1.

NWS FAILURE/HARDOVER

NWS failure may be detected by the NWS FAIL
caution light or uncommanded NWS inputs with nocaution light. If NWS FAIL caution light is on, do not
engage NWS. If the NLG strut is overextended, theNWS cannot engage. If the NLG strut overextends
after NWS engagement, NWS becomes disengaged
and the ARlNWS light goes off.

I WARNING I
NWS malfunctions at any speed may
cause an abrupt turn, tire skidding or
blowout, aircraft tipping, and/or
departure from the prepared surface.

Treat any leak as a hydrazine leakuntil investigation proves otherwise.

1.

NWS - Disengage.

OXYGEN - 100%.

2.

AR/NWS light - Verify off.

3.

Rudder and brakes - As required.

When on the ground:

~12'

-

To prevent sitting in a sealed cockpit(hot) without ECS, consider waiting
for ground crew to arrive with ladderand oxygen bottle prior to shutting
down the engine.

IWARNING I

~

~

NOTE

AIR SOURCE knob - OFF (if required).
Consider turning the ECS off to prevent thepossibility of hydrazine fumes or EPU
exhaust gases entering the cockpit.

If AIR SOURCE knob is placed to OFF,also turn off nonessential avionicequipment as electronic equipment
may be damaged.
3.

Taxi to designated isolated parking area (ifrequired) and park aircraft with left wing intowind if possible.

4.

Insure all nonessential personnel are clear.

5.

EPU switch - OFF.

6.

Shut down the engine (after left main wheel ischocked).

TAKEOFF EMERGENCIESDELAYED ROTATION

Several factors can cause the airspeed at whichrotation occurs to be greater than that determinedfrom T.O. IF-16A-l-1. As the weight of external stores
carried increases, more nose down moment must be
overcome to rotate for takeoff. Another factor is theapplication of roll stick force in addition to aft stickforce. Applying a roll input reduces the maximum
trailing edge up position for one horizontal tail andincreased airspeed may be required to compensate.
The last and most significant factor is improper
servicing of the nose gear strut. Improper servicing
may not be detectable during preflight inspection andmay cause rotation speed to increase by up to 15
knots. All of these factors combined may add up to 25
knots to the computed airspeed for rotation. If
pre-takeoff flight control checks were normal and theengine is operating normally (acceleration checknormal), the aircraft will rotate above computed
rotation speed. Therefore, takeoff should not beaborted due to delayed rotation until at least takeoff
speed is attained. Notify maintenance after flight if asignificantly delayed rotation occurred.

I I 1 1f~c~~""rF ~"IIIr ,I
roo 1

ABORT

The decision to abort or continue takeoff tkpends dn
many factors. Considerations should include. but not
be limited to, the following:• Runway factors: Runway remammg, surface
condition (wet, dry, etc.), type and/or number ofbarriers/cables available, obstructions alongside orat the departure end, wind direction and velocity,
and weather and visibility.
• Aircraft factors: GW, stores, nature of theemergency, speed at decision point, and importanceof becoming airborne.
• Stopping factors: Maximum antiskid braking,speedbrakes, aerodynamic braking, hook, and [@INQ] drag chute.

Aborts abo\"t; 100 KC.AS requ ire diligvl1t adhf. 1'1 :!lCe tothl: procedures in this section for the aboit to be
successful. (f aborting after roLation, retard tb l'ottle to
IDLE and maintain two-point attitude while applying
maximum wheel braking (maximum pedal pressure(antiskid on) consistent with maintaining directional
control). vVhen wheel brakes become effective, the nose
automatically lowers. After the nose wheel is on the
runway, use maximum effort braking (full aft stick, fullopen speedbrakes, and maximum wheel braking). If
aborting before rotation, retard throttle to IDLE,
maintain three-point attitude and apply maximum
efIort braking if stopping distance is critical. NWS
should be engaged if directional control is a problem.
Consider following hot brake procedures after any
abort. Taxiing after an abort will further increase braketern pera ture.

I WARNING I• When braking absorbs a high amountof energy, do not shut down engine
until firefighting equipment is available and do not use the parking brake.

Aborting takeoff at high speed with ablown tire may be more dangerous
than continuing takeoff. For heavy
GW takeoffs, an abort at high speedwith a blown tire is extremely dangerous because braking and directional
control are impaired.

• Hot wheels and brakes may ignite fueldrained overboard during engine shutdown or leaking hydraulic fluid. Wheel
fusible plugs may relieve tire pressure
within 15 minutes after stop.
1.

Normally, with the short takeoff distances of theaircraft, abort is not a problem unless directional
control is a factor (e.g., blown tire). An early decision toabort provides the most favorable circumstances. If
there is any doubt about the ability to stop on the
runway, lower the hook.
Consider aborting after becoming airborne only whensufficient runway is available and flight to a key
position is not possible.

Throttle - IDLE.

I WARNING I• At high speed (prior to WOW), forward
stick pressure in excess of approximately
2 pounds results in full trailing edge
down deflection of the horizontal tails.
This causes excessive loads on the NLG
which can lead to nose tire failure andpossible structural failure of the NLG.
• Failure to use full antiskid braking or
applying brakes with engine above idle
thrust significantly increases the wheel
brake temperature and probability of awheel brake fire.

6A-l

When the throttle is retarded to IDLE
from MAX AB, the thrust and rpm
decay to idle can take up to 2-4 seconds.
Do not mistake high thrust/rpm for
failure of the engine to respond to theidle command. Engine shutdown fromMAX AB may result in a tailpipe fire.
2.

!frn INQ] DRAG CHUTE switch - DEPLOY (ifrequired).

3.

Wheel brakes - Apply (as required).

4.

HOOK switch - DN (if required).
• The hook should be lowered at least 1500 feetfrom the cable to allow adequate time for hook
to stabilize and for full hold down force to be
developed by the hook actuator.
• Refer to CABLE ARRESTMENT, this section.

I
I
I
I
I
I
I
I
I
I
I

,,-'111111~ff'11111~F.
r.o. 1 f-··16A-l

I

IWARNINGIJ\B MALFUNCTION ON TAKEOFF
l\n AB m81function can be detected by a thrust loss

The hook may miss the cable if theaircraft is not slow enough to compress
the MLG struts sufficiently to makeWOW or if forward stick pressure is held.
Ifon fire:5.

Throttle - OFF.
NOTE

With engine shut down, NWS is lostand EPU does not activate automatically. After hydraulic pressure drops,
braking is available using the brakel
JFS accumulators only. Stop straightahead and engage parking brake.
6.

FUEL MASTER switch - OFF.

ENGINE MALFUNCTION ON TAKEOFF

An engine malfunction on takeoff presents ademanding situation where critical actions must be
accomplished quickly with little time for analysis. If
takeoff is continued, a straight ahead climb isgenerally preferred over an immediate turn to low
key. This action provides more favorable ejection
parameters and an increase in analysis time. If
necessary, use only shallow turns to avoid
aggravating the situation. Jettison stores if required
to reduce GW.
ENGINE FAILURE ON TAKEOFF

Engine failure shortly after lift-off may not permit
time for analysis or corrective action. The primaryconcern should be to trade any excess airspeed for
altitude and to eject prior to allowing a sink rate todevelop. Jettisoning stores may aid in gaining
altitude but must not delay the ejection decision. If
the failure occurs later in the takeoff phase, time
may be available for analysis or corrective action.
If conditions permit:
1.

Abort.

If conditions do not permit an abort:
1.

Zoom.

2.

Stores - Jettison (if possible).

3.

Eject.

3-36

Change 7(md nozzle' closun~ or failure of AB to light withinallowed time or stalls accompanied by a loud bang or
pop. An AB failure (other than a slowlno light) may
indicate other engine problems. If possible, abort the
takeoff. If takeoff is continued, the throttle should be
retarded to ~IL. If normal thrust is not available in
lv1IL, refer to LOW THRUST ON TAKEOFF OR AT
LOW ALTITUDE (NON-AB), this section. AB
operation should not be reattempted unless required
to sustain flight.
If decision is made to stop:
1.

Abort.

If takeoff is con tin ued:1.

Throttle - MIL.

2.

Stores - Jettison (if required).

LOW THRUST ON TAKEOFF OR AT LOW ALTITUDE(NON-AB) I PW200 I

Low altitude, for engine malfunction purposes, isgenerally defined as 10,000 feet AGL or below.
Low thrust can be the result of EEC or UFC related
failures, an rpm rollback, a start bleed strap failing toclose during the normal start cycle, or improper
nozzle position as indicated by a nozzle position
greater than 30 percent at stabilized MIL (approximately 5 seconds). Nozzle position greater than 95
percent at stabilized MAX AB is indicative ofimproper nozzle position which results in low thrustat MIL or below. These situations may result insignificant thrust loss and the inability to take off orsustain level flight. The following emergency
procedure provides the simplest response to a low
thrust problem with a variety of possible causes.
Consider jettisoning stores at anytime if satisfactory
airspeed cannot be maintained. When time andconditions permit, the more extensive procedures in
ABNORMAL ENGINE RESPONSE 1PW200 I, thissection, should be used.
If low thrust occurs during takeoff and conditions
permit, abort. If the takeoff must be continued or if in
any critical phase of flight and MIL thrust is
insufficient, place the EEC BUC switch to OFF. If
thrust is still insufficient, AB should be attempted. If
the AB does not light, the throttle should be placed to
MIL and the EEC BUC switch placed to BUC.
Because the BUC bypasses the UFC and many of its
inputs, engine thrust should increase. If thrust is stillinsufficient, move the EEC BUC switch to OFF andcontinue attempts to light the AB.

""""",,~~~[A] rnrl EEC BUC switch - OFF.
With nozzle confirmed failed open or missing,remain in EEC OFF for the most available
thrust. EEC BUC switch may be positioned to
OFF in MIL or AB.

One low thrust condition requires additional
considerations. If the nozzle is failed open, damaged,
or missing, MIL thrust is low. Thrust should be
sufficient with the EEC off to sustain flight except
when at low airspeeds and/or high GW's. If theexhaust nozzle is confirmed failed open or missing,remain in EEC OFF. In a partial thrust situation,
thrust available may increase as altitude decreases.
250 knots approximates the airspeed at which thrust
required for level flight is the lowest. If thrust is stillinsufficient with the EEC off, consider bypassing theattempt to obtain AB operation since an excessively
open nozzle may reduce the chance for a successful
AB light. Although transferring to BUC with an open
nozzle results in less thrust at MIL than with the
EEC off, air loads may force the nozzle closed andincrease thrust, especially at lower altitudes.
Generally, the more open the nozzle (greater than 50
percent) at BUC selection, the more difficult it is for
air loads to close the nozzle. If the nozzle does not
move toward close after transferring to BUC andthrust is still insufficient, move the EEC BUC switch
to OFF and attempt to light the AB.

If thrust is still insufficient:

If on takeoff and the decision is made to stop:3.

Throttle - MIL.

4.

[A] ~ EEC BUC switch - BUC.

5.

[A] ~ EEC BUC switch - OFF.

6.

Throttle - MAX AB.

7.

Stores - Jettison (if or when required).

1.

Abort.

1.

~

NOTE

If thrust is still insufficient:

2.

Throttle - MAX AB.

If takeoff is continued and/or thrust is insufficient:

IWARNING I• The decision to jettison stores must be
executed at anytime it becomes necessary to insure sufficient flying time isavailable to complete actions designed
to restore usable thrust.
• Failure to monitor sink rate and height
above terrain while applying low
thrust recovery procedures can resultin an ejection outside the seat performance envelope.

When time and conditions permit:

8.

Refer to ABNORMAL ENGINE RESPONSE

I PW200 I, this section.

~~~ ~ ~ ~ ~ ~ 6 ' 6 ' 6 ' ~h;;; ~-~-;;a~LOW THRUST ON TAKEOFF OR AT LOW ALTITUDE(NON-AS) LPW220

NOTE

In a partial thrust situation. thrustavailable may increase as altitudedecreases. 2,)0 knots approximates theairspeed at which thrust required forlevel flight is the lowest.

Low altituck. for engine malfunction purposes, IS
gpnerally defined as 10,000 feet AGL or below.
Low thrust can be the result of DEEC-related
failures; a failed open, damaged, or missing nozzle; oran engine rpm rollback. A failed open, damaged, or
missing nozzle may result in significant thrust loss
and the inability to takeoff or maintain level flight.
For a description of failed open, damaged, or missing
nozzle, refer to NOZZLE FAILURE I PW2201, thissection. Low thrust can also be the result of the start
bleed strap failing to close during the normal start
cycle.

3.

If nozzle is failed open, damaged, or missing:
4.

If low thrust occurs during takeoff and conditions
permit) the takeoff should be aborted. If the takeoff
must be continued or in any critical phase of flight andMIL thrust is not sufficient, AB should be used. An
excessively open nozzle may reduce the chance forsuccessful AB light. If the AB does not light (allow the
DEEC to automatically resequence the AB if conditions
permit), place the ENG CONT switch to SEC.
If an automatic transfer to SEC occurs or SEC isselected manually, resulting thrust is 80-100 percent
of normal MIL thrust with no AB capability. If thrustis still low, consider jettisoning stores.

Airspeed - Climb to arrive at 250 knots ordescend at 250 knots to obtain level flight
above minimum recommended ejection altitude or minimum safe altitude, whichever isappropriate.
NOTE• With a missing nozzle, level flight may
not be attainable above 5000 feet MSL.
• If descent is required, maintain 250

knots with throttle set at 850°C FTIT.
Iflevel flight cannot be maintained by 1000 feet above
minimum recommended ejection altitude or minimum safe altitude, whichever is appropriate:5.

If on takeoff and the decision is made to stop:
1.

Stores - Jettison (if required).

Throttle - As required to maintain 250 knots inlevel flight.

Abort.

If takeoff is continued andior thrust is insufficient:1.

If airspeed drops below 250 knots,

Throttle - AB.
The chances for a successful AB light with thenozzle open more than 30 percent are reduced.

trade altitude to reacquire 250 knots.
Do not descend below minimum
recommended ejection altitude or
minimum safe altitude, whichever isappropriate.

If thrust is still insufficient or AB does not light:•

2.

~ rn:EJ ENG CONT switch - SEC .

\WARNINGJ

6.

Land as soon as possible. Plan a flameout
landing. Refer to FLAMEOUT LANDING, thissection.

ENGINE FIRE ON TAKEOFF

With nozzle loss, catastrophic engine
failure and fire are probable withprolonged high power settings above
850°C FTIT while in SEC.

An engine fire may be indicated by the ENG FIRE

warning andJor OVERHEAT caution lights, high FTIT,
smoke, or fumes. Refer to ENGINE FIRE, this section.

Change 14

~"""'II'II'I"II"I~ F~~~:::~~

6T: 1RETRACT

vVhen desin)d:~~If the LG handle warning light remains illuminatedafter the LG handle is placed to UP, the LG or LG~

door,; are not fully up and locked.

~

1.

Airspeed - 300 knots maximum.

~

2.

LG handle - DN. (Use DN LOCK REL button
ifrequired.l

~

~

If LG

h~n:eAd~e~~: :~er,

selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle"~

remains closed, resulting in higher
than normal landing thrust.

~~~ :f LG;'::S:eOd:::::::l::landing.
~~

C!!:!~!~~Do not cycle LG handle. Damage to LG
or LG doors may result.

~~

Go to ALTERNATE LG EXTENSION, thissection.

4.

~~~ ~;t~eA::L:~~L~:: ~::::

~

con:;

~

2.

~~

IfLG must be raised:knots maxim u m.

GW·- Reduce prior to landing.

1.

LG handle DN LOCK REL button - Depress.

2.

LG handle - UP.

~~

T.O.ILAND CONFlG light is illuminated if
left MLG WOW switch has failed.

~~~fails to the groundposition, the LG handle does not move out of the DN
position. In addition, the T.O.ILAND CONFIG
warning light and touchdown skid control system areaffected. The LG handle may be raised by first
depressing the LG handle downlock release button.

~~ :f :~o::e:e~:~o~

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed, resulting in higher
than normal landing thrust.
After touchdown:
4.

Brakes - Apply after wheels spin up.

Touchdown antiskid protection is notavailable. Landing with feet on the brake
pedals may result in blown tire(s).

If LG does not indicate down:~"

LG handle - DN. (Use DN LOCK HEL button
if requ ired.)

IWARNINGI

~~

~

3.

3-38

Change 13

BLOWN TIRE ON TAKEOFF

Tire failure on takeoff is difficult to recognize and may
not be noticed in the cockpit.
Possible indications of a NLG tire failure include aloud explosion, slight deceleration, vibrations, flying
debris, and at night, a flash or flame. These
characteristics can be mistaken for an engine stall.
Rubber debris may cause damage to the engine, NWS
wiring harness, WOW switch assembly and/or gear
position sensor wiring. NWS may not be available
even though the ARINWS light is on and the NWS
FAIL light is off.

IWARNING

1

Aborting takeoff at high speed with ablown tire may be more dangerous
than continuing takeoff. For heavy
weight takeoffs, an abort at high speedwith a blown tire is extremely dangerous because braking and directional
control are impaired.
NOTE

The decision to take off or abort depends
on the airspeed at the time of the failure,
GW, stopping distance required, andarresting gear availability.

~"""""""""'~~
10. IF-16A-l

Iftakf'offis continued. do not retract the LG, n·duce
G\V if practical, and prepare to land as soon aspractical.
Directional control during stopping is the primaryconcern when aborting with a blown tire. Heavy GW
and high speed aborts place greater demands on thebrakes and tires. This may cause damage to the NWS,
wheels, and struts which may result in loss of
directional control. In addition, heavy differential
braking may result in MLG tire failure.
If aborting with a blown MLG tire, leave antiskid on
to minimize possibility of skidding the good tire. If thewheel with the blown tire does not turn, the antiskid
system switches to the pulsating antiskid mode. Useroll control to relieve pressure on the blown tire andNWS to maintain directional control.
If aborting with a blown NLG tire, hold the nosewheel
off the runway (if able) and use two-point
aerodynamic braking until control effectivenessbegins to decay. Lower the nosewheel to the runway

~~

t()~
~

and immedi:ltely engage I\\VS, If a\Oailable.
maintain directional control. ese aft stick to rQduce
load on the NLG after brakes are applied. 1fN\\:8 is
not available, the aircraft tends to drift right.
Attempt to move to the left side of the runway before
rudder effectiveness is lost and maintain c.rectional
control with rudder and differential bra1{ing. Stop
short of the departure-end arresting cable ifrossible.
The small nosewheel rolling radius with the tire
missing may allow the cable to pass over top of thenosewheel and cause NLG collapse.

A NLG tire failure accompanied by complete tire
separation from the wheel may cause reverse
castering. The conditions for this to occur are theNWS disengaged or inoperative; the nose wheel rim
rolling on a deformable surface (i.e. asphalt); andlateral force applied to the nose wheel from either arudder input or differential braking. If reverse
castering occurs the nose wheel will turn in the
opposite direction of rudder and brake inputs making
it extremely difficult to maintain directional control.

Change 13

3-38.1/(3-38.2 blank) _

~

~
~
~..,.

~

~
...~
~
~
~

~

~
~
~
~
~
~
~~

I

WARNINGI

If a blown NLG tire occurred and NWS
is not avaiiable, it may not be possible toprevent departure from the runway. A
reverse castering effect may occur inwhich the nosewheel moves opposite to
the rudder or differential braking input.

CANOPY WARNING LIGHT ON

If CANOPY warning light illuminates:
1.

If CANOPY warning light remains on:
2.

With a blown tire, avoid centerline
lights .as they may cause wheel damage
and subsequent loss of directional
control. Failure to use full aft stickwith a blown NLG tire may lead towheel failure and directional control
problems.
Stop straight ahead and shut down the engine as soon
as firefighting equipment is available. Do not attempt
to taxi unless an emergency situation exists.

Canopy loss/penetration in flight results in disorientation and may result in structural damage
caused by the canopy striking the aircraft. Due to thepossibility of severe disorientation, vision loss, injury,
or incapacitation at high airspeed, immediateejection may be the only option. Slow to 180 knots orless and check for controllability. Wind blasts up to
180 knots can be coped with by leaning forward anddown behind the glareshield and HUD.

I

Abort.

LG - Do not retract.

2.

Airspeed - 300 knots maximum.

3.

Refer to LANDING WITH A BLOWN TIRE,
this section.

IN-FLIGHT EMERGENCIES
When preparing to activate backup systems which
rely on stored nitrogen pressure to function, consider
the potential for time-related failures and do not
activate the system earlier than required. For
example, the hydrazine mode of the EPU requires
nitrogen pressure to force hydrazine to the EPU. If anitrogen leak exists, turning the EPU on early could
lead to an inability of the EPU to function on
hydrazine. Similarly, alternate LG and hookextension use stored nitrogen pressure. If alternate
LG extension is used early and a nitrogen leak exists,
hydraulic system B could subsequently faiL Such aleak could also result in insufficient pressure to
maintain proper hook holddown force. Since nitrogenleaks are not apparent prior to system activation,
consider their potential existence and activate the

WARNINGI• Arms must be kept close to the body to
avoid letting wind blast pull arms outof the cockpit.

If takeoff is continued:
1.

Go to CANOPY LOSS/PENETRATION IN
FLIGHT, this section.

CANOPY LOSS/PENETRATION IN FLIGHT

If takeoff is not feasible:
1.

Canopy handle - Push outboard.

• HUD glass disintegration can be
expected following a medium to highenergy bird strike with or without
canopy penetration.
Wind buffet increases slightly with increased AOA.
Therefore, if fuel is not criticaL TEF's should be
extended using the ALT FLAPS switch or by placing
the LG handle to DN.
1.

Airspeed - 180 knots maximum.

2.

Seat - Full down.

3.

ALT FLAPS switch - EXTEND.

4.

Land as soon as possible.

DRAG CHUTE DEPLOYED IN FLIGHT [N§ M

If the drag chute is deployed in flight below 190 knots:
NOTE

If the drag chute is deployed below
approximately 190 knots, it will not
break away from the aircraft.
Change 1

,.,.,111111111111'1~'Tr~J"~r~~
~

~
~~

~~~
~~

~~

~

~
~
~
~
~
~~

~
~
~~

IO.lF-16A-l
1.

DRAG CHUTE switch - REL.

systems dependent on the ECS are operational:
If the drag chute does not release:2.
Throttle _ MAX AB.

5.

~

COCKPIT PRESSURE/TEMPERATURE MALFUNC-

liON
Loss of cockpit pressurization could be caused bycanopy seal, air-conditioning system, or cockpit
pressure regulator safety valve malfunctions or ECS
shutdown or failure.
Certain ECS equipment malfunctions result in
temporary shutdown of the ECS. These shutdowns
are more prevalent at high altitude during low speed
flight with high engine thrust settings. An ECS
shutdown is characterized by an oily, smokey smell,followed by loss of cockpit noise and airflow andgradual loss of pressurization. These temporary
shutdowns typically last from 20-45 seconds or, on
occasion, up to 2 minutes. The EQUIP HOT caution
light may illuminate if the shutdown lasts longer
than 20 seconds.

IWARNING I
With the ECS shut down or the AIR
SOURCE knob in OFF or RAM, theg-suit does not inflate m and PBG isdisabled.

corrected by use of the MAN position.

If cockpit pressure altitude exceeds ~ 22,500, LESS
00 27,000 feet, the CABIN PRESS caution lightilluminates.
If the cockpit temperature is excessive and does not".~

respond to AUTO or MAN temperature commands or~

cockpit pressure is lost, proceed as follows:~

1.

OXYGEN -:.

::::t:e::~Flight may be continued below 25,000 feet.

IfECS has failed or cockpit temperature control is notregained:"~ Most AUTO position temperature failures can be~

If cockpit pressure is not regained but all other

~ 4

100%.

~255~::::::t:a:::::~.

AIR SOURCE knob - OFFthen NORM.

~

3-40

Change 13(10-15 seconds),:.

:~: :~~:~:k:::b~O::~ (after cockpit isdepressurized).

NOTE

External fuel cannot be transferred in
OFF or RAM. Consider jettisoningtanks to decrease drag if range iscritical and ECS cannot be turned on
for short periods of time to transfer
fuel.
7.

Nonessential electrical equipment - Off.
NOTE

Ifin VMC and the ADI and HSI are notrequired for flight, the INS should beconsidered nonessential.
8.

Land as soon as practical.

9.

Check for failed essential dc bus(es). Refer to
EMERGENCY POWER DISTRIBUTION, thissection.

EQUIP HOT CAUTION LIGHT

If EQUIP HOT caution light illuminates:
NOTE• Certain ECS equipment malfunctions
result in temporary shutdown of the
ECS and illumination of the EQUIP
HOT caution light.
• An ECS shutdown and EQUIP HOT
caution light illumination for up to 2
minutes can occur either duringextended LG down flight between sea
level and 7000 feet MSL or during
operation above a line from 42,000 feet
MSL at 0.2 mach to 50,000 feet MSL at
0.95 mach. These ECS shutdowns arenormal, but may still require additional action if the EQUIP HOT lightremains on for more than 1 minute.

~""""""""""""'~~~NOTE

If cockpit temperature is pxces8ive,
refer to COCKPIT PRESSURE/TEMPERATURE MALFUNCTION, thissection.
1.

T. C ,F- oA- 1~~

II E(!l'IP HOT ('"utim, light rem ,inS ,'" atl,.,. 1~

minutp:..~4,~onessential

aVi,::::

~

Off.

V~1C

AIR SOURCE knob - Confirm in NORM if
smoke or fumes are not present.

Ifin
and the ADI and RSI are notrequired for flight, the INS should beconsidered nonessential.
5.

~

LESS ~ DEFOG lever - 3/4 forward.

EJECTION

3.

Throttle - 80 percent rpm minimum (in flight).

Refer to figure 3-4.1. Ejection should be accomplishedat the lowest practical airspeed.

F-16 ACES II Ejection Injury Risk Chart
245~~~~r=~~====~~~~~~FFr=~~==~~~~~==~~==~r-=~==~~===4240~~~~~~~--~~~~~--~~~~~~~-1-~~~-~~~~~~~~++++~~

Vi 200

CIl

~
~

::t

~ 180
LLI

3:
> 160

C

oCIl
140

120
103~:::::t:::I==

100

200

300

~

Land as soon as practical.

2.

o

~~~

400

500

AIR SPEED (KTS)

600

700
1F16X-1-4023X@

Figure 3-4.1.

Change 13

3-40,1/(3,40,2 blank)

I ~~
~

~
~
~
~
~
~
~
~
~~

_\

~"""""""""'':~
IO.lF-16A-l

IWARNINGI• The minimum altitude obtained fromEJECTION SEAT PERFORMANCE
charts, Section I, does not include any
allowance for pilot decision making,changing flight conditions, or hand
movement from the stick and/or
throttle to the ejection handle. Therefore, minimum altitude for ejection
decision could be significantly higher.
• When in a spin/deep stall or otheruncontrolled flight, eject at least 6000
feet AGL whenever possible. This is
the minimum altitude to initiate
ejection with minimal risk of injury
under the most adverse conditions.
The decision to eject must have been
made prior to this altitude. Delaying
ejection below this altitude may result
in serious injury or death.
• Under controlled flight conditions,eject at least 2000 feet AGL wheneverpossible. If below 2000 feet AGL,
attempt to gain altitude if airspeed
permits. Do not delay ejection below
2000 feet AGL for any reason which
may commit you to unsafe ejection.
• Failure to monitor sink rate and height
above terrain while performing anairstart or applying low thrust recovery procedures can result in an ejection
outside the ejection seat performanceenvelope.
• The ACES II ejection seat was designed
for body weights in the 140 to 211 pound
range. There are additional ejection
injury risks associated with bodyweights outside this range.

~~s• For body weights less than 140
pounds, limb flailing, less seatstability, and more severe drogue
chute opening shock (ejection modes
2 and 3) are concerns. The risk of
injury associated with limb flailingand drogue chute opening shockincreases for ejection above 420
knots. This injury risk also increases
as body weight decreases below 140~.....
...~S
S~.....
~

pounds.

S~• For body weights greater than 211
pounds, limb flailing, seat structural
failure, and parachute landings areconcerns. The risk of injury from limb
flailing is high for ejection above 400
knots. The seat leg braces frequently
deform during ejections above 500
knots; this deformation has led toseat side panel failures (and unsuccessful ejections) during 600 knot
ejection tests. The risk of injury
during parachute landing is threetimes the average. These injury risks
also increase as body weight In-

~

S

~
~

b
• w:::a:::s:::: :~:r::~d::m force on
the body up to 400 knots, severe forces~-sS

causing flailing and skin injuries
between 400-600 knots, and excessive
force above 600 knots.

S• During high altitude ejections (mode3), automatic pilot/seat separation andrecovery parachute deployment occur
between 16,000-14,500 feet MSL. Ifhigh terrain is a factor, manual seat
separation procedures must be used to~~

To eje:~;::::;e::i:nm::~cd:::::::e~ two-handed Ss
grip with thumb and at least two fingers of eachhand. Pull up on handle and continue holding until
pilot/seat separation. The ejection handle does notseparate from the seat.

~~~~

Refer to figure 3-5 for manual seat separation andmanual survival equipment deployment.

~~

Change 133-~~

{~~~"fff'ff""ff'ff'"
~~

~~

~

~~

~
~

~
~
~
~
~
~

Manual Survival Equipment Deployment!
Manual Seat Separation
A. IF EMERGENCY OXYGEN FAILS TO ACTIVATE AUTOMATICALLYUPON EJECTION, PULL THE EMERGENCY OXYGEN GREEN
RING LOCATED NEAR THE LEFT HIP.

~~."
;0 .,",:.~J."';' " ,-=~VISOR AND DISCARD OXYGEN MASK. IF SURVIVAL KIT DOESNOT DEPLOY AUTOMATICALLY, GRASP KIT RIPCORD HANDLE
WITH RIGHT HAND AND PULL. KIT RIPCORD HANDLE IS LOCATEDNEAR RIGHT HIP.

~11~~~~)rjfL

B. AFTER RECOVERY PARACHUTE ASSEMBLY DEPLOYMENT, RAISE,J-~~r6li~ii
~

I""

I

C. LlFERAFT INFLATION IS INITIATED WHEN THE DROP LlNE/
LANYARD IS FULLY EXTENDED AFTER SURVIVAL KITDEPLOYMENT. CHECK LlFERAFT AND IF NOT INFLATED,SNATCH PULL DROP LINE/LANYARD TO INFLATE.

c.

\~
J,~,,:I

n

~l

NOTE
If the survival kit is deployed after landing inwater, a snatch pull on the drop line/lanyard (nearCO 2 bottle) is required to inflate the liferaft.

MANUAL SEAT SEPARATIONTO PERFORM MANUAL SEAT SEPARATION AND DEPLOY THE RECOVERY PARACHUTE ASSEMBLY, PULL THE EMERGENCY MANUAL CHUTE HANDLE.

I WARNING

I• After ejection, the EMERGENCY MANUAL CHUTE
handle should only be used if the automaticsequence has failed or if high terrain is a factor.
Pilot/seat separation in modes 1 and 2 should
occur rapidly after pulling the ejection handle.
If the pilot has time to realize seat separation
has not taken place, a failure has probably
occurred and manual seat separation should be
performed. In mode 3, pilot/seat separation
occurs between 16,000-14,500 feet MSL. If
automatic pilot/seat separation does not occur
in this altitude range or if high terrain is a factor,
manual seat separation must be performed.
• Failure to fully pull the EMERGENCY MANUALCHUTE handle may result in ballistically deploying
the recovery parachute assembly without releasing
the lapbelt and inertia reel straps and unlatching
the seat pan lid.
• Do not attempt to open the lapbelt. If the lapbelt is opened, the seat will partially fall away.
The parachute risers remain attached to theinertia reel straps and the survival kit is retained
under the latched seat pan lid. The only way toseparate from the seat is to pull the EMERGENCYMANUAL CHUTE handle at least 6 inches.

~I~j,~·'11SURVIVAL EQUIPMENT (TYPICAL)

Figure 3-5.

3-42Change 1

1F-16X-1-0023D@

IIIIIIIIIIII~~~­
Eiection (Immediate)

IWARNINGI
1.

Ejection handle - Pull.
If canopy is jettisoned or manually
released/opened after pulling theejection handle, the ejection seatfunctions immediately after canopy
separation. Be prepared to immediately put arm back in ejection position
when the canopy starts to separate.

Eiection (Time Permitting)If time permits, descend to avoid the hazards of highaltitude ejection. Stow all loose equipment and direct
the aircraft away from populated areas. Sit with headagainst headrest, buttocks against back of seat, andfeet on rudder pedals.
1.

IFF MASTER knob - EMER.

2.

~ AIFF MASTER knob - EMER.

1.

Canopy - Open normally.

2.

Canopy - Jettison.

EMER is automatically selected upon seatejection.
3.

I WARNING I

a3 MASTER ZEROIZE switch (combat staPulling the CANOPY JETTISON Thandle other than straight out may
cause the handle to jam. If the
CANOPY JETTISON T-handle ismounted so that the words CANOPYJETTISON engraved on the T-handle
are upright, then an underhand grip
should be used. If the CANOPYJETTISON T-handle is mounted sothat the words CANOPY JETTISON
are inverted, then an overhand or
underhand grip may be used.

tus) - ZEROIZE.
4.

LESS ~ Secure voice (if installed) (combat
status) - ZEROIZE.

5.

Loose equipment and checklist - Stow.

6.

Lapbelt and helmet chin strap - Tighten.

7.

Night vision devices - Remove (if appropriate).

I WARNING I
Failure to remove night vision goggles(NVG) prior to ejection may cause
serious injury. If unable to remove
NVG, a proper ejection body position(head back against the seat headrest)
reduces the chance of injury from theNVG.
8.

Visor - Down.

9.

Throttle - IDLE.
Slow to lowest practical airspeed.

10. Assume ejection position.
11.

Ejection handle - Pull.

3.

MANUAL CANOPY CONTROL
crank - Push in and rotate ccw.

hand-

I WARNING I
Use of the CANOPY JETTISON Thandle or MANUAL CANOPY CONTROL handcrank may result in serious injury. To minimize chances of
injury, immediately release the handle
when the canopy starts to separate.
Eiection Seat Failure

If the ejection seat fails to function after the ejection
handle is pulled and the canopy has separated fromthe aircraft, there are no provisions designed into theescape system for manual bailout.

Failure of Canopy To Separate
DITCHING

If canopy fails to separate, remain in position forejection while keeping arms inboard and perform thefollowing:

Ditch the aircraft only as a last resort. All attempts toeject should be accomplished prior to ditching.

-

~

'"

~

,

IIIIIIIIIIIIII~ELECTRICAL SYSTEM FAILURES

2.

EPU switch - ON (if EPU run light is off).

Electrical system failures are indicated by illumination of the ELEC SYS caution light and one or more
ELEC control panel lights (in any combination). Afteraccomplishing the appropriate emergency procedures, refer to EMERGENCY POWER DISTRIBUTION, this section, to determine inoperativeequipment for any remaining ELEC control panel
lights.

3.

MAIN PWR swi tch - BATT, then MAIN PWR.
Verify that switch is returned to MAIN PWR.

If MAIN GEN light goes off:4.

EPU switch - OFF, then NORM.

5.

ADI - Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to TOTALINS FAILURE, this section.

FLCS PMG Failure

If all four branches of the FLCS PMG fail in flight, theFLCS PMG light illuminates. The converter/regulator automatically selects the power source with thehighest voltage from the available alternate sources.
Other FLCS power sources are the main generator,
the EPU generator, the EPU PMG, the aircraft
battery, and the FLCS batteries.

IWARNING@ID If only AUX flag is in view, pitch androll attitude information is likely to be
erroneous due to INS autorestart in
the attitude mode when other thanstraight and level, un accelerated
flight conditions existed.

NOTE

Since the hydraulic system A pumpand the FLCS PMG share the same
shaft, a failure of that shaft results inloss of both systems.

I

6.

SERVO ELEC RESET switch - ELEC.
Resets LEF's (if locked) and LE FLAPS,
CADC, and ADC caution lights (if on).

7.

I PW200 I ~ ~ EEC BUC switch - OFF, then

IfFLCS PMG light illuminates:
1.

Land as soon as practical.

EEC.

Main Generator Failure (in Flight)

Main generator failure is indicated by ELEC SYS
caution, MAIN GEN, EPU run, and EPU AIR lights.
Additional lights such as HYDRAZN, LE FLAPS,DUAL FC FAIL, AVIONICS, and I PW200 I EEC may
also illuminate. The EPU powers the EPU generator,and electrical power to the nonessential ac and dc
buses is lost.
If MAIN GEN light illuminates:

IPW200 I EEC stall protection may belost. Do not retard throttle below MIL
until subsonic. Set throttle at
midrange prior to cycling EEC BUC
switch.
8.

I PW2201 [!] rn!J AB RESET switch - AB RESET,
then NORM.

1.

AOA - 12 degrees maximum (200 knots
minimum).

I WARNING I
LEF's may be locked un til reset.
Exceeding 12 degrees AOA reduces
departure resistance. Limit rolling
maneuvers to a maximum bank angle
change of 90 degrees and avoid rapidroll rates.

I PW2201 DEEC stall protection may belost. Do not retard throttle below MIL
until subsonic.
9.

Land as soon as practical.

10. Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINE

////////////~~~
If MAIN GEN light remains on or comes on again:

Mter landing and aircraft is stopped:

4.

11.

Chocks - Installed (or parking brake set).

12.

EPU switch - OFF.

ADI - Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to TOTALINS FAILURE, this section.

I

WARNINGI~ If only AUX flag is in view, pitch androll attitude information is likely to be
erroneous due to INS autorestart in
the attitude mode when other thanstraight and level, un accelerated
flight conditions existed.
5.

6.

SERVO ELEC RESET switch - ELEC.
Resets LEF's (if locked) and LE FLAPS,
CADC, and ADC caution lights (if on).

I PW200 I [B] ~ EEC BUC switch - OFF, thenEEC.

I PW200 I EEC stall protection may belost. Do not retard throttle below MIL
until subsonic. Set throttle at
midrange prior to cycling EEC BUC
switch.
7.

I PW2201 rru rn:tl AB RESET switch - AB RESET,
then NORM.

I PW2201 DEEC stall protection may belost. Do not retard throttle below MIL
until subsonic.
8.

Land as soon as possible.

9.

Refer to ACTIVATED
LEAK, this section.

If chocks are not installed, be preparedto immediately engage the parking
brake ifit disengages when the EPU isshut off.
13. MAIN PWR switch - MAIN PWR (until
chocks are installed).
Main and EPU Generator Failure

Failure of both the main and EPU generators isindicated by illumination of the MAIN GEN lights
without an EPU run light. Other indications include
loss of all avionics, ADI OFF and AUX warning flags,
and uncontrollable cold airflow into the cockpit or
reduced airflow to the cockpit if the water separator
coalescer freezes up. The caution lights which
normally illuminate for a failure of just the main
generator (e.g., AVIONICS, LE FLAPS, IPW200 IEEe(EEC faults) do not illuminate, since they requireessential bus power to operate.
The EPU generator may be inoperative for several
reasons, two of which may be remedied from the
cockpit. If the EPU AIR light is off, the EPU may nothave received an automatic start command; manually turning the EPU on may correct this failure. If
the EPU GEN, EPU AIR, and HYDRAZN lights are
illuminated but the EPU run light is off, the EPU may
be underspeeding. If the EPU PMG light is on or
blinking, EPU speed is very slow. The underspeed
could be caused by failure of hydrazine to power the
EPU in conjunction with a low thrust setting and may
be corrected by advancing the throttle. If the EPU
generator operates, refer to MAIN GENERATORFAILURE (IN FLIGHT), this section.

EPU/HYDRAZINE

10. Ifhydrazine depletes or EPU run light goes off
at low thrust - Go to ABNORMAL EPU
OPERATION, this section.

If the EPU generator is still inoperative and the main

generator fails to reset and the FLCS PMG lightremains off, the primary concern is aircraft battery
life for communications, brakes, hook, and IB.E] IfmIdrag chute.

~
~
~
~

~

"

~
~
~
~
~
~
~
~

-~
-

TO.1F-16A-1

~
~

~
~
~

If the FLCS PMG light illuminates when the MAIN

GEN and EPU GEN lights are on, there are threepossible sources of FLCS power (EPU PMG, aircraft
battery, and FLCS batteries). If the EPU PMG, ACFT
BATT TO FLCS, and FLCS BATT lights are off, theEPU PMG is supplying power to the FLCS whether
the EPU run light is on or off.

1.

lights are on, the aircraft battery and/or FLCS
batteries are supplying power to the FLCS. The
ACFT BATT TO FLCS light may come on first. The
FLCS BATT lights may come on individually and may
be intermittent until FLCS and aircraft battery
voltages equalize. The FLCS BATT lights thenremain on steady. The FLCS batteries and aircraft
battery supply power for FLCS operation for
approximately 55 minutes after total generator
failure. As the FLCS batteries continue to deplete, theflight controls become increasingly unresponsive anduncommanded maneuvers occur with gradually
increasing severity.

AOA - 12 degrees maximum (200 knots
minimum).

IWARNING

1

LEF's may be locked and departure
susceptibility may be increased. Near
19 flight, 200 knots should keep AOA
less than 12 degrees. Limit rolling
maneuvers to a maximum bank angle
change of 90 degrees and avoid rapidroll rates.

If the ACFT BATT TO FLCS and/or FLCS BATT2.

EPU switch - ON (ifEPU run light is ofD.

3.

Climb if necessary.

4.

Throttle - As required to extinguish the
HYDRAZN light.

If EPU GEN light goes off:I

WARNINGI

Imminent loss of electrical power tothe FLCS is indicated by increasingly
degraded flight control response anduncommanded motions. Total loss of
FLCS power results in a pitchingmotion and complete loss of control.
NOTE~

If MAIN GEN and EPU GEN lights illuminate:landing configuration and near final
approach airspeed, the pitchingmotion is gradual and in the noseup

5.

If EPU GEN light is still on:

6.

'~46~

Change 11~~

MAIN PWR switch - BATT, then MAIN PWR.
Verify that switch is returned to MAIN
PWR.

If MAIN GEN light goes off:

7.

SERVO ELEC RESET switch - ELEC.
Resets LEF's (if locked) and LE FLAPS,
CADC, and ADC caution lights (if on).

8.

IPW200 I EEC BUC switch - OFF, then EEC.

~ The A:::~:f::;~ l:::~::::::::attery voltage

less than 20 volts. Brake operation is doubtful during
total generator failure with the ACFT BATT FAIL
light on. As aircraft battery voltage continues todecrease, the capability to operate the brakes and,lower the hook, and [N;] M deploy the drag chute is
lost. Lower the hook early since significantly higher
battery voltage is required to lower the hook than isrequired to keep it fully extended. Once lowered, thehook remains full down well past the point at which
the brakes are lost. An approach-end arrestment isrecommended, if conditions permit, because it isdifficult to ascertain brake operation. Relative
intensity of the warning and caution lights is not a
positive indication of battery voltage level.

Go to MAIN GENERATOR FAILURE (IN
FLIGHT), this section.

IPW200 I EEC stall protection may belost. Do not retard throttle below MIL
until subsonic. Set throttle at
midrange prior to cycling EEC BUC
switch.
9.

I PW2201 AB RESET switch - AB RESET, thenNORM.

I PW2201 DEEC stall protection may belost. Do not retard throttle below MIL
until subsonic.

""""",,~~~10.

EPU switch - OFF, then NORM.

11.

Land as soon as possible.

12.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

If MAIN GEN and EPU GEN lights remain on orcome on again:

I WARNING I• Plan to land within 30 minutes toinsure adequate electrical power forcommunications, brakes, hook, and ~INQ] drag chute.
• If the FLCS PMG and EPU PMG lights

are on in combination with the ACFT
BATT TO FLCS or one or more FLCS
battery lights, the aircraft battery or
FLCS batteries are powering theFLCS. With the aircraft battery
powering the FLCS in addition to thebattery bus, approximately 55 minutes flight time is available. Communications, brakes, hook, and ~ INQ]
drag chute are not available after
depletion of the aircraft battery.
• When the FLCS is powered bybatteries, remain alert for degraded
flight controls. At the first indication of
degraded response, reduce airspeed
and climb to safe ejection altitude.
Eject prior to complete loss of control.
7.

HOOK switch - DN.

8.

Minimize UHF transmissions.

If conditions permit:
9.

Land as soon as possible.
Fly airspeed for 11 degrees AOA approach
using fuel state when power was lost.

10. LG handle - DN (use DN LOCK REL button).

I WARNING I
If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle idle
area reset is not available resulting inhigher than normal landing thrust.

11. ALT GEAR handle - Pull (190 knots maximum).
• Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.
• WHEELS down lights and T.O.ILAND CONFIG warning light function are inoperative.
Monitor LG handle warning light to verify
that LG is down.

• NWS is not available following alternate LG extension.
• Do not depress the ALT GEAR reset
button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.
• Pulling the ALT GEAR handle withnormal system B hydraulic pressure
may result in system B hydraulicfailure within 15 minutes.
12. Consider an approach-end arrestment, if
conditions permit. Refer to CABLE ARRESTMENT, this section.

I WARNING I
As the aircraft battery depletes,
brakes may become inoperative withno cockpit indications. The hookremains fully extended well past the
point at which the brakes are lost. If
the HYD/OIL PRESS warning light ison, PTO shaft failure may be indicated.
If aircraft battery power is sufficient,braking is available using the brakel
JFS accumulators only.
13. Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINE

After landing:
14. Stop straight ahead and have chocks installed(or engage parking brake).
15. MAIN PWR switch - MAIN PWR (until
chocks are installed).

-EPU Malfunctions
UNCOMMANDED EPU OPERATION

Failures can occur which allow engine bleed air tospin the EPU turbine even though the EPU has notbeen commanded on. This may not be apparent if thethrust level and amount of bleed air are such that the
EPU is turning above or below the speed range whichturns on the EPU run light. During uncommanded
EPU operation on bleed air, EPU speed varies directlywith throttle position. High thrust settings are likely
to result in EPU failure.
The EPU may also activate for reasons not apparent
to the pilot (momentary main generator loss).
Although this is not an EPU malfunction, it may beinterpreted as uncommanded EPU operation.
If uncommanded EPU operation on bleed air issuspected:
1.

Throttle - Minimum practical thrust.

2.

Stores - Jettison (if required).
Only if required to maintain low thrust.

3.

Land as soon as possible.

If EPU is running with normal indications:

4.

EPU - Leave running.

5.

Land as soon as possible.

6.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINEABNORMAL EPU OPERATION

Abnormal EPU operation after a normal startcommand is indicated by one or more of thefollowing: EPU run light flashes, indicating EPU
operation in the tertiary speed control mode; EPU run
light does not come on or goes off after initial
illumination, indicating sustained underspeed oroverspeed operation; EPU HYDRAZN light does not
go off or EPU fuel quantity continues to deplete when
the throttle is advanced to assure adequate bleed air,indicating an EPU bleed air or hydrazine fuel control
system problem.
When tertiary speed control is functioning, the EPU
run light alternately cycles on and off (one to threetimes per second) as a function of EPU speed
fluctuating between the normal and slightly above

normal speed ranges. The MASTER CAUTION and
the EPU HYDRAZN lights may illuminate. Hydrazine use occurs regardless of available engine bleed
air. The hydrazine supply depletes in approximately
10 minutes. After hydrazine depletion, the EPU
continues to operate with available bleed air on
tertiary speed control.
If tertiary speed control cannot control EPU speed(constant overspeed), total EPU failure is imminent.
Excessive voltage generated by the EPU generator or
its PMG is regulated by the converter/regulators
before going to the FLCS inverters. Electrical andavionic equipment other than the FLCS inverters
may be damaged by the overvoltage condition. WhenEPU failure occurs, FLCS power is provided by the
FLCS PMG, main generator, aircraft battery, or the
FLCS batteries (whichever has the higher voltage).
Under some failure conditions, hydrazine may not be
available to the EPU or it may continue to deplete
even with adequate bleed air. If a hydrazine
malfunction or depletion occurs, landing must be
accomplished using an engine thrust setting
sufficient to maintain an adequate bleed air supply to
the EPU. Failure to maintain minimum engine rpm
can result in hydraulic pressure fluctuations orelectrical bus cycling. Advance throttle as required tomaintain adequate bleed air supply. (If the EPU is thesole source of hydraulic pressure, the throttle should
be set so as to keep the EPU run light on during mild
flight control inputs.) This action may result in athrust level that is higher than required for a normal
straight-in approach. Fully open speedbrakes or ashallower than normal approach may be required. A
straight-in approach followed by an approach-end
arrestment is recommended.
IfEPU was turned on for an ACFT BATT FAIL light:
1.

EPU switch - OFF, then NORM.

2.

Land as soon as possible.

3.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINE

If EPU was activated for other reasons:1.

Throttle - As required (75-80 percent rpm).
Keep thrust high enough to assure adequate
bleed air if EPU fuel usage continues above
80 percent rpm or ifEPU run light is flashing.
If EPU fuel is depleted or if EPU run lightgoes off at low thrust, set throttle to keep EPU
run light on.

•

~"""""'~"""""
lC' 'F-16:·.-1

EPU Fl7EL quantity - \Ionitor.
3.

lll\minen:~ ~:C~i:}Land as soon as possible.
Make an approach-end arrestment, if practical, if EPU fuel depletes before landing or if
EPU run light goes off at low thrust settings.
Refer to CABLE ARRESTMENT, this section.

~:
pmm tothe FLCS is indicated by increasindy
degraded flight control response anduncomrnandpd motions. Total lo~~ of
FLCS power re~u lts ina pitchingmotion and complete loss of control.

I WARNING I

,
,

NOTE
If total loss of FLCS power occurs in thelanding configuration and near final
approach airspeed, the pitchingmotion is gradual and in the noseup
direction for all configurations.

If PTO shaft or both hydraulic systems

are failed, underspeed of the EPU
results in loss of control. Do not retard
throttle completely to IDLE until after
touchdown.

If one FLCS BATT light comes on in flight:
1.

If EPU underspeeds, electrical bus

cycling may affect brake operation. Fora missed engagement, attempt CHAN
1, then CHAN 2 brakes. If no braking
is available, consider going around foranother engagement or making adeparture-end arrestment. The parking brake still operates.
4.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINEFLCS Battery Discharge

FLCS battery discharge is indicated by illuminationof one or more FLCS BATT lights. FLCS batteries
may supply power to the inverters (discharge) evenwhen other FLCS power sources appear to be
operating normally. If more than one FLCS BATT
light comes on, the EPU should be started to possibly
provide an alternate FLCS power source. If all FLCS
BATT lights come on, the EPU must be turned on
immediately, regardless of its apparent status. If the
FLCS BATT lights do not go off, remaining flight time
may depend on the charge state of the FLCS
batteries. Inverter output may be monitored bymoving the FLCS PWR TEST switch to TEST. A
nonilluminated FLCS PWR light indicates that
inverter output is bad and that the associated brakein that channel is inoperative. Also the P, R, and Y
lights illuminate when power fails in at least one
FLCS branch. The DUAL FC FAIL warning light also
illuminates if power fails in two or more branches.

Land as soon as practical.
The FLCS PWR TEST switch should be used
to determine which brake(s) may be inoperative. A nonilluminated FLCS PWR lightindicates an inoperative brake in the channel
indicated.

If more than one FLCS BATT lights corne on in flight:
1.

EPU switch - ON.

If FLCS BATT lights go off:
2.

Land as soon as practical.

3.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINE

If FLCS BATT lights remain on or the EPU runs
abnormally:

2.

Airspeed - 200-250 knots (6-8 degrees AOA)
and climb.
Climb to the highest practical altitude until
descent for landing is required.

\WARNING

1

Land as soon as possible.
The FLCS PWR TEST switch should be used
to determine which brake(s) may be inoperative. A nonilluminated FLCS PWR lightindicates an inoperative brake in the channel
indicated.
Change 1

~
~
~
~
~
~
~
~

I~

At higher airspeeds, the pitchingmot~,-:;n after loss of FLCS power could
be excessive and interfere with theability to eject.
3.

,

3-49

~,

_~

T{ef('r to ACTIVATED
LEAK, thi~ section.
"~

EPU/HYDRAZINE~At the fir:c;t indication of uncommanded or degraded
flight control response:

~

fi.

~

~I
~
~
~
"~

~

I

Loss of power to several systems or indicators without
any indications on the ELEC control panel may be theresult of wire harness chafing or the loss of power toone or more ac or dc buses. Refer to EMERGENCY
POWER DISTRIBUTION, this section, to determine
affected buses and equipment. If one item on a bus ispowered, then that bus should be considered
powered. Wire harness chafing can affect numerous
items on more than one bus without causing loss of
power to a bus.
The ELEC CAUTION RESET button on the ELEC
control panel is used to reset a tripped overcurrent
protection unit; however, the unit may not remain
reset if the fault persists.
The buses with resettable overcurrent protection
units are nonessential ac bus No.1 a3 and the nacellenonessential ac bus.
The items with nonresettable overcurrent protection
units are the radar ac bus; stations 3, 5, and 7; a3 andleft and right inlet stations.

2.

~

:f

Consider a gear up landing, refer to LANDING
WITH LG UNSAFEfUP, this section.
If power to the battery ~ buses, LESS ~ bus
is lost after the landing gear has been
extended, the landing gear cannot be raised.

6.

Refer to EMERGENCY POWER DISTRIBUTION, this section.

7.

Land as soon as possible.

If EPU was activated:

8.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINE

Aircraft Battery Failure

Aircraft battery system failure is indicated by the
ELEC SYS and ACFT BATT FAIL lights. The ELECSYS caution light may not be resettable. If the battery
fails and the main generator subsequently fails, only
the rotating main generator PMG may be available tomanually activate the EPU. Turn the EPU on
immediately after battery failure is indicated.
The ACFT BATT FAIL light illuminates only for low
battery voltage while in flight (approximately 20
volts). If a battery cell imbalance occurs in flight andis still present after landing, the ACFT BATT FAIL
light illuminates 60 seconds after WOW.

Land as soon as practical.

If the aircraft battery has failed (and EPU
is om, do not taxi except to clear runway.
Subsequent loss of the main generator
results in loss of all braking, NWS, hook,
radios, and ~ INQl drag chute.

PO::t~r:i::es:::edpower status of electricalbuses.
Refer to EMERGENCY POWER DISTRIBUTION, this section, to determine thepower status of individual buses. If one item
on a bus is powered, then that bus should beconsidered powered.

I
~

~

NOTE

~~

~

Determining the status of the battery~ buses, LESS ~ bus is critical for a~safe recovery of the aircraft.

~ • :f oneEo;~:t:i:::e:t~: ac buses are not powered:~5.

If power is restored:~~~~

ELEC CAUTION RESET button - Depress.
May reset overcurrent protection unites).

1.

Consider a net arrestment, refer to NET
ARRESTMENT, this section.

If net arrestment is not available:

Partial Electrical Power Loss~.

~

Eject.

4.

3-50

Change 13

1.

EPU switch - ON.
NOTE• Ifboth radios become inoperative afteran aircraft battery failure indication,refer to PARTIAL ELECTRICAL
POWER LOSS, this section.
• The ACFT BATT FAIL light may
subsequently extinguish. This should
not be interpreted to mean that thebattery has recharged. It may indicatethat the battery voltage is so low thatthe light cannot remain illuminated.

•, I I I I I I~~~~~~T1!~~
2.

3.

Land as soon as practical.
Refer to ACTIVATED
LEAK, this section.

6.
EPU/HYDRAZINE

If EPU runs abnormally:
4.

5.

EPU switch - OFF, then NORM.
If main generator subsequently fails andaircraft battery power is not available, EPU
turn-on power may not be available.
Land as soon as possible.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINE

Prior to shutdown:
7.

Loose items - Secure.

8.

Canopy - Open.

Emergency Power Distribution

Refer to figure 3-6.

~

"

~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~

IIIIIIIIIIIIII~

"""""""~~~
Emergency Power Distribution~

I ~MAIN GENERATOR FAILED
BUS ASSIGNMENTNONESSAC

INOPERATIVE EQUIPMENTSYSTEM

NO.1
ENGINEMAX POWER Switch

FUEL

Boost and Transfer Pumps

STORESMGTAVIONICS

NO.2

X*
X

ENG FEED KnobX*

FUEL HOT Caution Light

X*

Tank InertingX***

a3 AIM -7/-120*

Stations 3, 5 & 7 - EO, Radar-Guided
Weapons; ECM Pods

X**

Stations 4 & 6 - EO, Radar-Guided Weapons

X

~

DTU

X

~

ECM

X

LESS IBQI ECMX***LESS IBQI FCC

X

FCR

X

REO

X

X*

X

TISL
~ TWS

LIGHTS

LESS~TWSX

FLOOD CONSOLESXFLOOD INST PNL

X

FORM

X

X***
X

TAXI

X*

Chaff Dispenser
X

Seat Adjust
TI Probe Heater*a3 Nacelle dc bus.
**a3 Overcurrent protection panel No.1, LESS a3 ECM power panel.
***LESS a3

X
X

MAL & INDLTS
TestJBRT DIM Feature

OTHER

NONESS DC

X"~MAIN AND EPU GENERATORS FAILED~~(All equipment from sheet 1 plus the following:)

BUS ASSIGNMENTiSYSTEM

INOPERATIVE EQUIPMENT

NO.1
ENGINE

FLIGHT
INSTRUMENTS

ESSDC

ESSAC

I

NO.2

ANTI ICE Switch

X

I PW200i

EEC Caution Light (EEC faults)

X

IpW2201

ENGINE FAULT Caution Light

Fire/Overheat Detect and TestX

HYD PRESS IndicatorsX

NOZ POS IndicatorX

OIL Pressure Indicator

X

ADI

X

Altimeter (ELECT)XXAOA Indicator

X

HSI

X

Turn Needle

X
X

Automatic Forward Fuel Transfer
FUEL FLOW Indicator

X
X

FUEL LOW Caution Lights

X

FUEL Quantity Indicator~I

FLIGHT
CONTROLS

NO.2

X

AOA Indexer

FUEL

NO.1

Tank Inerting

X*

Autopilot

X

X

X

ECA
FLCP (all lights, reset, arm, and self-test
capability)
LEF's

X

X

X

LE FLAPS & ADC Caution Lights

X
X

Speedbrakes
Stick Trim

X

Figure 3-6. (Sheet 2)

",

3-52

, . , . ~~ ~ ~ ~,~~ ~ ~.I ~ ~ ~""""""~~~~
Emergency Power Distribution~
MAIN AND EPU GENERATORS FAILED - CONT(All equipment from sheets 1 and 2 plus the following:)

BUS ASSIGNMENT

NO.1
NAV/COMM

ESSDC

ESSAC

INOPERATIVE EQUIPMENTSYSTEM

NO.2

NO.1
X

[@] AIFF

X[@] AIFF (mode 4)

X

FCNP
LESS 1M!] IFF (mode 4)

X[@][8]X

HF Radio

X

ILS
X***

X

INS

X

LESS 1M!] Secure Voice

X

TACAN

AIM-9

X**
X

ALT REL Button

X

CIU*
Gun

X
X

VHF

[8]

.~

X

IFF

STORES
MGT

NO.2

X
X

X

MASTER ARM Switch*X

MSL STEP Switch

X

ArmlRelease, Stations 1,4,5,7&8

X

X

X

ArmlRelease, Stations 2,3,6&9
LESS M NUCLEAR CONSENT SwitchX

SCP (left half)X

SCP (right halD

X

STORES CONFIG Caution LightX

Stores Jettison (SEL and EMER)*

X

.~

X

* Indicates redundancy.
**GITj Nacelle essential ac bus.
***~~~~~~~;:·;)~~Ia'l-~

T. O. 1 F- 16A- 1

Emergency Power Distribution
MAIN AND EPU GENERATORS FAILED - CONT(}\ll equipment from sheets 1, 2, and 3 plu:-; the following:)
BUS ASSIGNMENTINOPERATIVE EQUIPMENTSYSTEM

ESSAC
NO.1

STORES
MGT(cont)[A]

mIl WPN REL Button~

WPN REL Button

AVIONICS

CADC

ESSDC

NO.2

NO.2

X
X
X
X

CADC Caution Light

X

DTU
~ FCC

X
X

Radar Altimeter

LIGHTS

NO.1

LESS ~TWSX

LESS [8Q] ECMX*ANTICOLLISION Strobe

X

AR (Flood)X[@] and ~ [fill [A] IdentificationX

PRIMARY CONSOLESXPRIMARY INST PNLXLANDINGX

LAND INGfrAXlfExternal Switches

X

X[@] MAL & IND LTS TestlBRT DIM
FeatureLESS ~ MAL & IND LTS TestlBRT DIM

X*

FeatureNAVIFREQ DISPXPOSITION
LGINWS/BRAKES~X

LG Handle Down Permission Button

X

LG Hydraulic IsolationX

LG Sequence (doors)X

LG UP-DN Command

X

Figure 3-6. (Sheet 4)~,5,;;e'~~~1'1'1'1'1'1'1'1

""""",,~~~-

Emergency Power Distribution
MAIN AND EPU GENERATORS FAILED - CONT(All equipment from sheets 1,2,3, and 4 plus the following:)
BUS ASSIGNMENT
ESSDC

ESSAC

INOPERATIVE EQUIPMENTSYSTEM

NO.1

NO.2

NO.1

LGINWSIBRAKES(cont)NWSX

WHEELS Down Lights

X

OTHER

Air Data Probe Heaters (nose & fuselage)X

AOA Probe Heaters

X

NO.2

X

AR System

X

AVTR

X

Battery Charger PowerCABIN PRESS Caution LightXCAMERA/GUN Trigger

X

Cockpit Pressure Dump Capability

X
X

Cockpit Temperature Control
Engine Bleed Air Valves (close capability)

X

EQUIP HOT Caution Light

X

HUD

X
X

HUD/CTVSX

OXY LOW Caution Light[A) [If] LIQUID OXYGEN Quantity IndicatorX~ Probe Heat Monitor

X

PROBE HEAT Switch

X

SEAT NOT ARMED Caution Light

Figure 3-6. (Sheet 5)

X

-

-

Emergency Power Distribution
OPERATING EQUIPMENT - MULTIPLE GENERATOR FAILURE

SYSTEMOPERATING EQUIPMENT

BUS ASSIGNMENTBATTERY ~BATTERY LESS ~NO.1

IpW2001

ENGINE

NO.2

Anti-Ice*
BUC*
EEC (no supersonic stall protection)*UFC*

I PW2201SEC*

ENGINEDEEC (no supersonic stall protection)*

INSTRUMENTS

AirspeedJMach Indicator*
Altimeter (PNEU)*
FTIT IndicatorX

RPM Indicator

X
X

SAl

X

X

X

X

X

Vertical Velocity Indicator*
FUEL

External Fuel Transfer*
FUEL MASTER Switch
FFP*

FLIGHT
CONTROLS

Functional (except LEF's, speedbrakes,
autopilot, and stick trim)*

NAV/COMMIntercom

X

X

X

X

Magnetic Com pass *UHF Radio
LIGHTS

LGINWS/BRAKES

Spotlights

X

X

Utility Light

X

X

X

X

Alternate LG extension*
AntiskidiChannel 1 Brakes
AntiskidiChannel 2 Brakes
LG U plockIDownlock

X
X* Indicates items that do not require power through the battery~ buses, LESS~ bus.

Figure 3-6. (Sheet 6)XX

""""""':~~~
Emergency Power Distribution~

OPERATING EQUIPMENT - MULTIPLE GENERATOR FAILURE - CONTSYSTEMOPERATING EQUIPMENT

BUS ASSIGNMENTBATTERY ~BATIERY LESS ~
NO.1

NO.2

LG/NWSI

MLG WOW (branches A & B)

BRAKES(cont)NLGWOW

X

X

Parking BrakeXX

WARNING
LIGHTSCAUTION
LIGHTSCANOPYXX

DUALFC FAIL

X

X

ENGINEXX

HYD/OIL PRESSX

X

LG Warning (handle)

X

X

ANTI SKID

X

X

IPW2001 BUC

X

X

IPW200i EEC (function ofEEC BUC switch

X

X

ELEC SYSX

X

FLTCONTSYS

X

X

position only does not indicate EEC faults)

HOOK

X

X

MASTER CAUTION

X

X

SEC

X

X

IPW2201

OTHER

X

X

Canopy Activation*~~ Drag Chute

X

EPU

X

Hook

X

X

X

X

JFS

X

X

MAIN PWR Switch

X

X

VMS

X

X

* Indicates items that do not require power through the battery ~ buses, LESS ~ bus.

Figure 3-6. (Sheet 7)

Change 7

3-56.1/(3-56.2 blank)~~~~~~~~~~~~

~
~
~
~
~
~
~
~
~
~
~
~

-

Emergency Power Distribution
PARTIAL ELECTRICAL POWER LOSSNONESSENTIAL AC BUS NO. 1

t----+i

NONESSENTIAL AC BUS NO.2
FLOOD InstrumenVConsole Lights
Seat Adjustment

FCC
REO

ESSENTIAL AC BUS NO. 1

ESSENTIAL AC BUS NO. 2
HSI (off flag)
HYD PRESS (frozen)
OIL Pressure (frozen)
Critical items lost:• ADI, HSI, Fuel Quantity

AOA Indicator (off)FUEL FLOW (frozen)
Critical items lost:• CADC• INS (60-90 sec of use)

• TACANESSENTIAL DC BUS NO. 1

ESSENTIAL DC BUS NO. 2

AOA Indexer
Speedbrakes
VHF Radio
Critical items lost:• LG DOWN Permission
Button IDN LOCK REL
button required)• LG WHEELS DOWN Lights• TACAN

Autopilot
HUD
SEAT NOT ARMED Caution Light
Critical items lost:

IAQJ BATTERY BUS NO. 1• LG UP-DN Command(ALT GEAR extension
required).....~ AIRCRAFT BATTERY[@J BATTERY BUS NO. 2
Intercom*
RPM lindicates < zero)
UHF Radio
Critical items lost:• Channel 2 Brakes andParking Brake• Hook• LG Warning Light(handle)EPU Fuel (indicates < zero)
FTIT (indicates < 200)
SAl (off flag)
Critical items lost:• Channel 1 Brakes• LG WHEELS DOWN Lights

LESS fAQl BATTERY BUS
EPU Fuel (indicates < zero)
FTIT (indicates < 200)
Intercom*
RPM (indicates < zero)SAl loff flag)
UHF Radio

Critical items lost
• Brakes• Hook• LG WHEELS DOWN Lights• LG Warning Light (handle)• Parking Brake*VHF radio is also inoperative because the ability to key either radio is lost.
lE-J6A-J-216JX

Figure 3-6. (Sheet 8)®

Fires can also occur in the exhaust nozzle area whenusing AB. These fires are the result of portions of thenozzle failing which allows the AB plume to burnthrough the nozzle. Ventilation should inhibit
forward movement of the fire into and through theengine bay. Since these fires are aft of the detection
circuit, the ENG FIRE warning light will not
illuminate. Additionally, the nozzle position indications are normal, and there are no vibrations or
instrument fluctuations. In most cases, theseAB-related nozzle fires are detected by someone
outside the aircraft (wingman, tower, etc.). When
operating in AB and a fire is reported at the rear of theaircraft, retard throttle below AB immediately. Thisaction should extinguish a nozzle fire withinapproximately 30-45 seconds and minimize damage
to the aircraft skin, speedbrakes, nozzle, and flightcontrols; however, nozzle damage may result in a
noticeable thrust loss.
A failure that causes an oil leak may also result in an
oil-fed fire in the AB section. The fire may continue forseveral minutes after the engine fails or is shut down
(until the oil supply is exhausted). Since this fire is
likely to be contained within the engine, the ENG
FIRE warning light does not illuminate.
If on takeoff and conditions permit:
1.

Abort.

If takeoff is continued:
1.

Climb.
Maintain takeoff thrust until mInImum
recommended ejection altitude is attained
and then throttle to minimum practical.

2.

Stores - Jettison (if required).

procedures, this section.

At a safe altitude:
Generally, the first indication of fire in the engine
compartment is the ENG FIRE warning light.
Abnormal fuel indications (quantity/flow) may also
be present. FTIT probably will not be higher thannormal. Explosions, vibrations, or engine instrument
fluctuations are usually indicative of a serious engine
problem; engine failure may be imminent. Immediate
action should be taken to reduce thrust to theminimum practical level after attaining safe ejection
parameters. If within gliding distance of a suitablerunway, consider shutting the engine down. Sufficient time should exist to analyze the situation andmake an ejection versus land decision. The ejection
decision should be based on visual and/or cockpit
indications that the fire is persisting. Cockpit
indications include continued illumination of the
ENG FIRE warning light and subsequent illumination of FLCS status lights/degraded flight controls or
subsequent loss of either hydraulic system.
3-58

Change 8

3.

Throttle - Minimum practical.
If fire occurred in AB, ENG FIRE warning
light may not illuminate. Fire should
extinguish after throttle is retarded; however, nozzle damage may result in lower thannormal thrust.

If ENG FIRE warning light goes off:
4.

FIRE & OHEAT DETECT button - Depress.
Determine if fire detection circuit is functional.

If fire persists:
5.

Eject.

If fire indications cease:
5.

Land as soon as possible.

"""""",~~
When airspeed
depressurized:

OVERHEAT Caution Light I PW200 I

Detection of an overheat condition in the engine
compartment, ECS bay, MLG wheel wells, or EPU
bays illuminates the OVERHEAT caution light.
Accomplish as many of the following as required toextinguish the caution light. If the light goes off,
verify the integrity of the detection circuit bydepressing the FIRE & OHEAT DETECT button andland as soon as possible.

7.

Throttle - Minimum practical.

2.

FIRE & OHEAT DETECT button - Depress.
Determine if fire detection circuit is functional.

If OVERHEAT caution light remains on (or detectcircuit checks bad) and EPU is running:
3.

EPU switch - OFF (if feasible).
If the EPU was manually turned on, consider
turning it off to determine if it is the source ofthe overheat condition. If the OVERHEAT
caution light remains on, the EPU should beturned back on.

1

If the EPU was activated for FLCS
BATT lights, be prepared to reactivate
the EPU if the battery discharge lights
reilluminate and remain on for longer
than 15 seconds.

OXYGEN - 100%.

5.

AIR SOURCE knob - OFF.
External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tanks to decreasedrag if range is critical and the ECS cannot be
turned on for short periods of time to transfer
fuel.

I

WARNING

I

With the ECS shut down or the AIR
SOURCE knob in OFF or RAM, the g-suit
does not inflate rn and PBG is disabled.
6.

Descend to below 25,000 feet and reduce
airspeed to below 500 knots.

cockpit

is

8.

_

_
WARNING

I

_

~

,

Nonessential electrical equipment - Off.
NOTE

If in VMC and the ADI and HSI are notrequired for flight, the INS should beconsidered nonessential.
If OVERHEAT caution light still remains on (ordetect circuit checks bad):_

TANK INERTING switch - TANK INERTING
even if Halon is not available.

10. LG handle - DN (300 knotslO.65 mach maximum). (Use DN LOCK REL button if required.)I

WARNINGI

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed, resulting in higher
than normal landing thrust.

11.

~

,

With the ECS shut down or the AIR
SOURCE knob in OFF or RAM, the g-suit
does not inflate rn and PBG is disabled.

If OVERHEAT caution light remains on (or detectcircuit checks bad):4.

and

AIR SOURCE knob - RAM (below 25,000 feet).
External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tanks to decreasedrag if range is critical and the ECS cannot be
turned on for short periods of time to transfer

I

9.

IWARNING

reducedfu~

If OVERHEAT caution light illuminates:1.

isLand as soon as possible.

Engine Vibrations I PW200 I

Some engines exhibit low frequency vibrations which
are non-damaging to both the airframe and engine.
The vibrations should disappear if engine rpm iseither increased or decreased.
Vibrations that change in intensity with throttle
movement and are present across the throttle/rpm
range may indicate a potential engine malfunction.
If vibrations persist:1.

Throttle - Minimum practical.

2.

Land as soon as possible.

_

-

-

~

,
~

,

~

,
,

Oil System Malfunction I PW200 I

An oil system malfunction is characterized by apressure (including fluctuations) below 15 psi at
IDLE or 30 psi at MIL, a pressure above 80 psi at any
thrust setting, pressure fluctuations greater than ± 5
psi at IDLE or ± 10 psi above IDLE, or by a lack of oilpressure rise when the throttle is advanced. The OIL
pressure indicator can be used as an early indicatorof oil loss. An indication of excessive oilloss is the lackof oil pressure rise when the throttle is advanced in
IDLE to MIL range. These conditions may not occur
until approximately one-half the usable oil is lost. The
HYD/OIL PRESS warning light may not illuminate
until most of the usable oil is lost. At the firstindication of an oil system malfunction, take
immediate action to land as soon as possible.

Climbing to a higher altitude allows a higher cruise
and increases glide range. However, if the oil
malfunction is caused by an internal engine oil leak,the rate of oil loss is decreased at low altitudes andlow throttle settings. Usually it is advisable to climb
to a reasonable cruise altitude. Once at altitude,
retard throttle to approximately 80 percent rpm anddo not move the throttle unless absolutely required.
With zero oil pressure, any throttle movement may
cause the engine to seize. Minimize maneuvering g tominimize loads. Plan an approach which allows aflameout landing from any position should engine
seize. Refer to SIMULATED FLAMEOUT (SFO)
LANDING and FLAMEOUT LANDING, this section.
The EPU should be manually activated; otherwise, if
the EPU does not start automatically when theengine seizes, the short time remaining before loss ofcontrol may be inadequate for recognition of the EPU
failure and corrective action. Monitor hydrazine use
after activating the EPU. If consumption rate is too
high, cycle EPU switch to OFF, and then NORM toconserve hydrazine. Be prepared to place the EPU
switch back to ON if the engine seizes.
If an oil pressure malfunction is suspected:
1.

Attain desired cruise altitude.
The rate of oil loss is decreased at low
altitudes and low throttle settings.

2.

Stores - Jettison (if required).

3.

Throttle - Approximately 80 percent rpm.

4.

EPU switch - ON.
Monitor hydrazine use. If consumption rate is
too high, cycle EPU switch to OFF, thenNORM to conserve hydrazine. Be prepared toplace EPU switch back to ON if the engine
seizes.

5.

Throttle - Do not move until landing
assured.

IS• Throttle movement/rpm change may
cause engine seizure.
• Do not start the JFS if engine seizure
has occurred or is anticipated. Starting the JFS may result in no brake/JFS
accumulator pressure for the brakes.
6.

Land as soon as possible.
Plan to fly an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING and FLAMEOUT LANDING, this section.

7.

Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.

EEC Caution Light I PW200 I

The EEC caution light comes on in response to an
EEC failure, electrical power interruption, an EEC
input failure, or detection of an engine overspeed
condition. Positioning the EEC BUC switch to OFF or
BUC also turns the EEC caution light on. If the EEC
caution light is on as a result of an engine overspeed,
the EEC is automatically turned off and cannot be
reset.
If the EEC caution light comes on while supersonic,
retard the throttle to MIL and use speedbrakes as
required to decelerate to subsonic. If the throttle isretarded below MIL at supersonic speeds, inlet buzz
may result which produces severe cockpit vibrations
and a probable stall or stagnation. When subsonic, set
the throttle approximately halfway between MIL and
IDLE (midrange). This normally corresponds to 80-85
percent rpm. If the CADC caution light is also
illuminated, the EEC is not receiving valid machdata. In this case, all EEC functions are working
properly except supersonic stall protection. The
CADC caution light must reset before the EEC
caution light resets. If the lights do not reset, leave
the EEC BUC switch in EEC and avoid supersonicflight. If only the EEC caution light is illuminated,attempt to reset the EEC and clear the light by cycling
the EEC BUC switch from EEC to OFF and back to
EEC. If the light remains on, return the EEC BUC
switch to OFF and land as soon as practical. With the
EEC off or inoperative, engine stall protection
features at MIL and IDLE are lost. Avoid rapid
throttle movements. Do not exceed 88 percent rpm or
use AB unless required to sustain flight. Do not allow
rpm to decrease below 80 percent above 15,000 feet or
70 percent below 15,000 feet to avoid a stall or
stagnation, particularly if the throttle is advanced.

During landing with the EEC off, IDLE thrust is
higher than normal because nozzle idle area reset is
not available. The EEC BUC switch should not be
returned to EEC after landing in an attempt to open
the nozzle and decrease thrust.
If EEC caution light illuminates more than once, it
may indicate an engine problem. The EEC caution
light may be reset, but consideration should be given
to landing as soon as practical.
If EEC caution light illuminates or an EEC
malfunction is suspected while supersonic:
1.

Throttle - Retard to MIL.

4.

IfEEC caution light resets:

5.

IfEEC caution light remains on:
5.

IAllmEEC BUC switch - OFF.

6.

Throttle - As req uired.

7.

Land as soon as practical.
During landing with EEC ofT, idle thrust is
higher than normal because nozzle idle area
reset is not available.

Retarding the throttle below MIL
while supersonic may induce inlet
buzz which produces severe cockpit
vibration and a probable stall or
stagnation.
When subsonic or if the EEC caution light illuminates
or an EEC malfunction is suspected while subsonic:
2.

Throttle - Midrange.
80-85 percent rpm.

FTIT Indicator Failure I PW200 I

If engine rpm is erroneously zero:3.

Go to ZERO RPM, this section.

If CADC and EEC caution lights are on:
4.

5.

SERVO ELEC RESET switch - ELEC.
CADC caution light must reset before EEC
caution light will reset.
EEC BUC switch - OFF, then EEC.

If CADC and EEC caution lights reset:
6.

Continue normal operation.
Repeated illumination of the EEC caution
light may indicate an engine problem.

If CADC and EEC caution lights remain on:

Zero RPM/Erroneous RPM Indication I PMOO I

-~ ~

~

~

streak from the engine; however, the AB should beselected only if required to sustain flight. Hot streak
ignition of the AB may result in a stall or stagnation.
If the engine is shut down, there may be no way torestart it.
If ENGINE warning light is not illuminated and tests
good:Land as soon as practical.

1.

If ENGINE warning light is illuminated or fails to
test:
Throttle - Midrange.
Maintain 700°-750°C FTIT until landing isassured. EEC may be inoperative and EEC
caution light may not be illuminated.

1.

I

WARNINGI• Assume engine alternator is inoperative or malfunctioning. If the engine isshut down, an airstart may not be
possible.
• Select AB only if required for flight.
Use of AB may result in a stall or
stagnation. If engine is shut down, an
airstart may not be possible.
2.

Land as soon as practical.

Abnormal Engine Response I PW200 I
Refer to LOW THRUST ON TAKEOFF OR AT LOW
ALTITUDE I PW200 I, this section.
Improper nozzle position is considered an abnormal
engine response. Improper nozzle position isindicated by a nozzle open greater than 30 percent
during a MIL thrust takeoff or 95 percent during a
MAX AB takeoff; uncommanded nozzle opening
during stable engine operation; abnormal, hesitant,or stepped nozzle movement during engine transients; failure to close after AB cancellation; or failure
to close with throttle advance during go-around after
a low approach. Any of these failures may be causedby the EEC, UFC, or CENC. Selecting EEC OFF or
BUC progressively eliminates these sources offailure.
In general, if abnormal engine response occurs, theEEC BUC switch should be positioned to OFF. The
EEC can be turned off regardless of throttle position.
If satisfactory engine response is obtained, keep the
EEC off and land as soon as practical.

3-62

Change 6

L

thrl

n;'

c,;"

d.

agl'd

'n 1."Jd ;wh,

HI ,\'

EJ' off'

Je ,.
k,\',

lY;.
.,.\ 11

Ie fa. open.
rust \'. t h the

1e~lt

'1St

tl)::lstain·;pc!',,:nei/or :iI. highG\r'..,. If". 'xtl.~lOll.
con f . :' 'd fa ,kd open
or fnis::;il,'i~m;\i"EFi! !\,'F. I! ;: partial thrust
situation ~,)'ust ,; .J :db!t 'iY iri( ~l';i-;e as altitude
dccrease~ ~~;)O In. ; '. april ",·natE'''': ! he airspeed atwhich th!':!.! reqi! t,,·j fin' \,';i.:l t1kht IS tlll> lowest.
MIL thrll';' :'1 BI 't' ith ~ ,. ·z)zzle hiled open morethan 60 pl"'i. •.:nt Pi,,' ,;I:CS ; .• ~ t hrU";I l~l;m required tornaintaini '\ el n \,'j: ( k ; h , :nfigl1 r'l~ ;, If'. ;)000 feetnil'.t ex(MSL, and ',:IJO kn,ll") Mll. i{.rust ill HUC with thenozzle fnil (lpen i~, :q)prOX1!!ldLdy l'(jual to UFC idle
thrust. If the noz:.dl' f<lils t( i cl~)se after transferring to
RUe andn!' if j I'\l !i~i is ~t jll in~1.lmci(>nt, move theEE(~ BU<' .,\vilc!1 I.,j ()FV

If selecti !If.' EEl..' I ; FF dol"; IiI)!, prr ,vide satisfactory
engine response and t.hril~t is insuffll'it~nt, attempt tolight the AB by ~n(jppjllf.: Ltie throttle to MAX AB.
Several attt.'mpb to O;)l<lln J ;;uccessful AB light may
be made. Chances of a ::iw:cessful AB light with a fullopen nozzle are minimal. /\.fter light-off, throttle
mnvemeT',t in AB i:,:, alh\v(IdAn UFC malfunct inn i:i indicated by insufficientthrust or abnormal engine response with normal
nozzle position (less than 30 percent) and EEC off.
Cockpit indications at MIL may show high rpm and
FTIT with extremely low thrust (nozzle is open at
MIL with EEC on and is closed with EEC om. If thisoccurs, transfer to BUe, if the BUe caution lightremains off (transfer does not occur), advance throttle
until transfer o('cur~;A throttle link<1ge prohldn should be suspected if
throttle movements in bot h UFC and BUC produce
either no rpm change ur an rpm increase but no rpm
decrease. In either ca~e, the OFF position does notshut down the engine. If the throttle is stuck andthrust is ,;uitablt' for su~tainpd t1ight~ attempts to free
the throttle shcnld be delayed until within gliding
distance of a suitable landing field. If throttle is stuckor otherwise prevelltEd from normal movement,
controi might be regained by depres~i:lg the cutoff
release, rotatmg t ht' throttle outboard, andapplying necessary foret'.
If thrust is too Inw to sustain level flight, turnimmediately toward t hl' r:earest suitable runway andestablish 250 knots v·rhiel! approximates the airspeed
at which thrust rpquired for level flight is the lowest.
Consider jetti~ii)njng stores to increase flying time
available to complete t1nior.s designed to restore
usable thrust and impron· range in the event thoseactions are unsuccessfuLNOTE
In a partial thrust situation, thrustavailable may increase as altitudedecreases. 250 knots approximates theairspeed at which thrust required forlevel flight is the lowest.
If thrust is too high to permit a safe landing, use
excess thrust to climb and maneuver toward thenearest suitable airfield. Once high key for a flameout
landing is assured, follow procedures as outlined in
FLAMEOUT LANDING, this section. Activate the
JFS and EPU and then shut down the engine as soon
as landing is assured (normally high key) by placing
the throttle to OFF or, if necessary, by placing the
FUEL MASTER switch to OFF.

With EEC off. stall protection at IDLE
and MIL is lost. lVlaintain 80-88
percent rpm above 15,000 feet and
70-88 percent rpm below 15,000 feet
until landing is assured.
If thrust is insufficient:

2.

NOTE
In a partial thrust situation, thrustavailable may increase as altitudedecreases. 250 knots approximates theairspeed at which thrust required forlevel flight is the lowest.

NOTE
During landing with the EEC off, idle
thrust is higher than normal. The EEC
BUC switch should not be moved after
landing in an attempt to open thenozzle and decrease thrust since thismay result in reoccurrence of theoriginal malfunction.
If abnormal engine response occurs:

I WARNING I

I• Failure to monitor sink rate and height
above terrain while applying low
thrust recovery procedures can resultin ejection outside ejection seat performance envelope.
• If the throttle is stuck and thrust issuitable for sustained flight, attempts
to free the throttle should be delayed
until within gliding distance of asuitable landing field.

• Jettison stores when necessary toincrease flying time available toaccomplish actions designed to restore
thrust.
1.

~ ~ EEC BUC switch - OFF.

With nozzle confirmed failed open or missing,remain in EEC OFF for the most available
thrust. The EEC BUC switch can bepositioned to OFF regardless of throttle
setting.

Airspeed - 250 knots (if thrust is too low tosustain level flight).

3.

Throttle - MAX AB.
Several attempts may be made to obtain asuccessful AB light. Chances of a successful
AB light with a full open nozzle are minimal.
After light-off, throttle movement in AB isallowed.

If thrust is still insufficient or if serious engine
problem exists:

4.

Throttle - In MIL to BUC IDLE range.

5.

[8J ~ EEC BUC switch - BUC (in limits).

If transfer does not occur, advance throttle
until BUC caution light illuminates.

• If nozzle is open greater than 50
percent while in BUC, there may not be
sufficient aerodynamic loading for thenozzle to close. To increase aerodynamic loading, decrease altitude,
increase airspeed above 250 knots, andincrease throttle setting.

• Retel to Section V for BUC operating
limits.
• Maintain 80 percent rpm minimum
above 15,000 feet and 70 percent rpm
minimum below 15,000 feet until
landing is assured. Do not exceed 96
percent rpm.

Change 13

,1"",,"",""""1~

T.O. 1F- 16A- 1~

E!}~!!J• Refer to Section V for BUC operating
limits.

"~• Avoid rapid throttle movements.

"~• Use 5 seconds minimum for throttle
movements between BUe IDLE andl\1]L and MIL to BUe IDLE.

~~• AB operation is prohibited in BUe.

~~

~

"~
~~

~

• An SFO is not recommended with theengine operating satisfactorily in
BUe.

~

If nozzle fails to close after transferring to BUC or if~~

thrust is still insufficient:

~

::

~

~h~~~7:~~;t:~::~OOI~:t

the AB may~~ :fthc~n;n;:;:t~~2.

Throttle - In MIL to BUe IDLE range.
If the engine is stuck in AB, a flameout
landing is required.

.

~

~

~

When landing is assured (normally high key):

I WARNING I:;tch _ BUC (in limits).
If transfer does not occur, advance throttle
until BUe caution light illuminates. During

~~

".~
~~

Do not start the JFS if engine seizure
has occurred or is anticipated or ifengine failure is a result of fuelstarvation. Starting the JFS may
result in no brake/JFS accumulator
pressure for the brakes.

If throttle is stuck, control might beregained by depressing the cutoff
release, rotating the throttle outboard,and applying necessary force.

~~~

landing,

3-64

Delaying engine shutdown can resultin a long, fast landing. Wheel braking
is less effective due to lack of WOW andthere is an increased probability of a
missed cable engagement.

G:~~~r than nonnal.

• If nozzle is open greater than 50
percent while in BUe, there may not be
sufficient aerodynamic loading for thenozzle to close. To increase aerodynamic loading, decrease altitude,
increase airspeed above 250 knots, andincrease throttle setting.

Change 1

Plan a flameout landing. Refer to FLAMEOUTLANDING, this section.

IWARNING I

NOTE

~~

~
~

4.

to permit a safe landing:~"~If thrust remains too high to permit a safe landing:Land as soon as possible.
During landing, idle thrust is higher thanIf~~

• Use 5 seconds minimum for throttle
movements between BUe IDLE andMIL.

• An SFO is not recommended with theengine operating satisfactorily in BUe.

Stores - Jettison (if or when required).

~ thrus~~:::l~igh~~• Avoid rapid throttle movements.

• AB operation is prohibited in BUe.

result in increased thrust.

~~~. 8 ,.
~ 9• Maintain 80 percent rpm mInImum
above 15,000 feet and 70 percent rpm
minimum below 15,000 feet until
landing is assured. Do not exceed 96
percent rpm.

5.

Throttle - OFF.
If throttle is stuck or engine does not respond,
shut down the engine with the FUELMASTER switch. At MIL, the engine flamesout in approximately 6 seconds; at IDLE, theengine flames out in approximately 45
seconds.

1 1 1 1 1 1 1 1 1 1 1 1~~~~
6.

HOOK switch - DN (ifrequired).

IWARNINGIThe hook may miss the cable if theaircraft is not slow enough to compress
the MLG struts sufficiently to makeWOW or if forward stick pressure is
held.
AS Blowout/Failure To Light IPW200 I

If an AB blowouUfailure to light occurs, the nozzle
may open 30-50 percent and then close. Retard thethrottle to MIL. Refer to figure 5-3 for AB limitations
before further AB use is attempted.
ENGINE STALLS I PW200 I

Refer to figure 5-3. The three primary causes of a stall
are inlet flow distortion, AB instabilities, andhardware malfunctions. During normal aircraft
operation, inlet flow distortion severe enough tocause engine stall is not expected. However, undersome departure conditions, inlet flow distortion may
induce engine stalls. Hardware-associated stalls may
result from a failed nozzle, control system malfunctions, or FOD.
The first indication of an engine stall at high thrust
settings may be a loud bang or pop. At lower thrust
settings, the first indication may be loss of thrust,lack of throttle response, or decreasing rpm. Throttle
reduction is appropriate as a first response to clear
any engine stall.
If the engine stalls at a low altitude, an immediateclimb should be initiated. At low altitude, time may
not be available to recover the engine if it stagnates.
If engine response at low altitude is not sufficient tomaintain or gain altitude and a suitable landing field
is not available, ejection may be required.
AB-Associated Engine Stalls I PW200 I

Types of AB stalls are:• AB initiation - Stall at AB light-off.
• AB sequencing - Stall during AB sequencing,either as the AB stages light with the throttle in
MAX AB or as the throttle is advanced in the AB
range.
• AB cancellation - Stall during throttle movement
fromAB.

• AB blowout/relight - Stall during relight after a
blowout in stabilized AB. May be preceded by AB
rumble.
AB-associated stalls are normally accompanied by aloud bang or pop and a series of fireballs from theengine exhaust and occasionally the engine inlet.
This is followed by an erratic flame from the engine
exhaust if the engine stagnates. These characteristics
could be mistaken for an aircraft fire.
When a stall occurs while operating in AB, the EEC
reduces AB fuel flow to minimum. The throttle shouldbe immediately snapped out of AB to MIL. In MIL,
stall recovery logic is most effective and provides thebest protection from engine stagnation. This action
usually clears the stall and restores normal
operation; however, stalls may continue at MIL andmay be severe. They may be characterized by bangsor pops of low intensity or engine vibrations severe
enough to preclude reading engine instruments.
Refer to NON -AB STALLS, this section.
Refer to figure 5-3. Depending upon flight conditions,
AB stalls may be expected. Stalls should not occur inregion 1. If a stall occurs in region 1, the engine is safe
to operate for the rest of the flight in the IDLE to MIL
range provided no other abnormal engine indications
are observed. Stalls may occur during throttle
transients in region 2. In region 3, stalls may occur
during steady-state AB operation (AB blowouU
relight). If a stall occurs in region 2 or 3, if FTIT limits
were not exceeded, and if the AB operates normally inregion 1, the flight may be continued and the AB may
be used.
Non-AB Engine Stalls I PW200 I

Non-AB stalls may occur if the UFC or EEC ismalfunctioning, particularly during throttle transients near IDLE. Non-AB stalls are often a symptomof a serious engine problem. Non-AB stalls may be
inaudible particularly at high altitudes; the firstindication may be a lack of throttle response which
may be difficult to differentiate from abnormal engine
response. However, non-AB stalls can be severe andmay be characterized by bangs, pops, low intensity or
severe engine rumble or vibration. An erraticorange-yellow flame from the engine exhaust may bepresent. This exhaust flame should not be mistaken
for an engine fire. If the stall(s) is confirmed, thethrottle should be immediately retarded to IDLE. If
the engine recovers, maintain the throttle at
midrange or below but above 80 percent rpm above
15,000 feet and above 70 percent rpm below 15,000
feet until landing is assured. If stalls continue at
IDLE, a stagnation may occur.
Change 7111111111111

3-65

~

~
~
I~
~

I~
~
~
~
~

-

I~1_
~
"

-':~~IIIIIIIIIIII

~
~
~
~
~
~
~
~
~
~
~
~

-

Engine Stagnation I PW200 I

Stagnations are usually characterized by either
rising FTIT and decreasing rpm or rpm less than 60
percent and a lack of rpm response to throttle
commands or illumination of the ENGINE warning
light. These indications are usually accompanied bybangs, pops, low intensity engine rumble or vibration,
and/or an erratic orange-yellow flame from theengine exhaust. This exhaust flame should not be
mistaken for an engine fire. FTIT can either spike,steadily increase, or even decrease immediately
following a high thrust stagnation. During low thruststagnation, FTIT can stabilize in the engine normal
operating range of less than 970°C. Since FTIT can bedeceptive, low rpm (less than 60 percent) with noresponse to throttle movement is generally the bestindication. Once a stagnation occurs, there is no way
to recover normal engine operation except to shut
down the engine and perform an airstart. Allowing astagnated engine to run results in decreasing rpm,
increasing FTIT, a loss of thrust and altitude, andengine damage. As soon as a stagnation is confirmed,
immediately retard the throttle to OFF. There is
every reason to expect a normal airstart and normal
engine operation if the engine is shut down to clear astagnation, especially if the engine was shut down
without allowing FTIT to remain at high temperatures for an extended period.

I WARNING I
Prolonged engine operation with FTIT
in excess of 990°C can result insignificant engine damage and may
cause a nonrecoverable engine failure.

If stalls continue or progress to a stagnation:
A positive indication of a stagnation is rpm lessthan 60 percent with no rpm response to throttle
movement.
3.

Throttle - OFF. Initiate airstart. Refer to
AIRSTART UFC/BUC, this section.

I WARNING I
Shutting down the engine with anengine alternator failure (indicated byzero or erroneously low rpm, ENGINE
warning light illuminated, and normal
thrust) results in no ignition for anairs tart.
If stall(s) clears:
4.

Throttle - Midrange.
80-85 percent rpm.

5.

IN lBtI EEC BUC switch - OFF, then EEC.
NOTE
Cycling EEC BUC switch resets EEC
logic to prevent an overspeed.

If stall(s) occurred at MIL or below:
6.

Land as soon as possible.
Maintain throttle at midrange or below but
above 80 percent rpm above 15,000 feet andabove 70 percent rpm below 15,000 feet until
landing is assured.

Engine Stall Recovery I PW200 I

If stall(s) occurred in AB:

If an AB stall(s) occurs:

7.
1.

Throttle - Snap to MIL.

If AB stalls do not clear or stall(s) occurs below AB:

NOTE
Non-AB stalls may be inaudible.
2.

Throttle - IDLE.

Throttle - As required.
If stall(s) occurred in:• Region 1, do not use AB unless necessary tosustain flight.
• Region 2 or 3, FTIT limits were not exceeded,and AB operates normally in region 1,
continue flight.

INLET BUZZ ~200

Inlet buzz occurs <it ;:;upersonic airspeeds if the engine
con trol system fai Is to maintain adequate engine rpm

when the throttlt, is retarded below MIL. Inlet buzz
causes moderate to severe vibration within the
cockpit and probably results in multiple engine stalls
and/or stagnation.
If inlet buzz occurs, the throttle should not be moved
until subsonic. Decrease airspeed to subsonic asquickly as possible by opening the speedbrakes andincreasing g. If engine stalls occur and persist, thethrottle should be retarded to IDLE when subsonic. If
the stalls do not clear, the engine must be shut down
and restarted.
ENGINE FAILURE OR FlAMEOUT I PW200 I

Engine failures can result in rpm decrease with noabnormal vibration or sound (flameout), rpm
decrease with abnormal vibration and/or stalls, or
stable rpm with abnormal vibration and/or low
thrust.
If the engine flames out, fuel starvation ormechanical failure has occurred. A flameout isindicated by a decrease in FTIT and engine rpm
decaying below approximately 60 percent. Loss ofthrust and lack of response to throttle movement
confirm the flameout. The ENGINE warning lightilluminates when engine rpm goes below 55 percent.
Additionally, the MAIN GEN light illuminates below
45 percent rpm and the EPU should start running. Do
not mistake a loss of ECS noise as an engine flameout.
If the reservoir tanks do not contain fuel, an airstartis impossible.
Main fuel pump failure or tower shaft geartrain
failure also causes flameout. Both present similar
symptoms: an abrupt decrease of indicated fuel flow
to less than 500 pph, loss of main generator and FLCS
PMG, EPU activation, and illumination of the BUC
caution light even though the EEC EUC switch is in
EEC or OFF.
If only the main fuel pump has failed, the cockpit rpm
indication reflects a gradual spooldown. The JFS canbe started and the engine can be motored at
approximately 30 percent rpm. If the BUC caution
light remains illuminated (with the EEC BUC switch
still in EEC or OFF and rpm above 15 percent), theengine probably cannot be restarted; therefore, place
primary emphasis on a flameout landing while
continuing airstart attempts. If unable to make aflameout landing or restart the engine, refer to
EJECTION, this section.

Tower Shah Failure -PW200~

Failure of tlw engint.: tower shaft or its associatedgeartrain results in loss of all rotation to the l'nhrine
gearbox and the ADG. Loss of rotation to the {·ngine
gearbox renders the engine alternator, main fuelpump, and oil pump inoperative resulting in a zerorpm indication, zero oil pressure, illumination of the
ENGINE warning and BUC caution lights, and engine
flameout due to fuel starvation. The initial symptoms
are similar to main fuel pump failure; however, theprimary difference is that the rpm and oil pressure
indications drop immediately to zero with a tower
shaft failure since the engine alternator and oil pump
are not being driven. Additional symptoms caused byloss of rotation to the ADG include loss of hydraulicsystems A and B, main generator, and FLCS PMG andsubsequent activation of the EPU. It may be possible
to regain engine operation using the JFS andperforming a BUC airstart. The JFS drives the ADG
and the engine gearbox (through the PTO shaft),restoring rotation to both hydraulic pumps, FLCS
PMG (at a reduced output), main fuel pump (BUC
caution light goes om, engine alternator (cockpit rpm
signal, EEC power, and engine ignition), and oil pump(oil pressure increases). Without the load of the engine,
the JFS produces an rpm indication fluctuating
between 30-50 percent which is the speed of the engine
alternator, not the actual engine rpm. This rpm may
be high enough to restore main generator power;
however, main generator power may cycle on and offline with the rpm fluctuations. The BUC caution lightgoes off when fuel pump pressure is restored; however
a UFC airstart is not possible since the rpm signal to
the EEC is in error. Perfonn a BUC airstart. Since the
JFS is not preserving engine rpm, maintain 250 knots
minimum during the airs tart attempt, which should
assure adequate actual engine rpm for the airstart. A
BUC airstart is more difficult because of the lack of atrue rpm indication.
Low Altitude Engine Failure or Flameout I PW200 I

Refer to figures 3-7 and 3-8. Initial reaction to any
malfunction at low altitude should be to trade excess
airspeed for altitude. Higher altitude translates
directly to either additional time to achieve anairstart or to additional glide range to reach a suitablelanding field. Above 310 knots, more time is available
by performing a zoom climb using a 3g pullup to
30-degree climb until approaching the desired
airspeed (use approximately 50 knots lead point) andthen initiating a zero g pushover. Below 310 knots
and above the minimum recommended ejection
altitude, more time is available by performing a
constant altitude deceleration to the desired
airspeed. If below the minimum recommended
ejection altitude and below 310 knots, primaryconcern should be to trade excess airspeed for altitudein preparation for ejection.

~~~~~~~~~'·~'I':":~~~"-."-."~~ g; Low Altitude Zoom Capability~-....""DATA BASIS ESTIMATEDENGINE Fl00-PW-200

CONFIGURATION:
e GW = 23,000-25,000 LB

CONDITIONS:e WINDMILLING OR SEIZED ENGINEe 30-DEGREE CLIMB MAINTAINED TO 170/250 KIAS

=

eDI
0-50
eLG-UP
12~ ...~UTI] 111111111 In±±IllIl~lIllIEn11'fIf FBJJ lIE 111"'EJ FEklllllllfffl
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mtt NU 11 rtti;Ntt tt It III rtM;NjJ'I'N~~,......

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3G PULLUP TO 30°
ZOOM CLIMBTO ACHIEVE:

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160

180

200

220

240

260

280

300

320

340

INITIAL AIRSPEED -

360

380

400

420

440

KIAS

460

480

500
1 F-16X-l··l028X@

Figure 3-7.

.

.

~

"II1II

Low Altitude Airstart Capability
DATA BASIS ESTIMATEDCONFIGURATION :• GW _ 23,000-25,000 LB• OJ = 0-50
• LG - UP

ENGINE F100-PW-200

CONDITIONS:• 30 0 DIVE TO DESCENT KIAS OR 3G PULLUP
TO 30 · ZOOM CLIMB INITIATED FROM THE
AIRSPEED/ALTITUDE EXISTING AT FIRSTRECOGNITION OF ENGINE FAILURE (60PERCENT RPM )• 60 SECONDS ASSUMED AFTERTHROTTLE ADVANCE TO ACHIEVE
USABLE THRUST.~• DESCENT AIRSPEED IS 170 KIAS
I.JFS RUN LIGHT ONI• AIRSTART INITIATED WHEN RPM REACHES
2 ~O PERCENT (10 SECONDS AFTERINITIATION OF DIVE OR ZOOM )10 I ::;:::::r:r::;~"'t;:"
g
o

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INITIAL AIRSPEED - KIAS
1F_ 1I:IA.. 1_1154....

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Figure 3-8.

~
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1,',''''~'ZF''

If required, jettison stores as soon £1:- possiblp to aid ingaining or sustaining altitude and lnaneuver towarda suitable landing field, if available. If the zoom climb
results in an altitude below 5000 feet AGL, there willprobably be insufficient time to achieve an airstartprior to minimum recommended ejection altitude. Inthat case, primary consideration should be given to
preparing for ejection; do not delay ejection below
2000 feet AGL. If a suitable landing field is within
gliding distance, then all attention should be directed
toward successful interception of an appropriate
flameout landing pattern key position. Uponreaching a proper key position, all attention should bedirected toward a successful landing. If the zoom
climb results in an altitude between 5000-10,000 feet
AGL, there is probably time for one airstart attemptprior to minimum recommended ejection altitude.
This attempt will be in BUC. Above 10,000 feet AGL,
there is sufficient time for a UFC airstart attemptprior to switching to BUC.
If low altitude engine failure or flameout occurs:
1.

Zoom.

2.

Stores - Jettison (if required).
If stores jettison is attempted after main
generator drops off line but before EPU
generator powers the SMS (approximately 5
seconds delay), stores will not jettison.
NOTE

3.

Perform BUC airstart (if altitude permits).

IWARNING~I

AIRSTARTS ~W200 I

Refer to figure 3-9. Factors such as altit ucie, air::;peed,
weather, etc .. must be considered in determining
whether to try an airstart, accompli:-;h a flameout
landing, or eject. Jettisoning stores reduces altitude
loss during an airstart and improves glide ratio
during a flameout landing.
Oil pressure is directly related to rpm. Do not confuse
a low oil pressure indication due to windmilling rpm
as an oil system malfunction.

I

If the engine has seized due to an oil system

malfunction or flamed out due to fuel starvation ormechanical fail ure, a flameout landing or ejection isrequired.
There are two airstart options available in either
UFC or BUC. The primary option is a spooldown
airstart, for which the throttle is advanced from OFF
to IDLE as rpm is decreasing between 40-25 percent.
The recommended spooldown airspeed is 250 knots
minimum for either UFC or BUC. The secondary
option is a JFS-assisted airstart which differs from aspooldown airstart in that once the JFS RUN light ison, airspeed can be reduced to achieve maximum
range or maximum endurance (200 or 170 knots,
respectively, plus 5 knots per 1000 pounds offuel/store weights over ~ 3000, [ID 2000 pounds).
Under normal conditions the JFS will motor theengine at a minimum of 22 percent.
There are critical requirements which apply to any
airstart attempt (UFC or BUC). The most importantis engine rpm. During a spooldown airstart, if thethrottle is advanced from OFF to IDLE after rpm goes
below 25 percent, light-off may not occur before rpm
decreases through 15 percent. Below 15 percent, themain fuel pump does not supply sufficient fuel toeffect an airstart and engine ignition is not available.
In general, rpm spooldown rate can be decreased withincreased airspeed. As much as 350 knots or more
may be required to prevent rpm from decaying below
15 percent. If rpm is allowed to drop to near zero,
400-450 knots for 20-25 seconds may be required toregain 15 percent. If this situation occurs and theaircraft is not within gliding distance of a suitablelanding field, a 50-degree dive angle or greater shouldbe established to accelerate to 450 knots.

~,7,;;e,~~~~~~~~~~~IIIIIIIIIIIIIIIIIIIIII~
T.O. 1 F-16A-1

Engine RPM and FTIT Response During
Spooldown and A irs tart
ENGINE F100-PW-200

CONDITIONS:e GW = 17,000 LB
eNO WINDe DRAG INDEX = 0
e KIAS
250

=

--..

50
POINT B TOSL
9 MINI=

A......

30° DIVE...
............

48 NM

-

",

40

B

....w
....w

MIN KIAS(WITH JFS RUN
LIGHT OFF)UFC-250
BUC-250

o 30
o
o
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UFC AIRSTARTENVELOPE~/_ _ ~ BUC AIRSTART/~ENVELOPE-

I.

~¥.~~~~~~~~~-------375
KIAS,..-----;:).... 20

5<t~---MARGINALII
MIN KIASI(WITH JFSI
RUN LIGHT
I
ON) ____,'

10

UFC-170
BUC-170I:

I-

O~------~~--~~------~~------~------~------~------~

o

0.2

-

0.4

0.6

0.8

1.0

1.2

1.4

MACH NUMBER-

OPTIMUM FLIGHT PATH DURING AN AIRSTART (TYPICAL)Engine out descent flight path maintains the aircraft in the required airstart envelope.
A 3O-degree dive to descent KIAS is used to approach the airstart envelope(Point A to B).

1 F-16X-1-0032X @Figure 3-9. (Sheet 1)

Change 4

3-71-

~""""""""""~__

T.O. 1F-16A-l~

Engine RPM and FTIT Response During~ Spooldown and A irs tart
ENGINE F100-PW-200

--

~

1200

--

1100

~

1000

;> 900t: 700
600

--~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~

~

800°C
- - - - 700°C

I 800

l-

500
400
100
80

I

:e

60

a.

40

a::

20

1

THROTTLE
OFF

THROTTLE
IDLE

AIRSTART
THROTTLEOFF

STABILIZEDIDLE THRUST

I

IDLE

--=-~----------------1----o~--~----~--~--~----~--~----~--~--~----~--~--~

o

10

20

30

40

50
60
70
TIME - SECONDS

80

90

100

110

120

ENGINE RPM AND FTIT RESPONSE DURING A HOT AIRSTART FOLLOWINGA HIGH THRUST STAGNATION (TYPICAL)
A hot airstart is characterized by FTIT reaching 800°C or above or by FTIT
approaching 800°C so rapidly that there is no doubt that it will exceed
800°C. The only way to clear a hot airstart is to retard the throttle OFF,
increase airspeed within limits and then attempt another UFC airstart. If
the second airstart attempt is unsuccessful and all parameters were within
limits, retard the throttle to OFF and initiate a BUC airstart. The throttle
should always be placed to IDLE when engine rpm reaches 25 percent
regardless of all other parameters to insure light-off occurs before engine
rpm decreases below 15 percent.

1 F-16X-1-0033X @Figure 3-9. (Sheet 2)

3-72

Change 1

".~~~~~~~~~~~~~~
to. 1F-16A-l

Engine RPM and FTIT Response During

Spooldown and Airstart
ENGINE F100-PW-200

1000
(.)

800

...~I 600

400

THROTTLE
IDLE

200
100

STARTING FUELSWITCH RICH

THROTTLE
OFF

80
60
40

----==-::.~

40%- - - - -25%

20o~--~~--~----~----~----~----~----~----~--~
40
50
60
70
90
20
80
o
10
30
TIME - SECONDSENGINE RPM AND FTIT RESPONSE DURING A HUNG AIRSTARTFOLLOWING A HIGH THRUST STAGNATION (TVPICAL)
A hung airstart is characterized by engine rpm hung and FTIT less thanaoooe. Hung airstarts usually occur in the 40-50 percent engine rpm

range but may occur at any engine rpm. The lower the altitude, the lower
the engine rpm at which the airstart hangs. At low altitude, the airstart
may hang shortly after light-off. Increasing airspeed may increase engine
rpm or positioning the STARTING FUEL switch to RICH increases fuelflow by 100 pph allowing the engine rpm to increase through idle. If
the airstart is still hung, a BUC airstart should be attempted.

1F-16X-1-0034xe

Figure 3-9. (Sheet 3)

I
I
I

I
I
I
I
I
I
I
I
I
I
I

1200

°

I

I
I

I

"T~1r~~~~~~~~~~~~-Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-200

-

1200
1000
()
o

800
600
400
THROTTLE
IDLE

200
THROTTLE
OFF

100

STABILIZEDIDLE THRUST
LlGHT-OFF

80

I
:i
0.

a:

60
40%

40
20

o o~--------------------------------------~----------~
10
20
30
40
50
70
80
60
TIME - SECONDSNORMAL AIRSTART ENGINE RPM AND FTIT TIME TRACES FOLLOWING A
HIGH THRUST STAGNATION (TYPICAL)Light-off occurs within 15 seconds after the throttle is advanced to IDLE.
However, engine rpm and FTIT turnaround are slow, making light-off subtle
or difficult to detect. Engine rpm stabilizes momentarily after light-off and
FTIT begins to increase and then decreases and stabilizes as the engine rpm
increases normally to idle.

-,,

1 F-1 6X- 1-0035X @Figure 3-9. (Sheet 4)

3-74

Change 1

r ~ ~ ~ ~ ~ ~ ~ I' I'l'l'

I'r1!l1!,

Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-200

1000

800
(J

°

~ 600
~

250 KIAS
10,000 FEET

400

350
KIAS
35,000
FEET

200
100
350 KIAS
10,000 FEET
80~

250 KIAS
35,000 FEET

60

I
::i
11-

a:

40

20

o

250 KIAS
10,000 FEET

I

o

I
20
40
TIME - SECONDS60

RATE OF ENGINE RPM AND FTIT DECAY
AS A FUNCTION OF ALTITUDE AND
AIRSPEED (TYPICAL)At low altitude, regardless of airspeed,engine rpm decreases rapidly which
decreases the time available in theairstart window (FTIT less than 700°C
and engine rpm between 25-40 percent).

1F-16X-l-Q036X ®

Figure 3-9. (Sheet 5)

,
,
,
,
,
,
,
,
,
,
,
,
,
,
,
,
,

~IIIIIIIIIIIII~

"T.~!!r '" '" '" ~ ~ ~ ~ ~ ~ ~ ~ ~--Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F10o-PW-200
1200

1200

1000

CONDITIONS:.250 KIAS.25,000 FEET

1000

u
° 800
I

~ 800

t-

Ii: 600

It-

Ii: 600

400
200~--------~------~--------~

400

.

100

1-THROTTLE OFF
1000°C
2 -THROTTLE OFF
12000C

2

I

200
80

"II
:I

B:

100
60
THROTTLEADVANCEAT 250/. RPM
REQUIRED

80
40
60

20

o~--------~--------~--------~
o
10
20
30
TIME- SECONDS-l---~

40_ _ _ _ 1_2 _ _ _ ~20

SPOOLDOWNAIRSTART WINDOWENGINE RPM AND FTIT RESPONSE FOLLOWING A HIGH THRUST STAGNATION (TYPICAUooFollowing a high thrust stagnation, engine rpm
rapidly decreases and FTIT varies within theinitial few seconds as shown by three examples.
Example 1 shows a rapid decrease in engine
rpm and steady increase in FTIT.
Example 2 shows a rapid decrease in engine
rpm and a momentary step increase in FTIT
which, within a few seconds, turns into a
steady increase in FTIT.
Example 3 shows a rapid decrease in engine
rpm and an immediate drop in FTIT of up to
200 0 e as a result of momentary main
combustor blowout. Light-off always reoccurs
within a few seconds causing FTIT to rise
rapidly.

20
40
60
TIME - SECONDS80

RATE OF ENGINE RPM AND FTIT DECAY AS A
FUNCTION OF THROTTLE MOVEMENT DURING
AND FOLLOWING A HIGH THRUST STAGNATION(TYPICAl)
Allowing the engine to remain stagnated reduces
the spooldown airstart window as shown by two
examples.
Example 1 shows the throttle retarded to OFF when
FTIT reached 1000 oe. The spooldown airstart
window indicates there is no significant engine
turbine distress.
Example 2 shows the throttle retarded to OFF when
FTIT reached 1200 o e. There is no spooldown
airstart window since minimum engine rpm wasreached before FTIT cooled below 700 o e. Possible
engine turbine distress may affect airstart success.
1F-18X-1-OO37xe

Figure 3-9. (Sheet 6)

3-76

Change 1

,
,,
,,
,

IIIIIIIIIIIIIIIIIIIIII~
T. o. 1F- 16A- 1

Engine RPM and FTIT Response During
Spooldown and A irs tart
ENGINE F100-PW-200
1200

1200
1000

CONDITIONS:.250 KIAS• 25,000 FEET

1000

(J

° 800
I

,,
,,
,,

~ 800

~

t: 600I~

t: 600400

2

1 - THROTTLE OFF

200~----------------~----------------~----------------~

900°C

400

2 -THROTTLE OFF

100
200
80

I
~

a.

a::

100

60
THROTTLEADVANCEAT 25% RPM
REQUIRED

80

40

I

20

60

,,
,,
,

~

a.

01....-----------....1".---------------'
o
10
20
30
TIME -

CC

40%-------

40

1

SECONDS

20ENGINE RPM AND FTIT RESPONSE DURING
AND FOLLOWING A LOW THRUSTSTAGNATION (TYPICAL)
During a low thrust stagnation, engine rpm
decreases while FTIT increases. The rate at
which engine rpm decreases and FTIT increases
varies with altitude and airspeed. Engine rpm
decreases at the same rate regardless of FTIT
response. However, FTIT following a low thruststagnation often levels off at 800°C - 1000°C.

--~-1

1

o

I
o

I

25%

2----SPOOLDOWN[AIRSTART WINDOW

I

I

20

40

I

I

I

60

I
80

I

I
100

TIME - SECONDS,,RATE OF ENGINE RPM AND FTIT DECAY AS A
FUNCTION OF THROTTLE MOVEMENT DURING
AND FOLLOWING A LOW THRUST STAGNATION(TYPICAL)
Allowing the engine to remain stagnated reduces
the spooJdown airstart window as shown by two
examples.
Example 1 shows the throttle retarded to OFF withengine rpm below 60 percent and when FTIT reached
900°C. The spooldown airstart window is shown above.
Example 2 shows the throttle retarded to OFF
when FTIT reached 1000°C. There is no spooldown
airstart window. The operations are the same as
those in RATE OF ENGINE RPM AND FTIT DECAY
AS A FUNCTION OF THROTILE MOVEMENTDURING AND FOLLOWING A HIGH THRUSTSTAGNATION, example 2.
1F-16X-1-D038X®

Figure 3-9. (Sheet 7)

Change'

3-77

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,

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...

~

...
~

~

...
~
...
~

~

IIrr....
~

~
~
~
~
~
~
~

~
~
~
~
~
~
~
~
~

~

10. 1 F- 16A- 1

Once at 450 knots, the dive angle may be reduced toapproximately 20 degrees to maintain airspeed. As
the engine rpm increases to near 15 percent, an
airstart may be attempted. If the dive is started from
20,000 feet MSL at 200 knots and zero engine rpm,
16,000-18,000 feet is required to regain 15 percent
engine rpm. If the engine is allowed to stop rotating,it may thermally seize after which it will not rotate
even with high airspeeds or by engaging the JFS. The
second requirement is engine temperature. Hightemperatures during airs tart may cause a stagnation. High temperatures may result if the airstart isinitiated before the FTIT is allowed to decrease below
700°C or if an airstart is attempted with the
STARTING FUEL switch in RICH.

To meet these requirements, an airs tart should beinitiated at 40-25 percent rpm with FTIT below
700°C. These parameters can usually be achieved bymaintaining 300 knots minimum above 30,000 feet or
250 knots minimum below 30,000 feet. These
airspeeds do not maintain rpm above 15 percent;however, they provide the best tradeoff between therate of rpm spooldown and loss of altitude. If it
appears that engine rpm will drop below 25 percent,increase airspeed if feasible and advance the throttle
to initiate the airstart even if FTIT is above 700°C.
Generally, higher airspeeds increase start reliability
by slowing rpm spooldown which allows the FTIT todecrease more rapidly.
At low altitudes, however, higher airspeeds do notsignificantly affect the rpm spooldown. Unlessrequired to recover rpm from below 15 percent,airspeed during a UFC airstart without the JFS RUN
light on should be at or just above 250 knots. If a BUC
airstart without the JFS RUN light on is anticipated,
airspeed must be 250-375 knots. Once the JFS RUN
light is on, airspeed can be reduced to achieve
maximum range or maximum endurance.
High Altitude Airstart Considerations I PW200 I

Refer to figure 3-9. At high altitudes, dive atapproximately 30 degrees to gain or maintain 300
knots minimum (250 knots minimum below 30,000feet). Once established, approximately 5-10 degrees
of dive should maintain airspeed. Note that airspeed
can be reduced to less than 250 knots after a JFS RUN
light is confirmed. Unless an airstart is obviously
impossible (total lack of fuel, engine seizure, etc.), do
not become tempted to establish a maximum range or
maximum endurance glide. The first consideration
should be an immediate spooldown airstart attempteven if the engine failed for no apparent reason. If
airstart airspeed is not maintained, rpm decreases ata faster rate. Time constraints due to EPU fuel3-78

Change 4

consumption must be considered. A maximum range
or maximum endurance glide from above approximately 35,000 feet may exhaust EPU fuel prior tolanding. (Refer to T.O. IF-16A-l-l, figure AS-3.) At
least one UFC start should be attempted.
If all UFC airstart parameters (rpm, FTIT, andairspeed) are met and the start is unsuccessful,
perform a BUC start below 25,000 feet. When below
20,000 feet MSL, turn the JFS on. Activating the JFS
above 20,000 feet MSL is prohibited since successful
JFS start/motoring of the engine is unlikely and thebrake/JFS accumulators will be depleted. If the JFS
RUN light is on, airspeed may be reduced to best I
range or best endurance airspeed. With the JFS
running, EPU fuel consumption is also reduced.
Low Altitude Airstart Considerations IPW200 I

Due to the limited time available and the rapid rpm
spooldown rate at low altitude, some additional
considerations are required. Below approximately
10,000 feet MSL, rpm decreases rapidly regardless ofairspeed and remains between 40-25 percent for only
5-10 seconds; therefore, rpm should be closely
monitored. Advance the throttle to initiate theairs tart before rpm goes below 25 percent regardlessof FTIT indication. This action should insure thatlight-off occurs prior to 15 percent rpm. Start the JFS
immediately after advancing the throttle (if airspeed
is below 400 knots).
Following a zoom climb, plan to arrive at 250 knots
without JFS confirmed running or with tower shaft
failure. After the JFS RUN light is on (and tower I
shaft failure has not occurred), airspeed may be
reduced to achieve maximum range or maximum
endurance. If a higher airspeed is maintained or an
attempt is made to gain airspeed to delay the rpm
decay, available time may be reduced to the point thatan airstart is not possible. During any low altitude
airstart attempt, constantly evaluate AGL altituderelative to airstart success. Do not delay ejection
below 2000 feet AGL unless the engine is producing
thrust capable of sustaining level flight or safely
controlling the sink rate or unless a flameout landing
can be accomplished.
Airstart - UFC IPW200 I

Refer to figure 3-9. To perform a UFC airstart,
advance the throttle from OFF to IDLE when airstart
conditions are met.
If the throttle is retarded to OFF to clear astall/stagnation, the rpm decreases rapidly and the
FTIT begins to decrease. The throttle should be~"""""""""'" S
T. O. 1F- 16A- 1

maintained in OFF for a few seconds even if rpm and
FTIT are within airstart limits to allow the stall toclear. If the airstart attempt is not due to astagnation, FTIT will probably be well below 700') C
when the throttle is retarded to OFF. If it appe3rs thatengine rpm will drop below 25 percent, the throttle
must be advanced to IDLE to preserve rpm regardlessof FTIT or airspeed. A hot start will probably occur,and the throttle must again be retarded to OFF.
However, rpm is preserved permitting a subsequent
airstart within parameters.
Start the JFS below 20,000 feet MSL and below 400
knots immediately after advancing the throttle to
IDLE to initiate the spooldown airstart.
When the throttle is advanced to IDLE, rpm and FTIT
may continue to decrease until engine light-off occurs
which takes up to 15 seconds. Do not rush the airstart
attempt. Increasing rpm is normally the firstindication of an airstart. The light-off is subtle since
rpm and FTIT turnaround are very slow. If a light-off
is not attained in 15 seconds, position the STARTING
FUEL switch to RICH for 5 seconds. If no light-off
occurs after 5 seconds in RICH, retard the throttle to
OFF and initiate a BUC airstart. From the time thethrottle is advanced from OFF, it may take up to 1
minute for 1ight-off and acceleration through idle.
Engine acceleration may be slow around 40-50
percent rpm during the airstart attempt. Do notconfuse this slow acceleration with a hung start. If ahung start occurs (stabilized FTIT below 800°C, rpm
hung and definitely stabilized below 60 percent),
position the STARTING FUEL switch to RICH. If thehung start continues, increase airspeed to a
maximum of 400 knots. If a hung start still persists,
move the throttle to OFF and initiate a BUC airstart.
If a hot start occurs (FTIT above 800°C) and UFC
airstart parameters were met and altitude is still
sufficient, increase airspeed (if possible) and attemptanother UFC airstart. If a hot start occurs and UFC
airstart parameters were not met and maintained,
attempt another UFC airstart when inside parameters. If a second hot start occurs or if conditions
dictate, such as low altitude or events which indicate
a UFC airstart will not be possible, retard the throttle
to OFF and initiate a BUC airstart.
If the JFS stops running or fails to run within 30
seconds, do not reattempt a JFS start until thebrake/JFS accumulators have time to recharge. Allow
1 minute of engine rotation (either windmilling or
JFS assisted) at 12 percent rpm or above to insurethat the brake/JFS accumulators are fully recharged.
Recharging begins 3-4 seconds before the JFS RUN
light illuminates or 30 seconds after selecting a start
position (in the event of a JFS failure to run).
Recharging begins regardless of JFS switch position.

In the event of a ,JFS :'lhutdo\'. n, the JFS ;:,witch does
not relatch in either ~tlrt p(Jsition while the JFS isspooling down. Spooldown from full gov~rned spe~dtakes approximately 17 seconds. The JFS switch
must be cycled to OFF and then to START 2 toreinitiate a JFS start. It is possible to complete thespooldown before the brake/JFS accumulators arerecharged if the JFS ran for only a short time.

~~

111....'"...~~

When the airstart is completed, turn the JFS off. ..~Cycle the EPU switch to OFF and then back to NORM~
and verify that STARTING FUEL switch is in AUTO~~
LEAN.

III....

To accomplish a UFC airstart:
1.

~

Throttle - OFF.

~FTIT should decrease rapidly whenthrottle is 0 FF. If FTIT does notdecrease rapidly, verify that thethrottle is OFF.
2.

~

~

Airspeed - As required.
Maintain a minimum of 300 knots above
30,000 feet, 250 knots below 30,000 feet.

~~

111..... .
~~enT::t::e4~-:;:;rcent with FTIT below 700°C: ~G:!!!!~~~.~

If it appears rpm will go below 25
percent, advance the throttle to IDLEo
4. J:~::~~::S:::::T ; :::::e;~:OOO

feet MSLand below 400 knots.
IVlaintain maximum range or maximum
endurance airspeed (200 or 170 knots,
respectively, plus 5 knots per 1000 pounds offuel/store weights over [ru 3000, [ID 2000
pounds) with the JFS RUN light on.

If maximum gl:::E range is not a
factor, consider maintaining 250 knots
above 10,000 feet AGL to reduce rpm
spooldown rate (in case of JFS failure).
Below 10,000 feet AGL with the JFS
RUN light on (where only one airstart
attempt is likely), maintain a minimum of 170 knots plus 5 knots per 1000
pounds of fuel/store weights over [ru3000, 1!l2000 pounds.
Change 13

~~

~

~

~~

~~

~

~

~

3-7~ ~NOTE

IWARNINGI• If the .JFS switch is erroneously placed
to START 1, leave it there.
• If the JFS RUN light does notilluminate or goes off once illuminated,
place the .JFS switch to OFF andreattempt START 2 when the brakel
JFS accumulators are recharged. The
JFS switch does not relatch in either
start position while the JFS is spooling
down.

5.

Stores - Jettison (if required).
If stores jettison is attempted after main
generator drops off line but before EPU
generator powers the SMS (approximately 5
seconds delay), stores will not jettison.
NOTE@) If only AlJX flag is in view, pitch androll attitude information is likely to be
erroneous due to INS autorestart in
the attitude mode when other thanstraight and level, un accelerated
flight conditions existed.
13.

STARTING FUEL switch - AUTO LEAN.

14.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINE

Airstart - Hot (UFC) I PW200 I

If a hot start occurs:

1.

Throttle - OFF.

If altitude is still sufficient:

Visually confirm the stores havejettisoned and jettison again if required.

2.

I WARNING I

If hun gino start:
6.

STARTING FUEL switch - RICH (5 seconds).
Insure the second start is accomplished in parameters. Increased airspeed (400 knots maximum) may
improve airstart capability.

If hung start persists:
7.

Airspeed - Increase (400 knots maximum).

If hunglno start persists:
8.

Initiate AIRSTART - BUC, this section.

Ifbelow 10,000 feet AGL or a second UFC airs tart isunsuccessful:

Ifhot start:
FTIT above BOO°C.
9.

IWARNINGI

Go to AIRSTART - HOT (UFC), this section.
Below 5000 feet AGL, there may beinsufficient time to perform an airstartprior to minimum recommended ejection altitude.

When start is completed:
Observe EEC off operating limits.
If engine does not respond normally after airstart iscompleted:
10. Refer to FLAMEOUT LANDING, this section.
If engine responds normally:
10. JFS switch - OFF.
11.

EPU switch - OFF, then NORM.

12. ADI - Check for presence of OFF andlor AUX
warning flags.
If warning flag(s) is in view, refer to TOTALINS FAILURE, this section.

3-80

Reinitiate AIRSTART - UFC, this section.

Change 13

2.

Initiate AIRSTART - BUC, this section.

Airstart - BUC I PW200 I

During a BUC airstart, engine airflow (related toairspeed) and fuel flow (regulated manually bythrottle) are critical to prevent an overtemperatureand subsequent stagnation. For a BUC airstart,
airspeed must be 250-375 knots with altitude below
25,000 feet. To preserve rpm, the JFS should bestarted anytime during the airstart sequence if thealtitude is less than 20,000 feet MSL. Once the JFS
RUN light is on, the minimum airspeed is 170 knots

,
,
,
,,
,

IIIIIIIIIIIIIIIIIIIIII~
T.o. 1F-16A- 1

plus 5 knots per 1000 pounds of fuel/store weights
over [!] 3000,
2000 pounds. If the JFS stops
running or fails to run within 30 seconds, do not
reattempt a JFS start until the brake/JFS accumulators have had time to recharge. Allow 1 minute ofengine rotation (either windmilling or JFS assisted)at 12 percent rpm or above to insure that thebrake/JFS accumulators are fully recharged.
Recharging begins 3-4 seconds before the JFS RUN
light illuminates or 30 seconds after selecting a start
position (in the event of a JFS failure to run).
Recharging begins regardless of JFS switch position.

rm

In the event of a JFS shutdown, the JFS switch does
not relatch in either start position while the JFS isspooling down. Spooldown from full governed speed
takes approximately 17 seconds. The JFS switch
must be cycled OFF and then to START 2 to reinitiatea JFS start. It is possible to complete the spooldown
before the brake/JFS accumulators are recharged if
the JFS ran for only a short time.
NOTE~ A successful BUC airstart is notpossible without ~ pilot assistance.

Throttle movement in and out of OFF
and movement past the BUC IDLE
detent require ~ pilot assistance.
Th initiate the airstart, retard the throttle to OFF andselect BUC (BUC caution light illuminates). The
STARTING FUEL switch must be in AUTO LEAN or
RICH. (If the switch is in LEAN, the engine willprobably light off; however, it may stagnate as thethrottle is advanced near the BUC IDLE detent.)

Advance the throttle to IDLE when rpm is 40-25
percent and FTIT is below 700°C. If the throttle wasin OFF for less than 5 seconds prior to being advanced
above IDLE, excess fuel flows into the engine and a
hot start will probably occur. This may be prevented
by insuring that the throttle is advanced to IDLE andnot above, even momentarily, or waiting approximately 5 seconds in OFF prior to advancing thethrottle to IDLE. Waiting may not be practical since,in addition to the 5 seconds required for fuel to drain,
an additional 10-15 seconds is required to fill the fuelmanifolds to reinitiate the airstart after the throttle
is advanced to IDLE. The best option is to insure thethrottle is carefully advanced around the IDLE
detent without allowing it to advance above IDLE,
even momentarily. The throttle can be advanced to
IDLE above 25,000 feet to preserve rpm, but theaircraft must be flown into the BUC start envelope
before further advancing the throttle.
When the throttle is advanced to IDLE, rpm and
FTIT may continue to decrease for a brief period. It
may take up to 15 seconds to obtain a light-off. The
BUC light-off is more subtle than in UFC due to lower
minimum fuel flow. The best indication of a light-off
is a slow leveling of rpm or an rpm increase.
Depending on the type of failure that necessitated a
BUC airstart (flameout, stagnation, etc.), FTIT may
begin to rise either before or after rpm. The greatest
stall margin and probability of a successful airstart
result if airstart actions are based on rpm increase
rather than FTIT. If rpm is unreliable (e.g., tower
shaft failure), FTIT must be used. If light-off does notoccur in 15 seconds, continue to very slowly advancethe throttle until it does occur. Mter light-off, stop any
throttle movement and allow rpm to increase andbegin to stabilize (approximately 10 seconds). Thisaction is critical to increase the stall margin andsignificantly increases the probability of a successful
airstart. If the throttle is advanced immediately after
light-off, an engine stall may occur which will
preclude a successful airstart since rpm will not
accelerate past approximately 45 percent. After rpm
begins to stabilize, slowly advance the throttle tomaintain a steady rpm rise similar to a normal UFC
airstart. Since fuel flow is directly controlled by thethrottle, a concentrated effort is required to insurethat throttle movement is slow and deliberate; aminimum of 30 seconds is required to advance thethrottle from IDLE to BUC IDLE (based on gliding at
250 knots or less). A BUC airstart takes approximately 15-30 seconds longer than a UFC airstart.
Place primary emphasis on controlling the throttle
movement to produce a steady rpm increase andcarefully monitor FTIT. FTIT does not normallyexceed 500°C above 15,000 feet (below 15,000 feet,temperatures as high as 600°C may be required toachieve normal acceleration and starting times).
Rapidly advancing the throttle in an attempt toachieve these FTIT values will probably result in a
hot airstart. If it appears that FTIT will exceed thesevalues, stop advancing the throttle and wait for FTIT
to stop increasing; then continue the throttle
advance. In some cases (such as after a stagnation)
where FTIT may be relatively high when the engine
lights off, FTIT may remain fairly constant in the
500°-600°C range. As long as rpm continues toincrease at a steady rate and FTIT does not rise
rapidly, continue advancing the throttle. Do notretard the throttle unless aborting the airstart.

Change 1

3-81
__

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..411

~""""""""""~
~

~
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~~

~

~
~

~

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~

~

~
~
~
~
~
~

~.
~
~
~
~
~
~
~
~
~
~
~

10. 1 F-16A-l
At the backside of the BUC IDLE detent, pause; thenrotate the throttle outboard to reduce resistance (thethrottle can be pushed through the detent; however,this will probably result in a rapid throttle
advancement and a hot airstart). Rate of rpm
increase may slow down slightly while passing thedetent; however, do not be tempted to advance thethrottle faster. Continue advancing the throttle justas before and monitor FTIT (FTIT is more sensitive
to throttle advancement during this phase of theairstart). Once past the BUC IDLE detent, normal
FTIT limits apply; however, use the same rate ofthrottle advancement until reaching 80 percent rpm
minimum above 15,000 feet and 70 percent rpm
minimum below 15,000 feet. If at anytime during thestart FTIT exceeds 800 ° C or rpm stops increasing as
the throttle is advanced (stall/stagnation), retard thethrottle to OFF and reinitiate the airstart if altitude
permits.
If valid rpm is not available (e.g., tower shaft failure),
the airstart is more difficult to control. After engine
light-off, the throttle should be slowly and smoothly
advanced from IDLE to BUC IDLE over a 30-45
second period. Maintain a minimum of 250 knots
since the JFS is not assisting engine rotation.
Advance the throttle slowly to produce a steady FTIT
increase; if FTIT reaches 500°-600°C prior toreaching BUC IDLE, use the above technique tojudge how fast to advance the throttle. Do not shut the
JFS off since the JFS is all that is providing fuelpressure. After restarting, engine thrust, throttle
response, fuel flow, and FTIT should be normal for
BUC operation. Indicated rpm will remain 30-50
percent as long as the JFS RUN light is on.
After the airstart is complete, set the throttle tosustain level flight (about 80 percent rpm or 700°C
FTIT without valid rpm indication). Throttle may be
used as required in the landing pattern; however,throttle movements must be smooth and slow. Do not
reduce the throttle to BUC IDLE until landing isassured.

NOTE
During landing in BUC, idle thrust is
higher than normal. The EEC BUC
switch should not be moved out ofBUC
after landing in an attempt to open thenozzle and decrease thrust since thismay result in reoccurrence of theoriginal malfunction.
After the airstart is complete, cycle the EPU switch
OFF and back to NORM and turn the JFS off unlessindicated rpm is below 60 percent with adequate
thrust.
3-82

Change 4

To accomplish a BUC airstart:
1.

Throttle - OFF.

FTIT should decrease rapidly whenthrottle is OFF. If FTIT does notdecrease rapidly, verify that thethrottle is OFF.

2.

Airspeed - 250-375 knots.

3.

r:ru ~ EEC BUC switch - BUC.
NOTE
The proximity of the EEC BUC switch
to the JFS switch makes the JFS
switch susceptible to being bumped to
OFF when selecting BUC.

When rpm is 40-25 percent with FTIT below 700°C:

4.

Throttle - IDLE.

• If it appears rpm will go below 25
percent, advance throttle to IDLE
regardless of FTIT or airspeed.

• Advancing the throttle above IDLE,
even momentarily, may add excess fuelto the engine which will probably
result in a hot start.

5.

JFS switch - START 2 below 20,000 feet MSL
and below 400 knots.
Maintain maximum range or maximum
endurance airspeed (200 or 170 knots,
respectively, plus 5 knots per 1000 pounds offuel/store weights over ~ 3000, lm 2000
pounds) with the JFS RUN light on.

NOTE
If maximum gliding range is not a
factor, consider maintaining 250 knots
above 10,000 feet AGL to reduce rpm
spooldown rate (in case of JFS failure).
Below 10,000 feet AGL with the JFS
RUN light on (where only one airstart
attempt is likely), maintain a minimum of 170 knots plus 5 knots per1000 pounds offuellstore weights over[!] 3000, [ID 2000 pounds.

•I"~'~I~r~~i~v~~,r I l l
1 o. 1F - I oA- 1

NOTE• Monitor FTIT during the start; FTIT should
not l'xceed 500 C above 15,000 feet or 600 a C
below 15,000 feet. If the throttle is rapidly
advanced to obtain these FTIT's, a hot start
may result. If FTIT reaches these values, stop
throttle advance, wait for FTIT to stop
increasing, and then continue the throttle
advance.

~~r• If the JFS switch is erroneously plaCt'd

to START 1, leave it there.
• If the JFS

RUN light does notilluminate or goes off once illuminated,
place the JFS switch to OFF andreattempt START 2 when the brake/
JFS accumulators are recharged. The
JFS switch does not relatch in either
start position while the JFS is spooling
down.

~

~

When at the backside of the BUC IDLE detent:

B.

Throttle - Pause, rotate outboard, andsmoothly advance past the BUC IDLE detent.

After light-off:
NOTE
NOTE

Light-offis indicated by a slow levelingof rpm or an rpm increase. Do not use
FTIT as an indicator of engine light-off
unless rpm is obviously invalid (e.g.,
tower shaft failure). If light-off does
not occur within 15 seconds, very
slowly advance the throttle until it
does occur; then stop all throttle
movement.
6.

Allow rpm to increase and begin to stabilize(approximately 10 seconds).

• Pause (2-3 seconds minimum) at thebackside of the BUC IDLE detent toallow FTIT and rpm to stabilize; thenrotate the throttle outboard andadvance slowly into BUC IDLE.
• Total time to advance the throttle fromIDLE to BUC IDLE will be a minimumof 30 seconds.
For a hot start (FTIT above BOOaC) or for astall/stagnation:
9.

Throttle - OFF.

10. STARTING FUEL switch - AUTO LEAN or
RICH.
If the throttle is advanced earlier, theengine may stall, precluding a successful airstart since rpm does not
accelerate past approximately 45
percent.

11.

Reinitiate BUC airstart.

Mter usable thrust is regained:

As rpm begins to stabilize:
7.

Throttle - Advance slowly to produce a steady
rpm increase (without valid rpm indication,
monitor FTIT).

Do not turn JFS or EPU off if indicated
rpm is below 60 percent with adequate
thrust (e.g., tower shaft failure).

• If above 25,000 feet and/or above 375 knots,
maintain throttle at IDLE until BUC airstartenvelope is entered.

13. EPU switch - OFF, then NORM.

• Advance the throttle slowly and smoothly tothe backside of the BUC IDLE detent toproduce a steady rpm increase similar to anormal UFC start.

14. ADI - Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to TOTALINS FAILURE, this section.

12. JFS switch - OFF.

~

".

~
~
~
~
~
~
~

~~~~.I.1.1I'I'I'I'I'I'I~

~""""IIIIII~~~

T.0.1F-16A-l

I

WAR N IN GI

1 AUX fl . . .
. h d
on. y
. ag IS I? VI~W,. PItc anroll attItude InformatIOn IS lIkely to ~eerroneo.us due to INS autorestart m
the .attItude mode when other thans~ra1ght ~~d lev~l, un accelerated
flIght condItlons eXIsted.
. d
15 . Th ro ttl e - A s reqUIre .
@ If• If nozzle is open greater than 50
percent while in BUC, there may not be
sufficient aerodynamic loading for thenozzle to close. To increase aerodynamic loading, decrease altitude, increaseairspeed above 250 knots, and increase
throttle setting.

• Refer to Section V for BUC operating
limits.
• Maintain 80 percent rpm mInImum
above 15,000 feet and 70 percent rpm
minimum below 15,000 feet or 700°C
FTIT minimum with tower shaft
failure until landing is assured. Do not
exceed 96 percent rpm.
• Avoid rapid throttle movements.
• Use 5 seconds minim urn for throttle
movements between BUC IDLE andMIL or MIL and BUC IDLE.
• AB operation is prohibited in BUC.
• An SFO is not recommended with theengine operating satisfactorily in
BUC.
16.

Land as soon as possible.

17. Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINEENGINE MALFUNCTIONS I PW2201

The EDU compares expected versus actual engine
operation. The purpose of the EDU MFL is to provide
maintenance personnel with an early indication of anengine condition which requires correction. No action
is required for an engine MFL at anytime during aflight.

Low altitude, for engine malfunction purposes, isgenerally defined as 10,000 feet AGL or below. If an
engine malfunction is suspected, the initial reaction
should be to trade excess airspeed for altitude. Unlessa suitable airfield is within gliding distance, turns
should be avoided as they decrease the amount oftime/altitude available to successfully recover engine
performance or prepare for ejection. Optimizing theexchange of airspeed for altitude must be a priority
action for any engine malfunction. Above 350 knots,
more time is available by performing a zoom climb.
usmg
a 3g pu 11 up to 30 -d egree cl'1m b untl'1
approaching the desired airspeed (use approximately
50 knots lead point) and then initiating a zero-g
pushover. Below 350 knots and above the minimum
recommended ejection altitude, more time isavailable by performing a constant altitude deceleration to the desired airspeed. If below the minimum
recommended ejection altitude and below 350 knots,
primary concern should be to trade excess airspeed
for altitude in preparation for ejection. If appropriate,
jettison stores as soon as possible.
For any situation where automatic activation of the
EPU is relied upon, verify that the EPU run light is
on to insure that the EPU has started. If the EPU run
light is off, position the EPU switch to ON.
Idle thrust in SEC during ground operation isapproximately twice that in PRI. After landing in
SEC, consider minimizing taxi distance and consider
following HOT BRAKES procedures, this section.

Engine Fire I PW2201

Generally, the first indication of fire in the engine
compartment is the ENG FIRE warning light.
Abnormal fuel indications (quantity/flow) may also
be present. FTIT probably will not be higher thannormal. Explosions, vibrations, or engine instrument
fluctuations are usually indicative of a serious engine
problem; engine failure may be imminent. Immediate
action should be taken to reduce thrust to theminimum practical level after attaining safe ejection
parameters. If within gliding distance of a suitablerunway, consider shutting the engine down. Sufficient time should exist to analyze the situation andmake an ejection versus land decision. The ejection
decision should be based on visual and/or cockpit
indications that the fire is persisting. Cockpit
indications include continued illumination of the
ENG FIRE warning light and subsequent illumination of FLCS status lights/degraded flight controls or
subsequent loss of either hydraulic system.

IIIIIIIIIIII~~~
Fires can also occur in the exhaust nozzle area whenusing AB. These fires are the result of portions of thenozzle failing which allows the AB plume to burnthrough the nozzle. Ventilation should inhibit forward
movement of the fire into and through the engine bay.
Since these fires are aft of the detection circuit, the
ENG FIRE warning light will not illuminate.
Additionally, the nozzle position indications arenormal, and there are no vibrations or instrument
fluctuations. In most cases, these AB-related nozzle
fires are detected by someone outside the aircraft(wingman, tower, etc.). When operating in AB and afire is reported at the rear of the aircraft, retard
throttle below AB immediately. This action should
extinguish a nozzle fire within approximately 30-45
seconds and minimize damage to the aircraft skin,speedbrakes, nozzle, and flight controls; however,nozzle damage may result in a noticeable thrust loss.
A failure that causes an oil leak may also result in an
oil-fed fire in the AB section. The fire may continue forseveral minutes after the engine fails or is shut down
(until the oil supply is exhausted). Since this fire is
likely to be contained within the engine, the ENG
FIRE warning light does not illuminate.

Iffire indications cease:5.

OVERHEAT Caution Light I PW2201

Detection of an overheat condition in the engine
compartment, ECS bay, MLG wheel wells, or EPU
bays illuminates the OVERHEAT caution light.
Accomplish as many of the following as required toextinguish the caution light. If the light goes off,
verify the integrity of the detection circuit bydepressing the FIRE & OHEAT DETECT button andland as soon as possible.
If OVERHEAT caution light illuminates:
1.

Throttle - Minimum practical.

2.

FIRE & OHEAT DETECT button - Depress.
Determine if fire detection circuit is functional.

2.

3.

'"

3.

Throttle - Minimum practical.
If fire occurred in AB, ENG FIRE warning
light may not illuminate. Fire should
extinguish after throttle is retarded; however, nozzle damage may result in lower thannormal thrust.

If the EPU was activated for the FLCS

BATT lights, be prepared to reactivate
the EPU if the battery discharge lights
reilluminate and remain on for longer
than 15 seconds.
If OVERHEAT caution light remains on (or detectcircuit checks bad):OXYGEN - 100%.

5.

AIR SOURCE knob - OFF.
External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tanks to decreasedrag if range is critical and the ECS cannot be
turned on for short periods of time to transfer
fuel.

I

FIRE & OHEAT DETECT button - Depress.
Determine if fire detection circuit is functional.

Eject.

I

m

6.
5.

WARNING

With the ECS shut down or the AIR
SOURCE knob in OFF or RAM, the g-suitand PBG is disabled.
does not inflate

If fire persists:

_

_

-.~

"

4.

If ENG FIRE warning light goes off:4.

EPU switch - OFF (if feasible).
If the EPU was manually turned on, consider
turning it off to determine if it is the source ofthe overheat condition. If the OVERHEAT
caution light remains on, the EPU should beturned back on.

IWARNING I

Climb.
Maintain takeoff thrust until mInImum
recommended ejection altitude is attained
and then throttle to minimum practical.

At a safe altitude:'"

IIIIIIIIIII

Abort.

Stores - Jettison (if required).

IIIIIIIIIII

circuit checks bad) and EPU is running:If takeoff is continued:
1.

-

If OVERHEAT caution light remains on (or detect

If on takeoff and the conditions permit:
1.

Land as soon as possible.

Descend to below 25,000 feet and reduce
airspeed to below 500 knots.

-

~

"

When airspeed is reduced and cockpit is depressurized:
7.

AIR SOURCE knob - RAM (below 25,000 feet).
External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tanks to decreasedrag if range is critical and the ECS cannot be
turned on for short periods of time to transfer
fuel.

I WARNING I
With the ECS shut down or the AIR
SOURCE knob in OFF or RAM, theg-suit does not inflate m and PBG isdisabled.
8.

Nonessential electrical equipment - Off.

NOTE
If in VMC and the ADI and HSI are notrequired for flight, the INS should beconsidered nonessential.
If OVERHEAT caution light still remains on (ordetect circuit checks bad):
9.

TANK INERTING switch - TANK INERTING even if Halon is not available.

10. LG handle - DN (300 knots/O.65 mach maximum). (Use DN LOCK REL button if required.)I

WARNINGI

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed, resulting in higher
than normal landing thrust.
11.

Land as soon as possible.

Engine Vibrations I PW2201

Some engines exhibit low frequency vibrations which
are non-damaging to both the airframe and engine.
The vibrations should disappear if engine rpm iseither increased or decreased.
Vibrations that change in intensity with throttle
movement and are present across the throttle/rpm
range may indicate a potential engine malfunction.

Oil System Malfunction I PW2201

An oil system malfunction is characterized by apressure (including fluctuations) below 15 psi at
IDLE or 30 psi at MIL, a pressure above 80 psi at any
thrust setting, pressure fluctuations greater than ± 5
psi at IDLE or ± 10 psi above IDLE, or by a lack of oilpressure rise when the throttle is advanced. The OIL
pressure indicator can be used as an early indicationof oil loss. An indication of excessive oil loss is the lackof oil pressure rise when the throttle is advanced in
the IDLE to MIL range. These conditions may notoccur until approximately one-half the usable oil islost. The HYD/OIL PRESS warning light may notilluminate until most of the usable oil is lost. At thefirst indication of an oil system malfunction, take
immediate action to land as soon as possible.

Climbing to a higher altitude allows higher cruise
airspeed and increases glide range. However, if the oil
malfunction is caused by an internal engine oil leak,the rate of oil loss is decreased at low altitude and low
throttle settings. Usually it is advisable to climb to areasonable cruise altitude. Once at altitude, retard
throttle to approximately 80 percent rpm and do not
move the throttle unless absolutely required. With
zero oil pressure, any throttle movement may cause
the engine to seize. Minimize maneuvering g tominimize loads. Plan an approach which allows aflameout landing from any position should engine
seize. Refer to SIMULATED FLAMEOUT (SFO)
LANDING and FLAMEOUT LANDING, this section.
The EPU should be manually activated; otherwise, if
the EPU does not start automatically when theengine seizes, the short time remaining before loss ofcontrol may be inadequate for recognition of the EPU
failure and corrective action. Monitor hydrazine use
after activating the EPU. If consumption rate is too
high, cycle EPU switch to OFF, and then NORM toconserve hydrazine. Be prepared to place the EPU
switch back to ON if the engine seizes.
If an oil pressure malfunction is suspected:
1.

Attain desired cruise altitude.
The rate of oil loss is decreased at low
altitudes and low throttle settings.

2.

Stores - Jettison (if required).

3.

Throttle - Approximately 80 percent rpm.

4.

EPU switch - ON.
Monitor hydrazine use. If consumption rate is
too high, cycle EPU switch to OFF, and thenNORM to conserve hydrazine. Be prepared toplace EPU switch back to ON if the engine
seizes.

5.

Throttle - Do not move until landing isassured.

If vibrations persist:1.

Throttle - Minimum practical.

2.

Land as soon as possible.

~~8~ c;;e~ ~ ~ ~ ~ ~ .., ~ ~ ~ ~ ~'"'""~""'"'-'-"'-'-'-"~~~~','''~'~

T.O. 1F- ~ 6A-l~~

ENGINE FAULT Caution Light LESS Iilli!, LESS tB.~• Throttle movement/rpm change may
caust' engine seizure.
• Do not start the ,IFS if engine seizure
has occurred or is anticipated. Starting the JFS may result in no brake/JFS
accumulator pressure for the brakes.
6.

7.

•

Land as soon as possibIE~.
PI an to fly an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING and FLAMEOUT LANDING, this section.
Refer to ACTIVATED
LEAK, this section .

The ENGINE FAeLT caution light illuminates as aresult of either loss of signals required by the DEEC
to control the AB or loss of mach number signal fromthe CADC to the DEEC. The AB may be partially(only segments 1 and 2 available) or completely
inhibited if the light illuminates as a result of AB
control signal failures; AB operation is not affected if
mach data is not received from the CADC.

SS~~::~~~~ FAULT caution light illuminates while

~s

1.

...~

Throttle - Retard to MIL.
Maintain until subsonic.

EPU/HYDRAZINE

Retardingr:!~::~IOW MIL ~ENGINE FAULT Caution Light [8NJ tL) I PW2201

Illumination of the ENGINE FAULT caution lightindicates that an engine PFL item was detected.
If ENGINE FAULT caution light illuminates:
1.

FALT ACK button - Depress and note PFL
display(s) on FCNP.

while supersonic may induce inlet
buzz which produces severe cockpit

If CA~~ra:~o: :;~;~:ab~:::~nec::::~~ lights~lum~n:~::h~:;~b;:~c~T switch -ELEC.
CADC caution light must reset before
ENGINE FAULT caution light will reset.

NOTE

If ENG 003 PFL is displayed or has been
displayed, MUX communication with
the EDU is no longer possible. Subsequently, if an engine PFL occurs, the
ENGINE FAULT caution light illuminates but cannot be reset and that PFL
cannot be displayed on the FCNP.
If ENGINE FAULT caution light does not reset when
the fault is acknowledged:
2.

Throttle - Midrange.

3.

Land as soon as practical.

3.

[BJ mEl AB RESET switch - AB RESET, then
NORM.

4.

Continue normal operation.

Refer to PILOT FAULT LIST I PW2201, this section.

3.

[8J rn:EJ AB RESET switch - AB RESET, then

ENGINE

NORM.
This action resets the DEEC and may clear
the engine PFL.
4.

FALT ACK button - Depress for fault recall.
The PFL is not displayed if it has cleared.

~

S

sS

s~~

...~

~

IfCADC and ENGINE FAULT caution lights do not ~reset:~

4.

~

Refer to CADC MALFUNCTION, this section.
All DEEC functions are working properly
except for supersonic stall protection.

:UbSO;CIMl AB RESET switcb _AB RESET, then
NORM.

2.

~~

IfCADC and ENGINE FAULT caution lights reset:

If ENGINE FAULT caution light illuminates when

If ENGINE FAULT caution light resets when thefault is acknowledged:

S
S
...~

O'vmOl

~
~

~

~

IL~

:r EN;::~n::~~:~:~:::r~~::rcsets:~

If ENGINE FAULT caution light does not reset:

~

3.

~

4.

Throttle - As required.
AB operation may be partially or totally
inhibited.
Land as soon as practl'cal.

~~___________ :n:2 _':87_~

~"IIIIIIIIIIIIIIIIII".
"~

TI)

~
~~

IF-loA-l

SEC Caution Light lpw2201

~

~

~

~

~~

~

~

~~

~

If the rpm indication is also zero or
erroneously low, the engine alternatormay have failed. If the engine is shut
down, an airstart may not be possible.

Automatic transfers to SEC after an engine
alternator failure may also cause engine stalls. The
com bination of stalls and an erroneously low rpm
indication may be incorrectly interpreted as astagnation. Confirm that a stagnation actually exists
before shutting down the engine.

Below 15,000 feet MSL, maintain 70
percent rpm minimum until landing isassured.

If the SEC caution light illuminates above 40,000 feet
MSL, minimize throttle movement until below 40,000
feet MSL. If the SEC caution light illuminates below
15,000 feet MSL and engine rpm is below 70 percent,slowly advance the throttle to achieve a minimum of
70 percent rpm. If the SEC caution light illuminates
while supersonic, do not retard throttle below MIL
until subsonic. During landing in SEC, idle thrust isapproximately twice that in PRI with a normal nozzle
because the nozzle is closed. Refer to figure 5-3 forthrottle limitations.

~
~

~
"~
~

~
~
:~
~ :f SE~::~::il:n~i~:t ~~:m:::::: Wb~:::u:;~on~:til
NOTE

The ENG CONT switch should not be
returned to ~ rn£] PRI, ~ NORM after
landing in an attempt to open thenozzle and decrease thrust.

~

subsonic.

~~

G~~

~

Retarding the throttle below MIL
while supersonic may induce inlet
buzz which produces severe cockpit
vibration and probable engine stalls.

"~

When subsonic or if SEC caution light illuminates

.,~

~~~

~

"~
~~

~

I

IWARNINGI

Illumination of the SEC caution light indicates thatthe engine is operating in SEC. If the ENG CONT
switch is in ~ mEl PRI, [ID3J NORM and the SEC
caution light is on, an automatic transfer to SEC has
occurred. The transfer may be due to a DEEC
malfunction, the DEEC sensing the loss of a critical
input signal to the DEEC, or loss of power to the
DEEC (engine alternator failure).

while subsonic:
2.
Throttle - Verify engine responds to throttle
movement; set as required.
AB operation is inhibited. Above 40,000 feet
MSL, minimize throttle movement. Below
15,000 feet MSL, ifrpm is below 70 percent,slowly advance throttle to achieve a
3-88mini~:o~:~:: percent rpm

3.

ENG CONT switch - SEC.

4.

Land as soon as practical.
During landing in SEC, idle thrust isapproximately twice that in PRI with anormal nozzle. [N§ ~ Consider using thedrag chute.

If engine is operating abnormally in SEC:

5.

Refer to ABNOR1'vIAL ENGINE RESPONSE

I PW2201, this section.
FTIT Indicator Failure I PW2201

Certain failures can cause erroneous FTIT indications above 1000°C and illumination of the ENGINE
warning light. If all other engine indications arenormal and the engine responds normally to throttle
movement, the engine should not be shut down.
Routine missions should not be continued since FTIT
cannot be monitored.
Zero RPM/Erroneous RPM Indication I PW220 I

If the RPM indicator displays a zero or erroneous
indication while other engine instruments indicate
normal operation, the cause is loss of power to theindicator or indicator failure. Loss of power to theindicator causes the ENGINE warning light toilluminate. RPM indicator failure may not cause
ENGINE warning light illumination.

IWARNINGIAssume engine alternator is inoperative or malfunctioning. If the engine isshut down, an airstart may not be
possible.

~"~":~''l~'~~," "~~"' "' "' "" ' ' '' "' "~~~~.."f'....

lfth\ RPl\I II1dic:ttor ell ~pb:,:- a Z('I"O or (rronc'()us]y
low indication aC(,f)mpanied b:- an automat ic transfer
to SEC (SEC caution light illuminated!, the ('ngine
alternator may have failed. Sincp rpm cannot bemonitored. roubne missions :"hould not be continued.
If SEC caution light is illuminated:
1.

Go to SEC CAUTION LIGHT lEW.i~ol, thissection,If SEC caution light is not illuminated:

1.

Land as soon as practical.

Abnormal Engine Response I PW220!

Refer to LOW THRUST ON TAKEOFF OR AT LOW
ALTITUDE (NON-AB) I PW2201, this section, if
appropriate.
Abnormal engine response is varied and generallyindicated by abnormal thrust in relation to throttle
position, engine oscillations (either continuous,
momentary, or recurring), a complete lack of engine
response to throttle movement, autoaccelerationldeceleration, exhaust nozzle failure, or insufficientthrust. The DEEC detects and automaticallyattempts to take corrective action for engine
malfunctions. This action may result in partially or
totally inhibited AB operation or in engine control
being transferred to SEC. The action taken by the
DEEC may be indicated by illumination of either orboth the ENGINE FAULT and SEC caution lights ~fIB and by ENG 087 PFL being displayed on the FCNP.
If thrust is too low to sustain level flight, turnimmediately toward the nearest suitable runway andestablish 250 knots airspeed. Consider jettisoning
stores to increase flying time available to complete
actions designed to restore usable thrust and improverange in the event those actions are unsuccessful.

NOTE
In a partial thrust situation, thrustavailable may increase as altitudedecreases. 250 knots approximates theairspeed at which thrust required forlevel flight is the lowest.
A throttle linkage problem should be suspected if
throttle movements in both PRI and SEC produce
either no rpm change or an rpm increase but no rpm
decrease. In either case, the OFF position does notshut down the engine. If the throttle is stuck andthrust is suitable for sustained flight, attempts to

fn'0 tllP tlm,ttk ,houlrl bp delay,·d unt il \\ ithin

gliding distance of;l suitable landing field. If throttle
i~ stuck in AB, placing thl' ENG COKT switch to SEC
terminates AB and provides SEC MIL thrust. If
throttle is stuck or otherwise prevented from normal
movement, control might be regained by depressing
the cutoff release, rotating the throttle outboard,SS

S::: ::;:::n~:ye::~ls::~:O::iCh prevents the engine S
SSSfrom reaching normal rpm and FTIT levels when thethrottle is advanced. This rollback is generally
caused by the DEEC sensing an out-of-limits
condition and may not be accompanied by a SEC
caution light. For this situation or any abnormal
engine response below AB, follow the procedures ofthis section until the situation is corrected.
If thrust is too high to permit a safe landing, use
excess thrust to climb and maneuver toward thenearest suitable airfield. Once high key for a flameout
landing is assured, follow procedures as outlined in
FLAMEOUT LANDING, this section. Activate the
JFS and EPU and then shut down the engine as soon
as landing is assured (normally high key) by placing
the throttle to OFF or, if necessary, by placing the
FUEL MASTER switch to OFF.
If the decision is made to manually select SEC while
subsonic below 40,000 feet MSL, set the throttle tomidrange before positioning the ENG CONT switch
to SEC. If autotransfer to SEC occurs below 15,000
feet MSL and rpm is below 70 percent, slowly advancethe throttle to achieve a minimum of 70 percent.
During landing in SEC, idle thrust is approximately
twice that in PRI with a normal nozzle because thenozzle is closed.

~

~
~
~S

I~

S

IWARN'NGI• Failure to monitor sink rate and height
above terrain while applying low
thrust recovery procedures can resultin ejection outside ejection seat performance envelope.
• If the throttle is stuck and thrust issuitable for sustained flight, attempts
to free the throttle should be delayed
until within gliding distance of asuitable landing field.

• Jettison stores when necessary toincrease flying time available toaccomplish actions designed to restore
thrust.

I~~~~__________ :n :9_~g:_

1'"11111111111111~1~11~~~.

~1.0. 1F- 16A- 1
NOTE

~~

• Transfer to SEC removes stall recovery
logic. If SEC is selected while theengine is stalling, a stagnation may
occur.

~

landing in an attempt to open thenozzle and decrease thrust.

9.

ENG CONT switch - SEC.
Transfer to SEC while supersonic should beaccomplished with the throttle at MIL; if thethrottle can not be retarded to MIL, transfer
to SEC is permissible with the throttle in AB.
Subsonic transfers to SEC below 40,000 feet
MSL should be accomplished with thethrottle at midrange or above.

~~1.

Throttle - MIL.

Retarding the throttle below MIL
while supersonic may induce inlet
buzz which produces severe cockpit
vibration and probable engine stalls.

~~ ;~~~~~~~:~:;, ~~~~s~~~~: is suspected to be failed
2.

Refer to NOZZLE FAILURE I PW2201, thissection.

~~ If problem still exists:
3.

AB RESET switch - AB RESET, then NORM.

4.

Airspeed - 250 knots (if thrust is too low to

~

~~
"~
~

~~

Below 15,000 feet MSL, maintain 70
percent rpm minimum until landing isassured.
10.

Throttle - Minimum practical.

If current SEC thrust will allow a safe landing:11.

Land as soon as practical.
During landing in SEC, idle thrust isapproximately twice that in PRI. ~ [f!Q]
Consider using the drag chute.

sustain level flight).
Ifproblem still exists::. ::;;t:~~ ::~:h _OFF

Stalls may be caused by the anti-ice valve
failing to close at high throttle settings (above

7.

~
~~

~

sufficient thrust.

3-90

An SFO is not recommended if engine
is operating satisfactorily in SEC.
When landing is assured:
12.

midrange).
Throttle - Slowly advance to minimum practical.
Attempts to establish a minimum practicalthrottle setting that provides sufficientthrust may result in repeated stalls that clear
when the throttle is retarded. Note stalled
RPM/throttle position and attempt to establish a lower throttle setting that provides~~

If suitable thrust cannot be attained or thrust is toohigh to permit a safe landing:Throttle - Midrange.

If in AB or supersonic:~

Land as soon as possible .

8.

~

"~

8.

• The ENG CONT switch should not be
returned to [K] ill!] PRI, ~ NORM after

"~

~
~

If current thrust will allow a safe landir ~:Change 14

Throttle - Verify engine responds to throttle
movement; set as required.

If suitable thrust cannot be attained:11.

ENG CONT switch - [K] mEl PRI, ~ NORM.

12.

Throttle - AB (if required to sustain level
flight).

13.

Land as soon as possible.

I:~r.~~~-~~~~,
IO. 1 F-16A- 1

If thrust i;:; too hi;:h tn !,crmd a safl.' land,ng:
NOTE
If throttle is stuck, control might beregained by depressing the cutoff
release, rotating the throttle outboard,and applying necessary force.
11.

Plan a flameout landing. Refer to FLAMEOUTLANDING, this section.

II

WARNING

Do not start the JFS if engine seizure
has occurred or is anticipated or ifengine failure is a result of fuelstarvation. Starting the JFS may
result in no brake/JFS accumulator
pressure for the brakes.
When prepared to land (normally high key):I

I

WARNING

Delaying engine shutdown can resultin a long, fast landing. Wheel braking
is less effective due to lack of WOW andthere is an increa~ed probability of a
missed cable engagement.
12. Throttle - OFF.
If throttle is stuck or engine does not respond,
shut down the engine with the FUEL MASTER
switch. At MIL, the engine flames out inapproximately 6 seconds; at IDLE, the engine
flames out in approximately 45 seconds.
13. Hook switch - DN (if required).

I WARNINGI

The hook may miss the cable if theaircraft is not slow enough to compress the ]\ILG struts sufficiently tomake WOW or if forward stickpressure is held.

Nozzle Failure Ll'w22!iExhaust nozzle malfunctions and nozzle control
system malfunctions can result in the nozzle being
too far open or too far closed. These malfunctions
can result in loss of AB capability, engine stalls, or
low thrust. Separation of the nozzle assembly fromthe engine is also possible and results in low thrust.
The ENG 088 PFL is displayed for failed
open/missing nozzle events.
A failed closed nozzle results in normal thrust below
AB and stalls when AB is attempted.

-....~

~
....
S~

S
S
s··
..
~

..

S

Low or insufficient thrust can be caused by a failed~
open, damaged, or missing nozzle or a nozzle control
system malfunction. If thrust is too low to sustain
level flight, turn immediately toward the nearest
suitable runway and establish 250 knots. With a
missing nozzle, level flight may not be possible.....'111
above 8000 feet MSL.~
1IIr..~
Thrust available should increase as altitudedecreases. The airspeed at which thrust required
for level flight is the lowest is approximately 250
knots.

..:~

Indications of a nozzle loss are as follows:~• An initial loud bang or pop, similar to acompressor stall, but rpm is stable above 60
percent; in PRI MIL, engine rpm is approximately
5 percent lower than normal and FTIT isapproximately 250°C lower than normal; fuelflow is lower than normal; the nozzle is likely
indicated in the full closed position; and thrust is
decreased. Malfunctions of the exhaust nozzle
control system may have symptoms similar to a
missing nozzle, but the nozzle may indicate fullopen since the nozzle actuation system is intact.
• Presence of the ENG 088 fault indicates that the
DEEC has detected the malfunction and has
activated logic to increase the thrust available in
PRJ. AB is inhibited. Remain in PRI if possible, as
it should provide a sufficient level of thrust while
also maintaining safe engine operation.
• If level flight cannot be attained by 1000 feetabove minimum safe ejection altitude or minimum safe altitude with the ENG CONT switch in
PRI, select SEC. Set the throttle as required tomaintain 250 knots. Continuous operation above
850° C in SEC is likely to result in catastrophic
engine failure and fire in as little as 5 minutes.

~
~

~
~
~
,~

_________Ch:e: -. 3:1-.~

~~<~:A~f II~IIIIIIIIIIIII'"~~~If thrust is low and a failed open, damaged, or
missing nozzle is suspected:~

1.

Throttle - MIL or below.

~

2.

Stores - Jettison (if required).

3.

Airspeed - 250 knots.

~
."~

~~
~~

~
~
~

~~

~

field:
4.

~~~

Land as soon as possible. Plan a flameout
landing. Refer to FLAMEOUT LANDING, thissection.

7.

Land as soon as possible. Plan a flameout
landing. Refer to FLAMEOUT LANDING, thissection.

If unable to reach a suitable landing field and level

flight cannot be maintained by 1000 feet above
minimum recommended ejection altitude or minimum safe altitude, whichever is appropriate:5.

ENG CONT switch - SEC.

onl~:::elected

SEC should
when itbecomes apparent that sufficientthrust cannot be achieved in PRI. SEC
eliminates the additional thrust andthe engine protection benefits provided by the DEEC in PRI. The nozzle
loss logic holds the engine in PRI for~"~

trade altitude to reacquire 250 knots.
Do not descend below minimum
recommended ejection altitude or
minimum safe altitude, whichever isappropriate.

If thrust is sufficient to reach a suitable landing~~

With nozzle loss, catastrophic engine
failure and fire are probable withprolonged high power settings above
850°C FTIT while in SEC.

If airspeed drops below 250 knots,~~~~

IWARNING I

6.

T:~~~:I:~a:nr~qUired

to mamtain 250 knots inlevel flight above minimum recommended
ejection altitude or minimum safe altitude,
whichever is appropriate.

3-90 2

Change 14

AB Blowout/Failure To Light I PW2201

An AB blowout is indicated by the nozzle opening,then closing, after the throttle is advanced to AB. If
an AB blowout occurs and the throttle is left in AB,
the DEEC automatically recycles the AB up to three
additional times (each cycle indicated by the nozzle
opening and closing). An AB no-light is indicated by
the nozzle failing to start open within 5 seconds ofadvancing the throttle to AB (nozzle remains closed
or shows minimal movement). If an AB no light occurs
and the throttle is left in AB, the DEEC automaticallyattempts to relight the AB up to 3 times. The initialattempt and 3 subsequent no lights could take up to
20 seconds. A combination of no lights and blowout
recycles could take longer. If further AB attempts are
required and the DEEC has completed all recycle
attempts, then the throttle must be retarded to MIL
or below and advanced to AB.
If theAB blowout/failure to light was notaccompanied by an ENGINE FAULT caution light,
flight may be continued. If an ENGINE FAULT
caution light also occurred, refer to ENGINE FAULT
CAUTION LIGHT I PW2201, this section.

I I I I I~~~~~~~~!!~~ENGINE STALLS I PW2201

The three primary causes of a stall are inlet flow
distortion, AB instabilities, and hardware malfunctions. During normal aircraft operation, inlet flow
distortion severe enough to cause engine stall is notexpected. However, under some departure conditions,
inlet flow distortion may induce engine stalls.
Hardware-associated stalls may result from a failed
nozzle, control system malfunctions, anti-ice systemfailed on, or FOD. Stalls may also result if rapid
engine acceleration is attempted from an abnonnally
low SEC idle rpm (less than 70 percent).
Stalls may be caused by an anti-ice valve failed in the
open position at high thrust settings (throttle above
midrange). The engine should be operable with thiscondition by limiting throttle position to midrange or
below. If flight conditions permit, place the ANTI ICE
switch to OFF.
The first indication of an engine stall at high thrust
settings may be a loud bang or pop. At lower thrust
settings, the first indication may be loss of thrust,lack of throttle response, or decreasing engine rpm.
When a stall is sensed, the DEEC cancels the AB (ifthrottle is in AB range), opens the nozzle, anddecreases fuel flow until the stall clears. FTIT andNOZ POS may fluctuate in response to the stallrecovery signal. If the engine auto transfers to SEC,
automatic stall recovery and overtemp protection are
not available. A malfunction such as engine internal
damage, or primary control system failure could
result in a stall, an automatic SEC transfer, andpossible FTIT overtemp. Throttle reduction isappropriate as a first response to clear any engine
stall.
If the engine stalls at low altitude, an immediateclimb should be initiated, and stores jettison should
be considered. Retarding the throttle may clear thestall. During a high thrust stall that is self recovering,
there will be an immediate thrust loss. In PRI, theDEEC gradually restores thrust to the original level.
If engine response at low altitude is not sufficient tomaintain or gain altitude and a suitable landing field
is not available, ejection may be required.

I

If a stall occurs at MIL or below, retarding the throttle

may clear the stall. Further throttle movement
should be limited to midrange or below.
AB-Associated Engine Stalls I PW2201

AB-associated stalls are normally accompanied by aloud bang or pop and a series of fireballs from theengine exhaust and occasionally the engine inlet. Thisis followed by an erratic flame from the engine exhaust

if the stall is nonrecoverable. These characteristics
could be mistaken for an aircraft fire. Whenever a stall
occurs while operating in AB, the DEEC automaticallycancels AB and activates stall recovery. This may be
accompanied by a nozzle swing to full open for a few
seconds and an associated temporary reduction ofthrust. The throttle should be snapped out of AB to
MIL. This action usually clears the stall and restores
nonnal operation; however, stalls may continue at MIL
and can be severe. They may be characterized by bangsor pops of low intensity or engine vibrations severe
enough to preclude reading engine instruments. Refer
to NON-AB ENGINE STALLS I PW2201, this section.

II

If a self-recovering AB sequencing stall occurs whentransitioning from region 3 while operating withapproved fuels other than JP-4, NATO F-40, or JET B
and no other abnonnal engine indication is observed,
the engine is safe to operate from IDLE to MAX AB.

~
~
~
~
~
~
~
~
~

Non-AB Engine Stalls I PW2201

Non-AB stalls may occur if the engine is malfunctioning, particularly during throttle transients near
IDLE. Non-AB stalls are often a symptom of a serious
engine problem. Non-AB stalls may be inaudible; thefirst indication may be a lack of throttle response
which may be difficult to differentiate from abnormal
engine response. However, non-AB stalls can also be
severe. They may be characterized by bangs, pops,
low intensity or severe engine rumble or vibration. A
momentary nozzle swing to near full open may occur,causing a temporary reduction in thrust, as the
DEEC activates stall recovery. An erratic orange-yellow flame from the engine exhaust may be present.
This exhaust flame should not be mistaken for anengine fire. If the stall is confirmed, the throttle
should be immediately retarded to IDLE which may
clear the stall. Further throttle movement should be
limited to midrange or below. If a stall occurs while
operating in SEC, do not advance the throttle until
engine rpm is stable at SEC idle.
Engine Stagnation I PW2201

Stagnations are usually characterized by either
rising FTIT and decreasing rpm or rpm less than 60
percent and a lack of rpm response to throttle
commands or illumination of the ENGINE warning
light. These indications are usually accompanied bybangs, pops, low intensity engine rumble or vibration,
and/or an erratic orange-yellow flame from theengine exhaust. This exhaust flame should not be
mistaken for an engine fire. FTIT can either spike,steadily increase, or even decrease immediately
following a high thrust stagnation. During low thruststagnation, FTIT can stabilize in the engine normal
operating range of less than 980°C. Since FTIT can bedeceptive, low rpm (less than 60 percent) is generally

~

~
~

~

-

IIIIIIIIIIIIII~

the best indication. Once a stagnation occurs, there is
no way to recover normal engine operation except toshut down the engine and perform an airstart.
Allowing a stagnated engine to run results indecreasing rpm, increasing FTIT, a loss of thrust, lossof altitude, and engine damage. As soon as stagnationis confirmed, immediately retard the throttle to OFF.
There is every reason to expect a normal airstart andnormal engine operation if the engine is shut down toclear a stagnation, especially if the engine was shut
down without allowing FTIT to remain at hightemperatures for an extended period.

IWARNING

Engine Stall Recovery I PW2201

If an AB stall(s) occurs:
Throttle - Snap to MIL.

If AB stalls do not clear or stall(s) occurs below AB:
NOTE

Non-AB stalls may be inaudible.
2.

Throttle - IDLE.

3.

ANTI ICE switch - OFF when conditions permit.
NOTE

Stalls may be caused by anti-ice valve
failing to close at high thrust setting(throttle above midrange).
If stalls continue at idle and progress to a stagnation(engine rpm less than 60 percent with no rpm
response to throttle movement):
4.

Throttle - OFF. Initiate airstart. Refer to
AIRSTART PROCEDURES I PW2201, this section.

1

Shutting down the engine with anengine alternator failure (indicated byzero or erroneously low rpm, illuminated
SEC caution light, illuminated ENGINE
warning light, and normal thrust)
results in no ignition for an airstart.
If non-AB stall(s) clears:
5.

Throttle - Midrange or below.
'If a non-AB stall clears, maintain throttle at
midrange or below unless required to sustain
flight.

6.

Land as soon as possible.

1

Prolonged engine operation with FTIT
in excess of 1000°C can result insignificant engine damage and may
cause a nonrecoverable engine failure.

1.

IWARNING

If AB stall(s) clears:
2.

Throttle - As required.
• If a self-recovering AB sequencing stall occurs
when transitioning from region 3 while
operating with approved fuels other than JP-4,
NATO F -40, or JET B and no other abnormal
engine indication is observed, the engine is safe
to operate from IDLE to MAX AB.

• If an AB stall clears under conditions otherthan described in the preceding bullet, theengine is safe to operate in the IDLE to MIL
range, provided no other abnormal indication
is observed. Attempt further AB operation
only if needed to sustain flight.
INLET BUZZ I PW2201

Inlet buzz occurs at supersonic airspeeds if the engine
control system fails to maintain adequate engine rpm
when the throttle is retarded below MIL. Inlet buzz
causes moderate to severe vibration within the
cockpit and probably results in multiple engine stalls.
If inlet buzz occurs, the throttle should not be moved
until subsonic. Decrease airspeed to subsonic asquickly as possible by opening the speedbrakes andincreasing g. If engine stalls occur and persist, thethrottle should be retarded to IDLE when subsonic. If
the stalls do not clear, the engine must be shut down
and restarted.
ENGINE FAILURE OR FLAMEOUT I PW2201

Engine failures can result in rpm decrease with noabnormal vibration or sound (flameout), rpm
decrease with abnormal vibration andJor stalls, or
stable rpm with abnormal vibration andJor low
thrust.

If the engine names out, fuel starvation ormechanical failure has occurred. A flameout isindicated by a decrease in FTIT and engine rpm
decaying below approximately 60 percent. Loss ofthrust and lack of response to throttle movement
confirm the flameout. The ENGINE warning lightilluminates when engine rpm is below 55 percent.
Additionally, the MAIN GEN light illuminates below
45 percent rpm and the EPU should start running. Do
not mistake a loss of ECS noise as an engine flameout.
If the reservoir tanks do not contain fuel, an airstartis impossible. If fuel starvation was due to atemporary lack of fuel, restart should be possible. If
fuel quantities appear normal, the flameout may
have been caused by fuel contamination. In this case,retarding the throttle to OFF may clear thecontaminated fuel and allow an airstart.
Main fuel pump failure or tower shaft geartrain
failure also causes flameout. Both present similar
symptoms: an abrupt decrease of indicated fuel flow
to less than 500 pph; loss of main generator and FLCS
PMG and EPU activation; no throttle response; andillumination of the SEC caution light even though the
ENG CONT switch is in [!] rn:El PRI, rnBl NORM.
If only the main fuel pump has failed, the rpm
indication reflects a gradual spooldown. The JFS canbe started and the engine can be motored at
approximately 30 percent rpm. If the SEC caution
light remains on (with ENG CONT switch in ~ I:m
PRI, ~ NORM and engine rpm at 12 percent orabove), the engine probably cannot be restarted;
therefore, place primary emphasis on a flameout
landing while continuing airstart attempts. If unable
to make a flameout landing, refer to EJECTION, thissection.
Tower Shaft Failure I PW2201

Failure of the engine tower shaft or its associatedgeartrain results in loss of all rotation to the engine
gearbox and the ADG. Loss of rotation to the engine
gearbox renders the engine alternator, main fuelpump, and oil pump inoperative resulting in a zerorpm indication, zero oil pressure, illumination of the
ENGINE warning and SEC caution lights, andengine flameout due to fuel starvation. The initial
symptoms are similar to main fuel pump failure;
however, the primary difference is that the rpm andoil pressure indications drop immediately to zero witha tower shaft failure since the engine alternator andoil pump are not being driven. Additional symptoms
caused by loss of rotation to the ADG include loss ofhydraulic systems A and B, main generator, andFLCS PMG and subsequent activation of the EPU. It
may be possible to regain engine operation using the

JFS and performing a SEC airstart. The JFS drives
the ADG and the engine gearbox (through the PTO
shaft), restoring rotation to both hydraulic pumps,
FLCS PMG (at a reduced output), main fuel pump
(SEC caution light goes off in PRI until SEC isselected), engine alternator (cockpit rpm signal,
DEEC power, and engine ignition), and oil pump (oilpressure increases). Without the load of the engine,
the JFS produces an rpm indication fluctuating
between 30-50 percent which is the speed of theengine alternator, not the actual engine rpm. Thisrpm may be high enough to restore main generator
power; however, main genera tor power may cycle on
and off line with the rpm fluctuations. If the ENG
CONT switch is still in ~ rn:El PRI, [J!BJ NORM, theSEC caution light goes off when fuel pump pressure
is restored; however, a PRI airstart is not possible
since the rpm signal to the DEEC is in error. Perform
a SEC airstart. Since the JFS is not preserving rpm,
maintain 275 knots minimum during the airstart
attempt, which should assure adequate actual engine
rpm for the airstart.
Low Altitude Engine Failure or Flameout I PW2201

Refer to figures 3-10 and 3-11. Initial reaction to any
malfunction at low altitude should be to trade excess
airspeed for altitude. Higher altitude translates
directly to either additional time to achieve anairstart or to additional glide range to reach a suitablelanding field. At low airspeed, the climb may be only
enough to insure a safe ejection altitude. Above 350
knots, more time is available by performing a zoom
climb using a 3g pullup to 30-degree climb until
approaching the desired airspeed (use approximately
50 knots lead point) and then initiating a zero gpushover. Below 350 knots and above the minimum
recommended ejection altitude, more time isavailable by performing a constant altitude deceleration to the desired airspeed. If below the minimum
recommended ejection altitude and below 350 knots,
primary concern should be to trade excess airspeed
for altitude in preparation for ejection.
If required, jettison stores as soon as possible to aid ingaining or maintaining altitude and maneuver
toward a suitable landing field, if available. If thezoom results in an altitude below 5000 feet AGL,
there will probably be insufficient time to achieve anairstart prior to minimum recommended ejection
altitude. In that case, primary consideration should
be given to preparing for ejection; do not delayejection below 2000 feet AGL. If the zoom results inan altitude between 5000-10,000 feet AGL, there is
probably time for one airstart attempt prior tominimum recommended ejection altitude. This
attempt shall be performed in the control mode
selected by the DEEC.

n:T

TIONS:o

IR SEIZED ENGINEITAINED TO DESIRED KIAS

12

- - 2 7 5 KIAS

10170 KIAS
LINEAR INTERFOR AIR:DS BETWEENKIAS AND 275(USE ONLY FOR
AL AIRSPEEDS~75 KIAS OR

....I

~TION

c:I

cs: 8

I-

u.

I

o
o
o

~TER).

~ 6
:::I

I-

5

cs:
cs:

....I

E4
~

2

o.

o

160

200

220

240

260

300
320 340 360
INITIAL AIRSPEED - KIAS

280

Figure 3-10.

380

400

420

440

460

480

500
'F-'6X-'~028X®

s

Low Altitude Airstart Capability
DATA BASIS ESTIMATEDCONFIGURATION:• GW _ 23,000-25,000 L8

ENGINE F100-PW-220

CONDITIONS:• 30 ° DIVE TO DESCENT KIA5 OR 3G PULLUP
TO 30 G ZOOM CLIMB INITIATED FROM THE
AIRSPEED/ ALTITUDE EXISTING AT FIRSTRECOGNITION OF ENGINE FAILURE (60PERCENT RPM)

.01 _ 0-50
-lG - UP

• AIRSTART INITIATED WHEN RPM REACHES
25-50 PERCENT (10 SECONDS AFTERINITIATION OF DIVE OR ZOOMI

'0

~

'"4
~
o

8 6
w

g
~

~
4

4

~

E
~

2

o .j::i;J:::
INITIAL AIRSPEED - KIA5

• 45 SECONDS ASSUMED AFTERTHROTTLE ADVANCE TO ACHIEVE
USABLE THRUST• DESCENT AIRSPEED IS 200 KIAS
(SEC) tJFS RUN LIGHT ON)•

2.

Stores - Jettison (if required).
If stores jettison is attempted after main
generator drops off line but before EPU
generator powers the SMS (approximately 5
seconds delay), stores will not jettison.

NOTE
Visually confirm the stores havejettisoned and jettison again if required.
3.

Perfonn airstart (if altitude pennits). Refer to
AIRSTART PROCEDURES I PW2201, this section.

I WARNING I
Do not fly slower than 200 knots for
SEC JFS-assisted airstarts.

AI RSTARTS I PW2201
Refer to figure 3-12. Factors such as altitude,
airspeed, weather, etc., must be considered indetermining whether to try an airstart, accomplish aflameout landing, or eject. Jettisoning of stores
reduces altitude loss during an airstart and improves
glide ratio during a flameout landing.
Oil pressure is directly related to rpm. Do not confuse

a low oil pressure indication due to windmilling rpm
as an oil system malfunction.
If the engine has seized due to an oil system
malfunction or flamed out due to fuel starvation ormechanical failure, a flameout landing or ejection isrequired.
The most likely reason to perform an airs tart is thatthe engine has shut down due to a PRI system failure,
hardware failure, or stagnation. The DEEC assesses
any faults or internal failures and automaticallytransfers to SEC, if required. The first airstart
attempt should be made in the engine control mode
selected by the DEEC. When a tower shaft failure issuspected, perform the airstart in SEC. Procedures
for SEC and PRI airstarts are identical except for
ENG CONT switch position and airspeed.

3-96Change 8

There are critical requirements which apply to any
airstart attempt. The most important is engine rpm.
During a spool down airstart, if the throttle isadvanced from OFF to IDLE after rpm goes below 25
percent, light-off may not occur before rpm decreases
through 12 percent. Below 12 percent, the main fuelpump does not supply sufficient fuel to effect anairstart and engine ignition is not available. In
general, rpm spooldown rate can be decreased withincreased airspeed. As much as 350 knots or more isrequired to prevent rpm from decaying below 12
percent. If rpm is allowed to drop to near zero,
400-450 knots for 20-25 seconds may be required toregain 12 percent. If this situation occurs and theaircraft is not within gliding distance of a suitablelanding field, a 50-degree dive angle or greater shouldbe established to accelerate to 450 knots. Once at 450
knots, the dive angle may be reduced to approximately 20 degrees to maintain airspeed. As theengine rpm increases to near 12 percent, an airstart
may be attempted. If the dive is started from 20,000
feet MSL at 200 knots and zero engine rpm,
16,000-18,000 feet is required to regain 12 percent
engine rpm. If the engine is allowed to stop rotating,it may thermally seize after which it will not rotate
even with high airspeeds or by engaging the JFS. The
second requirement is engine temperature. Hightemperatures may result if the airstart is initiatedbefore the FTIT is allowed to decrease below 700°C.

IIIIIIIIIIIIIIIIIIIIII~
10. 1 F- 16A- 1

Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-220

CONDITIONS:- DRAG INDEX- KIAS
250==0- GW = 17,000 LB-NO WIND

... A

,,

50

,,

,,

,

30° DIVE

POINT B TO"
SL = 9 MIN/52 NM ------.... " ,40B~~~~~~~~~

PRI/SEC AIRSTARTENVELOPE

MIN KIAS(WITH JFS RUN
LIGHT OFF)PRI-250
SEC-275

I-

w

W

LI.

o 30
o

o....

w

C

:)

I- 20

MIN KIASI(WITH JFSI
I
RUN LIGHT
I
I
ON)
I
PRI-170 ----.t
SEC-200

5c:r:~---MARGINAL

I
I
I----L /

10~---- 400

/~'

KIAS

I
I

I

I

I

I

I

I
I

I

O~------~~~------------~~----~--------~------~------~

o

0.2

0.4

0.6

0.8

1.0

1.2

1.4

MACH NUMBER
OPTIMUM FLIGHT PATH DURING AN AIRSTART (TYPICAL)Engine out descent flight path maintains the aircraft in the required airstart envelope.
A 30-degree dive to descent KIAS is used to approach the airstart envelope(Point A to BI.

1 F-16X-1-00398 @Figure 3-12. (Sheet 1)

Change 4

3-97

II:
-

-

~""""""""""~~T. O. 1F- 16A- 1~

Engine RPM and FTIT Response During
Spooldown and Airstart~

ENGINE F 1OO-PW-220

~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~

--

1000

u

t

~

- - - - - - -- - - -- -

800

°
I

-----

....

t 600

DEEC TRIMS FUEL FLOW TOPREVENT OVERTEMPERATURE

-

-

-

-

-

-

- 800°C--

700°C----- - - - - - - - -THROTTLE
OFF

400
THROTTLEOFF

STABILIZEDIDLE THRUST100
THROTTLE
IDLE

80

I

60

:E
~

40

THROTTLE
IDLE
50%

25%
20
o~~--~--~~--~--~--~~--~--~~--~--~~--~~

o

20

40

60

80

100

120

140

160

TIME - SECONDSENGINE RPM AND FTIT RESPONSE DURING A HUNG AIRSTART FOLLOWINGA HIGH THRUST STAGNATION (PRI ONLY) (TYPICAL)DEEC airstart overtemperature protection logic limits FTIT to approximately 700°C
to prevent a hot airstart by trimming fuel flow which results in engine rpm
decreasing, hung, or slowly increasing. During a normal airstart, FTIT may be
700°C but engine rpm is increasing normally.

1 F-16X-1-o040X @Figure 3-12. (Sheet 2)

3-98

Change 1

,. ~ ~ ~ ~.I.I ~ ~ I I I I!!!!,Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-220

1200
1000 ---~- ..

~ 800

...I

700°C

fE 600400
THROTTLE
OFF

200

THROTTLE
IDLE
STABILIZEDIDLE THRUST100
LIGHT-OFF80

DELAYED FTIT
AND ENGINE RPMRESPONSE

60

50%

40
25%
20

o~----~----~~----~----~------~----~----~----~

o

10

20

30

40

50

60

70

80

TIME - SECONDSNORMAL AIRSTART ENGINE RPM AND FTIT TIME TRACES FOLLOWING A
HIGH THRUST STAGNATION (TYPICAL)Light-off occurs within 15 seconds after the throttle is advanced to IDLE.
However, engine rpm and FTIT turnaround are slow, making light-off subtle
or difficult to detect. Engine rpm stabilizes momentarily after light-off and
FTIT may increase, stabilize, or decrease as the engine rpm increases
normally to idle.

1F-16X-1-0041 X ®

Figure 3-12. (Sheet 3)

,
,
,
,
,
,
,
,
,
,
,
,
,
,
,
,
,

f";~!r~~ ~.~ I I ' I ' I ' I ' ~ ~ . .

--Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-220

1000

800
CJ

°
I 600
t::

t:

350
KIAS
35,000FEET

250 KIAS
10,000 FEET

400

200
100
350 KIAS
10,000 FEET

80

"f1.

250 KIAS
35,000 FEET

60

~

_____
500!

I
:E
~

cc

40

20

o

350 KIAS
35,000 FEET
25%

250 KIAS
10,000 FEET

I

o

I
20
40
TIME - SECONDS60

RATE OF ENGINE RPM AND FTIT DECAY
AS A FUNCTION OF ALTITUDE AND
AIRSPEEDAt low altitude, regardless of airspeed,engine rpm decreases rapidly which
decreases the time available in theairstart window (FTIT less than 700°C
and engine rpm between 25-50 percent).

1F-16X-1-o042X@

Figure 3-12. (Sheet 4)-'ii"IILlILIII~

r~~~~~~~~~~~~'-'-111111
T.O. 1F-16A-l

Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-220
1200

1200

1000

1000

u

°1

CONDITIONS:
250 KIAS
25,000 FEET

800

~ 800

l-

t 600

700°C

....I

t: 600400
200~--------~--------~--------~

1-THROTTLE OFF
1000°C
2 -THROTTLE OFF

400

2

100
200
80
o~
1

100
60

:!
~

THROTTLEADVANCEAT 25% RPM
REQUIRED80
40
o-:1!.

20

o~----------~----------------------~
o
10
20

30

___j____

60

~O~

40

TIME - SECONDS20
ENGINE RPM AND FTIT RESPONSEFOLLOWING A HIGH THRUST STAGNATION

25%------1-2 - - - - SPOOLDOWNAIRSTART WINDOWooFollowing a high thrust stagnation, engine rpm
rapidly decreases and FTIT varies within theinitial few seconds as shown by three examples.

20
40
60
TIME - SECONDS

80

Example 1 shows a rapid decrease in engine
rpm and steady increase in FTIT.

RATE OF ENGINE RPM AND FTIT DECAY AS A
FUNCTION OF THROTTLE MOVEMENT DURING
AND FOLLOWING A HIGH THRUST STAGNATION

Example 2 shows a rapid decrease in engine
rpm and a momentary step increase in FTIT
which, within a few seconds, turns into a
steady increase in FTIT.

Allowing the engine to remain stagnated reduces
the spooldown airstart window as shown by two
examples.

Example 3 shows a rapid decrease in engine
rpm and an immediate drop in FTIT of up to
200°C as a result of momentary main
combustor blowout. Light-off always reoccurs
within a few seconds causing FTIT to rise
rapidly.

Example 1 shows the throttle retarded to OFF
when FTIT reached 1000 0 C. The spool down
airstart window indicates there is no significant
engine turbine distress.
Example 2 shows the throttle retarded to OFF
when FTIT reached 1200°C. There is nospooldown airstart window since minimum engine
rpm was reached before FTIT cooled below 700°C.
Possible engine turbine distress may affect airstartsuccess.
1F-16X-1-o043X@

Figure 3-12. (Sheet 5)

,

-

-

-

-

-

-

-

~IIIIIIIII""~

,,-;~~~~~~~~~~~~~~-Engine RPM and FTIT Response During
Spooldown and Airstart
ENGINE F100-PW-220

1200

1200
1000

1000

CJ

°

I

800
CJ

° 800

l-

t:
600
u.

I

~

~ 600

400
200--------------------~--------~

2

1-THROTTLE OFF

900°C

400

2 -THROTTLE OFF

100
200
80

I

:!:
~

100

60

THROTTLEADVANCEAT 25% RPM
REQUIRED

80
40
60

20
o----------~--------~--------~o

10

TIME -

20

30

40

SECONDS

20ENGINE RPM AND FTIT RESPONSE DURING
AND FOLLOWING A LOW THRUSTSTAGNATION
During a low thrust stagnation, engine rpm
decreases while FTIT increases. The rate at
which engine rpm decreases and FTIT increases
varies with altitude and airspeed. Engine rpm
decreases at the same rate regardless of FTIT
response. However, FTIT following a low thruststagnation often levels off at 800 0 e - 1000o e.

25%------12------SPOOLDOWNAIRSTART WINDOW

o

o

20

40

60

80

100

TIME - SECONDSRATE OF ENGINE RPM AND FTIT DECAY AS A
FUNCTION OF THROTTLE MOVEMENT DURING
AND FOLLOWING A LOW THRUST STAGNATION
Allowing the engine to remain stagnated reduces
the spooldown airstart window as shown by two
examples.
Example 1 shows the throttle retarded to OFF
with engine rpm below 60 percent and when FTIT
reached 900 o e. The spooldown airstart window isshown above.

-

,,
,

CONDITIONS:
250 KIAS
25,000 FEET

Example 2 shows the throttle retarded to OFF
when FTIT reached 1000°C. There is no spool down
airstart window. The operations are the same as
those in RATE OF ENGINE RPM AND FTIT DECAY
AS A FUNCTION OF THROITLE MOVEMENTDURING AND FOLLOWING A HIGH THRUSTSTAGNATION, example 2.
1F-16X-1--o044X@

Figure 3-12. (Sheet 6)

3-102

Change 1

"""IIIIII~~~~To meet these requirements, an airstart should beinitiated at 50-25 percent rpm with FTIT below
700°C. These parameters can usually be achieved bymaintaining 250 knots minimum for PRIor 275 knots
minimum for SEC below 40,000 feet MSL. This
airspeed will not maintain rpm above 12 percent;however, it provides the best tradeoff between therate of rpm spooldown and loss of altitude. If it
appears that engine rpm will drop below 25 percent,increase airspeed if feasible and advance the throttle
to initiate the airstart even if FTIT is above 700°C. A
hot start may result; however, rpm is preserved. If a
hot start does result, retard the throttle to OFF andreattempt the airstart when rpm and FTIT are within
limits. Generally, higher airspeeds increase airstart
reliability by slowing rpm decay which allows the
FTIT to decrease more rapidly.
At low altitudes, however, higher airspeeds do notsignificantly affect the rpm decay. Maintain airspeed
at 250 knots minimum for PRIor 275 knots minimumfor SEC for spool down airstarts and at maximum
range or maximum endurance airspeed for JFSassisted airstarts (200 knots minimum for SEC).
High Altitude Airstart Considerations I PW2201

Refer to figure 3-12. At high altitudes, dive atapproximately 30 degrees to gain or maintain 250
knots in PRIor 275 knots in SEC below 40,000 feet.
Once established, approximately 5-10 degrees of diveshould maintain airspeed. Note that airspeed cannot
be reduced to less than 250 knots for PRIor 275 knots
for SEC until the JFS RUN light is on. Unless anairstart is obviously impossible (total lack of fuel,engine seizure, etc.), do not become tempted toestablish a maximum range or maximum endurance
glide. The first consideration should be an immediatespool down airstart attempt even if the engine failed
for no apparent reason. If airstart airspeed is not
maintained, rpm decreases at a faster rate. The only
airstart option available is then a JFS-assisted
airstart. Time constraints due to EPU fuel consumption must be considered. A maximum range or
maximum endurance glide from above approximately
35,000 feet may exhaust EPU fuel prior to landing.
(Refer to T.O. 1F-16A-1-1, figure C6-3.)
If the first attempt is in the control mode selected by
the DEEC and all parameters (FTIT, rpm, andairspeed) are met and the start is unsuccessful,reattempt airstart with the ENG CONT switch in
SEC.
When below 20,000 feet MSL, turn the JFS on.
Activating the JFS above 20,000 feet is prohibitedsince successful JFS startimotoring of engine isunlikely and the brake/JFS accumulators will be
depleted. If the JFS RUN light is on, airspeed may be
reduced to achieve maximum range or maximum
endurance (200 knots minimum for SEC airstarts).
With the JFS running, EPU fuel consumption is also
reduced.
Low Altitude Airstart Considerations I PW2201

Due to the limited time available and the rapid rpm
spooldown rate at low altitude, some additional
considerations are required. Below approximately
10,000 feet MSL, rpm decreases rapidly regardless ofairspeed and remains between 50-25 percent for only
5-10 seconds; therefore, rpm should be closely
monitored. Advance the throttle to initiate theairstart before rpm goes below 25 percent regardlessof FTIT indication. This action should insure thatlight-off occurs prior to 12 percent rpm. Start the JFS
immediately after advancing the throttle (if airspeed
is below 400 knots).
Following a zoom climb, plan to arrive at 250 knots for
PRIor 275 knots for SEC; airspeed may be reduced toachieve maximum range or maximum endurance(200 knots minimum for SEC airstarts) only after the
JFS RUN light is on. Maintain 275 knots minimum
with tower shaft failure. If a higher airspeed ismaintained or an attempt is made to gain airspeed todelay the rpm decay, available time may be reduced
to the point that an airstart is not possible. During
any low altitude airstart attempt, constantly
evaluate altitude above the ground relative toairs tart success. Do not delay ejection below 2000 feet
AGL unless the engine is producing thrust capable ofmaintaining level flight or safely controlling the sinkrate or unless a flameout landing can be accomplished.
Airstart Procedures I PW2201

~
~
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~

,
,
,
,
~
,
~
~

To perform an airstart, retard throttle to OFF, obtainairs tart conditions, and advance the throttle to IDLE.

".

If the throttle is retarded to OFF to clear a stagnation,
the rpm decreases rapidly and the FTIT decreases.
The throttle should be maintained in OFF for a few
seconds to allow the stagnation to clear. If the airstart
attempt is not due to a stagnation, FTIT may be well
below 700°C when the throttle is retarded to OFF.

".

Ifit appears that rpm will drop below 25 percent, thethrottle must be advanced to IDLE to preserve rpm
regardless of FTIT or airspeed. If FTIT was above
700°C, a hot start will probably occur, and the throttle
must again be retarded to OFF. However, rpm ispreserved permitting a subsequent airstart withinparameters.

~
~

~
"
~

~
"

~~~~~~~~~~~III~

~~~IIIIIIIIIII~

~

~

~

Regardless of altitude, all initial airstart attempts
shall be performed in the control mode selected by the
DEEC. In PRI, maintain 250 knots minimum for aspooldown airstart or maximum range or maximum
endurance airspeed with the JFS RUN light on. In
SEC, maintain 275 knots minimum for a spooldown
airstart or 200 knots minimum with the JFS RUN
light on.
Start the JFS below 20,000 feet MSL and below 400
knots immediately after the throttle is advanced to
IDLE to initiate the spool down airstart.
When the throttle is advanced to IDLE, rpm and FTIT
may continue to decrease until light-off occurs whichtakes up to 15 seconds. Increasing rpm is normallythe first indication of an airstart. The light-off is
subtle since rpm and FTIT turnaround are very slow.
If light-off is not attained within 20 seconds, retard
the throttle to OFF and reattempt an airstart with
the ENG CONT switch in SEC. From the time thethrottle is advanced from OFF, the engine takes
approximately 45 seconds to reach idle rpm. Engine
acceleration may be slow around 40-50 percent rpm
during the airstart attempt. Do not confuse the slow
acceleration with a hung start. DEEC airstart
overtemperature protection logic attempts to limit
FTIT to approximately 700°C, which results in hung,decreasing, or slowly increasing rpm. If FTIT isstabilized between 700°C and BOO°C and rpm isslowly increasing, a successful airstart results.
If FTIT is stabilized between 700°C and 800°C and
rpm is definitely hung or decreasing, retard thethrottle to OFF and reattempt the airstart. If all
airstart parameters were in limits, increase airspeed(275 knots minimum), and reattempt an airstart with
the ENG CONT switch in SEC; if not, attain airstart
parameters and reattempt in the control mode
selected by the DEEC. If airstart was still
unsuccessful with the airstart conditions met and the
ENG CONT switch in SEC, place the ENG CONT
switch back to PRI, maintain the airstart parameters
and reattempt the airstart. If the engine autotransferred to SEC, repositioning the ENG CONT
switch to PRI will have no effect. However, if SEC was
selected manually and not selected by the DEEC, theengine should transfer back to PRI.

If a hot start occurs (FTIT above BOO°C) and start
parameters were met and altitude is still sufficient,increase airspeed (if possible) and reattempt anairstart with the ENG CONT switch in SEC. If a hot
start occurs and start parameters were not met andmaintained, attain airstart parameters and reattempt airstart in the control mode selected by the
DEEC.

If the JFS stops running or fails to run within 30
seconds, do not reattempt a JFS start until thebrake/JFS accumulators have time to recharge. Allow
1 minute of engine rotation <either windmilling or
JFS assisted) at 12 percent rpm or above to insurethat the brake/JFS accumulators are fully recharged.
Recharging begins 3-4 seconds before the JFS RUN
light illuminates or 30 seconds after selecting a start
position (in the event of a JFS failure to run).
Recharging begins regardless of JFS switch position.

In the event of a JFS shutdown, the JFS switch does
not relatch in either start position while the JFS isspooling down. Spool down from full governed speed
takes approximately 17 seconds. The JFS switch
must be cycled to OFF and then to START 2 toreinitiate a JFS start. It is possible to complete thespooldown before the brake/JFS accumulators arerecharged if the JFS ran for only a short time.
When the airstart is completed, turn the JFS off (iftower shaft failure is not suspected). Cycle the EPU
switch to OFF and then back to NORM.
To accomplish an airstart:
1.

Throttle - OFF.

FTIT should decrease rapidly whenthrottle is OFF. If FTIT does notdecrease rapidly, verify that thethrottle is OFF.
2.

Airspeed - As required.
Maintain 250 knots minimum for PRIor 275
knots minimum for SEC below 40,000 feet fora spooldown airstart.

When rpm is 50-25 percent with FTIT below 700°C:
3.

Throttle - IDLE.

If it appears rpm will go below 25
percent, advance throttle to IDLE
regardless of FTIT or airspeed.
4.

JFS switch - START 2 below 20,000 feet MSL
and below 400 knots.
Maintain maximum range or maximum
endurance airspeed (200 or 170 knots,
respectively, plus 5 knots per 1000 pounds offuel/store weights over
3000, [ID 2000
pounds) with the JFS RUN light on (200knots minimum for SEC airstarts).

rru~li;;,~~~~I'I'I'I'I'~~

~""""""""""""'~~
S
10. 1 F- 1 i:lA- 1• If the JFS s\vitcb is erroneou<y pJ. H:ed
to START 1. leavp it thvre .

• If the JFS RUN light does not illummf'lte
or goes out once illuminated, placE' the
JFS switch to OFF and reattpp'pt
START 2 when the brakelJFS accumulators are recharged. The JFS switch
does not relatch in either start position
while the JFS is spooling down.

Do not fly slower than 200 knots for
SEC JFS-assisted airstarts.
NOTE• If maxim urn gliding range is not a
factor, consider maintaining 250 knots
minimum for PRIor 275 knots
minimum for SEC above 10,000 feet
AGL to reduce rpm spooldown rate (incase of JFS failure). Below 10,000 feet
AGL with the JFS RUN light on (where
only one airstart attempt is likely),
maintain maximum range or maximum endurance airspeed (200 knots
minimum for SEC airstarts).

5.

Stores - Jettison (if required).
If stores jettison is attempted after main
generator drops off line but before EPU
generator powers the SMS (approximately 5
seconds delay), stores will not jettison.

Visually confir:O::e stores havejettisoned and jettison again if re-

~s

~

~s~

....~

S

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~

~

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~

quired.

Change 13~""lIII

3- 104.1/(3- 104.2 blank)

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~"""'"''''''''''''''T. O. 1F- 16A-1

6.

Throttle - OFF.

7.

Reattempt airstart in mode selected by the
DEEC.

~~

IWARNING I

If hungfhotlno start and airstart conditions were notmet:~@ If only AUX flag is in view, pitch andIf hunglhotlno start and airstart conditions were met:
8.

Throttle - OFF.

9.

ENG CONT switch - SEC.

~~

roll attitude information is likely to be
erroneous due to INS autorestart in
the attitude mode when other thanstraight and level, unaccelerated
flight conditions existed.
17. Throttle - As required.

NOTE
When operating in SEC below 15,000
feet MSL, maintain 70 percent rpm
minimum until landing is assured.

• Place the ENG CONT switch to SEC
prior to placing the throttle to IDLE,
otherwise a start anomaly may result.

NOTE• The proximity of the ENG CONT
switch to the JFS switch makes the
JFS switch susceptible to being
bumped to OFF when selecting SEC.
10. Throttle - IDLE.
If still hunglhotlno start and airstart conditions were
met:

11.

Throttle - OFF.

12.

ENG CONT switch - PRI.

If the SEC caution light is on, refer to
SEC CAUTION LIGHT I PW2201, thissection.

18.

Land as soon as possible.

19.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINEFLAMEOUT LANDING
The decision to eject or make a flameout landing rests
with the pilot. Considerations for attempting aflameout landing must include:

13. Throttle - IDLE.
If engine does not respond normally after airstart iscompleted:

14. Refer to FLAMEOUT LANDING, this section.
If engine responds normally:• Nature of the emergency.
• Weather conditions.
• Day or night.
• Proximity of a suitable landing runway.

Do not turn JFS or EPU off if indicated
rpm is below 60 percent with adequate
thrust (e.g., tower shaft failure).
14. JFS switch - OFF.
15. EPU switch - OFF, then NORM.
16. ADI - Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to TOTALINS FAILURE, this section.

• Proficiency in performing simulated flameout(SFO) landings.
Due to the capabilities of the ejection seat, the entire
approach is within the ejection envelope; however,ejection should not be delayed in an attempt to salvageaa questionable approach. When performing
flameout landing, the aircraft can safely stop (dryrunway without arresting gear ~ [NQ] or drag chute)in approximately twice the computed ground rolldistance (8000-foot minimum runway length recommended), assuming a touchdown no more than 1/3 ofthe way down the runway at 11-13 degrees AOA.

Change 13

3-105

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"~

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"~
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T. 0 1F- 16A- 1

To perform a flameout landing, turn immediately
toward the desired runway. Jettison stores andestablish maximum range airspeed. Maximum range
airspeed may be less than the minimum airstart
airspeed. If range to the desired runway is critical, thedecision to attempt an airstart or a flameout landing
rests with the pilot.
NOTE• During an airstart attempt, do notslow below the minimum airstart
airspeed.
• If the engine is still running, but thrustis insufficient to sustain level flight,
treat it as a flameout situation.
Maximum range airspeed varies only with GW and is
not affected by drag index. Maximum range airspeed
is 200 knots for a GW of 20,000 pounds and increases
5 knots per 1000 pounds of actual GW above 20,000
pounds. For most circumstances, sufficient accuracy
is obtained by adding 5 knots per 1000 pounds offuel/store weights over!Al 3000, [ID 2000 pounds.
NOTE• This formula is based on the average
aircraft operating weight. Refer to T.O.
1F-16A-1-1, PART 1, DRAG INDEXESAND WEIGHTS - BASIC AIRCRAFT.
If range to desired runway is critical,
maximum range airspeed may be
calculated using actual GW in excess of
20,000 pounds.
• For a 10,000-foot descent (LG up), each
10 knots above or below maximum
range airspeed decreases glide range up
to 114 nm.
The maximum range airspeed equates to approximately 7 degrees AOA (any GW or drag index) andprovides a glide ratio of approximately 7 nm per 5000
feet AGL (a no wind condition). Retaining stores or
flying into a headwind decreases glide range
significantly.
The EPU should be on and, if aircraft fuel is available,
the JFS should be started using START 2 when below
20,000 feet MSL and below 400 knots unless theengine is either seized or anticipated to seize. The
EPU should provide a minimum operation of 10
minutes (HYDRAZN light on) with normal flightcontrol demands before EPU fuel depletion. Operating time can be extended to as much as 15 minutes if
the JFS is running and flight control inputs are

3-106

minimized. If f'xpectf'd time to landing exceeds
expected EPU operating time and excess energy isavailable, a steeper/faster descent may be flown. The
JFS also provides hydraulic pressure for normal
braking and NWS after landing.

When bleed air is no longer available to operate the
EPU, confirm that the EPU is operating on hydrazine(EPU run and HYDRAZN lights on) since the JFS
alone does not provide adequate hydraulic pressure
to land the aircraft. If the EPU is inoperative,
maneuver the aircraft as necessary on JFS-assisted
hydraulic pressure to a more favorable ejection
envelope and initiate ejection .
There are two basic types of flameout landing
patterns: the overhead approach (figure 3-13) or thestraight-in approach (figure 3-14). The overhead
approach is preferred as it affords the mostopportunities to properly manage available energy
while providing the best visual cues for patterncorrections. The overhead approach may be enteredat any position, provided the proper altitude for thatpoint in the pattern can be obtained. The mainconcern is to reach high key, low key, or base key at or
above the recommended minimum key altitudes. A
straight-in approach is an alternate approach when
the overhead approach cannot be attained. For both
approaches, the initial aimpoint should be approximately 1/3 of the way down the runway.
Overhead Approach

Refer to figure 3-13. Plan to arrive over the landing
runway (high key) at 7000-10,000 feet AGL. The highkey position may be approached from any direction .
The recommended key altitudes are based on flying a
360-degree descending turn from high key with the
LG down. The altitudes vary with GW and withadditional drag due to stores. The recommended highkey altitude is 7000 feet AGL plus 500 feet per 1000
pounds of fuel/store weights over !Al 3000, [ID 2000
pounds. The recommended low key altitude is 3000
feet AGL plus 250 feet per 1000 pounds of fuel/store
weights over [BJ 3000, [ID 2000 pounds. These formulas
include compensation for stores drag effects; thus, noadditional correction is required.
If altitude will be significantly higher at high key,some form of altitude dissipating maneuver such as
a dive, gentle S-turns, or a 360-degree descending
turn should be used. Speedbrakes also may be used tolose excess altitude. However, if the speedbrakes are
not closed when a satisfactory flightpath is reached,
the added drag may preclude a successful flameout
approach.

to. 1F- 16A- 1
Mter departing high key, all attention should bedirected toward a successful landing. If actual
altitude at high key was below the recommended
altitude, fly maximum range airspeed with the LG up
until a satisfactory flightpath is reached and thenlower the LG. Optimum LG down airspeed is 10 knots
less than maximum range (LG up) airspeed.
Minimum LG down airspeed is 20 knots less thanmaximum range (LG up) airspeed and provides
sufficien t maneuverability to arrest the high sinkrate associated with a flameout approach. Optimum
angle of bank is 50 degrees with the LG up and 55
degrees with the LG down. Bank angles more than 10
degrees abovelbelow optimum result in a significant
increase in altitude loss per degree of turn and may
preclude a successful flameout approach.

Straight-In Approach

Refer to figure 3-14. If one of the overhead approachkey positions cannot be reached, a straight-in
approach may be flown. The clean glide at maximum
range airspeed should be continued until the initial
aimpoint is 11-17 degrees below the horizon; then the
LG should be lowered. Seventeen degrees is below theforward field of view. A good visual reference for 15
degrees is when the initial aimpoint is at the bottomof the HUD (just above the radome). Optimum LG
down airspeed is 10 knots less than maximum range(LG up) airspeed. Minimum LG down airspeed is 20
knots less than maximum range (LG up) airspeed andprovides sufficient maneuverability to arrest the highsink rate associated with a flameout approach.

NOTE

NOTE

• Delaying LG extension until low key
allows successful completion of theoverhead approach from as low as 1500
feet below the recommended high key
altitude.

For a 10,000-foot descent (LG down),
each 10 knots above optimum LG down
airspeed decreases glide range up to 112
nm.

• Altitude loss for a 360-degree descending turn with the LG down increases
up to 500 feet for every 10 knots above
optimum LG down airspeed.
• Altitude loss for a 360-degree descending turn with the LG down increases
up to 500 feet for each 5 degrees
abovelbelow the optimum bank angle.
The ground track of a flameout/SFO overhead
approach is approximately the same as that of anormal overhead approach except the final approach
is approximately 3/4 nm long.
Avoid rapid flight control inputs which use excessive
EPU fuel and may exceed the emergency hydraulicpump capability.

I WARNING I
If EPU fuel quantity is below 25
percent at high key (20 percent with
the JFS running), a flameout landing
should not be attempted since adequate hydraulic pressure may not be
available through the landing.

IMe Penetration

Should IMC be encountered during a flameout
approach to the intended runway and no alternate
runway is available, an alternate descent/penetration may be flown which should allow maneuvering
airspeed after penetrating the undercast.

I WARNING I
IMC penetration should not be
attempted unless present position is
known and navigation can be performed throughout the descent, andhigh terrain or other hazards are not a
factor.
The stores should be jettisoned and the aircraft glided
at maximum range airspeed until a 1:1 ratio between
altitude in thousands offeet and range to the runway(e.g., 20,000 feet AGL at 20 nm, 15,000 feet AGL at 15
nm, etc.) is attained. The descent angle should thenbe increased and airspeed allowed to increase tomaintain the 1:1 ratio. This equates to a 9-10 degree
descent angle. This 1: 1 glide ratio must be
maintained until sufficient airspeed is attained tomaneuver after penetrating the undercast.

Change 5

3-107

IT.O. 1 F-16A-1

Flameout Landing Pattern (Typical)(OVERHEAD APPROACH)NOTES:
1. Jettison stores (if required).

4. Altitudes:• High Key - 7000 - 10,000 feet AGL
Recommended altitude is 7000 feet AGL plus
500 feet per 1000 pounds of fuel/store
weights over 1A130oo,
2000 pounds.

2. Maximum range (LG up) airspeed is 200 knots. Optimum airspeed (LG down) is 190 knots. Minimum LG
down airspeed is 180 knots. Increase airspeeds by
5 knots per 1000 pounds of fuel/store weights over~3000,
2000 pounds.

rnJ

00

• Low Key - 3000 - 5000 feet AGL
Recommended altitude is 3000 feet AGL plus
250 feet per 1000 pounds of fuel/store
weights over 1A13000,
2000 pounds.

3. Maximum range (LG up) airspeed equates to approximately 7 degrees AOA (any GW or drag index) andprovides a glide ratio of approximately 7 nm per
5000 feet AGL. If stores are retained, glide ratio
decreases.

rnJ• Base Key - 2000 feet AGL minimum
5. Optimum bank angles are 50 degrees (LG up) and 55
degrees (LG down) for least altitude lost per degreeof turn.

A.

HIGH KEY -Above a point approximately 1/3 of the way down the runway

B.

LOW KEY -Abeam point of rollout on final

C.

BASE KEY -Midpoint of turn from downwind to final

1

3/4 nm~==Point of intended
touchdown[ XL - ~"'--=::;LJ
D.\r---,--========
X
Initial aimpoint shouldbe 1/3 runway length
1 nm

Low key posItIOnvaries with winddirection/velocity~==========:::::::

1======::3:::<Figure 3-13. (Sheet 1)

3-108

IF-16A-1-1180X@

Flameout Landing Pattern (Typical)(OVERHEAD APPROACH)NOTES:WARNING

6. With FCC off, HUD continues to compute flightpath
marker and to position scales for use during flameout approach.
7. Frost or condensation on the canopy could restrict
visibility during flameout approach. Place AIR
SOURCE knob to RAM and place DEFOG leverforward below 25,000 feet MSL.
8. Time constraints due to EPU fuel consumption must
be considered as well as distance to be covered.
To estimate required EPU fuel for a nonstandard
approach, use 15 percent per minute as a basis
for computation.

•

The JFS alone does not provide adequate
hydraulic pressure to land the aircraft.

•

Do not allow airspeed to decrease below
minimum LG down airspeed.

• Eject if it becomes obvious that a safe
landing cannot be made. Ejection can beaccomplished at any point in the pattern; however, do not delay ejection
below 2000 feet AGL in an attempt
to salvage a questionable approach.

9. Starting JFS reduces load on EPU, conserves EPU
fuel, and partially restores hydraulic system B.
10. If alternate LG extension is used, the NLG may notindicate down and locked until airspeed is reduced
below 190 knots.

HIGH KEY
Do not extend LG unlessbase key is assured.

I WARNING

I

EPU fuel quantity should be at least 25
percent (20 percent with JFS runningl
at high key to insure adequate hydraulicpressure throughout landing.
ROLLOUT• ~~ Drag ChuteDEPLOY (if requiredL

FLARE------..
Touch down 11-13 degrees
AOA optimum. Speedbrakes
as required.

LG down. Increase airspeed
and/or open the speed brakes
to move touchdown closer toapproach end of runway.

Do not extend LG unlessbase key is assured.

3-109~~~~~~~~~~~~~~~

~~~IIIIIIIIIII~
Flal1leout Landing Pattern (Typical)
NOTES:

1. Jettison stores (if required).
2. Maximum range (LG up) airspeed is 200 knots. Optimum airspeed (LG down) is 190 knots. Minimum LG
down airspeed is 180 knots. Increase airspeeds by
5 knots per 1000 pounds of fuel/store weights over
~ 3000, [ID 2000 pounds.

6. With FCC off, HUD continues to compute flightpath
marker and to position scales for use during flameout approach.
7. Frost or condensation on the canopy could restrict
visibility during flameout approach. Place AIR
SOURCE knob to RAM and place DEFOG leverforward below 25,000 feet MSL.

8. lime constraints due to EPU fuel consumption must
be considered as well as distance to be covered.
To estimate required EPU fuel for a nonstandard
approach, use 15 percent per minute as a basis
for computation.

9. Starting JFS reduces load on EPU, conserves EPU
fuel, and partially restores hydraulic system B.
5. After lowering LG, glide range decreases by approximately 30 percent. Airspeed greater than optimum
LG down airspeed significantly increases energy
loss rate and decreases glide range.

10. If alternate LG extension is used, the NLG may notindicate down and locked until airspeed is reduced
below 190 knots.

ROLLOUT.~~Drag ChuteDEPLOY (if required) .
• Speedbrakes - Open .
• Hook - DN (if required).

FLARE
Touch down 11-13 degrees
AOA optimum. Speedbrakes
as required.
AREA C
4 - 0 nm(no wind)
4 nm (no wind)4000 - 8000
feet AGL~

POINT B

I........
I... POINT A~~, ' , ~ ~ I I ~L L L L . I " " ~1F-16A-1-11S2X®

Figure 3-14. (Sheet 1)

T.O.1F-16A-1

Flameout Landing Pattern (Typical)(STRAIGHT-IN APPROACH)I WARNINGI• EPU fuel quantity (points A, 81, and 82)
should be sufficient to insure adequate
hydraulic pressure through landing.

•

Do not allow airspeed to decrease below
minimum LG down airspeed.

•

If the aimpoint on the runway moves up in thefield of view while maintaining maximum range(LG up) airspeed, the runway probably cannot
be reached. This path corresponds to a glide
angle of about 7 degrees between the horizon
and the aimpoint.

••

Eject if it becomes obvious that a safe landing
cannot be made. Ejection can be accomplishedat any point in the approach; however, do notdelay ejection below 2000 feet AGL in an
attempt to salvage a questionable approach.

The JFS alone does not provide adequate
hydraulic pressure to land the aircraft.

*17 0

*

OVERHEAD
APPROACH*110

08 POINT A

AREA C_.:=::;:::,:::-- POINT 81 --------------- ---- -------~"!--~--·;.-::::::::::::::::__.:;~~=-----h-

1-----1
--I

+

2000
FEET
AGL

* Angle between horizon and aimpoint

POINT A

POINT 8AREA C

8 nm (no wind), 7000 feet AGL,
continue glide until initial aimpoint is 11-17 degrees below
horizon. Then lower LG andestablish optimum LG down
airspeed. As a guide, no windminimum EPU fuel is 45 percent(40 percent with JFS running).

4 nm (no wind), 4000-8000 feet
AGL, airspeed and LG as required.
As a guide, no wind minimum EPU
fuel is:4 - 0 nm (no wind), initial aimpoint is more than 17 degrees
below horizon (under nose of
aircraft and not visible). Normal
straight-in approach is not feasible.

POINT 81
6 nm - 35 percent (30percent with JFS running).

Options are:

POINT 82

• Delay LG lowering. Plan an
overhead approach from a highkey altitude but below thenormal recommended altitude.

4 nm - 25 percent (20percent with JFS running).

• Delay LG lowering. Plan a modified flightpath to low key.

WARNINGI• Lower LG, open speedbrakes,
and dive and maneuver aircraft
to intercept a point on thenormal straight-in glidepath.

Do not delay lowering LG
below 2000 feet AGL.
1F-16A-1-1183X®

Figure 3-14. (Sheet 2)

3- 111

T.O. 1F- 16A-1
NOTE

A 90 degree level turn at 50 degrees bankangle with the LG and speedbrakes
retracted will dissipate 65-85 knots. A
180 degree turn will dissipate 145-250
knots. Airspeed dissipation increases
with increasing GW and DI. A glide
angle at a 1: 1 ratio begun frommaximum range airspeed will result inan airspeed of 260-320 knots after a
10,000-foot descent. Higher airspeed at
the start of the glide, additional descent
altitude, heaver gross weight, or lower
drag index will result in higher airspeed
at the completion of the glide.

I

At 3000 feet AGL, the aircraft should be 3 nm from thetouchdown point. If the runway is not in sight by base
key altitude, the aircraft may be zoomed for acontrolled ejection. When VM:C is attained and the
runway is in sight, the aircraft should be glided to an
attainable key position for an overhead approach or
to a straight-in approach and the LG should belowered. Excess airspeed above optimum LG down
airspeed not required to maneuver to the flameout
landing approach should be dissipated by use of
speedbrakes or early LG extension.

I

Landing Phase

The LG should be lowered no later than 2000 feet
AGL to allow adequate time for alternate LG
extension. Establish a glidepath to achieve the initial
aim point while maintaining optimum LG down
airspeed. Once wings level on fmal approach, beaware of the tendency to slow below minimum LG
down airspeed.

I WARNING I
Do not attempt to stretch a glide byallowing the airspeed to decrease below
minimum LG down airspeed. A slower
airspeed decreases the maneuverability
available to arrest the high sink rate
associated with the flameout approachand may preclude a successful flameout
landing.
Once landing is assured, the recommended procedure
is to shift the aimpoint from 1/3 of the way down the
runway to a position short of the intended touchdown
point. Speedbrakes may be used to help control
airspeed. The higher the airspeed, the shorter the
aimpoint should be to allow for additional float (from
3-112

Change 5

flare to touchdown). The aircraft is easiest to control
in the flare if the flare is begun between optimum andminimum LG down airspeeds. The point at which theflare is begun depends upon airspeed, sink rate, andglide angle. The flare should be started high enoughto allow a smooth gradual reduction in glide angle but
not so high as to run out of airspeed prior totouchdown. Under a no wind condition, the aircraft
floats 3000-4000 feet after beginning the flare, if theflare is begun at the optimum LG down airspeed.
Once the sink rate is arrested, attempt to slow to anormal touchdown airspeed and AOA. If excess
airspeed exists after arresting the sink rate, the best
method to slow the aircraft is to stay airborne un til
normal touchdown airspeed is reached.
After Touchdown

After touchdown from a flameout landing, use anormal or short field stopping technique as required by
the stopping distance available. If the JFS and EPU
are running, normal braking and NWS are available(NWS is inoperative if the LG was lowered with thealternate LG system). If the JFS is not running, only
the brakelJFS accumulators are available to supply
hydraulic pressure for braking. Stop the aircraft bymaking one steady brake application just short ofantiskid cycling. If there is any doubt about stoppingon the remaining runway, lower the hook. When theaircraft is fully stopped, have chocks installed or set
parking brake. Leave the battery on line until chocks
are installed. If JFS START 2 was attempted but was
unsuccessful, no braking is available for stopping or
directional control unless the brake/JFS accumulators
are recharged. Use flaperons and rudder as required
to maintain directional control. As the aircraft slows
below 70 knots, directional control is reduced and theaircraft may drift right.
Flameout Landing Procedures

If the engine has flamed out or if flameout isimminent, turn toward a suitable runway andaccomplish either an overhead approach or astraight-in approach, as appropriate.
• Altitudes (overhead approach):• High key - 7000-10,000 feet AGL.
Recommended altitude is 7000 feet AGL plus
500 feet per 1000 pounds of fuel/store weights
over ~ 3000, lID 2000 pounds.
• Low key - 3000-5000 feet AGL.
Recommended altitude is 3000 feet AGL plus
250 feet per 1000 pounds of fuel/store weights
over ~ 3000, lID 2000 pounds.
• Base key - 2000 feet AGL minimum.

""""",,~~~• If the JFS RUN light does notilluminate or goes off once illuminated,
place the JFS switch to OFF andreattempt START 2 when the brake/
JFS accumulators are recharged. The
JFS switch does not relatch in either
start position while the JFS is spooling
down.

~
~
~
~
~
~
~

AIR SOURCE knob - RAM (below 25,000 feet
MSL).

,

6.

DEFOG lever - Forward.

~

7.

LG handle - DN. (Use DN LOCK REL button
ifrequired.)• Do not start the JFS if engine seizure
has occurred or is anticipated or ifengine failure is a result of fuelstarvation. Starting the JFS may
result in no brake/JFS accumulator
pressure for the brakes.

• Altitudes (straight-in approach):• 8 nm - 7000 feet AGL minimum.
The minimum altitude is based on an LG up
glide at maximum range airspeed to 2000 feet
AGL followed by an LG down glide at
optimum LG down airspeed to the runway fora drag index of 100. A lower drag index
slightly reduces the minimum altitude
required. A higher drag index slightly
increases the minimum altitude required.

NOTE• If engine is not operating, consider
placing the FUEL MASTER switch to
OFF if a fuel leak exists. This action
may conserve fuel for the JFS.

• 4 nm - 4000-8000 feet AGL.
Delay lowering the LG until the initial
aim point is 11-17 degrees below the horizon.

I

I

WARNING• If the JFS is erroneously placed to
START 1, leave it there.

I

Eject if a safe landing cannot be made.
Ejection can be accomplished at any
point in the pattern but do not delayejection below 2000 feet AGL in an
attempt to salvage a questionableapproach.
1.

Stores - Jettison (if required).

2.

Airspeed - 200 knots.
Increase airspeed 5 knots per 1000 pounds offuel/store weights over [8] 3000, [ID 2000
pounds. This airspeed equates to approximately 7 degrees AOA.

5.

IWARNINGINOTEDuring an airstart attempt, do notslow below the minimum airstart
airspeed.
3.

EPU switch - ON.

4.

JFS switch - START 2 below 20,000 feet MSL
and below 400 knots.

I WARNING I• EPU fuel quantity should be at least 25
percent (20 percent with JFS running)at high key for an overhead approachor 45 percent (40 percent with JFS
running) at 8 nm for a straight-in
approach to insure adequate hydraulicpressure through landing .
• The JFS alone does not provide
adequate hydraulic pressure to land
the aircraft.

• Do not delay lowering LG below 2000
feet AGL.
• If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed, resulting in higher
than normal landing thrust.
8.

ALT GEAR handle - Pull (if required) (190knots maximum, if practical).
Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

NWS is not available following alternate LG extension.

~

I'

~
~
~

_

_

-

10. 1F- 16A- 1
JETTISON
Selective Jettison

Do not depress the ALT GEAR reset
button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.
9.

Airspeed - 190 knots optimum in pattern.
Increase airspeed by 5 knots per 1000 poundsof fuel/store weights over ~ 3000, [ID 2000
pounds.

I WARNINGIDo not allow airspeed to decrease
below 180 knots plus 5 knots per 1000
pounds of fuel/store weights over ~3000, [ID 2000 pounds.

Selective jettison is used to release selected store(s)(except air-to-air missiles) or suspension equipment
in an unarmed or unguided condition.
NOTE~ Selective jettison cannot be accom-

plished.
1.

GND JETT ENABLE switch - ENABLE (ifLG
is down).

2.

MASTER ARM switch - MASTER ARM.

3.

~ ARMT CONSENT switch - On.

4.

SMS PWR switch - SMS PWR.

5.

DOG FIGHT switch - Center.

6.

SEL JETT button - Depress.

7.

SCP - WPN/RACK select.

After touchdown:
10. ~ INQI DRAG CHUTE switch - DEPLOY (ifrequired).
11.

HOOK switch - DN (if required).

If brake/JFS accumulator braking is used:
12.

Stop straight ahead and engage parking
brake.

• Brakes should be applied in a single,
moderate, and steady application
without cycling the antiskid.
• Brake pedal deflection of 1116 inch
activates the brakes and bleeds thebrake/JFS accumulators. To avoid
brake activation and loss of accumulator fluid, do not rest feet on the brake
pedals.
• Do not attempt to taxi clear of the
runway. Loss of brake/JFS accumulator pressure results in the inability tostop or steer the aircraft.
13.

Refer to ACTIVATED
LEAK, this section.

EPU/HYDRAZINE• Jettison of an inboard shouldermounted store from a TER-9/A atstation 4 or 6 with MLG down may
result in LG and store(s) collision. To
avoid this, select RACK for jettison
instead of WPN.
• Jettison of external wing fuel tankswith stores/suspension equipment at
stations 3 and/or 7 with MLG down
may result in LG and external wing
fuel tank collision.
• Failure to load the actual storesconfiguration into SMS inventory
could cause damage to the aircraft byinhibiting the selective jettison release time delay used to insure safe
370-gallon fuel tank separation whena store is present at station 3 or 7.
NOTE• To jettison store(s) from auxiliary
racks (TER-9/A, LAU-88/A), selectWPN. To jettison the auxiliary racks orstore(s) mounted directly on a
MAU-12C/A, D/A, select RACK.
• To jettison a bomb mounted directly on
a MAU-12C/A, D/A, either WPN orRACK may be selected.

~,lli;;g,~~~~~~I'I'I'~~

~~~~~r~".~~~T. O. 1 F- 16A- 1
NOTE\\'Iwn 30 1}-g'allon :mcl 370-gallon fud
tanks an' carried ~imultan('ously, the~)OO-gallon fuel tank must bp separated
prior to tlw 370-gallon fuel tanks.
8.

Station(s) - Select.

9.

WPN REL or [Al ALT REL button - Depress.
NOTE

When jettisoning tanks from stations 4
and 6, hold release button depressedfor 1 second.
Emergency Jettison

Emergency jettison is a one-step operation whichclears all expendable stores and racks except air-toair missiles. All weapons are released in an unarmedor unguided condition. If the SMS is off, depressing
the EMER STORES JETTISON button supplies
electrical power to the SMS. In addition, while thebutton is depressed, the avionic system enters theNAV mode. When the button is released, the avionicsystem returns to the previous operating mode.
Emergency jettison is not available unless the mainor EPU generator is operating.
1.

2.

GND JETT ENABLE switch - ENABLE (ifrequired),
Use EMER STORES JETTISON on the
ground only as a last resort.
EMER STORES JETTISON button - Depress(1 second).
NOTE~

FLCS FAILURES

ADC/Air Data Malfunctions

A single failure in the air data system (static or
impact pressure or AOA) is indicated by the ADC
caution light. The CADC caution light may also
illuminate for a single AOA malfunction.
Contamination of the pitot probe can cause
blockage of the passage which provides totalpressure signals. This causes erroneously low
airspeed indications in the HUD and on theairspeed/mach indicator with an ADC caution lightat approximately 300 knots. An erroneously low
airspeed value is used for FLCS gain schedulingwhich may result in pitch oscillations.
If erroneously low airspeed indications are present
both in the HUD and on the airspeed/machindicator during takeoff, consider aborting. If
takeoff is continued, maintain airspeed below 275
knots (use INS groundspeed or wingman indications to determine airspeed) and land as soon aspractical. Pitch oscillations may be experienced at
higher speeds. If oscillations are encountered,minimize stick inputs and slow to less than 275
knots.

A dual failure of static or impact pressure systems isindicated by the FLT CONT SYS, ADC, and LE
FLAPS caution lights. STBY GAINS light isilluminated and cannot be reset in flight. The DUAL
FC FAIL warning light does not illuminate. LEF's are
zero degrees with the LG handle in UP and the ALT
FLAPS switch in NORM. LEF's are 15 degrees downwith the LG handle in DN or ALT FLAPS switch in
EXTEND.

If the initial actuation of the EMERSTORES JETTISON button fails tojettison all aircraft stores, subsequentattempts may successfully release theremaining stores.
3.

A dual failure of AOA sources is indicated by the ADC,FLT CONT SYS, LE FLAPS, and CADC caution
lights. The DUAL FC FAIL warning light also
illuminates and can be reset if the malfunction clears.

SCP display - Confirm release.
NOTE

If a CRIU is not installed at station 5,
a plus (+) symbol does not appear on
the SCP to indicate store present when
the EMER STORES JETTISON button is depressed.

NOTE

The DUAL FC FAIL, CADC, ADC, FLT
CONT SYS, and LE FLAPS lights may
illuminate for high AOA and/or sideslip maneuvers. After recovery, resetand continue normal operations.

,~~~~~~~~~~~III~1.

Do not attempt SERVO or ELEC reset.

~~~

I WARNINGIDo not attem pt SERVO or ELEC reset
if DUAL FC FAIL warning and P, R, or
Y light(s) are on. This action may
cause affected control surface(s) to be
inoperative.
2.

Retarding the throttle below 1\l1L
while supersonic may induce inlet
buzz which produces severe cockpit
vibration and probable engine stalls or
I PW200 I stagnation.
If CADC caution light goes off:

Land as soon as possible.
3.

Continue normal operation.

If DUAL FC FAIL warning light is off or if DUAL FC
FAIL warning light is on with no P, R, or Y lights:
3.

LESS

rn PROBE HEAT switch - Verify in

PROBE HEAT.
4.

If CADC caution light remains on:

3.

AOA - Cross-check with airspeed.
Use AOA indications with caution.

4.

Land as soon as practical.
Final approach airspeed [8] 125, [ID 129 plus 4
knots/1000 pounds of fuel/store weights
equals 13 degrees AOA (add 8 knots for 11
degrees AOA).

SERVO ELEC RESET switch - ELEC.

If all lights reset:

5.

Continue normal operation.

If all lights do not reset or LEF's are not functioning
normally:
6.

AOA - 12 degrees maximum.

7.

Land as soon as practical.
If operating on standby gains, do not exceed
650 knots; do not slow below 240 knots with
the LG up.

If DUAL FC FAIL warning light remains on:
8.

Land as soon as possible.

CADC Malfunction I PW200 I orl PW2201 LESS~, LESSCADC Malfunction I PW2201 [8N] tID•

A failure of any of the CADC electrical inputs (AOA,
total temperature, altimeter barometric reference,etc.) or a detected failure internal to the CADC causes
illumination of the CADC caution light. If the CADC
caution light does not reset, systems dependent on
CADC information should be checked for proper
operation.

til
A failure of any of the CADC electrical inputs (AOA,
total temperature, altimeter barometric reference,etc.) or a detected failure internal to the CADC causes
illumination of the CADC caution light. If the CADC
caution light does not reset, systems dependent on
CADC information should be checked for proper
operation.

If CADC caution light illuminates and ENGINEFAULT caution light is also on:

If CADC caution light illuminates:
1.

Retarding the throttle below MIL
while supersonic may induce inlet
buzz which produces severe cockpit
vibration and probable engine stalls.

SERVO ELEC RESET switch - ELEC.

If I PW200 I EEC, I PW2201 ENGINE FAULT caution
light is on:2.

Go to IPW200 I EEC, IPW2201 ENGINE FAULT
CAUTION LIGHT, this section.

1.

Depress FALT ACK button and check for an
ENG 086 PFL.

~,l,;;g,~~~~~~~~~~~IIIIIIIIIII'~~~•

After fault acknowledgement or if ENGINE FAULT
caution light was off:
2.

SERVO ELEC RESET switch - ELEC.

If CADC caution light goes off:

3.

[BJ mIl AB RESET switch - AB RESET, then
NORM.

4.

Depress FALT ACK button and check for an
ENG 086 PFL.

If ENG 086 PFL is still present:
5.

Continue flight and observe throttle limitation if supersonic. Refer to PILOT FAULT
LIST - ENGINE, this section.

If CADC caution light remains on:
3.

AOA - Cross-check with airspeed.
Use AOA indications with caution.

4.

Land as soon as practical.
Final approach airspeed 125 ([ID 129) plus 4
knots/l000 pounds of fuel/store weights
equals 13 degrees AOA (add 8 knots for 11
degrees AOA).

IWARNING I

BRAKES channel switch - Change channels(if required).

,If the DUAL FC FAIL warning light also illuminates,
power failure in more than one branch may haveoccurred. Either internal FLCC branch failures or
FLCS power supply branch failures cause the P, R,
and Y lights to come on simultaneously. FLCS power
supply branch failures may be identified by moving
the FLCS PWR TEST switch to TEST and observing
which FLCS PWR lights fail to illuminate (the effectof the failure on brakes may be noted at the sametime).
If P, R, and/or Y light(s) illuminates:

Airspeed - 400 knots maximum (subsonic).

If DUAL FC FAIL warning light is on:

2.

Do not attempt SERVO or ELEC reset.

IWARNINGI

If P, R, and/or Y light(s) is reset morethan once, subsequent failure inanother branch may cause loss of thataxis and loss of control.

Do not attempt to reset P, R, and/or Y
light(s) if DUAL FC FAIL warning
light is on. This action may cause
affected control surface(s) to be inoperative.

NOTE

The P, R, and/or Y light(s) may come on
in proximity to certain HF antennas(ground or airborne) giving a false
indication of FLCS malfunctions. The
light(s) should be resettable after
moving away from the HF source.
Record the altitude, location, and timeof the occurrence to aid postflightanalysis.

FLCS PWR TEST switch - TEST.
Determine brake and brake channel affected.

To reset P, R, and/or Y light(s), momentarily position
the SERVO ELEC RESET switch on the FLCP to
ELEC. If the light(s) resets, a minor flight control
transient may be felt.

1.
P, R, and/or Y Malfunctions-

Electrical failures in the FLCS are indicated byillumination of the P, R, and/or Y lights on the FLCP.
Branch failures (all three lights) can result in
momentary, mild movement in the pitch and roll axes
and possible illumination of the ADC and LE FLAPS
caution lights. If a branch failure cannot be reset, one
brake in channell or 2 may be inoperative. If an axis
failure (single light) occurs during maneuvering
flight and can be reset, a passive failure has occurred.
If the same light illuminates again, it should not be
reset.

3.

Land as soon as possible.

If P, R, and Y lights are all on:

4.
5.

IIIIIIIIIIJ"

IIIIIIIIIIJ-

If DUAL FC FAIL warning light is off:I

WARNINGI

If flight parameters that existed at thetime the P, R, and/or Y light(s)illuminated cannot be duplicated, do
not reset light(s). If the light(s) is reset,
a subsequent failure in another branch
may cause loss of that axis and loss ofcontrol.

2.

SERVO ELEC RESET switch - ELEC.

If P, R, and/or Y light(s) resets:

3.

Maneuver aircraft in affected axis.

If P, R, and/or Y light(s) remains off:

4.

Continue normal operation.

If P, R, and/or Y light(s) does not reset or comes on
again:

5.

ISA lockout occurs following a second failure of a
servo which previously was armed. This condition isindicated by illumination of the DUAL FC FAIL
warning light. Rudder ISA lockout causes the rudder
to center. Flaperon ISA lockout causes the affected
flaperon to move to 1.5 degrees trailing edge up andcauses the other flaperon to act only as an aileron (toprevent asymmetric TEF extension when the LG
handle is placed to DN). Horizontal tail ISA lockout
causes the affected horizontal tail to move to 2
degrees trailing edge up.
If SERVOS light(s) illuminates:
1.

Airspeed - 400 knots maximum (subsonic).

If a hydraulic failure is confirmed:
2.

Do not reset, arm, or disarm any servos.

3.

Go to SINGLE/DUAL HYDRAULIC FAILURE, this section.

If hydraulic pressures are normal and DUAL FC
FAIL warning light is off:
4.

SERVO ELEC RESET switch - SERVO while
depressing FCS CAUTION RESET button.

Land as soon as practical.
NOTE

If P, R, and Y lights are all on:
6.

FLCS PWR TEST switch - TEST.
Determine brake and brake channel affected.

7.

BRAKES channel switch - Change channels(if required).

Servo Malfunction

A servo failure in a rudder, flaperon, or horizontal tail
ISA is indicated by illumination of the appropriate
SERVOS light on the FLCP. If a single SERVOS lightilluminates during maneuvering flight and can bereset, a passive failure has occurred. If the same
SERVOS light illuminates again, it should not be
reset. If all five SERVOS lights are illuminated, one
hydraulic system is failed or a momentary drop in
hydraulic pressure has occurred (e.g., wake turbulence encounter, air in hydraulic system).
To reset any FLCS SERVOS light, momentarily
position the SERVO ELEC RESET switch to SERVO
while simultaneously depressing the FCS CAUTIONRESET button. If the SERVOS light(s) resets, aminor flight control transient may be felt.

Successful reset of a SERVOS light(s)may also cause illumination of the FLT
CONT SYS caution light with no lights
on the FLCP. If this occurs, reset the
FLT CONT SYS caution light bydepressing the FCS CAUTION RESET
button.
If SERVOS light(s) resets:

5.

Continue normal operation.

If SERVOS light(s) does not reset or comes on again:
NOTE

Multiple SERVOS lights may illuminate again if air is in a hydraulicsystem. Additional servo resets may be
performed only if multiple SERVOS
lights are on and hydraulic pressure
indications remain normal.
6.

Arm the appropriate servo(s).

7.

Land as soon as practical.

If DUAL FC FAIL warning light comes on:•

8.

Do not attempt SERVO reset.

IWARNINGI

Do not attempt to reset SERVOS light(s)if DUAL FC FAIL warning light is on.
This action may cause affected control
surface(s) to be inoperative.
9.

Certain LEF malfunctions do not illuminate the LE
FLAPS caution light. The presence of higher thannormal buffet levels during maneuvering flight andreduced directional stability in the high AOA region
are indications that the LEF have failed to schedule
properly.
If LE FLAPS caution light illuminates or amalfunction is suspected (without LE FLAPS caution
light on):1.

I WARNING I

Exceeding 12 degrees AOA reduces
departure resistance. Limit rolling
maneuvers to a maximum bank angle
change of 90 degrees and avoid rapidroll rates.

Landing at greater than 12 degrees AOA
may result in inadequate control authority.
10.

Stores - Selectively jettison asymmetric stores.
Refer to SELECTIVE JETTISON, this section.

11.

Land as soon as possible.

2.

SERVO ELEC RESET switch - ELEC.

I WARNING I

NOTE• If possible, avoid landing in a crosswind with a rudder or flaperon ISA
lockout. For a locked-out rudder ISA,
lower the nose after touchdown anduse flaperon and NWS for directional
control. With a locked-out horizontal
tail ISA, pitch commands result in roll
inputs which are easily compensatedfor by using flaperons.
• With a locked-out flaperon ISA, TEF's
do not extend. If ISA lockout occurs with
TEF's extended, TEF's retract. Therefore, final approach airspeed is approximately 20 knots faster than normal. The
T.O.ILAND CONFIG warning lightilluminates during approach.
12.

Do not attempt SERVO or ELEC reset
if DUAL Fe FAIL warning and P, R, or
Y lights are on. This action may cause
affected control surface(s) to be inoperative.
If LE FLAPS caution light resets:
3.

4.

Airspeed - Decelerate to subsonic, if supersonic.

5.

LE FLAPS switch - LOCK (after LG is down).
Lock LEF's in landing configuration at final
approach airspeed at a safe altitude. This
makes final approach and landing as normal
as possible and protects against uncommanded LEF excursions close to the ground.

LEF Malfunction (Symmetric)

A symmetric LEF malfunction may be indicated by
the LE FLAPS caution light. This indicates that one
or both of the LEF branches have malfunctioned. LEF
may stop and remain fixed in position when themalfunction occurs. An LE FLAPS caution light may
also indicate that an asymmetry was detected and theasymmetry brakes have locked the LEF's. LEF's
should remain symmetrical (within 10 degrees).

Continue flight.

If LE FLAPS caution light does not reset or amalfunction is suspected (without the LE FLAPS
caution light):

Go to CONTROLLABILITY CHECK, this section.

6.

•
,

AOA - 12 degrees maximum.

I WARNING I

AOA - 12 degrees maximum.

-Land as soon as practical.
With the LEF's at or near full up, there are no
unique control inputs required. A small
increase in airspeed may be noted compared
to a normal landing approach at 11 degrees
AOA. With the LEF's at or near full down, theaircraft may tend to float in ground effect anda slight forward stick force may be required.

-

-

•

,

~

~
~
~
~

,

'LLIIIILLLL.lII~

~.~!r~~~~~~~~~~~~~~

During engine shutdown:
7.

MAIN PWR switch - Do not place to OFF until
engine rpm has reached zero.
I::;'A:V~!:;:;::S

l_~~~:_~~:~~~~:':~

Placing MAIN PWR switch to OFF
before hydraulic pressure is lost may
cause damage to two LEF shafts.

approach at approximately 8 degrees AOA (fly no
lower than 6 degrees AOA) with minimum roundout
for touchdown. Immediately prior to touchdown, use
rudder as required to align the aircraft with the
runway. Reduce the rate of descent somewhat prior totouchdown, as required, but do not flare or raise thenose above 10 degrees AOA because available roll
control is reduced and heading drift will increase a
AOA increases. Lower the nose immediately afteltouchdown. Directional control should not be a
problem.
If LEF asymmetry occurs:

The most likely cause of an asymmetric LEF
malfunction is a mechanical disconnect in one of the
LEF drive trains accompanied by a failure of theasymmetry brake. This failure may not illuminate
the LE FLAPS caution light. The first indication of anasymmetry is an uncommanded roll. The failed LEF
may be as much as 90 degrees up or down. Adequateroll control is available below 10 degrees AOA at
subsonic speeds. Use lateral stick for roll control. Useroll trim to reduce lateral stick force as required. Do
not attempt to achieve coordinated flight. Avoid using
rudder except to reduce sideslip when jettisoning
stores or to aid in maintaining' desired ground track
during the final part of landing approach. Do not use
rudder trim. If the yaw is away from the failed LEF(Le., nose left yaw with right LEF failed up), rudder
inputs to reduce resulting sideslip actually aggravatethe situation by increasing roll control requirements.
Accepting some sideslip reduces roll control requirements. To prevent excessive sideslip, maintain AOA
as low as practical. Banked flight reduces the amountof heading change due to sideslip-induced heading
drift. Lock the good LEF as close to symmetrical as
possible to aid in roll control and to prevent transients
caused by automatic scheduling. Monitor fuelconsumption since significantly higher thrust isrequired to compensate for the increased drag.
Selectively jettison stores to reduce asymmetry andsideslip and reduce fuel weight as necessary to reduce
approach speed. Perform a controllability check. The
aircraft tends to roll into the wing with the least lift
(i.e., the heavy wing). If the LEF is failed up, lift onthat wing is less. If the LEF is failed down, lift on thatwing is more or less depending on the failed LEF
position and the position at which the other LEF islocked. If there is a significant crosswind, diminish
crosswind effects, if possible, by landing with theheavy wing upwind. Fly a shallow, straight-in

~
~

3-120

1.

AOA - 6-10 degrees.

I WARNING I• Exceeding 10 degrees AOA may resultin insufficient roll authority. Limitrolling maneuvers to gentle roll in witha maximum bank angle of 30 degrees.
• Flying a fast approach (lower than 6
degrees AOA) presents additional
control difficulties caused by a change
in the path of the disturbed airflow
coming off the failed LEF.
2.

Lateral stick/roll trim - As required.

I WARNING I
Minimize rudder inputs. Use rudder as
required to reduce sideslip when
jettisoning stores or to aid in maintaining desired ground track during thefinal part of landing approach. Do not
use rudder trim.
3.

LE FLAPS switch - LOCK.
Lock operating LEF as near symmetrical as
possible.

4.

Stores - Jettison (if required).
Consider selective jettison of stores from theheavy wing as a means to reduce roll control
requiremen ts. Refer to SELECTIVE JETTISON, this section.

5.

Fuel weight - Reduce (if feasible/required).

Change 2

L'LLLLLLLLLLIL~

TEF Malfunctions

Reduce fuel weight ifpilot arm fatigue
is not a factor. Fuel flow is significan tly
higher with an LEF failed full up or
down and must be considered during
recovery.
6.

Controllability - Check.
Lower LG at a safe altitude and check
handling qualities at 6-8 degrees AOA.

7.

Land as soon as practical.

IWARNING

I• Prior to landing with a significant
asymmetric LEF condition, consider
aircraft configuration, pilot experience level, pilot arm fatigue, airfield
facilities, weather, winds, and light
conditions (day/night). If conditions
are not favorable, a controlled ejection
is recommended.
• If crosswind com ponen tis grea ter than
10 knots, choose a runway, if possible,
which allows landing with the heavy
wing upwind. Fly a shallow,
straight-in approach at approximately
8 degrees AOA (fly no lower than 6
degrees AOA) with minimum roundout for touchdown. Use rudder, as
required, to align aircraft with the
runway immediately prior to touchdown.
8.

Stick - Lower the nose immediately after
touchdown.

Until WOW, forward stick pressure inexcess of approximately 2 pounds
results in full trailing edge down
deflection of the horizon tal tails with
reduced directional control and wheel
braking effectiveness.
If departure-end arrestment is required:

9.

HOOK switch - DN.

The most likely cause of a TEF malfunction is amechanical failure of the flaperon connection to thewing. This malfunction usually results in theflaperon remaining attached at the fuselage with theflaperon bending 60-90 degrees up and damage to theaffected flaperon ISA. If this situation occurs,
perfonn a controllability check.

To precl ude asymmetric TEF extension, use al ternate
LG extension with the LG handle in UP position. The
nozzle idle area reset is not available resulting inhigher than nonnal landing thrust. NWS is notavailable, and BRAKES CHAN 2 must be selected.
FLCS remains in cruise gains. Primary consideration
should be given to landing as soon as possible due tothe possibility of loss of hydraulic fluid which may
result in an eventual dual hydraulic failure .
Trim Malfunction

1.

TRIM/AP DISC switch - DISC, then NORM.

If normal operation is not restored:

2.

TRIM/AP DISC switch - DISC.
Autopilot cannot be engaged.

Stick Interference

I WARNING I
Prior to any ejection seat movement,
clear the area around the stick.
Stick interference can occur at anytime for a numberof reasons. Known hazards include intentional orunintentional input by the passenger/pilot not incontrol, the utility light/adjustable sliding holder, or
the right lapbelt buckle.
Con tact between the right leg/knee and the stick can
result in an unintentional right roll command. The
resulting right roll may be perceived as a flightaircraft when acontrol problem especially in thepassenger/pilot not in control unknowingly interfereswith the stick. The probability of interference isincreased with feet on the floor versus feet on therudder pedals, bulky personal equipment, or g-suit
inflation.

rru

~~~IIIIIIIIIII~~I~ARNINGI

FUEL MALFUNCTIONS
Fuel Management System PFL

rn.J The passenger/pilot not in control
must take care not to interfere with the
stick as a result of leg/knee movement
or g-suit inflation. If stick interferenceis suspected, the pilot in control should
depress the paddle switch to eliminate
undesired inputs.
Contact between the utility light/adjustable sliding
holder and the stick can result in unintentional stickcommands. Unintentional stick commands may beperceived as a flight control problem. If
uncommanded stick inputs are encountered, checkfor interference between the stick and utility
light/adjustable sliding holder.

I WARNING I
Failure to properly secure utility
light/adjustable sliding holder can
result in stick interference. The
adjustable sliding holder may become
loose and come in contact with the
stick. If rn.J stick interference isconfirmed, undesired inputs may beeliminated by placing the STICKCONTROL switch to the appropriate
position and depressing the paddle
switch.

~I

If the seat is moved after an object (especially theright lapbelt buckle) becomes lodged between the seatand stick, unintended stick inputs can occur.
Reversing the direction of the initial seat movement
should correct the situation.

I

WARNINGI• The lapbelt should remain fastened at
all times. If the lapbelt is opened inflight, caution must be taken to insure
the right lapbelt buckle does not
become lodged between the ejection
seat and stick.

An FMS FAIL PFL indicates that the fuel reference
voltage supplied to the FCC is out of tolerance. Fuelsystem effects associated with the PFL range fromdegraded FCC fuel computations (e.g., BINGO fuel)to degradation/failure of the fuel quantity indicating
system. If an FMS FAIL PFL occurs, monitor the
FUEL quantity indicator for proper operation.
Fuel Leak

A fuel leak may first be noticed by visual means, fuelimbalance, an unexpected FWD or AFT FUEL LOW
caution light, or an unusually high fuel flow
indication. Monitor the totalizer to determine
whether or not a leak exists.
If a fuel leak is suspected (indicated by abnormallyhigh fuel flow, by totalizer decreasing at abnormal
rate, or by visual means):

1.

Range - Maximize.
If a suitable landing field is not within gliding
distance, consider increasing airspeed andal ti tude (wi thou t the use of AB) to maximize
range by using fuel which would otherwise be
lost.

I

WARNING

Avoid negative g flight when either
reservoir is not full.
If fuel flow is abnormally high:2.

ENG FEED knob - OFF.
Leak is in the engine feed line or engine
components .

3.

Land as soon as possible.
Consider stores jettison if range is critical.
Consider an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING, this section.

If fuel flow is normal:• Do not move the seat with the lapbelt
disconnected.

~~

3- 122

Change 2I

2.

ENG FEED knob - NORM.

"""""""""""...

T.O.1F-16A-l~

If leak is from the forward system:

3.

FUEL QTY SEL knob - Out of NORM.
This action stops automatic forward fueltransfer.

FWD

400 pounds

F\VD

250 pounds

AFT

250 pounds

AFT

400 pounds

If external tanks contain fuel:If FWD FUEL LOW and/or AFT FUEL LOW caution
light illuminates:

4.

1.

TANK INERTING switch - TANK INERTING to reduce internal tank pressurization.

5.

AIR REFUEL switch - OPEN.

6.

Land as soon as possible.
Consider stores jettison if range is critical.

7.

AOA - 15 degrees maximum.

I WARNING I
Aft fuel heavy (red portion of AL
pointer showing) results in increased
susceptibility to departure and deep
stall conditions. Limit AOA and avoid
maximum command rolling maneuvers.

Fuel Low

A fuel low caution light may be caused by a fuel leak,trapped external fuel, a fuel imbalance between theforward and aft systems, prolonged AB operation, ora fuel sensing problem.

~

,~

.~

~

2.

ENG FEED knob - NORM.

3.

FUEL QTY SEL knob - RSVR.
Leave FUEL QTY SEL knob out of NORM if
FUEL quantity indicator displays erroneous
information.

If either or both reservoir tanks are low:
NOTE

Fuel flow indications may fluctuatewith either reservoir empty.
4.

If two-point aerodynamic braking isused with an aft CG, pitch overshoots
may occur and the nozzle, speedbrakes, and ventral fins may contact
the runway.

~::b::16:0: ,Ph.

Limit fuel flow to the minimum
required to sustain flight while thecause of the fuel low light(s) isdetermined. Avoid negative g flightwhen either reservoir is not full.

If aft fuel imbalance exists (aft CG):~

Fuel flow - Reduce to the minimum required

to sustain

If external tanks are not installed or when they are
empty:~

Land as soon as possible.

~
~
~
~
,

Consider stores jettison if range is critical.
Consider an SFO. Refer to SIMULATEDFLAMEOUT (SFO)LANDING, this section.

~

...~

~,.

~

If a fuel leak is suspected (indicated by abnOrmallY,
high fuel flow, by totalizer decreasing at abnormal
rate, or by visual means):
5.

Go to FUEL LEAK, this section.

If external fuel has not transferred:
6.

Go to TRAPPED EXTERNAL FUEL, thissection.

~,~:IIIIIIIIIIIIII

If forward and aft fuselage fuel is not properly

The FWD FUEL LOW and AFT FUEL LOW caution
lights indicate reservoir tank quantities are lessthan:~~lan~:d~o FUEL IMBALANCE, this section.
Change 1

~

,
3-123

~

"""'~'l'IIIIIIIIIIII~"~~

10. 1 F- 16A- 1

If fuel is properly balanced:~..

If FUEL HOT caution light illuminates or gravity
feed situation exists:

NOTE

~
~

Afuel line between the reservoir and

IWARNING

FFP may be ruptured, causing fuel tocycle between tanks in the same
system.

"~• Engine flameout may occur at low fuelflow rates when in a hot fuel situation .

~ ;; res~:::r:::::;i~::::::::·~ :~:~r:~p::~e::::s:b::'S:
~~

:; AL
indicator is inoperative:• Engine flameout may occur wheneither reservoir tank empties if agravity feed condition exists.

FUEL

quantity~~ :fALI::~:u::int::Ites::::~ities _Check and
~
"~ 5.
~~

Land as soon as possible.
Consider an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING, this section.

compare with totalizer.
Monitor reservoir tanks to insure they are
maintained full.

"~~ :01 F~::::~:::i:s::::tical.

"~ Gravity feed from the reservoirs to the engine occurs~~~ after loss of the main generator and failure of either

hydraulic system A or the FFP. Failure of the FFP
may be detected by improper fuel balance. Fuelcontinues to be transferred to both reservoirs bysiphoning action. Fuel distribution cannot be
manually or automatically controlled during gravity
feed. Minimize aircraft maneuvering for duration of"~ flight. Due to the ingestion of air into the engine fuel~~~ :::::~:::c::~::::e::::: :::~one::::.
~~

may result from high speed flight or fuel system/heatexchanger malfunctions. Excess fuel temperatures
may result in engine malfunctions. Engine flameout
may occur at low flow rates associated with the"~ landing pattern due to hot fuel.
~~ Fuel flow above 4000 pph minimizes fuel temperature

~

~

rise.

~ 3-124~

Change 13

1

1.

AIR REFUEL switch - Check CLOSE.

2.

TANK INERTING switch - Check OFF.

3.

Altitude -10,000 feet maximum (if practical).
Minimize aircraft maneuvering for durationof flight.

4.

Fuel flow - 4000 pph minimum until landing isassured when in a hot fuel situation.

If FUEL HOT caution light goes off:
5.

Land as soon as practical.

If FUEL HOT caution light remains on or gravity
feed situation exists:
5.

Land as soon as possible.
Consider an SFO. Refer to SIMULATEDFLAMEOUT (SFO) LANDING, this section.

Fuel Imbalance

NOTE

rn Compliance with the fuelmanagement in Section II may resultin forward fuel distributions which
uncover the red position of the AL
pointer. These fuel distributions are
not considered improper.
LESS

Refer to figure 3-15. A fuel imbalance when notcarrying an external tank(s) indicates a system
malfunction. A fuel imbalance when carrying an
external fuel tank(s) may be the result of normal
system operating tolerances.

•

NOTE
• Any correction required per total fupl
quantity usage with internal fuel only
indicates a system malfunction.
• More than one correction per total fuelquantity usage with either a 300-galIon fuel tank or two 370-gallon fueltanks indicate a system malfunction.
• More than two corrections per total
fuel quantity usage with three external fuel tanks indicate a system
malfunction.
A fuel imbalance is indicated by the red portion of the
AL pointer. This may be caused by an FFP
malfunction, fuel leak, uneven or partial refueling(either ground or AR), or a malfunction of theautomatic forward fuel transfer system.

4.

5.

If imbalance is not corrected:

6.

An unexpected FWD or AFT FUEL LOW caution
light may also be an indication of fuel imbalance;
however, verify that forward fuselage fuel and aftfuselage fuel (as indicated by AL and FR pointerswith FUEL QTY SEL knob in NORM) are notproperly balanced and that a leak does not exist
before selecting FWD or AFT ENG FEED.
If fuel imbalance is indicated by AL and FR pointerswith FUEL QTY SEL knob in NORM:
1.

Fuel flow - Reduce to the minimum required tosustain flight below 6000 pph.

I WARNING I
Limit fuel flow to the minimum
required to sustain flight while thecause is determined. Avoid negative gflight when either reservoir is not full.
If aft fuel imbalance exists (aft CG):

2.

AOA - 15 degrees maximum.

I

WARNINGI

Aft fuel heavy (red portion of AL
pointer showing) results in increased
susceptibility to departure and deep
stall conditions. Limit AOA and avoid
maximum command rolling maneuverso

ENG FEED knob - FWD or AFT.

If two-poin t aerodynamic braking isused with an aft CG, pitch overshoots
may occur and the nozzle, speedbrakes, and ventral fins may contact
the runway.
If proper distribution is attained:

6.

ENG FEED knob - NORM.

7.

Fuel balance - Monitor.

Trapped External FuelI

WARNINGI• LESS ~ A TRAP FUEL indication in
the HUD may be a symptom of an
external fuel leak. If a fuel leak issuspected (indicated by abnormallyhigh fuel flow, by totalizer decreasing
at abnormal rate, or by visual means),refer to FUEL LEAK, this section.

• With trapped external fuel, the totalizer does not indicate total usable fuel.
U sable fuel is the totalizer quantity
less the external fuel quantity.

r;!/~-161r ~ ~ ~ ~ ~ . I. I. I. I. I~~ Fuel Imbalance Indications

FUEL IMBALANCE WARNING tAFT CGIFUEL IMBALANCE WARNING (A FT CG)

Fuel imbalance warning (red portion of AL
pointer) shows when forward fuselage fuel(FR) indication is less than aft fuselage
fuel (AU indication.

Fuel imbalance warning (red portion of AL
pointer) shows when forward fuselage fuel
(FRJ indication is more than 1350 pounds
less than aft fuselage fuel (Al) indication.

Figure 3-15.

Accomplish steps 1 through 8 and 9 (if required)without delay;

external wing tanks, the centerline
tank, or all three external tanks. Witha three tank configuration, the fustindication that the centerline tank isfeeding is after the external wingtanks are emptied.

NOTE

1.

Fuel flow - Minimize.

2.

AIR REFUEL switch - Confirm in CLOSE.

3.

ENG FEED knob - NORM.

8.

Stick - Pulse aircraft in pitch several times bJ
applying differential g forces of approxim atelJ±2g.

If the AIR REFUEL switch was initially found in
CLOSE (step 2), perform step 9. If the AIR REFUEL
switch was initially found in OPEN (step 2), omit step
9.

4.

5.

TANK INERTING switch - TANK INERTING to reduce internal tank pressurization.

6.

EXT FUEL TRANS switch - WING FIRST.

NOTE
Selecting wings first bypasses electrical components that, if malfunctioning. can prevent fuel transfer from

~

7.

3-126

9.

AIR REFUEL switch - OPEN (1 second), then
CLOSE.
Open or close AR door at or below 400
knots/0.S5 mach.

10.

External tank fuel quantity - Monitor.
The time required to observe fuel transfer if
the problem is corrected can vary from 1-3
minutes (for a full centerline tank) to 10-12
minutes (for three external tanks with 500
pounds fuel in each) if reservoir tanks are full(i.e., both air ejectors are om.

""""ffff~~~
IWARNING

NOTE

1

If a trapped external fuel condition is
not discovered until either reservoir
tank is less than full or a fuel low lightis on, sufficient fuel transfer from theexternal tank(s) may not occur even if
the malfunction is corrected. Consider
fuselage fuel to be the only usable fuel.

Since the hydraulic system A pumpand the FLCS PMG share the same
shaft, a failure of the shaft results inloss of both systems.
1.

Do not reset ISA SERVOS.

2.

Land as soon as practical.
Make smooth control inputs and plan to fly astraight-in approach.

~
~

~

"
~

NOTE

If trapped external fuel occurs after air
refueling and completion of checklist
steps does not correct the problem,
consider descending well below thefreezing level to unfreeze the externalpressurization and vent valve. Cycling
the AR door at lower altitude may
restore normal operation.

11.

3.

System B HYD PRESS indicator - Monitor.

4.

Fuel balance - Monitor.
Fuel distribution must be controlled manually.

SYSTEM B FAI LU RE
NOTEEPU RUN light on may indicate a dual
hydraulic or PTO shaft failure.

Stores - Jettison (if required).

HYDRAULIC MALFUNCTIONS

A hydraulic system failure is indicated by illumination of the HYD/OIL PRESS, FLT CONT SYS, and all
five SERVOS lights. The HYD/OIL PRESS warning
light illuminates whenever either hydraulic systempressure drops below 1000 psi. The SERVOS lights
may illuminate prior to the HYD/OIL PRESS
warning light. Do not attempt to reset ISA SERVOS.
With system B hydraulic failure, perform alternate
LG extension with the LG handle up. This action
reduces the possibility of failing to unlock a LG door
actuator or the NLG extend/retract actuator due tolow or fluctuating system B hydraulic pressure.
Single Hydraulic Failure

I WARNING I
If hydraulic failure is due to structural
damage (e.g., battle damage, midair
collision, bird strike, fire, or hardlanding), the other system may bedamaged and failure can occur withlittle warning. The HYD PRESS
indicator may show normal pressure
until system fluid is depleted.
SYSTEM A FAI LU RE

The FLCS ISA's are operating in the nonredundantmode and the speedbrakes and FFP are inoperative.

The FLCS ISA's are operating in the nonredundantmode and normal braking, NWS, AR door operation,
gun operation, and normal LG extension are lost. ~lNQ] Drag chute operation is normal using drag chuteaccumulator pressure. Low hydraulic pressure may
cause one or more LG actuators to remain locked in
the LG up position. LG extension should beattempted in sufficient time to prepare for possible
LG up landing. Braking is available using brake/JFS
accumulator pressure. The fully charged brake/JFS
accumulators contain sufficient fluid for at least 75
seconds of continuous brake application. Use
aerodynamic braking to the maximum extent
possible. A single moderate and steady brakeapplication without cycling the antiskid should thenbe applied. Mter stopping, set the parking brake. If
there is reason to believe that the brake/JFS
accumulators are depleted or that directional control
may be a problem, an approach-end arrestment
should be considered.

~

~
~
~
~
~
~
~
~

1.

Do not reset ISA SERVOS.

~

2.

Land as soon as practical.
Make smooth control inputs and plan to fly astraight-in approach.

".

3.

ALT GEAR handle - Pull (190 knots maximum,
if practical).
Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

~~~

T.O. 1F- 16A- 1

~
~
~
~
~
~
~
~
~
~
~

III""
~
III""

~
~
~

• NWS is not available following alternate LG extension.
• Do not depress the ALT GEAR reset
button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.
4.

LG handle - DN. (Use DN LOCK REL button
if required.)

IWARNING

1inputs, decreasing pressure readings on both HYD
PRESS indicators, and associated warning andcaution lights. The EPU automatically provides
hydraulic pressure for system A when pressure ofboth hydraulic systems drops below 1000 psi. The
systems affected by dual hydraulic system failure
after the EPU is running are the same as thoseaffected by system B failure. Refer to SYSTEM B
FAILURE, this section.
1.

EPU switch - ON (if EPU run light is off).

2.

System A HYD PRESS indicator - Checkpressure increasing.

If hydraulic pressure does not increase or control

response is lost:
If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed resulting in higher
than normal landing thrust.
5.

HOOK switch - DN (if required).
Braking is available using brake/JFS accumulators only. To avoid brake activation andloss of brake/JFS accumulator pressure, do
not rest feet on brake pedals. If the brake/JFS
accumulators are depleted or if directional
control may be a problem, consider anapproach-end arrestment. Refer to CABLE
ARRESTMENT, this section.

3.

Eject.

If system A hydraulic pressure is restored:3.

Do not reset ISA SERVOS.

4.

EPU run light - Check light on at idle thrust.
NOTE

Before landing, confirm that the EPU
operates (EPU run light is on) with thethrottle in IDLE. If the EPU run lightgoes off, refer to ABNORMAL EPU
OPERATION, this section.

After landing:
6.

Stop straight ahead and engage parking brake.

G!:!!~~J

5.

Land as soon as possible.
Make smooth control inputs and plan to fly astraight-in approach.

6.

ALT GEAR handle - Pull (190 knots maximum,
if practical).
Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

• Brakes should be applied in a single,
moderate, and steady application
without cycling the antiskid.
• Brake pedal deflection of 1/16 inch
activates the brakes and bleeds thebrake/JFS accumulators. To avoid
brake activation and loss ofbrake/JFS
accumulator pressure, do not rest feeton brake pedals.

• NWS is not available following alternate LG extension.

• Do not attempt to taxi clear of the
runway. Loss of brake/JFS accumulator pressure results in the inability tostop or steer the aircraft.

A dual hydraulic system failure can be detected bysluggishness or lack of response to flight control• Do not depress the ALT GEAR reset
button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

7.

LG handle - DN. (Use DN LOCK REL button
ifrequired.)~ 3-128
Change 11
III""~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~""""""'£.~IWARNING,

2.

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed resulting in higher
than normal landing thrust.

8.

HOOK switch - DN (if required).
Braking is available using brake/JFS accumulators only. To avoid brake activation andloss of brake/JFS accumulator pressure, do
not rest feet on brake pedals. If the brakel JFS
accumulators are depleted or if directional
control may be a problem, consider anapproach-end arrestment. Refer to CABLE
ARRESTMENT, this section.

Mter landing:
9.

(200 knots

IWARNING,
LEF's may be locked until reset.
Departure susceptibility is increased
with LEF's locked. Limit rolling
maneuvers to a maximum bank angle
change of 90 degrees and avoid rapidroll rates.
If EPU run light is off and control response is lost:3.

Eject.

If EPU run light is on:

4.

Throttle - As required.

Stop straight ahead and engage parking brake.

• Brakes should be applied in a single,
moderate, and steady application
without cycling the antiskid.
• Brake pedal deflection of 1/16 inch
activates the brakes and bleeds thebrake/JFS accumulators. To avoid
brake activation and loss ofbrake/JFS
accumulator pressure, do not rest feeton the brake pedals.

Stall protection may be lost. Do notretard throttle below MIL until subsonic.
5.

Do not reset ISA SERVOS.

6.

ADI - Check for presence of OFF and/or AUX
warning flags.
If warning flag(s) is in view, refer to TOTALINS FAILURE, this section.

IWARNING,• Do not attempt to taxi clear of the
runway. Loss of brake/JFS accumulator pressure results in the inability tostop or steer the aircraft.

~ If only AUX flag is in view, pitch androll attitude information is likely to be
erroneous due to INS autorestart in
the attitude mode when other thanstraight and level, unaccelerated
flight conditions existed.

10. Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.
System B and Main Generator Failure (PTO Shaft)

A PTO shaft failure is indicated by failure of
hydraulic system B, the main generator, and theFLCS PMG. The EPU should start automatically toprovide emergency hydraulic and electrical power.
Mter accomplishing the appropriate emergency
procedures, refer to EMERGENCY POWER DISTRIBUTION, this section, to determine inoperativeequipment.
1.

AOA - 12 degrees maximum
minimum).

EPU switch - ON (ifEPU run light is ofD.

7.

8.
9.

SERVO ELEC RESET switch - ELEC.
Resets LEF's (if locked) and LE FLAPS,
CADC, and ADC caution lights (if on).

I PW200 I

~~

~

,

-

~

"

~

EEC BUC switch - OFF, then
EEC with throttle at midrange.

"

I PW2201 ~ rn::EJ AB RESET switch - AB RESET,~

then NORM.

"

10. Fuel balance - Monitor.

~

11.

"

EPU run light - Check light on at idle thrust.

NOTE

Before landing, confirm that the EPU
operates (EPU run light is on) with thethrottle in IDLE. If the EPU run lightgoes off, refer to ABNORMAL EPU
OPERATION, this section.
12. Land as soon as possible.
Make smooth control inputs and plan to fly astraight-in approach.
13. ALT GEAR handle - Pull (190 knots maximum,
if practical).
Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

• Brakes should be applied in a single,
moderate, and steady application
without cycling the antiskid.
• Brake pedal deflection of 1/16 inch
activates the brake/JFS accumulators.
To avoid brake activation and loss ofaccumulator fluid, do not rest feet on
the brake pedals.
• Do not attempt to taxi clear of the
runway. Loss of brake/JFS accumulator pressure results in the inability tostop or steer the aircraft.
17. EPU switch - OFF.
18. Refer to ACTIVATED EPU/HYDRAZINE
LEAK, this section.
INS FAILURES• NWS is not available following alternate LG extension.
• Do not depress the ALT GEAR reset
button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

Most INS failures are apparent and affect flight
operations significantly.
If the autopilot is engaged when an INS failure occurs

or during realignment to the attitude mode after an
INS failure, uncommanded pitch and roll flightcontrol inputs may occur.

14. LG handle - DN. (Use DN LOCK REL button
if required.)

IWARNINGIIWARNING I

The autopilot does not automaticallydisengage with INS failures. Failure to
manually disconnect the autopilot
may result in an unusual aircraft
attitude and disorientation.

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed resulting in higher
than normal landing thrust.
15.

HOOK switch - DN (if required).
Braking is available using brake/JFS accumulators only. To avoid brake activation andloss of brake/JFS accumulator pressure, do
not rest feet on brake pedals. If the brakel JFS
accumulators are depleted or if directional
control may be a problem, consider anapproach-end arrestment. Refer to CABLE
ARRESTMENT, this section.

After landing:
16. Stop straight ahead and engage parking
brake.

When the INU data is degraded, frozen, bad, or varies
from the rate sensor unit (RSU) data, the HUD
displays a flashing INU advisory.
Advisories are:
• INU BAD - INU data is erroneous. The INS
internal self-test has detected a failure.
• INU FRZ - INU data is frozen. The HUD has
detected unchanging INS navigation data.
• INU VEL - INU and RSU normal acceleration data
does not agree. The HUD has detected a significant
difference between INU and RSU accelerations.
The HUD INU advisory clears automatically after thecondition is corrected; the advisory may be manually
reset by momentarily positioning the DRIFT C/O
switch to TEST.

~~1~ ;;g~ ~ ~ ~ ~ ~ ~ ~ ~ ~""""",,~~~
G!:!!!~J

INS Computer Failure LESS @)A failure of the INS computer with the inertial
platform operating properly is indicated by the
AVIONICS caution light; AUX warning flag on the
ADI; loss of the heading scale, pitch ladder, FPM, androll scale in the HUD; and possibly a compass card
jump. The INS normally switches automatically tothe alternate (attitude) mode. In this condition, theHSI compass card is operative but must be oriented
to the best available magnetic heading using the ~rnfJ INSTR HDG knob. Full TACAN operation isavailable and attitude may be displayed normally on
the ADI even though the AUX warning flag is in view.

LESS @) The INS does not apply gyrobraking during any failure or normal
shutdown. The INU should not berestarted for at least 1 minute after an
INU power interrupt to prevent
damage to the INU. There are no
additional
aircraftmaneuvering
restrictions during this spin down
period.

2.

I WARNING I

If the INS computer fails:
1.

Cross-check attitude indications.

2.

~ rnfJ INSTR HDG knob - Slew HSI to match

Until the ADI OFF flag is out of view
after the INS is aligned in straight andlevel un accelerated flight, the ADI and
HUD display a false attitude.

best available magnetic heading.
Total INS Failure

A total INS failure is normally indicated by the
AVIONICS caution light, any INS fault code withseverity of 1 on the FCNP, and the ADI OFF and AUX
warning flags in view. The RSI compass card andbearing pointer and the ADI freeze. The HUD pitch
ladder, heading scale, FPM, and roll scale also blank.
Magnetic heading is displayed only on the magnetic
compass. Attitude reference is available only on the
SAL When the INSTR MODE knob is in TCN or
ILSrrCN, the HSI course deviation indicator, range
indicator, and TO-FROM indicator are operative, andthe capability to fly an inbound (or outbound) radial
to (or from) a TACAN station is available.

I WARNING I
It is possible for the displayed ADI
and/or HUD attitude to be in error withno ADI OFF or AUX warning flags inview and without an INS or HUD PFL.
Displayed RSI and/or HUD headings
may also be in error with no HSI OFF
or ADI AUX warning flags in view andwithout an INS or RUD PFL. Momentary warning flags may indicate
impending failure. To detect thesefailures and maintain proper flightorientation, basic and backup instruments must be cross-checked.
1.

INS PWR pushbutton - Off for @ID 10 seconds,
LESS @ID 1 minute.

FCNP FUNCTION knob - ATTD (cw).

3.

INS PWR pushbutton - On.

4.

Attitude - Straight, level, and unaccelerated
until ADI OFF warning flag goes out of view
after approximately@ID 10, LESS ~ 35 seconds.

5.

ADI and RUD - Verify attitude information
correct.

6.

~ rnfJ INSTR RDG knob - Slew RSI to match

best available magnetic heading.
OUT-OF-CONTROL RECOVERY

Refer to OUT-OF-CONTROL CHARACTERISTICS,
Section VI, for a detailed discussion of flightcharacteristics and indications during departures,deep stalls, spins, and recoveries.
In order to prevent a departure, immediately initiate
recovery after the low speed warning tone comes on.
Prompt recovery is even more critical with heavyweight loadings or during hard maneuvering since
the airspeed bleedoff is more rapid. Additionally, ayaw departure may be prevented if controls arepromptly neutralized following an uncommanded
nose slice or roll hesitation.
In order to minimize time and altitude loss following
a departure, immediately release the stick andrudder controls. The aircraft should be allowed theopportunity to self-recover. Self-recoveries usuallyoccur within the first two postdeparture pitch
oscillations, and may take up to 10-20 seconds.
Recovery is indicated by the nose dropping and the
AOA remaining below 25 degrees. Fly the aircraft at
a low AOA until airspeed reaches 200 knots or more
(if altitude permits) and recover from the resulting
dive. If the departure does not result in self-recovery,
then the aircraft is in a deep stall or spin.
Change 11~131~

Departures at high altitude may result in an engine
stall or stagnation. Ifin AB during an out-of-control
situation, retard the throttle to MIL. If at MIL or
below, do not move the throttle. Do not advance thethrottle until beginning the dive recovery.
Upright deep stalls may be very stable with little orno pitch motions or may be very oscillatory withlarge pitch, roll, and yaw motions. Generally, aclean configuration results in a deep stall with anear wings-level pitching motion.
During upright deep stalls with a centerline store,
particularly a 300-gallon fuel tank, the aircraft
tends to roll and yaw right while pitching up, androll and yaw left while pitching down. During deep
stalls with 370-gallon fuel tanks, the aircraft nose
motion appears triangular. This motion is characterized by a roll and yaw right while pitching up,
followed by a pitch down, a hesitation, and yaw to
the left.
In an upright deep stall or spin, the yaw rate limiter
automatically provides antispin controls and therudder authority limiter prevents pilot yaw
commands. The yaw rate limiter is effective inpreventing spins with almost all CAT I loadings.
However, following a yaw departure above 25,000
feet, aircraft with CAT I loadings that have all thefollowing characteristics may spin:• Centerline store.
• Inlet mounted pod(s).
• Lateral asymmetry greater than 300 pounds at
stations 1, 2, or 3.
Upright spins following a yaw departure can bedisorienting. The initial portion of the spin ischaracterized by highly oscillatory motions and ahigh yaw rate (70 to 100 degrees per second).
Initially, the aircraft spins roughly around theaircraft's flight path at departure. As the spin
continues, the rotation axes eventually becomes
vertical. Very noticeable forward g (eyeballs out)
and sideforces are present.
In a spin, the yaw rate must be allowed to subside
before the aircraft can be recovered. This may
require 20 to 30 seconds. Pitch, roll, and yaw
oscillations associated with a deep stall should not
be confused with the continuous yaw rotation of aspin. When the yaw rotation subsides, the aircraft
will either recover or will settle into an upright deep
stall.
In an inverted deep stall or spin, antispin controls
must be applied using rudder opposite the yaw
direction until the rotation stops or is minimized
prior to positioning the MPO switch to OVRD. The~

3- 132

Change 4~~~~~~~

best indication of yaw direction is the aircraft nose
movement relative to the horizon. If unable todetermine yaw direction with outside references,
the turn needle may be useful. If the nose is yawing
to the pilot's left (turn needle - left), push rightrudder to stop the yaw motion.

I

The aircraft must be rocked out of a deep stall with
the MPO switch held in OVRD until recovery iscomplete. The MPO switch allows the pilot to use
the horizontal tail surfaces to reinforce pitch
oscillations until the pitch rates are sufficient forrecovery. When sufficient nosedown pitch rate is
generated to reduce the AOA below the deep stall
AOA, the aircraft will recover.
The MPO switch must remain in the OVRD position
during pitch rocking. If the MPO switch is released, I
the horizontal tails reposition to reduce AOA andmay negate any pitch oscillations. Additionally, if
the MPO switch is positioned to OVRD without any
stick commands, the horizontal tails streamline
and prevent recovery.

I

In an upright deep stall, begin pitch rocking
in-phase with nose movement; i.e., if the nose ispitching up, pull back on the stick. Maintain aft
stick until the maximum pitch attitude is reached,which is indicated by the nose stopping andreversing direction, and then push full forward on
the stick to generate a nosedown pitch rate. If thenosedown pitch rate is high enough to break the
deep stall, the aircraft will recover.
During some upright deep stalls, the aircraft may
be stable with essentially no pitching motion. In
these cases, pull full aft stick (away from theground) and monitor nose movement. If nose
movement occurs, continue stick cycling in-phase.
If nose movement is not apparent after 3-4 seconds,
then push full forward on the stick to generate a
nosedown pitch rate. This nosedown pitch rate may
be sufficient to reduce AOA below the deep stall
AOA and recover the aircraft. If the nose does notcontinue down but reverses and starts up, pull backon the stick and continue to reinforce these pitch
cycles. Proper pitch rocking is accomplished byallowing the nose to lead stick motion; i.e., whennose movement reverses, the stick should bereversed. When sufficient nosedown pitch rate is
generated to reduce the AOA below the deep stall
AOA, the aircraft will recover.
During upright deep stalls that are not stable, rolland yaw motions make it more difficult todetermine proper recovery inputs; however, pitch
attitude is still the best indication available. This
pitch attitude is determined by the nose position
with respect to the horizon. If unable to determine
pitch motions with outside references, the ADI may
be useful.

I
I

I

T.O. 1 F-16A-1

With proper stick cycling, the magnitude of thepitch oscillations progressively increases untillarge enough for recovery. Rapid fore and aft cyclingof the stick or cycling out of phase with the pitchingmotion of the aircraft will not be effective and may
prevent recovery. Pitch inputs must be abrupt andmaximum command. Pitch inputs that are smoothor less than maximum command do not generate
pitch rate as effectively, and may prevent recovery.
Normally, only one or two correctly applied cycles
are required to break a deep stall; however, thepresence of stores, particularly a 300-gallon fueltank or 370-gallon fuel tanks, may necessitate fiveor more properly executed stick cycles for recovery.
Altitude loss is approximately 1000-1500 feet per
pitch rock cycle.
Ifinverted, the same pitch rocking procedures apply
except if no pitch motion is apparent, the first stickcommand should be full forward (away from theground). Inverted deep stalls are generally stable,
regardless of the stores configuration.

I
I

I

If the pitch rate is still high as the aircraft recovers,
there may be a tendency for the aircraft to continue
pitching through to a deep stall in the otherdirection. Attempt to stop the nose in a near vertical
dive by tracking a spot on the ground. If the aircraft
does transition to an opposite AOA deep stall, it may
be very disorienting; however, pitch oscillations are
generally high and recovery should be rapid with afew properly executed stick cycles. Recovery is
confirmed by the nose remaining down and the AOA
remaining in the normal range. As the airspeed
increases above 200 knots, release the MPO switch,
maintain neutral roll and yaw commands, andapply pitch commands as required to recover fromthe resulting dive using MILlAB thrust.

IWARNING,• Recovery from a deep stall condition
will present a low airspeed situation inwhich the aircraft may require morethan 6000 feet of altitude to attainlevel flight.

• If recovery (pitch rate stopped, AOA
within -5 to +25 degrees, and airspeed
200 knots or greater) is not apparent
by 6000 feet AGL, eject.
The engine may stall when out of control. Also, FLCS
failure indications may occur. Ignore these indications and concentrate on recovery.

I

In the event of a departure from controlled flight,
accomplish as much of the following as required toeffect a recovery:
1.

Controls - Release.

2.

Throttle - MIL ifin AB.
If other than AB, do not move the throttle.

If in an inverted deep stall:
Negative g and AOA indicator pegged at -5
degrees.
3.

Rudder - Opposite yaw direction.
Neutralize rudder as rotation stops and thenuse as required to minimize rotation. Turnneedle deflection indicates yaw direction.

If still out of control:
Positive g, AOA indicator pegged at 32 degrees(upright deep stall) or negative g, AOA indicatorpegged at -5 degrees (inverted deep stall).
4.

MPO switch - OVRD and hold.
Maintain firm pressure.

IWARNING

1

The MPO switch must be held in the
OVRD position until the deep stall ispositively broken as evidenced by thepitch rate stopping, AOA in the normal
range (-5 to +25 degrees), and airspeed
increasing above 200 knots. Early
release of the MPO switch may delay
recovery.

Change 43- 132.1/(3-132.2 blank)

11111111111111111111111
10. 1 F- 16A- 1

IWARNING

If unable to descend immediately:1• If upright, applying rudder with the
MPO engaged may delay recovery.

I
5.

• Failure to adequately secure andtighten lapbelt may result in inability
to reach and operate the MPO switch
during out-of-control situations.

2.

Emergency oxygen - Activate.

3.

Oxygen hose - Disconnect.

PBG MALFUNCTION m

A malfunction of the oxygen regulator while in PBG
may cause excessive pressure or failure of pressure
to decrease when g is reduced.

If excessive pressure is experienced or high pressure
continues after g is reduced:

Stick - Cycle in phase.

1.

I

WARNINGI• Pitch rocking with a high sustained
yaw rate may prevent recovery. Delaystick inputs until yaw rotation stops oris minimized. Pitch, roll, and yaw
oscillations associated with a deep
stall should not be confused with thecontinuous yaw rotation associatedwith a spin.
• After recovery, do not attempt SERVOor ELEC reset if DUAL FC FAIL
warning and P, R, or Y lights are on.
This action may cause affected control
surface(s) to be inoperative.

OXYGEN mode lever - ON.

If pressure is not relieved:2.

Oxygen hose - Disconnect.

3.

Cockpit pressure altitude -10,000 feet maximum.

If unable to descend immediately:
4.

Emergency oxygen - Activate.

5.

Land as soon as practical.

SMOKE OR FUMESAll unidentified odors will be considered toxic. Do nottake off when unidentified odors are present. Do notconfuse ECS condensation for smoke.
If smoke or fumes are detected:
1.

OXYGEN REGULATOR - Check ON, 100%,
and
EMER, LESSEMERGENCY.

mm

OXYGEN MALFUNCTIONNOTEThe OXY LOW caution light indicates oxygen
quantity below 0.5 liter or pressure below 42 psi.
If OXY LOW caution light illuminates:
1.

Cockpit pressure altitude - 10,000 feet maximum.

The emergency oxygen bottle is notrecommended for use in the smoke andfumes environment unless aircraft oxygen supply contamination is suspected.
Activation of the emergency oxygen
bottle does not prevent cockpit smoke orfumes from entering the oxygen mask.

Change 4

3-'33

-

-

-_ .....-

~""""""""""~~T. O. 1 F- 1 6A- 1'IIIIIIIIIIIIII

2.

Altitude - 25,000 feet maximum.

STRAIGHT-IN LANDING

3.

Airspeed - 500 knots maximum.

4.

AIR SOURCE knob - RAM.

A straight-in landing is recommended for emergencies which dictate minimum maneuvering inputs
such as hydraulic, flight control, or electricalproblems or situations which result in a relatively
high thrust level being maintained to touchdown
such as a stuck or closed nozzle or when the engine isoperating satisfactorily in IPW200 I BUC, I PW2201
SEC. A controllability check should be accomplishedprior to commencing the approach if minimum flying
airspeeds or control difficulties are experienced or are
anticipated.

~
~
~
~

~
~

~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~
~

External fuel cannot be transferred in OFF or
RAM. Consider jettisoning tanks to decreasedrag if range is critical and the ECS cannot be
turned on for short periods of time to transfer
fuel.
5.

Nonessential electrical equipment - Off.

NOTE

NOTE
Ifin VMC and the ADI and RSI are notrequired for flight, the INS should beconsidered nonessential.

6.

Determine cause of smoke or fumes andcorrect (if possible).

NOTE• Smoke in the cockpit may be indicativeof an engine oil system malfunction. If
possible, retard throttle to lowest
setting possible to sustain flight andmonitor the OIL pressure indicator.
Refer to OIL SYSTEM MALFUNCTION, this section, if appropriate.
• Any odor that smells of burning flesh
may be indicative of bird ingestion into
the engine. Monitor engine instruments for signs of abnormal operation.
7.

Land as soon as possible.

If cockpit visibility precludes safe operation:

8.

Airspeed - 180 knots maximum.

9.

Seat - Full down.

10. ALT FLAPS switch - EXTEND.
11.

Canopy - Jettison.

LANDING EMERGENCIES
Generally, the type of pattern flown in an emergency
is either an SFO or straight-in approach and depends
on several factors:• Nature of the emergency.
• Weather conditions.
• Day or night.
• Proximity of a suitable landing runway.
• Fuel status.
3-134

Change 3

When landing in I PW200 I BUC, I PW2201
SEC, an increased ground roll distanceis required due to higher idle thrust.

SIMULATED FLAMEOUT (SFO) LANDING
Anytime engine failure is anticipated (abnormal
engine response, oil system failures, low fuel, etc.), an
SFO landing should be performed. At or just prior tohigh key, turn the EPU on, and if engine seizure is not
anticipated, turn the JFS on and verify theiroperation (EPU run and JFS RUN lights on). If theengine is still running at touchdown, the JFS shuts
down at WOW.
Fly the SFO pattern and landing in accordance withthe procedures for FLAMEOUT LANDING, thissection. If the engine fails, this action provides
sufficient energy to safely land the aircraft or to zoom
and eject if a safe landing cannot be made.
An SFO landing is not recommended when landing
with the engine operating satisfactorily in I PW200 I
BUC, I PW2201 SEC. The higher level of idle thrust
may result in a long and fast landing and difficulty
stopping the aircraft.

To simulate an engine out glide with the LG up, use
idle thrust and 30 degrees speedbrakes. From thefront cockpit, this equates to the intersection of the
top of the speedbrakes and a line drawn from the tipof the horizontal tail to the top of the vertical tail root
fairing. To simulate an engine out glide with the LG
down, use idle thrust and 20 degrees speedbrakes.
From the front cockpit, this equates to theintersection of the top of speedbrakes and a line
drawn from the tip of the horizontal tail to the base of
the vertical tail root fairing. The additional dragproduced during an engine-out condition is equivalent to retaining stores with a drag index of 170 with
the LG up or 70 with the LG down. If stores are
retained, adjust speedbrake deflection accordingly. If
the engine fails, close speedbrakes and jettison
stores.

""""",,~~~
When flying an SFO approach with the engine
operating at a higher thrust level than normal idle,
control descent rate and airspeed with the speedbrakes rather than adjust the ground track. If thrustis excessively high or if landing on a runway where
stopping distance may be critical, the procedures for
ABNORMAL ENGINE RESPONSE, this section,
should be considered.

4.

I WaRNING I
If a condition which might cause
asymmetric TEF extension exists,
consider alternate LG extension with
the LG handle in UP to preclude TEF
extension. If the LG handle remainsup:

I WARNING I
On runways with less than 8000 feet
and without arresting gear Irm IfmJ ordrag chute, there may be insufficientdistance to safely stop the aircraft.

• Final approach airspeed is 20
knots higher than normal.
• The T.O.ILAND CONFIG warning
light may illuminate.

Mter touchdown from an SFO landing, use a normal
or short field stopping technique as required by the
stopping distance available. Extend the hook if
required. If the engine rpm is greater than normal
with the throttle in IDLE or some other malfunction
requires excessive braking action to maintain a safe
taxi speed, the brakes may absorb a high amount ofenergy in a short period of time. If required, refer to
HOT BRAKES, this section.

• Nozzle remains closed resulting inhigher than normal landing
thrust.
• NWS is inoperative.
• BRAKES
selected.

CONTROLLABILITY CHECK

Attain safe altitude.

In the event that structural damage of
unknown extent is encountered or if
continued control of the aircraft is in
doubt, consider accomplishing applicable steps of EJECTION (TIMEPERMITTING), this section, prior toproceeding with CONTROLLABILITYCHECK.
2.

GW - Reduce (as required).
LE FLAPS switch - LOCK (if required).
If LEF damage is observed, consider locking
LEF's.

2

must

be

• The LG handle warning lightremains on to indicate the positionof the gear handle is not inagreement with the actual gear
position.

The following items should be accomplished:

NOTE

CHAN

• FLCS remains in cruise gains.
Consider positioning AIR REFUEL switch to OPEN to obtaintakeoff and landing gains.

When structural damage or any other failure that
may adversely affect aircraft handling characteristics is known or suspected, a controllability checkshould be performed.

1.

Determine optimum configuration available
for landing.

5.
6.

~

.,

-~
-~
~
~

~

Stores - Selectively jettison (if required).
Refer to SELECTIVE JETTISON, this section.

.,

Slow only to that AOA/airspeed which allows
acceptable handling qualities.

.,

I WARNING I
If the aircraft is not controllable down
to a reasonable landing speed (given
consideration to weather, runway
condition, facilities, pilot experience,
pilot fatigue, etc.), an ejection isrecommended.

~-CABLE ARRESTMENT
Refer to figure 5-6 for hook engagement limits. If
there is any doubt about stopping on the remaining
runway, lower the hook. Engage the cable as close tocenter as possible, nosewheel on the runway withbrakes off and aircraft aligned with the runway. Placethe HOOK switch to DN at least 1500 feet before
reaching the desired arresting cable and reduce speed
as much as possible; however, ifbrakes and NWS are
inoperative, use flaperons and rudder as required tomaintain directional control. As the aircraft slows to
below 70 knots, directional control is reduced and theaircraft drifts right.
For most approach-end arrestments, touchdown
should be at least 500 feet in front of the cable to allow
sufficient time to lower the nosewheel to the runway
prior to engagement. For an approach-end arrestment with one MLG up or damaged as described in
LANDING WITH LG UNSAFEIUP, this section,
maintain landing attitude after touchdown and prior
to engagement. Immediately after touchdown, retard
throttle to IDLE.
After engagement, rollback should be controlled by
the throttle. For cable disengagement, place HOOK
switch to UP and use approximately 10-15 percent
increase in rpm to allow a rollback disengagement.

• Up to five seconds (after activation)
are required to fully raise the BAK-14
cable.
1.

GW - Reduce (as required).

2.

HOOK switch - DN.
• Approach-end arrestment: Touch down atleast 500 feet in front of the cable.
• Departure-end arrestment: HOOK switch to
DN at least 1500 feet before reaching thecable.

3.

SHOULDER HARNESS knob - LOCKED.

4.

Consider options available
engagemen t occurs.

5.

Throttle - IDLE.

6.

NWS - Engage (if required).

7.

Engage cable as close to center as possible,
nosewheel on the runway (if required) andbrakes off.

I WARNING I
U sing forward stick pressure to keep
an abnormally fast aircraft on the
runway for cable engagement willprobably result in a missed engagement or failure of the nose tire/NLG.

• Cable arrestment at speeds greaterthan emergency arrestment speed,with offcenter distances greater than
35 feet, or with the nosewheel in the air
could result in structural failure of the
NLG, hook, and/or hook backup
structure.
• The hook may miss the cable if theaircraft is not slow enough to compress
the MLG struts sufficiently to makeWOW or if forward stick pressure is
held.

NOTE
• Under certain conditions, arrestment
may produce a bouncing motion whichis readily apparen t.
• Offcenter engagement results in aircraft yaw motions during cable runout.

missedPrior to cable engagement:

I WARNING I• To prevent hook bounce and possible
missed engagement, avoid runway
centerline lighting.

if a

Do not use brakes while the cable isstretched or while being pulled backward. This action can result in aircraft
tipping backward. Control rollbackwith the throttle.

NET ARRESTMENT
Refer to NET ARRESTMENT LIMITATIONS,
section 5. Engaging a net barrier requires minimalpilot action as there is no hook to lower and little rollback after the aircraft's forward motion stops.
1.

SHOULDER HARNESS knob - LOCKED.

2.

Brakes - Release prior to engagemen t.

II II I II I I I II I~~~_
3.

Throttle - Off prior to engagement.

4.

Engage net perpendicular, preferably in thecenter portion of the runway.

I WARNING I
The canopy should be retained
throughout the engagement to provide
pilot protection. Barrier netting will
not prevent subsequent canopy opening/jettison.

Engage net perpendicular to preclude
aircraft rotating sideways during thearrestmen t. Avoid steering back toward the center of the runway justprior to engagement as this could
result in a non-perpendicular engagement. Nosewheel steering is not
required; however, if engaged, it may
be left engaged. The throttle should be
retarded to off prior to engagement to
reduce the possibility of foreign object
damage.

LANDING WITH A BLOWN TIRE

When landing with a blown MLG tire, the landing
gear may collapse during landing roll if portions of
the tire remain and cause a wheel imbalance
condition. To avoid possible directional control
problems associated with landing gear collapse, anapproach end arrestment is preferred over a normal
approach and landing. To reduce the possibility ofdamage, the lowest practical landing GW andairspeed should be attained. Jettison stores if
possible. Retain empty external fuel tanks. Stores/
suspension equipment at stations 3 and/or 7 may
cause external fuel tanks at stations 4 and 6 to moveinboard when jettisoned. If an approach-end
arrestment is not available and a normal approachand landing is flown, leave the anti-skid system on
to minimize the possibility of skidding on the goodtire. If the wheel with the blown tire does not turnafter landing, the antiskid system switches to the mJ
alternate braking mode, LESS (FE pulsating antiskid
mode. Use roll control to relieve pressure on the blown
tire and NWS to maintain directional control.

C~:~!::~_

With a blown tire, avoid centerline lights
as they may cause wheel damage andsubsequent loss of directional control.

~,

Stop straight ahead and shut down the engine as soon
as firefighting equipment is available. Do not attempt
to taxi unless an emergency situation exists.

,

If the blown tire is on the NLG, directional control
may be a problem due to a reverse castering effect. An
approach-end arrestment with the nosewheel off the
runway is recommended. Depending on GW andspeed at the time of engagement, type of cable
engaged, and height of nosewheel above the runway,
it is possible for NLG strut failure and/or inlet
structural damage to occur. To reduce the possibilityof such damage, the lowest practical landing GW andairspeed should be attained. Jettison stores if
possible. Retain empty external fuel tanks. Stores/
suspension equipment at stations 3 and/or 7 may
cause external fuel tanks at stations 4 and 6 to moveinboard when jettisoned. Landing should be madewith the remaining internal fuel in the aft system.
Mter touchdown from a 13 degree AOA approach,
pitch attitude should be reduced to approximately 5
degrees prior to cable engagement. The HUD gun
borecross can be used to determine pitch attitude.
During the arrestment, the aircraft is likely to turnslightly right before stopping.
If an approach-end arrestment is not available, referto procedures for aborting with a blown nose tire in
BLOWN TIRE ON TAKEOFF, this section, or
consider an all LG up landing (refer to LANDING
WITH LG UNSAFEIUP, this section).
Landing With A Blown Main Gear Tire

Prior to landing:

~

-1_
_

1.

Stores - Jettison. Refer to JETTISON, thissection.
Retain empty external fuel tanks.

2.

GW - Reduce (if practical).

3.

TANK INERTING switch - TANK INERTING even if Halon is not available.

,

AIR REFUEL switch - OPEN, ifexternal fueltank(s) is installed.

,

4.

I WARNING I
Failure to depressurize external fueltank(s) significantly increases the probability of tank explosion and fire if theaircraft departs the runway.

IIIIIIIIIIIIII

IIIIIIIIIIIIII-

IIIIIIIIIIIIII~

NOTE

4.

TANK INERTING switch - TANK INERTING
even if Halon is not available.

Delay placing the AIR REFUEL switch
to OPEN un til all external tanks are
empty.

5.

AIR REFUEL switch - OPEN, if external fueltank(s) is installed.

5.

ANTI-SKID switch - ANTI-SKID.
Use of antiskid minimizes skidding on goodtire during braking.

IWARNING

6.

HOOK switch - DN.
An approach-end arrestment is recommended. Refer to CABLE ARRESTMENT,
this section.

Fail ure to depressurize external fueltank(s) significantly increases theprobability of tank explosion and fire if
the nose gear collapses during thearrestmen t.

7.

Final approach AOA - 13 degrees.

6.

HOOK switch - DN.
An approach-end cable arrestment with thenosewheel off the runway is recommended.
Refer to CABLE ARRESTMENT, this section.

7.

Final approach AOA - 13 degrees.

If a missed approach-end cable arrestment occurs or
no approach-end cable is available:
NOTE

If no approach-end cable is available,
land on the side of runway away fromthe blown tire.

8.
9.

10.

~ ~ DRAG CHUTE switch - DEPLOY (if8.

Stick - Lower nose to approximately 5 degrees
pitch attitude for arrestment.

After cable engagement:

NWS - Engage (if required).
The NWS light does not illuminate when
NWS is engaged if the AIR REFUEL switch
is in OPEN.

9.

Landing With A Blown Nose Gear Tire

Prior to landing:
1.

Stores - Jettison. Refer to JETTISON, thissection.
Retain empty external fuel tanks.

2.

GW - Reduce (if practical).
Plan to land with approximately 1500 poundsof fuel on board.

3.

After touchdown:

required).

Brake - As desired on good tire.

Fuel distribution - All fuel in aft tank system(if practical).
At 3000 pounds fuel remaining, place ENG
FEED knob to FWD. When forward reservoir
is empty, place ENG FEED knob to NORM.
(Emptying forward tank system takes
approximately ~ 15 minutes, rnJ 9 minutes if
fuel flow is 4000 pph. When forward tanksystem empties, the fuel in aft tank system isapproximately ~ 2000 pounds, rnJ 2400
pounds.)1

Stick - Apply aft stick after nose starts down
to reduce load on the NLG.

If a missed cable engagement occurs:
10. Maintain pitch attitude and go around.

IWARNING

1With a blown NLG tire and loss of NWS,
it may not be possible to prevent
departure from the run way. A reverse
castering effect may occur in which thenosewheel moves opposite to the rudder
or differential braking input.
NOTE

The maxim urn allowable fuel flow with
one reservoir empty is 25,000 pph.
LG EXTENSION MALFUNCTIONS

Malfunctions in extending the LG are normallyindicated by failure of the LG handle to lower or by
failure of one or more LG to extend accompanied bylack of the corresponding WHEELS down lights andcontinuous illumination of the LG handle warning
light. Refer to LG FAILS TO EXTEND, this section.

""""",,~~~I

Mechanical failures can have varying causes. Failureof an LG component usually affects only one LG andcannot be corrected in flight. Abnormally highmoisture content in the hydraulic fluid can cause
more than one LG to fail to extend after prolonged
operation at low ambient temperatures. Thissituation can usually be corrected after several cyclesof the LG handle at low altitude.

~

Abnormal indications after the LG handle is lowered
can also be caused by electrically related malfunctions (i.e., electrical shorts, electrical component
failures, or cannon plug problems). The malfunction
may result in an indication problem or an actual
failure of one or more LG to extend. A visualconfirmation of the LG position should be obtained if
possible before any action is taken. LESS ~ Anelectrically related failure can make the brakes, hook,
NWS, and ~ ~ drag chute inoperative.

LESS ~ The ANTI-SKID caution light may give an
indication of probable braking, hook, NWS, and ~~ drag chute operation since it uses the same powersource and circuitry. If the ANTI-SKID caution lightilluminates with the ANTI-SKID switch in OFF,power should be available for braking. (This checkcan be accomplished in either CHAN 1 or 2 with the
LG handle in DN or in CHAN 2 only with the LG
handle in UP.) If the ANTI-SKID caution light does
not illuminate and hook does not extend, brakes,
NWS, and ~ ~ drag chute may not be available.

A common failure mode is electrical shorting withina downlock switch caused by moisture intrusion. This
shorting typically occurs after the LG handle islowered and results in opening of the LG uplockJ
downlock circuit breaker. Opening of this circuit
breaker causes all WHEELS down lights to remain
off. The light in the handle may go off normally or it
may remain on if the shorting has also adversely
affected the warning light circuit. Another result of
this circuit breaker opening is that the LG will go into
hydraulic isolation immediately after the LG handle
is raised and the LG won't retract. Other effects
associated with an open LG uplockJdownlock circuit
breaker are as follows:• No AOA bracket in the HUD.
• Speedbrakes are not limited to 43 degrees.
• Landing/taxi light is inoperative.
• NWS is inoperative.
An indication failure may be distinguished from an
actual extension malfunction by the WHEELS down
lights and the LG handle warning light. The
WHEELS down lights and the LG handle warning
light have separate power sources and circuitry.
Thus, if either circuit operates correctly, the LG• should be down even though the other circuit may
indicate a problem. If possible, cycle the LG toreattempt extension while watching the WHEELS
down lights and the LG handle warning light.

The ANTI SKID caution light may give an
indication of probable braking. If the ANTI SKID
caution light illuminates with the ANTI-SKID switch
in OFF, power should be available for braking. (Thischeck can be accomplished in either CHAN 1 or 2 with
the LG handle in DN or in CHAN 2 only with the LG
handle in UP.) If the ANTI SKID caution light does
not illuminate, brakes may not be available.

Because of the number of possible malfunctions,specific procedures for every situation are notfeasible. If time and conditions permit, groundsupervisory and technical assistance should berequested. Evaluate available options prior to using
the ALT GEAR handle. Other options may include an
LG up landing, diverting to a more suitable landing
field, landing where the hazards of departing the
prepared surface are minimal, or using a cablearrestment.
LG Handle Will Not Lower

If the LG handle cannot be moved to the DN position
after depressing the LG handle down permission
button, the electrical circuitry or solenoid has
probably failed. [A] mEl The DN LOCK REL button
mechanically permits the LG handle to be moved to
the DN position. If the LG handle cannot be moved to
the DN position due to a mechanical failure, CHAN 1
brakes are inoperative, TEF's must be extended using
the ALT FLAPS switch, and the nozzle idle area reset
is not available. Without nozzle idle area reset, idle
thrust is higher than normal.
If LG handle cannot be lowered normally:

1.

DN LOCK REL button - Depress and lower
LGhandle.

IfLG handle still cannot be lowered:2.

ALT FLAPS switch - EXTEND.

3.

BRAKES channel switch - CHAN 2.

~
~
~
~
~

_

~

,

-

4.

Go to ALTERNATE LG EXTENSION, thissection.
Nozzle remains closed resulting in higher
than normal landing thrust.
NOTE

Mter a successful alternate gear
extension with the landing gear handle
still up, the LG handle warning lightremains on to indicate the position of
the gear handle is not in agreementwith the actual gear position.
LG Fails To Extend
NOTE

If alternate LG extension was performed and one or more LG indicate
unsafe, refer to ALTERNATE LG
EXTENSION, this section.

If one or more LG fail to extend, the LG handle may
be cycled multiple times to attempt to extend the LG
if no structuraI/battle damage exists and if normal
hydraulic pressure exists. Multiple cycle attempts
may release the locking pawls and result in normal
gear extension if the failure to extend was caused by
failure of the uplocks to release.

If the LG previously failed to retract,
do not cycle the LG handle. Damage tothe LG or LG doors may preclude
successful extension.

If cycling of the LG handle fails to correct thesituation, use of the alternate extension system must
be considered.
If possible, get a visual confirmation of LG position.
If the NLG WHEELS down light is off, confirmationof the NLG position can be made by checking
landing/taxi light operation. Illumination of either
light confirms that the NLG is down. With the NLG
WHEELS down light off, NWS may be inoperative(without a NWS FAIL caution light).

If one or more LG indicate unsafe:

I WARNING I• If no WHEELS down lights are
illuminated or no LG extends, verify
ANTI SKID caution light operation bycycling the ANTI-SKID switch. (Thischeck can only be accomplished with the
LG handle in DN or in CHAN 2 with the
LG handle in UP.) ~ If the ANTI SKID
caution light does not illuminate,
brakes may not be available. LESS ~ If
the ANTI-SKID caution light does notilluminate and hook does not extend,
brakes, NWS, and ~ !till drag chute
may not be available. Evaluate available options (e.g., LG up landing,diverting to a more suitable landing
field, or using a cable arrestment) prior
to using the ALT GEAR handle.
• If at anytime, an LG intermittently
indicates unsafe (i.e. WHEELS down
light off and LG handle warning lighton), the overcenter lock on the LG dragbrace assembly may not be functioning
properly. The LG may appear down,
but the LG may collapse during
landing. Plan on using the LG unsafe/
up procedures even if the LG eventually indicates normal. Refer to LANDING WITH LG UNSAFEIUP, thissection.

•If the LG previously failed to retract,
do not cycle the LG handle. Damage tothe LG or LG doors may preclude
successful extension.
NOTE

If the NLG WHEELS down light is off,
confirmation of the NLG position can bemade by checking landing/taxi light
operation. Illumination of either light
confirms that the NLG is down. With the
NLG WHEELS down light off, NWS
may be inoperative (without a NWS
FAIL caution light).

••

""""",,~~~I
1.

LG handle - Cycle and monitor LG handle
warning light and WHEELS down lights.

2.

If LG handle warning light illuminated when the LG
handle was lowered, then went off and tests good, or
if WHEELS down lights operated normally:

2.

4.

Go to ALTERNATE LG EXTENSION, thissection.

1.

I

WARNING

If LG handle does not lower, selectBRAKES CHAN 2 and position ALT
FLAPS switch to EXTEND. Nozzle
remains closed, resulting in higher
than normal landing thrust.

"

4.

Stop straight ahead on the runway.
Consider go-around capability in the eventthe brakes are found to be inoperative after
touchdown.

If the LG was alternately extended due
to failure of system B, only brake/JFS
accumulator braking is available andafter stopping the parking brakeshould be engaged un til chocks are
installed.
If LG indicates unsafe:

3.

_

I~

Land normally.
If possible, get visual confirmation of LG
position. ~ If all WHEELS down lights were
initially off with the LG handle down and useof the hook may be required after touchdown,
verify before landing that the hook extends.

I_

G!::~!~J3.

LG handle - DN. (Use DN LOCK REL, if
required.)~• NWS is not available following alternate LG extension.

If LG indicates safe:

Alternate LG Extension

Alternate LG extension should be accomplished at
the lowest practical airspeed below 190 knots. The• NLG may not indicate down until airspeed is reduced
below 190 knots. NWS is inoperative after alternate
LG extension even if system B hydraulic pressure isavailable.

"• Pulling the ALT GEAR handle withnormal system B hydraulic pressure,e.g., NLG fails to extend, may result insystem B hydraulic failure within 15
minutes.

Land normally.

If LG handle warning light did not illuminate or
remained illuminated after LG handle was lowered,and if one or more WHEELS down lights did notilluminate:• If an unsafe MLG indication exists and both
MLG are out of the wheel wells, pulling the
ALT GEAR handle is not recommended.

• Do not depress the ALT GEAR reset
button while pulling the ALT GEAR
handle. This action may preclude
successful LG extension.

If RMLG WHEELS down light is off,
speedbrakes may not be limited to 43
degrees.

3.

"• Alternate LG extension can be used up to 300
knots; however, the NLG may not fully extend
until 190 knots. Time above 190 knots should
be minimized in case there is a leak in thepneumatic lines.

Speedbrakes - Verify opening is less than 43
degrees.
From the front cockpit, the top of the
speedbrakes should be slightly above a line
drawn from the tip of the horizontal tail to the
top of the vertical tail root fairing.

~

ALT GEAR handle - Pull (if required) (190knots maximum, if practical).

Stick - Apply alternating g forces (-1.0 to +3.0g) to free LG.
Up to 300 knots may be required to provide
sufficient g force.

--

If LG indicates safe:
4.

Land normally.
If possible, get visual confirmation of LG
position.

5.

Stop straight ahead on the runway.

To accomplish the landing:1.

Retain empty fuel tanks and racks.

I

WARNINGI

If time permits, delay landing until
external fuel tanks are empty. If an
immediate landing is required, jettison all external fuel tanks.
If the LG was alternately extended due
to failure of system B, only brake/JFS
accumulator braking is available andafter stopping the parking brakeshould be engaged until chocks are
installed.
If LG still indicates unsafe:
4.

2.

Armament - Jettison.

3.

GW - Reduce.

4.

TANK INERTING switch - TANK INERTING even if Halon is not available.

5.

AIR REFUEL switch - OPEN.

Speedbrakes - Verify opening is less than 43
degrees.
From the front cockpit, the top of the
speedbrakes should be slightly above a line
drawn from the tip of the horizontal tail to the
top of the vertical tail root fairing.

I

WARNING

I

Failure to depressurize external fueltanks significantly increases the probability of tank explosion and fire.
NOTE

Delay placing the AIR REFUEL switch
to OPEN until all external fuel tanks
are empty.

If RMLG WHEELS down light is off,
speedbrakes may not be limited to 43
degrees.

5.

Go to LANDING WITH LG UNSAFEIUP, thissection.

Landing With LG Unsafe/Up

Prior to landing with any of the LG unsafe or up,
consider the following:• Airfield facilities.
• Hook engagement limits.
• Crosswind component.

6.

Radar - STBY or 0 FF.

7.

SMS/HDPT/ECM power - Off.

8.

SHOULDER HARNESS knob - LOCKED.

9.

Refer to figure 3-16.
NOTE

If either MLG is not extended, EPU
operation cannot be terminated with
the EPU switch after engine shutdown.
ANTISKID MALFUNCTION (LANDING)

If conditions are not favorable:

Illumination of the ANTI SKID caution light whenthe LG handle is lowered indicates one of thefollowing:1.

• (f£) Loss of power to one of the two brake channels• Runway and overrun conditions.

Refer to EJECTION (TIME PERMITTING),
this section.

or from the ANTI-SKID switch.

"""IIIIII~~~I(FE Built in Test (BIT) detected malfunction of oneof the two brake channels.

- LESS irE Failure in the antiskid system (pulsatingbrake pressure mode activated).

I - LESS irE Loss of electrical power to the CHAN 1
brake and antiskid control circuits.
(fE Cycling the ANTI-SKID switch to OFF and backto ANTI -SKID will not extinguish the ANTI SKID
caution light. Touchdown skid control protection
may not be available, so do not apply brakes before
touchdown.

mJ If the ANTI SKID light extinguishes on wheel
spin up, antis kid protection is available and one of
the two brake control channels is fully functional.
Normal braking may be used. If the light does not
extinguish on wheel spinup, or first illuminates on
wheel spinup, braking may be in the alternate
braking mode, detectable by a pUlsing sensation
when differential braking is applied. In the alternate
braking mode, symmetric braking will give the best
stopping performance.
LESS mJ Changing to CHAN 2 brakes may
extinguish the ANTI SKID caution light and provide
normal antiskid and braking functions. If the failure
was in the antiskid system, cycling the ANTI-SKID
switch may correct the problem. If the ANTI SKID
caution light remains on, antiskid protection is notavailable and degraded (i.e. pulsating) or inoperative braking should be expected.

If the ANTI SKID caution light illuminates (with the
ANTI-SKID switch in ANTI-SKID) when the LG
handle is lowered:
1.

BRAKES channel switch - CHAN 2.

irE If the ANTI SKID caution light remains on:
2.

Refer to ANTISKID(GROUND), this section.

MALFUNCTIONLESS m'l If the ANTI SKID caution light remains on:

2.

ANTI-SKID switch - OFF, then
ANTI-SKID.

back toIf the ANTI SKID caution light remains on, expect
pulsating or inoperative braking and no antiskid
protection during landing:3.

Refer to ANTISKID(GROUND), this section.

MALFUNCTION

ASYMMETRIC STORES (LANDING)

Refer to figure 3-17 for computation of asymmetricmoment. A modified approach is required if the net
asymmetry exceeds 10,000 foot-pounds. Successful
landings have been demonstrated with asymmetries
as large as W 25,020 foot-pounds, LESS W 22,650
foot-pounds. If the net asymmetry exceeds W 25,020
foot-pounds, LESS
22,650 foot-pounds, storesshould be selectively jettisoned from the heavy wing
to reduce the asymmetry. The decision to land withlarge asymmetries should consider such factors as
weather conditions, runway length/width, surface
conditions (RCR), arresting gear availability, crosswind component/gusts, and pilot experience. Avoid
abrupt control inputs. Limit maximum bank angle
changes to 90 degrees and do not exceed 10 degrees
AOA until the net asymmetry can be determined. A
controllability check should be performed with the
LG down to determine handling qualities and thefeasibility of landing. During the controllability
check, determine maximum maneuvering AOA by
slowly increasing AOA until roll authority is
insufficient to maintain wings level. Do not exceed 12
degrees AOA during this check. If loss of rollauthority is experienced, the maximum maneuvering
AOA is 2 degrees below that at which roll authority
was lost or 10 degrees AOA, whichever is less. If
landing is feasible, plan to fly a shallow, power-on,
straight-in approach. Use roll trim and lateral stick
as required. Full roll trim may not be enough tocompensate for large asymmetries. Trim rudder into
the heavy wing up to a maximum of two dots. Thisdecreases bank angle and permits a low bank angle,
low sideslip approach, minimizing pilot induced
lateral oscillations. This action increases the roll trim
necessary to hold up the heavy wing.

m

Fly a shallow, power-on, straight-in approach andreduce bank angle to wings level on short final. Do not
exceed the maximum maneuvering AOA, as determined during the controllability check, on final
approach or during the flare/touchdown. Two-point
aerodynamic braking should be performed at 10-11
degrees AOA until approximately 120 knots. The
nose should then be lowered to the runway andmaximum effort antiskid braking should be used. At
no time should 11 degrees AOA be exceeded while on
the runway. When the ARI switches out with wheel
spin-up, large rudder changes can occur which cause
yaw into the heavy wing. When landing with nocrosswind, this yaw helps align the aircraft with the
runway. With crosswind components greater than 10
knots (5 knots if the net asymmetry exceeds 20,000
foot-pounds), land with the heavy wing into thecrosswind even if this results in landing downwind.
Failure to do so may result in inadequate roll control.

-

-

-

,,
,
,
,

~~~IIIIIIIIIIII- *- ~A=====--==~~====~~- t- --It-LG Unsafe/Up Landing
APPROACH-END ARRESTMENT

ALL LG INDICATE UNSAFE
BUT APPEAR NORMALUNAVAILABLEAVAILABLECONFIGURATION

10.
11.

HOOK - DOWN.
APPROACH-END CABLE - ENGAGE.

10.

LAND NORMALLY.

ARRESTMENT NOT RECOMMENDED.
USE APPROACH-END ARRESTMENTUNAVAILABLE PROCEDURE.

10.
11.
12.
13.

EPU - ON.
ALT FLAPS - EXTEND.
LOW ANGLE APPROACH AT 13° AOA.
THROTTLE - OFF IMMEDIATELYPRIOR TO TOUCHDOWN.

ALT GEAR HANDLE - IN.
WAIT 5 SECONDS.
LG HANDLE - UP.
ALT GEAR RESET BUTTON - DEPRESS(2 SECONDS).
USE ALL LG UP PROCEDURE.
IF NLG DOES NOT RETRACT:o. HOOK - DOWN.
b. LOW ANGLE APPROACH AT 11 0
AOA.
c. ATTEMPT A FLY-IN ENGAGEMENT.
AFTER
ENd. THROTTLE - OFF
GAGEMENT.

10.
11.
12.
13.

ALT GEAR HANDLE - IN.
WAIT 5 SECONDS.
LG HANDLE - UP.
ALT GEAR RESET BUTTON - DEPRESS(2 SECONDS).
USE ALL LG UP PROCEDURE.
IF NLG DOES NOT RETRACT:
o. CONSIDER LANDING FROM A
LOW ANGLE APPROACH AT 13°
AOA IF WING FUEL TANKS ARECARRIED.
b. RECOMMEND EJECTION IF WING
FUEL TANKS ARE NOT CARRIEDOR IF CONDITIONS ARE NOTCONSIDERED FAVORABLE FOR
AN ATTEMPTED LANDING WITH
WING FUEL TANKS.

,I

i

l

ALL LG UP

10.
11.
12.
;;;~===~~ 13.
._~
14.
15.

80TH MLG UP OR UNSAFE

IPJ1~==::::::;;;;~---=-NLG UP OR UNSAFE

I WARNING I
IF THEENGAGEMENT ISMISSED, MAINTAIN WINGS LEVEL AND GO AROUND. IF A
GO-AROUND IS NOT ACCOMPLISHED, THE AIRCRAFT
MAY GROUND LOOP.

ARRESTMENT NOT RECOMMENDED.
USE APPROACH-END ARRESTMENTUNAVAILABLE PROCEDURE.

10.
11.
12.
13.

-

14.
15.

14.

EPU - ON.
LOW ANGLE APPROACH AT 13 0 AOA.
THROTTLE - OFF AFTER TOUCHDOWN.
LOWER NOSE TO RUNWAY BEFORECONTROL EFFECTIVENESS BEGINSTO DECAY.
EPU - OFF AFTER STOP.

Figure 3-16. (Sheet 1)

3-142

Change 7~IIIIIIIIIIIIIIIIIIIII~~~­LG Unsafe/Up Landing
APPROACH·END ARRESTMENTCONFIGURATIONAVAILABLE

ONE MLG AND NLG UP OR UNSAFE

ARRESTMENT NOT RECOMMENDED.
USE APPROACH-END ARRESTMENTUNAVAILABLE PROCEDURE.

~

----=

UNAVAILABLE'ONE MLG INDICATES UNSAFE
BUT APPEARS NORMAL

~-

14.
15.

10. HOOK - DOWN.
11 . LOW ANGLE APPROACH AT 11 0 AOA.
12. AFTER TOUCHDOWN, USE ROLL

13.

CONTROL, IF NECESSARY, TO HOLD
WING UP. IF ROLL CONTROL IS NEEDED TO HOLD WING Up, MAINTAINLANDING ATIITUDE FOR ENGAGEMENT. IF ROLL CONTROL IS NOTNEEDED TO HOLD WING Up, LOWERNOSE FOR ARRESTMENT.
THROTILE - OFF AFTER ENGAGEMENT.

IWARNING I
IF THE ENGAGEMENT IS MISSEDAND ROU CONTROL WAS NECESSARY TO HOLD WING Up,
MAINTAIN WINGS LEVEL AND
GO AROUND. IF A GO-AROUNDIS NOT ACCOMPLISHED, THE
AIRCRAFT MAY GROUND LOOP.

ONE MLG UP~

1l~:::F===;;;;~=====Qno-=~

10. ALT GEAR HANDLE - IN.
11. WAIT 5 SECONDS.
12. LG HANDLE - UP.
13. ALT GEAR RESET BUTION - DEPRESS

10. ALT GEAR HANDLE - IN.
11. WAIT 5 SECONDS.
12. LG HANDLE - UP.
13. ALT GEAR RESET BUTION - DEPRESS
14.
15.

(2 SECONDS).
USE ALL LG UP PROCEDURE.
IF LG DOES NOT RETRACT:o. HOOK - DOWN.
b. LOW ANGLE APPROACH AT 11 0
AOA.
c. AFTER TOUCHDOWN, USE ROLLCONTROL TO HOLD WING UP
AND MAINTAIN LANDING ATIITUDE FOR ENGAGEMENT.
d. THROTILE - OFF
AFTER
ENGAGEMENT.

IWARNINGI

IF THE ENGAGEMENT IS MISSED,MAINTAIN WINGS LEVEL AND
GO AROUND. IF A GO-AROUNDIS NOT ACCOMPLISHED, THE
AIRCRAFT MAY GROUND LOOP.

Figure 3-16. (Sheet 2)(2 SECONDS).
USE ALL LG UP PROCEDURE.
IF LG DOES NOT RETRACT:
o. CONSIDER LANDING FROM A
LOW ANGLE APPROACH AT 13 0
AOA IF EXTERNAL FUEL TANK(S) ISCARRIED.

NOTELAND ON THE SIDE OF THE
RUNWAY AWAY FROM THE UNSAFE MLG.

b. RECOMMENDEJECTION
IF
EXTERNAL FUEL TANK(S) IS NOTCARRIED OR IF CONDITIONS ARENOT CONSIDERED FAVORABLE
FOR AN ATIEMPTED LANDING
WITH EXTERNAL FUEL TANK(S).

-IIIIIIIIIIIIII~

Asymmetric Moment

LEGEND:

~

mm
©

1
2
@
3
WEIGHT - 1000 POUNDS

LESS

m

4
IF-l6A-l-Ol27A@

Figure 3-17.

1.

2.

3.

Compare weights on mirror stations (4 and 6,
3 and 7, etc.). Include wingtip AIM-120's;
exclude wingtip AIM-9's since their weight isoffset by the lift they generate.
Determine asymmetric moment for each set of
stations. Enter with weight difference atbottom of chart, proceed vertically to theappropriate line, and proceed horizontally left
to read the asymmetric momen t.
Add or subtract each asymmetric moment todetermine net asymmetry.

SAMPLE PROBLEM 1:•

Full 370-gallon fuel tank on station 4; empty
fuel tank on station 6.

•

Three MK 82 (SNAKEYE) bombs on station 3,
station 7 empty.

A.

Station 4/6 weight
difference= 2405 pounds

14,000 foot-pounds

B.

Asymmetric moment

C.

Station 317 weight
difference= 1650 pounds

D.

Asymmetric momen t= 16,500 foot-pounds

E.

Net asymmetry (B+D)

30,500 foot-pounds

SAMPLE PROBLEM 2:•

Full 370-gallon fuel tank on station 6; empty
fuel tank on station 4.

•

Three MK 82 (SNAKEYE) bombs on station 3;
station 7 empty.

Follow same procedures (A through D) as above;
however, since the asymmetric moments are on the
opposite wing, the moments are subtracted rather
than added.

= 2500 foot-pounds

E.

Net asymmetry (B-D)

1.

AOA - 10 degrees maximum.

,.. ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ,6T ,6T ~1~ ,I WARNING I

I WARNING I

Large asymmetric loads severely limit
lateral control when rolling away fromthe heavy wing. Until determining net
asymmetry, limit maximum bankangle change to 90 degrees, avoid
abrupt control inputs, and do not
exceed 10 degrees AOA.
2.

• With crosswind component greaterthan 10 knots (5 knots if the net
asymmetry exceeds 20,000 footpounds), land with heavy wing into thecrosswind even if this results inlanding downwind. Failure to do so
may result in inadequate roll control.

Determine net asymmetry.
Refer to figure 3-17.

• Do not exceed the maximum AOA, as
determined during the controllability
check, during final approach, flare,
touchdown, or two-point aerodynamic
braking.

If asymmetry is greater than OJ 25,020 foot-pounds,
LESS m 22,650 foot-pounds:
3.

Stores - Jettison (as required).
Selectively jettison stores from the heavy
wing to obtain a net asymmetry less than
25,020 foot-pounds, LESS
22,650 footpounds. Refer to SELECTIVE JETTISON,
this section.

mm

If asymmetry is greater than 10,000 foot-pounds:
4.

Controllability - Check.
• Lower LG at a safe altitude and check
handling qualities until roll authority is
insufficient or up to 12 degrees AOA
maximum.
• Maximum maneuvering AOA for approachand landing is 10 degrees AOA or 2 degrees
less than the AOA at which roll authority is
insufficient to maintain wings level, whichever is less.

Roll trim and lateral stick - As required.

7.

Rudder trim - Trim into the heavy wing (ifrequired).

If landing is not feasible:
5.

Go to EJECTION (TIME PERMITTING), thissection.

If asymmetry is less than 10,000 foot-pounds:
4.

Land normally.

NLG WOW SWITCH FAILURE

If the NLG WOW switch fails to the ground position,
the AIR DISC function on the stick is inoperative, thespeedbrakes are not limited to 43 degrees with theright MLG down and locked, and NWS can be
engaged in flight.

If landing is feasible:

I WARNING I

5.

6.

The decision to land with a large
asymmetry should consider such factors as weather conditions, runway
length/width and surface conditions(RCR), arresting gear availability,
crosswind component/gusts, and pilot
experience.

If ARlNWS light comes on:

2.

NWS - Disengage.

Fly a shallow, power-on, straight-in approach.

3.

ARlNWS light - Off.

1.

NWS - Engage.

-

-

~
,

_

-

~",,:;,~

:1
-

-

NOTE
Visually confirm speedbrake opening
is limited to 43 degrees to prevent thelower surfaces from striking the
runway during landing.

Insure thatARJNWS light is off prior tolanding so that the NWS does notfollow rudder commands when thenosewheel is lowered to the runway.

DRAG CHUTE FAILURE ~ !BQl
4.

Speedbrakes - Close to less than 43 degrees.
From the front cockpit, the top of the
speedbrakes should be slightly above a line
drawn from the tip of the horizontal tail to the
top of the vertical tail root fairing.

If decision is made to go-around:1.

Drag chute - Release.

2.

Throttle - MAX AB.

T. O. 1F- 16A- 1

SECTION IV
CREW DUTIES(Not Applicable)

4- 1/(4-2 blank)

T.O. 1F- 16A-1

SECTION V
OPERATING LIMITATIONSTABLE OF CONTENTS
Introduction ............................ .
Instrument Markings .................... .
Engine Limitations I PW200 I ............... .
Engine Limitations I PW2201 ............... .
Fuels ................................... .
System Restrictions ...................... .
Jet Fuel Starter Limits ................ .
Tire Spee.d Limit ...................... .
Brake Energy Limits .................. .
Fuel/Oil System Limitations ........... .
One Reservoir Empty ............... .
Negative G Flight ................... .
Airspeed Limitations .................... .
System Airspeed Limitations ........... .
Maximum Airspeed Operating
Limitations ........................ .
Low Airspeed Operating Limitations .... .
Prohibited Maneuvers ................... .
Gross Weight Limitations ................ .
CG Limitations .......................... .
Acceleration Limitations ................. .
Fully Internal Fueled Aircraft .......... .
Partially Internal Fueled Aircraft ...... .
AOA and Rolling Limitations ............. .
Stores Limitations ....................... .
Asymmetric Stores Loading ............ .
Miscellaneous Limitations ................ .
Net Arrestment Limitations ............ .
Crosswind Limits ......................
Aircraft Battle Damage Repair ..........

Page
5-1
5-1
5-1
5-1
5-1
5-2
5-2
5-2
5-2
5-2
5-2
5-2
5-2
5-2
5-2
5-3
5-3
5-3
5-3
5-3
5-4
5-4
5-4
5-4
5-4.1
5-4.1
5-4.1
5-4.1
5-4.1

ENGINE LIMITATIONS 1PW200 I
Refer to figure 5-2 for Engine Limitations, figure 5-3
for AB Operation and Light-Off Limits, and figure 5-4
for BUC - Operational Envelope.

ENGINE LIMITATIONS I PW2201
Refer to figure 5-2 for Engine Limitations, and figure
5-3 for Operational Envelope ..

FUELS
JP-4, JP-5, JP-8; NATO F -34, F -35, F -40, F -43, andF -44; and commercial JET A, JET A-I, and JET Bare
approved fuels. Except for freeze point and possible
icing and corrosion inhibitor differences, JP-4, NATO
F-40, and JET B are equivalent and the same
operating limitations apply.
Operating and throttle movement limitations forfuels other than JP-4, NATO F-40, and JET B, are thesame as for JP-4, NATO F-40, and JET B
except: Ground starts with temperature below -4°F(-20°C) may produce more smoke and require alonger time for engine light-off. Ground starts should
not be attempted with fuel temperature below -40°F(-40°C) except JET A which is -35°F (-37°C).
Airstart light-off times also may be slightly longer.

NOTE
1

•

1

INTRODUCTION
The aircraft and system limitations that must be
observed during normal operations are presented inthis section and T.O.IF-16A-I-3 (Block 10) or T.O.
IF-16A-I-4 (Block 15 ~ ). Those limitations that arecharacteristic of a special phase of operations such as
emergency procedures, etc., are not covered in thissection.

PW200 I Changing operations from
JP-4, NATO F-40, or JET B may
initially result in an increase in thenumber of AB-related compressor
stalls due to minor trim variations .

•PW2201 Changing operations from
JP-4, NATO F-40, or JET B may resultin an increase in the number of AB
recycles.

.1 PW2201 When using approved fuels
other than JP-4, NATO F-40, or JET B,
a self-recovering AB sequencing stall
may occur when transitioning fromregion 3 of the AB lightoff envelope as
the AB automatically sequences up to
the requested throttle position.

INSTRUMENT MARKINGS

PW2201 Fuels with very high flash points (JP-5,
NATO F-43, and NATO F-44) may leave visible
signature on AB cancellation at high altitude.

1

Refer to figure 5-1 for location and range of themarkings.

Change 35-1

10. 1 F- 16A-1

Approved fuels may be intermixed in any proportion
during ground or AR operations. No change in engine
operating limitations is required. I PW200 I AB
hangups in segment 4 may occur. Segment 5 AB may
not be usable and there may be a noticeable decreasein AB thrust.
Due to fuel freeze points, non JP-4 or NATO F -40 fuelin external tanks may not transfer after sustained
operation (5 minutes or longer) below 275 knots from25,000-30,000 feet or below 0.72 mach from30,000-42,000 feet (all non JP-4 or NATO F -40 fuels
except JET A) or below 300 knots from 25,000-30,000
feet or below 0.83 mach from 30,000-42,000 feet (JET
A).

NATO F -34, NATO F -40, and NATO F -44 may notcontain corrosion inhibitor and NATO F -35, NATOF -43, JET A, JET A-I, and JET B may not containicing or corrosion inhibitors. Restrict operation
without icing inhibitor to one flight. Restrict engine
operation without corrosion inhibitor to 10 consecutive hours.

• Initiating maximum effort braking
above maximum brake application
speed may result in loss of braking
before the aircraft is stopped.
• Danger zone procedures should be
followed for any event which requires
excessive braking.
• If brake energy absorption is in thedanger zone, wheel fusible plugs
release tire pressure within 3-15
minutes after the stop.

FUEL/OIL LIMITATIONSOne Reservoir Empty

The maximum allowable fuel flow with one reservoir
empty is 25,000 pph.
Negative G Flight

Certain fuels are heavier; refer to FUEL QUANTITY
INDICATION AND TANK ARRANGEMENT, Section I.

Negative g flight with both reservoir tanks full islimited to:

SYSTEM RESTRICTIONS• I PW200 I MIL thrust or below - 10 seconds (oilsystem limit).

JET FUEL STARTER LIMITS• I PW220 I MIL thrust or below - 30 seconds.

Refer to figure 5-5.

TIRE SPEED LIMITThe MLG and NLG tires are certified for use to 217
knots groundspeed.

• AB thrust - 10 seconds.

NOTE
Negative g flight should be avoided
when a low fuel condition exists(forward or aft reservoir not full) or
ENG FEED knob out of NORM.

AIRSPEED LIMITATIONSBRAKE ENERGY LIMITSSYSTEM AIRSPEED LIMITATIONS
Refer to T.O. IF-16A-l-l, Part 2 for brake energy
limits for maximum effort braking, taxi, aborted
takeoff, landing, and the effect on turnaroundcapability. The actual energy per brake may differ
considerably from the value found in Part 2. This iscaused by unequal energy distribution between thebrakes or residual heat from previous braking.
Maximum brake application speed is the maximum
speed from which the aircraft can be stopped using
maximum braking. This speed is based on thecapability of each brake to absorb a maximum of 18.2
million foot-pounds of energy. (Refer to ABORTEDTAKEOFF MAXIMUM BRAKE APPLICATIONSPEED, T.O. IF-16A-l-l, Part 2.)
5-2

Change 8

Refer to figures 5-3, 5-6, and 5-7.

MAXIMUM AIRSPEED OPERATING LIMITATIONS~ Refer to figure 5-3. Maximum operating airspeed

is 800 KIAS from sea level to 30,000 feet MSL. Above
30,000 feet MSL, the aircraft is limited to 2.05 mach.

LESS ~ Maximum operating airspeed is 600 KCAS
or 2.05 mach, whichever is lower.
Maximum operating airspeed/mach may be reduced
as a result of system restrictions or stores limitations.
Refer to figure 5-7 and T.O. IF-16A-I-3 (Block 10) or
T.O. IF-16A-I-4 (Block 15 and~ ).

10.lF-16A-1
LOW AIRSPEED OPERATING LIMITATIONS

2.

Repeated maximum rudder reversals.

Recovery should be initiated no later than activationof the low speed warning tone.

3.

Consecutive 360-degree maximum commandrolls.

I WARNING I

4.

Maximum command roll reversals above 670
knots with STORES CONFIG switch in CAT I.

5.

Maximum command rolling maneuvers above
1.8 mach and either above 3g or below 35,000
feet MSL.

6.

Rudder rolls or rudder-assisted rolls of morethan 90 degrees of bank angle change with any
store on station 3,4,6, or 7.

7.

With LG and/or TEF's down:

For CAT I configurations with dragindices greater than 120, delaying
recovery until activation of the low
speed warning tone may result in
departure regardless of recovery technique. Rapid airspeed decay may
reduce control authority to the pointthat recovery inputs have no effect.
Low thrust settings, external fueltanks, or inlet pods increase thepossibility of a departure.

• Flight above 15 degrees AOA with stores atstation 3, 4, 6, or 7.

To avoid departures due to roll coupling, do not
operate with category III loadings below 200 knots(except for takeoff and landing).

I

WARNINGI• Maximum command rolls of more than 90
degrees of bank angle change.
8.

Rapid rudder release or reversal above 300
knots/O.6 mach.

GROSS WEIGHT LIMITATIONS
Departures from controlled flight withasymmetric category III loadings may
result in fast, flat (possibly nonrecoverable) spins.

PROHIBITED MANEUVERSThe following maneuvers are prohibited:
1.

Intentional departures and spins with any of
the following:• Symmetric category I loading with suspension
equipment or missiles at station 3, 4, 6, or 7.
• Asymmetric category I loading.
• Category III loading.
• Altitude below 30,000 feet AGL.
• CG aft of aft limit for the configuration being
flown.
• Lateral fuel (internal and external) imbalance
greater than 200 pounds.

Block 10 aircraft and ~ lim ~ IHQ] Block 15 aircraft.
The maximum allowable GW for ground handling,taxi, takeoff, in flight, and landing is 35,400 pounds.
~ Block 15 aircraft with 18 ply NLG tire and 20 ply

MLG tire. The maximum allowable GW for groundhandling, taxi, takeoff, in flight, and landing is 35,400
pounds.

CG LIMITATIONS
Refer to T.O. IF-16A-5-2.
W Generally, the aircraft is within CG limits when the
red portion of the AL pointer is not visible (FUEL QTY
SEL knob in NORM).
LESS rn Refer to Section II for FUEL MANAGEMENT and Section III if the proper fuel distribution
is not established and/or maintained.

ACCELERATION liMITATIONS
Refer to figure 5-8 this section, and T.O. IF-16A-I-3(Block 10) or T.O.1F-16A-1-4 (Block 15 and 18Q/), I
figures 5-10, 5-11, or 5-12 ~) for acceleration
limitations.

Change 11

5-3

T.O. 1 F-16A-1

Load factor limits should not be intentionally
exceeded. Notify maintenance of a possible over-g ifsymmetric maneuvering on the g limiter results in aload factor greater than or equal to 9.5g/-3.2g or ifsymmetric non-g limiter maneuvering or asymmetricmaneuvering exceeds a positive or negative g limit
specified in this section. Provide details of theoccurrence (maximum g indication, airspeed, altitude, description of maneuver, fuel weight anddistribution, etc.) and HUD videotape if it isavailable.

AOA AND ROLLING LIMITATIONSNOTE• Category III loadings are not protectedfrom AOA or roll-induced departureswith the STORES CONFIG switch in
CAT I.

• SYM G limits apply to maneuvers
resulting from less than abrupt rollstick inputs and in which roll rate does
not exceed 20 degrees/second. ROLL G
limits apply to maneuvers resulting
from abrupt roll stick inputs or
maneuvers in which roll rate exceeds
20 degrees/second.
• A false maximum g indication may bedisplayed in the HUD due to INU
vibration while the aircraft is at
maximum g. G indications above 10(e.g., 0.2 for 10.2g) have been observed.
• ~ Due to the location of the accelerometer, it should not be used todetermine maximum g force.
• For evaluating a possible over-g,
determine the allowable CARRIAGEMAX ACCEL G for the actual store
configuration at the time of theoccurrence. Use of figure 5-10 NzW
curves (if applicable) is permitted.
Refer to T.O. IF-16A-I-3 (Block 10) or
T.O.IF-16A-I-4 (Block 15 and IBQ]).

•• G's experienced during a wingtip
vortex/wake turbulence encountershould be considered as asymmetricalwhen determining if a g limit has been
exceeded.
FULLY INTERNAL FUELED AIRCRAFT

Refer to T.O. IF-16A-I-3 (Block 10) or T.O.
IF-16A-I-4 (Block 15 and ~).
PARTIALLY INTERNAL FUELED AIRCRAFT

Refer to T.O. IF-16A-I-3 (Block 10) or T.O.
IF-16A-I-4 (Block 15 and ~).

5-4

Change 11

Refer to figure 5-9. With heavy wing loadings, it may
be necessary to cancel the roll command up to 90
degrees early to avoid exceeding the maximum bankangle limit. Except for emergency conditions, do not
fly category III loadings with the STORES CONFIG
switch in CAT I.

I

WARNINGI• Damage to or failure of wing internal
structure can occur if rolling maneuvers are performed with the STORESCONFIG switch in CAT I while
carrying a category III loading.

An asymmetric loading is any asymmetry that
requires roll and/or yaw trim. Refer to ASYMMETRIC LOADINGS, Section VI.
Nose slice and yaw departure may occur during
maximum command rolls on the CAT I AOA limiterat high altitude when carrying a centerline tank.
Refer to YAW DEPARTURE, Section VI.

I WARNING I• If the aircraft CG is near the aft limit,
departure may occur while performing
low airspeed, high AOA, maximum
command rolling maneuvers witheither of the following:• Asymmetric category I missile loadings (station 2,3, 7, or 8).
• Speedbrakes opened.

• m The indicated bank angle change

limit is particularly critical for category I loadings with 370-gallon fueltanks plus suspension equipment on
stations 3 and 7. Care is required withthese loadings to check the roll so as
not to exceed the indicated bank angle
change limit.

STORES LIMITATIONS
Refer to T.O. IF-16A-I-3 (Block 10) or T.0.IF-16A-I-4 I(Block 15 and ~ ), figure 5-11 for authorized storesloading configurations and the related limitations.

T. O. 1F- 16A- 1
ASYMMETRIC STORES LOADING

Refer to figure 3-17 for computation of asymmetricmoment. The maximum allowable asymmetric
Crolling) moment for ground handling, taxi, takeoff,
inflight, and landing is m25,020 foot-pounds, LESSm
22,650 foot-pounds.
Takeoff is prohibited when the roll trim necessary tocompensate for an asymmetric loading exceeds the
maximum roll trim available. Refer to T.O.
1F-16A-1-1, PART 2.

MISCELLANEOUS LIMITATIONSNET ARRESTMENT LIMITATIONSI Refer to figure 5-6 for cable/net arrestment limits.
61QSIIIBAK-15 multielement net barrier systems
have the following designations: 24A, 30F, 40A, HP,
and MEN. Any net barrier system with these
designations, e.g., 61QSIIIHP30 or MK-6/MEN, is
compatible with the aircraft.
The 61QSIIIBAK-15 arresting system consists of two
major components: an energy absorber and an
engagement device. The engagement device is the net

which is attached to the energy absorber. Theseabsorbers include heavy chains as used with the
MA-lA system, friction brake systems such as the
BAK-9 and BAK-12, and water twisters as used in
BAK-13 systems. Because of the variety of energy
absorbers in use, it is not possible to establish one
maximum engagement speed for the aircraft. In eachcase, the maximum engagement speed is determinedby the capabilities of the energy absorption system
attached to the net. For example, the HP30 is an all
nylon, three element net. It incorporates all the latest
technology and, coupled with a 1200-foot runout
BAK-12, should provide a 35,000-pound aircraft with
190-knot arrestment capability.
CROSSWIND LIMITS

Refer to T.O. 1F-16A-1-1, Part 2 for crosswind limits
for takeoff and landing.
AIRCRAFT BAnLE DAMAGE REPAIR

For aircraft and system limitations that should beobserved for flight with battle damage or after battle
damage repair, refer to T.O. 1F-16A-39.

Change 85-4.1/(5-4.2 blank)

T.0 . 1F-16A-l

Instrument Markings
ENGINE Fl00-PW-200__

680°C GROUND ENGINE START_

970°C MAXIMUM STEADY STATE

96% MAXIMUM STEADY STATENOTE: RPM is limited to 94 percent on takeoH roll.

RP M

FTlT

_

15 PSI MINIMUM_

2850-3250 PSI NORMAL_

80 PSI MAXIMUM_

3250 PSI MAXIMUMOIL PRESSURE

HYDRAULIC PRESSURE

IF-16A-l -"86)(.

FigureS-I. (Sheet 1)

5-5

T.O I F-16A-l

Instrument Markings
ENGINE Fl00-PW-220__

680C GROUND ENGINE START_

9BOC MAXIMUM STEADY STATENOTE:

96% MAXIMUM STEADY STATE

RPM is limited to 94 percent on takeoff ron

RPM

FTlT

_

1 5 PSI MINIMUM_

2850-3250 PSI NORMAL_

80 PSI MAXIMUM_

3250 PSI MAXIMUM

H YDRAULIC PRESSURE

OIL PRE SS URE

Figure 5 -1. (Soot 2)

5-6

10. 1 F-16A-1

Engine Limitations
ENGINE Fl00-PW-200
GROUND
FTIT
°C

RPM
%

OIL PSI

START

680

-

-

IDLE

575

-

15 (min)FTIT 750°C in BUC (stabilized idle rpm)

MIUAB

954

94

30-80

At MIL and above, oil pressure must increase 15 psi minimum above
IDLE oil pressure

CONDITION

REMARKS
During cold start, oil pressure may be 100 psi for up to 1 minute

From MIL to AB, a slight decrease of oil pressure (2-3 psi) can be experienced
TRANSIENT

970

94

30-80

Maximum temperature limited to 30 seconds. In BUC, throttle movements must be slower than normal (5 seconds or longer from BUC IDLE
to MIL or MIL to BUC IDLE)

FLUCTUATION

± 10

±1

±5 IDLE

Must remain within steady-state limits. In-phase fluctuations of morethan one instrument or fluctuations accompanied by thrust surges indicate engine control problems. Nozzle fluctuations limited to ± 2% at
and above MIL. Fluctuations not permitted below MIL± 10 above
IDLE

IN FLIGHT
FTIT
CONDITION

DC

RPM
%

OIL PSI

800

-

-

-

-

-

15 (min)

-

MIUAB

970

96

30-80

Oil pressure must increase as rpm increases. AB operation is prohibitedinBUC

MAX POWER

990

96

30-80

Do not use MAX POWER except in actual combat. Use of MAX POWER
limited to 6 minutes per application when over 957°C. Total of 60 minutes allowed before turbine inspection/overhaul. All MAX POWER timeshall be logged on AFTO Form 781

TRANSIENT

990

96

30-80

Maximum temperature limited to 10 seconds. BUC operation same as
ground operation

FLUCTUATION

±10

±1

±5 IDLESame as ground operation. Zero oil pressure is allowable for periods up
to 1 minute during flight at less than + Ig

AI RSTARTIDLE

REMARKS± 10 above
IDLE

Figure 5-2. (Sheet 1)
Change 13

5-7

I10. 1F- 16A- 1

Engine Limitations
ENGINE Fl00-PW-220
GROUND
FTIT
°C

RPM
%

OIL PSI

START

680

-

-

IDLE

575

-

15 (min)Maximum FTIT in SEC is 650°C

MIUAB

965

94

30-80

At MIL and above, oil pressure must increase 15 psi minimum above
IDLE oil pressure

TRANSIENT

980

94

30-80

Maximum temperature limited to 30 seconds

FLUCTUATION

±10

±1

±5 IDLE

Must remain within steady-state limits. In-phase fluctuations of morethan one instrument or fluctuations accompanied by thrust surges indicate engine control problems. Nozzle fluctuations limited to ± 2% at
and above MIL. Fluctuations not permitted below MIL

CONDITION

I± 10 above
IDLE

REMARKS
During cold start, oil pressure may be 100 psi for up to 1 minute

IN FLIGHT
FTIT
°C

RPM
%

OIL PSI

800

-

-

-

IDLE

-

-

15 (min)

-

MIUAB

980

96

30-80

Oil pressure must increase as rpm increases

MAX POWER

1000

96

30-80

Do not use MAX POWER except in actual combat. Use of MAX POWER
limited to 6 minutes per application. Total of 60 minutes allowed before
turbine inspection/overhaul. All MAX POWER time is recorded by the
EDU and shall be logged on AFTO Form 781

TRANSIENT

1000

96

30-80

Maximum temperature limited to 10 seconds

FLUCTUATION

± 10

±1

±5 IDLESame as ground operation. Zero oil pressure is allowable for periods up
to 1 minute during flight at less than +lg

CONDITIONAI RSTARTREMARKS± 10 above
IDLE

Figure 5-2. (Sheet 2)

5-8Change 8

T. O. 1F- 16A- 1

Engine -AB Operation and Light-Off Limits
ENGINE F100-PW-200

Region 1: Unrestricted AB operation. Mislights, blowouts, or stalls should not occur. Some AB rumble may occur insegment 5 just below the lockout line and may result in AB blowout or stalls. If a mislight, blowout, or selfclearing engine stall occurs coincident with AB operation in region 1, an engine problem associated with AB
is probable. Do not attempt further AB operation and write up occurrence. The engine is safe to operate forthe rest of the flight in the IDLE to MIL range provided no other abnormal engine indications are observed.
Region 2: Most throttle transients in AB can be expected to be successful; however, AS mislights, light-off stalls, or
rumble blowouts or stalls may occur during AB light-off or during throttle transients within the AB range.
These anomalies should not occur during steady-state AB operation. If blowout or stall occurs in region 2
and FTIT limits were not exceeded, progressing to region 1 should permit a successful AB relight and normal
use of AB in region 2. If a second mislight, blowout, or self-clearing stall occurs in region 2, further AB use
should be restricted to region 1 and the discrepancy should be written up.
Region 3: Steady-state AB operation or AB cancellations only in this region. AS mislights, light-off stalls, blowouts, andrumble-induced stalls are probable during transients within the AB range. AB rumble and blowout are possible
during steady-state operation. If a mislight, blowout, or stall occurs, progressing to region 1 should permit a
successful AB relight and normal use of AB.

70

...~

."-

60

."I,

r-- REGION 3 r-- STEADY-STATE AB
r-r-- OPERATION OR AB
r-- CANCELLATIONS
r-50 -r-- ONLY

,'I
l"\j
.'\j

"-

."'

...

-" L "'
./

."'~

~v

-

1'1 ~

~

."-

""

~

Iol'(

.""

5c(
.'\.

,'\.

.'\.

NOTES:*1. Time limited transient
operation - Operation
limited to 3 minutes.

~"'

2. Restrictions pertain toregion where AB
light-off occurs, notnecessarily the regionin which the throttle
is advanced to AB.

."'
~:::
l'!::'-":.:U/' ."'1

1r.:JI: 1'\.' .'Jj

REGION 1 UNRESTRICTED
AB OPERATION

"::J

20

~

V

TIME LIMITEDTRANSIENT
SEGMENT 5 AB LOCKOUT - - OPERATION"( ± 0,20 MACH CURVED\
PORTION ONLY)
I~
1

"
1\

i.-""

."'."'~.~~'f ~t-."f-r- f-u

./

l'l
i"-.

."-

1/
./

l""

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~~L':--:'REGION 2 POSSIBLE MISLIGHTS,
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1F-16X-1-0017A@

Figure 5-3. (Sheet 1)

Change 1

5-9

IT. O. 1F- 16A- 1

Engine -Operational Envelope
ENGINE F100-PW-220

NOTES:• Transfers to SEC when subsonic below 40,000 feet MSL should be performed with the throttle at midrange
or above.
• Transfers from SEC to PRI should be performed with the throttle at MIL or below.
• Transfers to SEC above 40,000 feet below 300 knots may result in a self-recovering stall.
• For supersonic transfers to SEC, do not retard throttle below MIL until subsonic.

I

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Figure 5-3. (Sheet 2)

5-10

T.O. 1F-16A-1

sue -Operational Envelope
ENGINE F100-PW-200

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Figure 5-4.

5-11

T.O. 1F-16A-1

Jet Fuel Starter Limits
LIMITS

CONDITION* Continuous motoring of the engine shall not exceed 4 minutes. Mter 4 minutes of continuous motoring, a nonnal engine start may be initiated after 5 minutes of coolingN onnal Ground Operation
A minimum wait of 1 minute is required after each JFS start attempt to allow fueldrainage from the JFS
Hot Start of Engine

Motor until FTIT is below 200°C

AirstartlIn Flight

Below 20,000 feet MSL and 400 knots. 3-minute maximum run time when the engine
is operating satisfactorily above 60 percent rpm; otherwise, unlimited* Motoring is defined as JFS rotating the engine at 22 percent rpm minimum with the throttle in OFF.
NOTES:• OAT between 20°F (--6°C) and 100°F (38°C). A minimum brakelJFS accumulator pressure of 3000 psi is required
for START 1 and 2800 psi is required for START 2.
• OAT between -25°F (-32°C) and 20°F (-B°C) or OAT above lOO°F (38°C). START 2 and a minimum brakelJFS accumulator pressure of 2800 psi are required.
• OAT below -25°F (-32°C). START 2 and a minimum brakelJFS accumulator pressure of 3200 psi are required.
• If one brake JFS accumulator is depleted, verify a minimum pressure of 3000 psi in the remaining brake/JFS accumulator before attempting START 2.

Figure 5-5.

5- 12

Change 8

T.O. 1 F-16A-l

Cable/Net Arrestment LimitsCompatible Cable Systems With Established Aircraft Limits
BAK-6
BAK-9
BAK-12 (Standard, Extended, and Dual)
BAK-13 (Navy designation is E-28)
BAK-14
BAK-15 (ND (Net barrier with cable for hook)
MAAS*44B-2L•

ROUTINE ARRESTMENT LIMIT - 135 KNOTS (*156 KNOTS).
A planned event. Operational conditions are such that each landing requires arrestment. Such
operational conditions include operating from highways or from runways that are too short fornormal landings. The standard factor of safety is used to determine the maximum engagement
speed.
EMERGENCY ARRESTMENT LIMIT - 150 KNOTS (*171 KNOTS).
An unplanned event. A reduced factor of safety is accepted and the corresponding maximum engagement speed is slightly higher than for a routine arrestment. With a ground level ejection capability, a reduced factor of safety for the arrestment is possible since the pilot can still eject if thearrestment fails.

Compatible Cable Systems Without Established Aircraft Limits

Use of arresting systems without established aircraft limits may resultin failure of the hook, hook backup structure, or nose landing gear.

BLISS 500S-6
TAGS BLISS 500S-6B Transportable
44B-3H/SPIWRMAGS 44B-3HJSPIWR Mobile
RHAGMK-l
RHAGMK-2
PUAG
PAAG
P-IVIBAK-12 Portable

Figure 5-6. (Sheet 1)

Change 115-12.1I
I

I

T. O. 1F- 16A- 1

Cable/Net Arrestment LimitsCompatible Strut Engagement/Cable Systems Without Established Aircraft Limits

Use of arresting systems without established aircraft limits may resultin failure of the hook, hook backup structure, or nose landing gear.
MA-IA Modified
MA-IAlE-5
MA-IAIBAK-9
MA-IAIBAK-12

NOTE
Attempting to engage an unmodified (non-hook capable) MA-IA will most
likely be unsuccessful

Compatible Net Barrier Systems
BAK-15
SAFE-BAR (Safeland Barrier)Stanchions
6IQSII
6IQSIIMMK-6

NOTE
The above stanchions can be rigged with the following nets: HP30, HP40,
MEN241, MEN301, 24, 30, or 40. In addition, they will be associated withthe following energy absorbers: Chain (E-5), BAK-9, BAK-12, BAK-13,
34B, 44B, 500S, BEFAB 6:3, BEFAB 21:2, and BEFAB 24:4.
Other Net Systems
ARZ-30/-40
BEFAB 21:2/MK VI-I stanchion
BEFAB 6:3/MK VI-1 stanchion
MK-12ARAFTYPEARAFTYPEB

Figure 5-6. (Sheet 2)5-12.2

Change 8*u.s. ~ PRINTING OFFICE: 2002-760-001/63123

10. 1F- 16A-1

Airspeed Limitations(Systems)

Acceleration
Limitations
LOAD FACTOR (g)SYSTEM OR CONDITIONKIASIMACHCONFIG U'RATI ON
SYMMETRIC ASYMMETRIC

Canopy Open or in Transit

70 (Includes
ground windvelocity)

TAKEOFFLG Extended or in Transit

30010.65, whichever is less

TEF's Operated With ALT
FLAPS Switch

40010.85, whichever is less

LANDING

40010.95, which-

LG RETRACTION*+4.0,0.0+2.0,0.0+2.0,0.0+2.0,0.0

AR Door Opening/Closing
AR Door Openever is less

Flight in Severe Turbulence
(+3g)

500

~ If!QI Drag Chute Deploy-

170

LG EXTENSION

ment* If the lG handle is raised near 2 g's approaching 300 knots, actua-I
tor power may be insufficient to completely retract the lG until 9 isreduced.
Figure 5-7.

Figure 5-8.

Change 13

5- 13

T.O.1F-16A-l

AOA and Rolling Limitations
AIRCRAFTLESSm

MAX BANK ANGLE CHANGE
FOR MAX ROLL MANEUVER

LOADING
CATEGORYSTORES CON FIG
SWITCHMAXAOAII

LIMITER360

LIMITER360 0 Below 15 0 AOA

LIMITER

180 0 Above 15 0 AOA

III

LIMITER

360 0

III

III

LIMITER

180 0

I

I

LIMITER

360

III

III

LIMITER

360 0

I

0

--

II

0

ffi
NOTES:

I

1.

Determine loading category from the appropriate line in T.O. 1F-16A-1-3, figure 5-11 (Block 10) or T.O.
1F-16A-1-4, figure 5-11 and 5-12 (Block 15 and ~), STORES LIMITATIONS, and LESSin figure 2-4, FUEl
MANAGEMENT.

2.

The roll command should be released in sufficient time to avoid overshooting the indicated bank angle change
limits.

m

Figure 5-9.

5- 14

Change 3

T. O. 1F- 16A- 1

SECTION VI
FLIGHT CHARACTERISTICS
TABLE OF CONTENTS

Page

Introduction .............................
Flight Control System ....................
FLCS Limiters. . . . . . . . . . . . . . . . . . . . . . . . .
Leading Edge Flaps ....................
Speedbrakes . . . . . . . . . . . . . . . . . . . . . . . . . . .
Autopilot. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
'rrim ..................................
Normal Flight Characteristics. . . . . . . . . . . . .
Category I Loadings . . . . . . . . . . . . . . . . . . . .
Category II Loadings LESS m ...........
Category III Loadings ..................
Flight With LG Down. .. . . . . . .. .. . . .. . . .
Landing Configuration .................
Factors Affecting Flying Characteristics . . . .
Center-of-Gravity Considerations. . . . . . . .
Effect of Thrust .. . . . . . . . . . . . . . . . . . . . . . .
Effect of Low Airspeed Maneuvering . . . . .
High Pitch, Low Airspeed . . . . . . . . . . . . . . .
Flight With Stores .....................
Limit Cycle Oscillation .................
Asymmetric Loadings ..................
Store Separation. . . . . . . . . . . . . . . . . . . . . . .
Out-of-Control Characteristics. . . . . . . . . . . . .
Yaw Departure. ..... . ... ...... .. .... ...
Pitch Departure . . . . . . . . . . . . . . . . . . . . . . . .
Deep Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Spin ..................................
Recoveries. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Self-Recovery. . . . . . . . . . . . . . . . . . . . . . . .
Deep Stall Recovery. . . . . . . . . . . . . . . . . .
Spin Recovery. . . . . . . . . . . . . . . . . . . . . . .
Engine Operation During Departures!
Ou t-of-Con trol . . . . . . . . . . . . . . . . . . . . . . .
Degraded Flight Controls .................
Leading Edge Flaps Locked(Symmetric) . . . . . . . . . . . . . . . . . . . . . . . . .
'rrailing Edge Flaps (Flaperons) .........
Locked Control Surfaces ................
Flaperon . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Horizon tal Tail ......................
Rudder .............................
Standby Gains . . . . . . . . . . . . . . . . . . . . . . . . .
One Hydraulic System. ..... ....... .. .. .
Speedbrakes . . . . . . . . . . . . . . . . . . . . . . . . . . .

6-1
6-1
6-1
6-1
6-2
6-2
6-2
6-3
6-3
6-3
6-3
6-4
6-4
6-5
6-5
6-5
6-5
6-5
6-6
6-6
6-6
6-7
6-7
6-7
6-8
6-9
6-10
6-10
6-10
6-10
6-11
6-12
6-12
6-12
6-12
6-12
6-12
6-12
6-13
6-13
6-13
6-13

Aircraft Damage. . . . . . . . . . . . . . . . . . . . . . . . . .
Horizontal Tail . . . . . . . . . . . . . . . . . . . . . . . . .
Flaperon ..............................
Rudder and Ventral Fins. . ...... ... ... . .
Leading Edge Flaps ................ ,. . .
Wing..................................
Radome ...............................
Dive Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6-13
6-13
6-13
6-13
6-13
6-13
6-14
6-14

INTRODUCTION
Information presented in this section reflects theflight characteristics with category I, LESS mII, andIII loadings. Refer to AOA AND ROLLING
LIMITATIONS, Section V, for information rega~ding
specific categories.

FLIGHT CONTROL SYSTEMThe FLCS is a four-channel fly-by-wire system. The
FLCC and ECA combine pilot commands along withaircraft motion and flight conditions to command
position of the flight control surfaces. Artificial
stability provided by the FLCS allows for relaxed
static stability which increases performance andmaneuverability by reducing trim drag and increasing maximum lift. Refer to FLIGHT CONTROLSYSTEM, Section I for a detailed system description.
FLCS LIMITERS

FLCS limiters may be defeated if maneuvering limits
are not strictly observed. Departure may result frommaximum maneuvering combined with maximum
permissible aft CG. The most critical maneuvers are
maximum command rolls coupled with either
maximum aft stick or exceeding the maximum bankangle change limits.
The AOAlg limiter depends on the horizontal tails tocontrol g and AOA. If the airspeed decreases until
there is not enough airflow over the tails to provide
this control, the limiter is defeated and a departure or
deep stall may result. This condition may occur in a
nose-high, decreasing speed maneuver. Refer to LOW
AIRSPEED OPERATING LIMITATIONS, Section V.
LEADING EDGE FLAPSThe LEF system is designed to optimize wing airflow.
It also provides special functions in the takeoff andlanding configurations.

Change 8

6-1

T.O. 1F- 16A-1

referenced altitude if AOA increases
above a certain value. This value
assumes that the pitch trim is centered. Any use of nosedown trim
reduces net autopilot authority andallows descent to occur at lower AONs.
Depending on thrust setting, highrates of descent can develop. With the
STORES CONFIG switch in CAT I, theaircraft will start descending from thereferenced altitude at approximately
16-19 degrees AOA and in CAT III at
approximately 8-10 degrees AOA.

At subsonic speeds, the LEF's move from 2 degrees up
to 25 degrees down as a function of mach number,
AOA, and altitude. This automatic operation
significantly reduces buffet and drag and improves
high AOA directional stability. If the LEF's fail toschedule properly during maneuvering flight, higher
than normal buffet levels occur and, in the high AOA
region, reduced directional and longitudinal stability
may also be noted. At supersonic speeds, the LEF's
are scheduled to minimize drag.
NOTE

The LE FLAPS caution light may
illuminate during maneuvering flightat approximately 1.0 mach at 20,000
feet MSL or greater. The LE FLAPS
caution light should be reset at anairspeed other than 1.0 mach.

Since the autopilot command is additive to stick commands, use of ALT
HOLD in conjunction with high gmaneuvering may result in aircraft
over-g.

SPEEDBRAKESThe speedbrakes provide deceleration over the entire
flight envelope. There are no trim changes associatedwith speedbrake operation and induced buffet is
negligible. A yaw oscillation may occur at approximately 1,4 mach with speedbrakes opened. The
oscillation is neutrally damped and no action isrequired. The oscillation may be eliminated by either
closing the speedbrakes, reducing mach, or increasing the g level.
AUTOPILOT

The aircraft can be trimmed about all three axes.
With pitch trim centered in cruise gains and no input
to the stick, the aircraft attempts to maintain 19
flight regardless of flight condition unless AOA
exceeds 15 degrees. Full noseup/full nosedown trim
corresponds to +3,4g or -l,4g in cruise gains.
NOTE

With the HDG SEL and ALT HOLD modes engaged,
the aircraft turns, climbs, or dives within the limitsof the autopilot to capture the heading reference andthe altitude reference regardless of aircraft attitude.
This autopilot-commanded flight may eventually
return the aircraft to a preselected heading andaltitude if airspeed and altitude permit.
If the ALT HOLD remains engaged as airspeedtransits 1.0 mach, a mild pitch transient may occur
and can be eliminated by depressing the paddle
switch until the altimeter has stabilized.

I

WARNINGI

Use of ALT HOLD during decelerating
flight can produce a descent from the

6-2

TRIM

Airspeed must be closely monitored
because there is little aerodynamic
indication of large changes in airspeed. Cues which normally indicate
airspeed changes, such as stick movement or trim changes, are absent.
Above 15 degrees, the FLCS commands an increasing
nosedown pitch attitude as a warning of decreasing
airspeed. A specific force applied to the stickcommands a specific g increment from the trim
condition. Moving the PITCH TRIM wheel changes
the hands-off trim condition.
In takeoff and landing gains, zero pitch trim
commands zero pitch rate until 10 degrees AOA. A
slight amount of nose up trim is required to zero stickforces during an 11-13 degree AOA approach.

T.O. 1F- 1 6A- 1

When properly trimmed and no input is applied to thestick, the aircraft attempts to maintain zero roll rate.
Moving the ROLL TRIM wheel changes the hands-off
trim condition. Maximum roll trim authority isapproximately one-fifth of maximum stick commandof cruise gains. However, precise trimming is difficult
using the stick TRIM button. Roll trim requirements
may change with stores, particularly at supersonic
speeds. For asymmetric configurations (asymmetrical stores or rudder mistrim), roll retrimming may berequired as flight conditions change. Roll trim inputs
also command rudder deflection through the ARI.
The ARI switches out with wheel spinup uponlanding. Likewise, the ARI switches in following
takeoff as the wheels spin down. This switching may
cause abrupt rudder inputs to occur if roll (due toasymmetries or crosswind) is being input via the stickor trim.
Rudder trim inputs command rudder deflection.
Rudder trim is required with asymmetrical configurations and frequently during supersonic flight,
especially with stores. Maximum trim authority is 12
degrees.

NORMAL FLIGHT CHARACTERISTICS

I WARNINGIThe capability of the aircraft to rapidly
attain and sustain high g levels, which
may cause g-induced loss of consciousness, should be considered during
heavy maneuvering.
The FLCS provides constant response for specificinputs regardless of flight conditions. Commanded
pitch responses are in g increments per stick force for
AOA below 15 degrees. Above 15 degrees AOA, stickforce increases as a cue of increasing AOA.
Conventional cues such as aircraft buffeting forces
are not always present as AOA and g limits areapproached. The commanded lateral response is rollrate per stick force. Rudder position is commanded byrudder pedal force.
The ARI provides coordinated rudder commands andreduces sideslip during rolling maneuvers. Additional pilot rudder inputs do not improve roll performancebut do increase departure susceptibility. When ARI is

not available during takeoff and landing (MLG wheel
speed above 60 knots), pilot rudder inputs may berequired to provide coordinated flight and to control
yaw.

IWARNINGI
Rolling g limits are not protected by
the FLCS and must be observed.
CATEGORY I LOADINGSThe FLCS minimizes the possibility of departures or
spins. Roll rate inputs command flaperons andhorizontal tails for roll power to provide a relatively
constant roll response.
Maximum command 360-degree rolls at subsonic
speeds may cause a slight g reduction on termination.
At supersonic speeds, maximum roll rates may causea slight increase in g. At high AOA and low airspeed
conditions, roll performance is reduced by the FLCS
to minimize pitch/roll coupling. Aft CG's, open
speedbrakes, asymmetric missiles, or centerline
stores decrease departure resistance.
CATEGORY II LOADINGS LESSm

Category II loadings increase departure susceptibility to pitch/roll coupling and rudder inputs during
rolling maneuvers; however, flight characteristics are
not seriously affected. Structural overstress muststill be avoided. With STORES CONFIG switch in
CAT I, requirements to adhere to category II limits
increase workload. With STORES CONFIG switch in
CAT III, workload is significantly reduced, butaircraft performance is also reduced to category III
limits. In CAT III position, category II loadings may
be flown symmetrically to the AOA limiter, and max
command 360-degree rolls may also be performed to
the AOA limi ter.
CATEGORY III LOADINGS

Aircraft response with most category III loadings
remains similar to that of the clean aircraft; however,large stores significantly increase total aircraft dragand reduce performance. Light buffeting may occur
during level flight at approximately 0.92 mach. In
addition, surging may occur near the store limit
airspeed, especially at low altitude. Neither condition
requires specific action.

6-3

10. , F- , 6A- ,With STORES CONFIG switch in CAT III, the AONg
limiter provides departure resistance for all categoryloadings. Except for the requirement to avoid
structural overstress, pilot workload is reduced to alevel comparable to that with category I loadings.
FLIGHT WITH LG DOWN

With the LG handle down, LG and TEF's are extended
and the FLCS operates in takeoff and landing gains.
Normally, this mode of flight is limited to takeoff,approach, and landing; however, circumstances canarise which require flight for an extended distancewith the LG down. If so, the LG should be left pinned
bu t the streamers should be removed to prevent
damage.
With the LG pins installed, it is preferable to raise the
LG handle once airborne. This action retracts the
TEF's and significantly reduces drag and the FLCS
switches to cruise gains. For cruise with only the LG
down, the best airspeed is 230-250 knots. A clean
aircraft can be flown at 25,000-30,000 feet with the
LG down and TEF's up and fuel flow is 3000-3400
pph. If the LG handle is left down, the TEF's remain
down and the best cruise altitude is less than 20,000
feet with significantly higher fuel flows.
LANDING CONFIGURATION

Two distinct techniques may be used when landing.
One technique is to trim for approximately 11 degrees
AOA and to fly that airspeed throughout the final
approach. Attitude/glidepath is controlled by the
stick, and airspeedlAOA is controlled by the throttle.
This technique allows better pitch control, betterover-the-nose visibility, and a more stable HUD
presen tation. In gusty wind conditions, the aircraft
wallows less, and during the flare, the sink rate iseasier to control. The aircraft floats approximately
800-1200 feet from flare initiation to touchdown.
Another technique is to trim for 13 degrees AOA andto fly that airspeed throughout the final approach.
The throttle is used primarily to control glidepath,and the stick controls airspeed through control of
AOA and direction through bank angle. This type ofapproach primarily allows better control of touchdown point and more efficient energy dissipation;
however, since the aircraft is already at 13 degrees
AOA, the flare is more difficult, and care must be
exercised to avoid scraping the speedbrakes or

6-4

landing firm. The aircraft floats approximately
500-700 feet from flare initiation to touchdown.
Regardless of the technique used, establish computed
final approach airspeed for the desired AOA early on
final and trim the aircraft. Airspeed changes result in
pitch changes, which may require retrimming andmake glidepath control more difficult. Small throttle
adjustments may be required as the I pW200i EEC,I PW2201 DEEC retrims the engine.
On short final, avoid premature or large thrust
reductions which may cause increased sink rates anda hard landing. Use thrust rather than back stick tocontrol undesirable sink rates. Increased back stick
may result in a tail strike in this situation. AOA
decreases slightly as the aircraft enters ground effect.
All normal landings should be made withspeedbrakes opened to the 43-degree position to avoid
a floating tendency when entering ground effect. A
touchdown at the desired point at 13 degreesAOA canbe achieved when flying final at either 11 or 13
degrees AOA by adjusting the initial aimpoint.
Increased control inputs to achieve normal aircraft
response as airspeed decreases are unnecessary.
Control inputs should be kept small to avoid
overcon trol.
Due to the aircraft light wing loading and the floating
tendency associated with ground effect, waketurbulence on final approach and during touchdown
presents a significant hazard. Increased spacing
between landing aircraft should be used when there
is little or no effective crosswind. Exercise caution
and be ready to initiate a go-around when waketurbulence is encountered. An early go-arounddecision may help avoid the need for a large roll
control input. Such an input retracts a flaperon,causing decreased lift and possibly a sink rate as well
as a roll. A large roll input at slow airspeed also causesa large horizontal tail split. A horizontal tail surface
could contact the runway while trying to counter
wake turbulence effects during touchdown.
Ifpitch trim is used during the turn to final, forward
stick/trim will be required upon rollout on final
approach to counter noseup motion. Floating
tendencies following a high flare or aircraft bounce
may be increased. Slight forward stick force may berequired to prevent a long or slow landing. Stick force
per degree AOA change is reduced and should not be
relied upon as a slow speed cue.

10. 1 F- 16A-l

FACTORS AFFECTING FLYING CHARACTERISTICS
NOTE

A mild pitch oscillation (a maximum of± 0.15g at 3 cycles per second) may
occur at 0.75-0.90 mach while in cruise
gains or at 330-400 knots while intakeoff and landing gains. The oscillation is caused by the normal responseof the aircraft and FLCS and does notcause a significant tracking problem.
CENTER-OF-GRAVITY CONSIDERATIONSMonitoring the forward and aft fuel distribution
provides an indication of the aircraft CG.
As CG moves aft, higher pitch rates are obtainable

and susceptibility to departure and deep stall
increases.
NOTE• ~ The most aft CG occurs withapproximately 2000 pounds of internalfuel remaining .

• 00 With external fuel tanks, the mostaft CG occurs when the external tanks
have just emptied.
EFFECT OF THRUST

Thrust changes result in little or no change in aircraft
trim or stability at all operational load factors and forall store loadings.
EFFECT OF LOW AIRSPEED MANEUVERING

Departures are possible at low airspeeds and low
pitch angles if large, simultaneous pitch and roll
inputs are made.
The FLCS requires adequate airflow over the control
surfaces to be effective, which means that airspeed isa critical factor in departure susceptibility during
maneuvering. Low airspeeds should, therefore, be
avoided during maximum performance maneuvering.

I

WARNINGI• FLCS limiters can be defeated at low
airspeeds (below 200 knots in a CAT I
configuration) during maximum pitch
and roll commands initiated frombelow limiter AONs.

• The aircraft can be departed (fromparameters outside the tone on portionof figure 1-49) with no low airspeed
warning tone present if abrupt or
uncoordinated FLCS inputs are made.
HIGH PITCH, LOW AIRSPEED

The low airspeed warning tone sounds to aid inrecognizing that critical high pitch, low airspeed
flight conditions are reached.

I

WARNINGI

Proper assessment of flight path angle(not pitch angle) is key to determining
the nearest horizon and performing aproper recovery. Differences between
flight path and pitch angle of up to 25
degrees, combined with the visual
illusion caused by a reclined seat canlead to an incorrect decision tocontinue the maneuver through thevertical. The risk of a departure/deepstall in this instance is very high.
Avoiding a departure under these conditions requires
specific control techniques. To recover, first release
aft stick pressure. This action unloads the aircraft
and reduces AOA so that the flightpath more closely
coincides with the longitudinal axis of the aircraft.
Smoothly roll inverted to the nearest horizon. Mterthe roll, smoothly apply the aft stick pressure
required to keep the nose moving toward the horizon.
As airspeed continues to decrease during therecovery, more aft stick pressure may be required tokeep the nose moving. Continue to smoothly increase
aft stick pressure up to the AOAlg limiter. If full aftstick is inadvertently released, do not reapply it
unless required to keep the nose moving.

I WARNING

I

Avoid large, simultaneous pitch androll inputs to preclude a roll-coupled
departure. Small lateral inputs can bemade as required to maintain wings
level, inverted flight. Do not abruptly
apply aft stick pressure at anytime
during the recovery. Rapid aft stickpressure generates excessive AOA,
overshooting the AOA limiter andcausing departure.
Change 3

6-5

T. O. 1F- 16A- 1

During a recovery where full aft stick is required,
nose movement toward the horizon may slow down
markedly as the AOAJg limiter tries to limit ADA. As
long as the nose continues to move, no further action
is required. If the nose of the aircraft does notcontinue to move toward the horizon, the aircraft has
departed, and out-of-control recovery proceduresshould be initiated.

Mter attaining a nosedown attitude with airspeed
increasing, continue to avoid abrupt inputs. The
aircraft may either be unloaded and rolled upright ora split-s recovery can be made at airspeeds above 200
knots, altitude permitting, before continuing tomaneuver. The split-s recovery is the simplest way torecover the aircraft. However, if altitude is a factor,
allow airspeed to increase to a minimum of 150 knots
unload the aircraft to less than 19, smoothly roli
upright, and recover to level flight.
FLIGHT WITH STORES

The major effects of stores are increased weight andinertia. A reduction in aircraft response and damping
should be expected as GW increases, particularly
when stores are carried. Stores generally reduce
longitudinal and directional stability and increase
inertial effects so that the pilot must anticipate
initiation and termination of maneuvers based on theloadings. High roll and pitch rates are attainable withfull force application of the stick. Avoid abrupt control
inputs which may cause AOA overshoots in excess of
the limitations specified in Section V and T.O.
1F-16A-1-3 (Block 10) or T.O. 1F-16A-I-4 (Block 15
and~ ).
Bank angle change limits must not be exceeded.
During rolling maneuvers with category III loadings,
the roll rate must be stopped prior to m 360-degree,
LESS m lBO-degree bank angle change. Removing
the roll input is not always sufficient (opposite stick
may be required). Refer to STORES LIMITATIONS,
T.O. 1F-16A-1-3 (Block 10) or T.O. 1F-16A-1-4 (Block
15 and ~ ), for carriage limits.
Certain store loadings may exhibit decreased
yaw/roll damping in supersonic flight and result in
mild yawing oscillations. Neutral and divergent yaw
and roll oscillations may occur during sideslip
maneuvers at supersonic airspeed. These oscillations
are aggravated when large stores are carried.
Excessive vertical tail loads may be generated if
oscillations become sufficiently large. If oscillations
are encountered during rudder inputs, release therudder input. Additionally, buffeting may occur intransonic flight with certain store loadings.

6-6

NOTE
A mild airframe vibration may be
experienced while supersonic whencarrying a centerline store.
LIMIT CYCLE OSCILLATION

A limited amplitude constant frequency oscillation(commonly referred to as limit cycle oscillation or
LCD) may occur with certain stores loadings. The
LCD (typically 5-10 cycles per second) may occur inlevel flight or during elevated g maneuvers. The LCD
may appear as buffeting or turbulence similar to thatexperienced during normal transonic buffet, but thebuffeting is a constant frequency, lateral accelerationfrom side-to-side or, in some cases, verticalaccelerations up and down. The magnitude generally
increases with increasing airspeed and/or load factor.
Other cues of LCD include significant verticalmovement of the forward area of wing stores,
especially wingtip launchers and missiles; thismotion is typically up and down, but may also follow
a circular pattern. In addition, cockpit instruments
may become difficult to read as the LCD amplitude
increases from moderate to severe. Within publishedcarriage limits, LCD is not detrimental to the aircraft.
LCD susceptible loadings include air-to-surface andair-to-air loadings and associated downloadings. If
LCD is encountered and is uncomfortable or
distracting, reduce airspeed and/or load factor. Refer
to STORES LIMITATIONS, T.O. 1F-16A-1-3 (Block10) orT.O. IF -16A-1-4 (Block 15 and~ ), for carriage
limits.
ASYMMETRIC LOADINGS
lfroll trim is used to hold up a heavy wing, the ARI
adds rudder in the direction of the roll trim, causinga yaw away from the heavy wing. If roll trim is used
for takeoff, yaw occurs when the wheel speed drops
below 60 knots groundspeed after takeoff, activating~he ARI. This yaw is easily contropable by rudder~puts. Yaw and roll trim requirements change fordifferent flight conditions.
Asymmetric loads increase departure and spin
susceptibility. Roll inputs/trim away from the heavy
wing is required to maintain the desired roll attitude.
Increasing g requires additional roll inputs/trim.
Therefore, aft stick inputs result in increased rollrequirements which, in turn, produce yaw away fromthe heavy wing due to ARI action.

T. O. 1F- 16A- 1

I WARNING I• With certain asymmetric category III
loadings (2000 pounds or greater on
station 3 or 7 with stores on stations 4,6, and/or 5), rapid or abrupt aft stick
inputs may result in sudden noseslicing departures.
• Departure with an asymmetric category III loading may result in a fast, flat
(possibly nonrecoverable) spin.
NOTE

Left-wing heavy asymmetries are
more susceptible to departure.

limits are exceeded or in certain circumstances when
the FLCS provides only marginal protection.
With a lateral asymmetry in excess of 300 pounds(including wing tip missile and internal/externalfuel), abrupt maneuvering on or near the CAT I AOA
limiter can result in a departure if the aircraft isconfigured with any of the following:• 300-gallon fuel tank .

• Inlet mounted pod(s).
• Combination of a centerline store plus stores (orsuspension equipment) at stations 3 and/or 7.
With these loadings, maneuvering at high altitude(above 25,000 feet) increases the probability of a yaw
departure. Even moderate control commands may
cause some aircraft with CAT I loadings to departabove 25,000 feet.
Abrupt maneuvering on or near the CAT I AOA
limiter at slow airspeeds (less than 200 knots) may
result in a departure.

STORE SEPARATIONYAW DEPARTURE

Symmetrical store releases and wingtip AIM-9
missile launches can be accomplished with nounusual aircraft responses. Separation of the
300-gallon fuel tank produces negative g on theaircraft. The magnitude of this response depends on
the amount of fuel remaining in the tank and themach number at release. Separating a full centerline
fuel tank at supersonic speeds produces the worst
response (up to -1.5g).

A yaw departure occurs when sideslip increases
beyond the normal controllable range (i.e., beyond
about 15 degrees). The primary indication of mostyaw departures is an abrupt nose slice. The aircraft
then fails to respond properly to pilot commands
and exhibits uncommanded motions. AOA is in thenormal range (-5 to +25 degrees) during the initial
phase of the nose slice. Immediately following a yaw
departure, a pitch departure usually occurs
resulting in AOA indication of -5 or +32.

OUT-OF-CONTROl CHARACTERISTICS
A departure is a loss of aircraft control that ischaracterized primarily by uncommanded aircraft
motions or failure of the aircraft to respond to control
commands. In a pitch departure, the AOA increases
beyond the normal controllable range. In a yaw
departure, the sideslip angle increases beyond thenormal controllable range first, although a pitch
departure may immediately follow. The automaticfeatures of the FLCS normally prevent departures.
However, departures may occur when Section V

I

•

I

• 370-gallon fuel tanks.

At high airspeeds, asymmetric loads exhibit some
unusual flight characteristics. Frequent trim reversals may occur during supersonic acceleration. At
airspeeds greater than 700 knots, yaw oscillation
may occur with significant lateral accelerations. Over750 knots, a high frequency directional shaking may
occur with loadings such as the ECM pod.

Separation of 370-gallon fuel tanks produces aminimal aircraft response. Separation of a single
370-gallon fuel tank initially produces aircraft
positive g response and roll away from the separated
tank (up to +lg and 15 degrees of bank).

I

It is possible for the sideslip to briefly exceed thenormal controllable range without the aircraft
experiencing uncommanded motions. In this situation, the pilot's only indication of a departure may be
noticeable sideforces. These brief departures typically self-recover within 5 seconds.
A yaw departure may occur while maneuvering on ornear the CAT I AOA limiter in the 0.80 to 0.95 mach
range, especially at high altitude (above 25,000 feet).
Maneuvering at high altitude is more critical than
low altitude because mach effects reduce directional
stability. These yaw departures usually result frommaximum command left rolls; but they may also
occur during symmetric maneuvering.
The possibility of a yaw departure is increased
whenever the aircraft is configured with stores orsuspension equipment, especially a centerline store.
Change 4

6-7

I

I

T.O. 1 F-16A-l

In general, the possibility of departure increases as
the number, weight, and size of such equipment orstores increases. Susceptibility to a yaw departure
increases significantly with lateral asymmetry, with
left-wing-heavy loadings being more likely to departthan right-wing-heavy loadings. In addition, aheavier GW aircraft is generally more likely toexperience yaw departures. CAT I loadings mostsusceptible to yaw departures have one or more of thefollowing characteristics:

I- Centerline store.

PITCH DEPARTURE

I

I

A pitch departure occurs when the AOA exceeds the
AOA/g limiter. A pitch departure is classified either
as upright if the AOA is positive or as inverted if the
AOA is negative. Although the AOA indicator
displays a maximum of 32 degrees and a minimum of-5 degrees, the actual AOA during a departure will
exceed these values. In highly oscillatory departures,
the AOA indicator may momentarily indicate an AOA
below 32 degrees. Airspeed indications are erroneousand generally oscillate between the minimum value
and approximately 150 knots.
•- Inlet mounted pod(s).
- Lateral asymmetry greater than 300 pounds at
stations 1, 2, or 3.

I

Yaw departures can be minimized by avoiding abrupt
maneuvers in the 0.80 to 0.95 mach range with acenterline store, especially above 25,000 feet. Either
unload the aircraft prior to making roll commands orcommand only minimum required roll rate when
operating near the CAT I AOA limiter. Withcenterline store loadings having lateral asymmetries
greater than 300 pounds and inlet mounted pod(s),avoid abrupt aft stick commands above 25,000 feet.

IWARNINGt
The probability of a yaw departuresignificantly increases above 25,000
feet for CAT I loadings having acenterline store, a lateral asymmetry
greater than 300 pounds, and inlet
mounted pod(s). With these loadings,
moderate, full-aft-stick inputs at
35,000 feet and 300 knots have caused
yaw departures and spins.
A yaw departure may also occur with large
air-to-surface lateral asymmetries with the STORESCONFIG switch in the CAT III position. These
departures can be avoided if abrupt control• commands are not used with lateral asymmetries inexcess of 1500 pounds at station 3 (or equivalent).

I-

A yaw departure results in one of the following:
A self-recovery. The self-recovery may occur quickly(within approximately 5 seconds). If a pitch
departure follows the yaw departure, the self-recovery may require 10-20 seconds. Random andpossibly abrupt pitch, roll, and yaw rates may occur.

I

A deep stall.

- An upright spin.

6-8

Change 4

An upright pitch departure occurs when the AOA
exceeds the positive AOA/g limiter. Above 25 degrees •AOA, both horizontal tails are commanded to fulltrailing edge down by the pitch axis of the FLCS to try
to reduce the AOA. If AOA exceeds 29 degrees, theyaw rate limiter provides antispin commands to therudder, flaperons, and horizontal tails. Pilot roll andrudder commands are inhibited and pitch stickcommands are ineffective without use of the MPO
switch.

I

NOTE

m
m

LESS
Aircraft are more susceptible
thanaircraft to departure and deep
stall. In addition, LESS
aircraftroutinely operate near the aft CG
limits and require fuel management todecrease susceptibility to departure.

m

An inverted pitch departure occurs at negative AOAs
when the AOA significantly exceeds the negative glimiter. During the departure, the pitch axis of the
FLCS commands the horizontal tails to full trailingedge up to try to return AOA to the normal range.
Pitch stick commands are ineffective without use of
the MPO switch.

There are no automatic antispin features for an
inverted departure. Roll and rudder commands
should be released to allow any lateral-directionalmotion to subside. If a sustained yaw rate develops,
rudder opposite the yaw direction (opposite the turnneedle indication if no outside references areavailable) must be input to enable an effective
recovery through pitch rocking.

An upright or inverted pitch departure can occur •when the aircraft is flown to airspeeds below thatindicated by the low speed warning tone. Inverted
pitch departures usually result from inverted flightat high pitch attitudes and low airspeeds, such as
those often encountered by going over the top at too
slow an airspeed. An upright or inverted pitch
departure may also occur at any pitch attitude if10. 1 F-16A-l•

abrupt stick commands are made at airspeeds below
200 knots. Simultaneous abrupt roll and aft stickcommands are especially likely to cause an upright
pitch departure. However, the aircraft can be safely
flown to the AOA/g and roll limiters below 200 knots
with smooth commands.

I

An upright pitch departure may also result from ayaw departure. In addition, recovery from aninverted pitch departure may cause the aircraft topendulum into an upright departure and vice versa.

The likelihood of a pitch departure increases if theaircraft is configured with stores (especially 370-galIon fuel tanks), if the speedbrakes are opened, or if the
CG is near the aft limit.
Pitch departure characteristics are strongly influenced by the airspeed and aircraft rates present at
departure. A low airspeed departure (below 200
knots) may have relatively benign uncommanded
pitch, roll, and yaw motions. Higher airspeed
departures are usually very dramatic with large
uncommanded pitch, roll, and yaw motions which
may persist for 10 seconds or more.
It is possible for the AOA to briefly exceed the
AOAlg limiter without the aircraft experiencing
uncommanded motions. In this situation, the pilot's
only indication of a departure may be a failure of theaircraft to respond to control commands. These brief
departures typically self-recover within 5 seconds.

A pitch departure results in one of the following:- A self-recovery which occurs within 5-20 seconds.
Random and possibly abrupt pitch, roll, and yaw
rates may occur.

I-

A deep stall.

- A spin.
DEEP STALL

If the aircraft does not self-recover following adeparture, a deep stall may have developed. A deep
stall is an out-of-control flight condition in which theaircraft stabilizes at an AOA of approximately 60
degrees (upright) or -60 degrees (inverted) with low
yaw rates. The FLCS attempts to return AOA to thenormal range by commanding full horizontal taildeflection. However, the full horizontal tail deflection
is insufficient to return AOA to the controllable
range. The aircraft has entered a deep stall if the AOA
remains outside the controllable range. In a deep stall
the AOA indicator will be pegged at 32 or -5 degrees.
Recovery to controlled flight requires that the pilot
pitch rock the aircraft with the MPO switch in OVRD.

The MPO switch allows the pilot to override the FLCS
and to manually control the horizontal tails.

I

Susceptibility to and characteristics of a deep stall aredependent upon the aircraft CG and configuration.
LESS m Aircraft are susceptible to the deep stall ata more forward CG thanaircraft. An inverted deep
stall occurs at a more forward CG than an uprightone. The susceptibility to a deep stall increases as
the CG moves farther aft and if stores (especially fueltanks) are loaded.

m

Airspeed indications are erroneous in a deep stall andfluctuate between 0-150 knots. Altimeter indications
should be considered reliable; however, aircraft
oscillations may cause momentary stabilized or evenslightly increased altitude indications. Sink rate in a
deep stall is usually between 10,000 and 15,000 feetper minute. The normal load factor is approximately
Ig or -lg for upright and inverted deep stalls,
respectively.
Upright deep stalls may be very stable with little orno pitch, roll, and yaw motions or may be very
oscillatory with large pitch, roll, and yaw motions.
Generally, a clean configuration results in a deep stallwith a near wings-level pitching motion. If stores arebeing carried, especially a 300-gallon fuel tank or
370-gallon fuel tanks, the deep stall may be very
oscillatory, masking the pitch motions.
In an upright deep stall, the nose of the aircraft
usually oscillates ± 15 degrees about a slightly
nose-low pitch attitude. Pitch oscillations may be as
high as ± 40 degrees and normally reverse direction
approximately every 3 seconds. Roll reversals up to± 90 degrees from wings level may occur and the yaw
rate tends to cyclically reverse back and forth andmay be as high as 40 degrees per second. A slow net
heading change, usually to the left, may occur.

I

I

Inverted deep stalls may be either stable or highly
oscillatory in pitch, depending upon the CG. If there
is little or no pitch motion, the nose is slightly above
the horizon and the wings are generally level. If the
deep stall is oscillatory in pitch, the nose may oscillate
above and below the horizon by as much as ± 20
degrees. Yaw and roll oscillations in an inverted deep •stall are normally smaller than those during an
upright deep stall.
The aircraft is most likely to stabilize in a deep stall
if it does not self-recover within two postdeparture
pitch oscillations. The likelihood of a deep stall
occurring after a departure is also dependent upon
the aircraft CG and configuration. An inverted deep
stall can occur at a more forward CG than an uprightone. The likelihood of a deep stall developing after adeparture increases as the CG moves farther aft and
if stores (especially 370-gallon fuel tanks) are loaded.
Change 4

6-9

T.O.1F-16A-l
SPIN

A spin is a deep stall with a significant sustained yaw
rate in one direction (greater than 30 degrees per
second). The pitch, roll, and yaw oscillations
associated with a deep stall should not be confusedwith the continuous yaw rotation associated with aspin. AOA, airspeed indications, and altitude loss are
similar to those during deep stalls. Spins can be either
upright or inverted, although inverted spins aremuch less likely to occur than upright spins.
The yaw rate limiter is effective in preventing an
upright spin with all CAT I loadings. However,following a yaw departure above 25,000 feet, aircraftwith CAT I loadings that have all the following
characteristics may spin:• Centerline store.
• Inlet mounted pod(s).
• Lateral asymmetry greater than 300 pounds at
stations 1,2, or 3.
Upright spins following a yaw departure can bedisorienting. The initial portion of the spin ischaracterized by highly oscillatory pitch and rollmotions and a high yaw rate (70 to 100 degrees per
second). Initially, the aircraft spins roughly aroundthe aircraft's flight path at departure. As the spin
continues, the rotation axis eventually becomes
vertical. Very noticeable forward g (eyeballs out) andsideforces are present.
Yaw rate usually decreases to near zero within 10 to
25 seconds. The g forces decrease noticeably as yaw
rate decreases, which may give the sensation thatyaw rate is lower than it actually is. Use outsidereferences to determine when the yaw rate has
stopped. A recovery may occur after yaw rate has
decreased to near zero. If recovery does not occur, theaircraft has settled into a deep stall and pitch rocking
must be used to recover the aircraft to controlled
flight. However, because pitch rocking is less effective
when a yaw rate is present, pitch rocking should not
begin until yaw rotation stops or is minimized (ifaltitude permits).
A spin may also occur following any departure with a
CAT III loading that has a large lateral asymmetry
from air-to-surface stores. These spins may be fast,flat, and possibly unrecoverable.
An inverted spin can be caused by pilot rudder androll commands if they are not released following an
inverted departure.

6- 10

Change 4

For an inverted spin, rudder commands opposite theyaw direction (opposite the turn needle indication)
are required to stop the rotation. Large lateralasymmetries on CAT III loadings increase the I
potential for unrecoverable inverted spins.
RECOVERIES
Self-Recovery

Recovery from most departures is automatic,
requiring only release of the controls. Once thecontrols are released, self-recoveries usually occur
within the first two postdeparture pitch oscillations(10-20 seconds). Recovery is characterized by the nose
pitching down to a steep dive angle, increasing
airspeed, and AOA and sideslip returning to thenormal range. Some postdeparture yaw and rolloscillation may be evident, particularly if thedeparture was in yaw. To prevent departure reentry,
the airspeed should be allowed to increase to 200
knots prior to dive recovery. Flight control failure
indications may be present after any departure.
Deep Stall Recovery

The aircraft should be allowed the opportunity to
self-recover if altitude permits. Initiating pitch
rocking too soon should be avoided because it canaggravate postdeparture roll and yaw motions, which
can significantly lengthen recovery time.
If recovery is not apparent after two postdeparture
pitch oscillations 00-20 seconds) or if altitude is afactor, the aircraft must be rocked out of a deep stall.
To pitch rock the aircraft, the MPO switch must be
firmly held in OVRn until recovery is complete.
Recognize any pitching motions and begin stick
cycling in-phase with these motions. If no pitch
motions are apparent, an abrupt maximum commandstick input to pitch the nose away from the ground(full aft stick for an upright deep stall, full forward for
an inverted deep stall) will reverse the horizontal taildeflection and should generate a noticeable pitch
rate. The best indicator for timing stick cycles is thenose position relative to the horizon, or if no outsidereferences are available, the ADI may be useful. In an
upright deep stall, an aft stick command increases
the AOA and pitch angle. When the nose reaches its
highest point and reverses direction, a full forward
stick command reinforces the nosedown pitch rate.
One complete pitch rocking cycle takes approximately 6 seconds, during which time the aircraft descends
1000 to 1500 feet.

•I
I

T. O. 1F- 16A- 1
If the nosedown pitch rate is large enough, theupright deep stall is broken and AOA returns to thenormal range (below +25 degrees). If the nosedown
pitch rate is insufficient, the nose stops its downward
motion and either begins to rise or stabilizes.
Promptly reapply the full aft stick command when thenose reverses or after 2-3 seconds if reversal is notapparent. Holding full forward stick more than 2-3
seconds with the nose stabilized can generate a rapid
yaw rate and delay recovery. If the aircraft has notrecovered after 2-3 seconds or if a yaw rate develops,
reapply full aft stick and complete another in-phasepitch rocking cycle.

•

Just prior to breaking the deep stall, the nosedown
pitch rate may decrease or even stop. Unless the nose
either definitely starts back up or stabilizes for longer
than 2-3 seconds, another pitchup cycle should not be
started. On the pitchdown cycle, the stick should beheld full forward (upright) or aft (inverted) while thenose hesitates as the stall breaks. Following a shorthesitation (less than 2-3 seconds), the nose continues
down to near vertical. There is frequently a distinct, lowmagnitude airframe shudder which occurs as the stallis breaking. This shudder is a favorable indication thatrecovery is occurring.
Once the deep stall is broken, aggressive pilot
commands are usually required to stop the pitch rate ina steep dive. If possible, find and track a feature on the
ground. Maintain firm pressure on the MPO switch
until airspeed reaches 200 knots. If a transition to an
opposite AOA deep stall does occur, reinforce thealready present pitch motion with the MPO stillengaged and recovery should be rapid. Transitions
during upright deep stall recoveries are most likely tooccur with 370-gallon fuel tank configurations.
Recovery to controlled flight is recognized by a steep
pitch attitude (usually within 30 degrees of vertical),pitch rate stopping, and AOA in the normal range (-5
to +25 degrees). As airspeed increases above 200
knots release the MPO switch, maintain neutral rolland yaw commands, and recover from the resultingdive.

300-gallon fuel tank) or with 370-gallon fuel tanks
may require more pitch rocking cycles for recovery.
These loadings usually have more oscillatory deep
stalls, and the roll and yaw motions make it moredifficult to determine proper stick cycling. Pitch
attitude is still the best indication for proper stick
cycling.
During upright deep stalls with a centerline store,
particularly a 300-gallon fuel tank, the aircraft tends
to roll and yaw right while pitching up, and roll andyaw left while pitching down. During deep stalls with
370-gallon fuel tanks, the aircraft nose motion
appears triangular. This motion is characterized by aroll and yaw right while pitching up, followed by apitch down, a hesitation, and a yaw to the left. These
alternating yaw oscillations should not be confusedwith a sustained yaw rotation in one direction
indicating the aircraft is in a spin.
An inverted deep stall recovery is similar to an
upright recovery. However, any sustained yaw rate
must be controlled by using opposite rudder. If thisyaw rate is not minimized, pitch rocking will be lesseffective and may not be sufficient to recover theaircraft. Only enough rudder to stop the yaw rateshould be applied. Once the yaw rate is stopped, pitch
rocking effectiveness is increased.

To recover from an inverted deep stall, position the
MPO switch to OVRD and begin stick cycling
in-phase with pitch motions. If no pitch motions areapparent, make the first pitch input away from theground by pushing full forward and monitor pitch
motion. One or two pitch rocking cycles are usually
sufficient to recover from an inverted deep stall.

I
I

During recovery from an inverted deep stall, atransition to an upright deep stall is likely to occur if
large pitch motions are present during inverted pitch
rocking and the MPO switch is released too early.
Early release of the MPO switch reduces the •horizontal tail authority available and delays
recovery. The MPO switch must be held in the OVRD •position until the deep stall is positively broken as
evidenced by the pitch rate stopping, AOA in normal
range (-5 to +25 degrees), and a steep pitch attitude(usually within 30 degrees of the vertical). As
airspeed increases above 200 knots release the MPO
switch, maintain neutral roll and yaw commands,
and recover from the resulting dive.

Stick commands during pitch rocking should be
abrupt and full command, and should reverse after
pitch motion reverses. Stick commands that are notabrupt and full command may not be effective. Rapid
fore and aft cycling of the stick out of phase with theaircraft motion is also not effective.

Spin Recovery

The number of pitch rocking cycles required forrecovery from an upright deep stall is dependent on
aircraft configuration. Generally, recovery occurs in
one or two pitch rocking cycles. However, configurations with a centerline store (particularly a

To recover from a spin, yaw rate must be stopped orminimized before the aircraft can be recovered. Dueto large nose-up moments caused by the inertial
properties of the aircraft and decreased horizontal
tail pitch effectiveness with sideslip, attempts toChange 4

6- 11

10. 1 F-16A-l
pitch rock out of a spin are usually not effective. Pitch
rocking during a spin is also likely to aggravate rolland yaw oscillations, which make recovery moredifficult.
The large forward g (eyeballs out) and sideforces
present during the initial portion of an upright spindecrease noticeably as yaw rate decreases, which may
give the sensation that yaw rate is lower than it
actually is. Use outside references to determine when
the yaw rotation has stopped. Pitch rocking should
not begin until the yaw rotation stops or is minimized.
Pitch rocking with a steady yaw rate greater than 30
degrees per second may prevent recovery.
Waiting for yaw rotation to stop or minimize may
require from 20 to 30 seconds. When the yaw rotation
stops or is minimized, the aircraft will either
self-recover or will settle into an upright deep stall. If
recovery is not apparent after yaw rate has stopped oris minimized, perform the appropriate recovery
procedures described in the Deep Stall Recovery
section.
Pitch, roll, and yaw oscillations associated with adeep stall should not be confused with the continuous
yaw rotation associated with a spin. Once thecontinuous yaw rotation of a spin has been arrested
and pitch rocking has begun, pitch rocking should
continue until a recovery is achieved.
Upright spins with CAT III loadings that have large
lateral asymmetries from air-to-surface stores may
be fast, flat, and possibly unrecoverable. There is an
option to jettison an asymmetric store as a last ditch
effort in case of a fast, flat spin. However,aircraft-store collision may occur.
In an inverted spin, yaw rate must be reduced byusing opposite rudder. If yaw rate is not minimized,
pitch rocking will be less effective and may not be
sufficient to recover the aircraft.
If recovery is not apparent after yaw rate has stopped
or is minimized in an inverted spin, the aircraft has
settled into an inverted deep stall. Perform theinverted deep stall recovery procedures.
ENGINE OPERATIONOUT-OF-CONTROLDURINGDEPARTURES/

Departures at high altitude may result in an engine

I stall or stagnation. If in AB during an out-of-control

Isituation, retard the throttle to MIL. If at MIL or
below, do not move the throttle. Do not advance thethrottle until beginning the dive recovery.
6-12

Change 4

If the engine stalls during a departure or
out-of-control situation, refer to ENGINE STALLS,
Section III, after recovery is complete. The engine
should be left running during an out-of-control
situation to insure adequate hydraulic pressure toflight control surfaces for recovery. Recover from theout-of-control condition; then concentrate on theengine. If the engine does not recover, it must be shut
down and restarted.

DEGRADED FLIGHT CONTROLSLEADING EDGE FLAPS LOCKED (SYMMETRIC)

Flight characteristics for landing and low AOA
maneuvering are not significantly affected by locked
LEF's. At high airspeeds, LEF's locked down cause
increased buffet. At high AOA, LEF's locked up
reduce stability, increase departure susceptibility
significantly, and cause increased buffet. Above 16-18
degrees AOA, an abrupt yaw departure may occur,
producing an uncommanded roll with little or noforewarning. Do not exceed 12 degrees AOA with the
LEF's inoperative. Locked down LEF's significantly
reduce cruise range. During landing, floating may
also be noticeable if LEF's are locked at or near full
down. The aircraft may float, sink rate may decrease,and a slight forward stick pressure may be needed to
fly through the ground effect.
TRAILING EDGE FLAPS (FLAPERONS)

The flaperons function as TEF's when lowered
simultaneously. The wing flap system is designed toprevent split-flap operation. If a malfunction causes
either flaperon to lock, then neither TEF lowers. If the
TEF's were extended at the time of the malfunction,
both TEF's retract before the malfunctioning surface
is locked. The increase in descent rate associated with
TEF retraction at final approach AOA is mild andeasily controlled. Landing without TEF's presents no
significant control problems.
LOCKED CONTROL SURFACES

Flaperon

The aircraft can be landed safely with one locked
flaperon; however, there are no TEF's available androll control power is significantly reduced.
Horizontal Tail

Simulator evaluation indicates that the aircraft canbe landed safely with one locked horizontal tail. While
pitch commands may cause a rolling tendency,
flaperon authority is adequate to counteract the roll
inputs.

I

T. O. 1F- 16A- 1

Rudder

FLAPERON

Simulator evaluation indicates that the aircraft canbe flown and landed safely with a locked rudder.

Separation of a flaperon from the wing (flaperon still
attached to the ISA in the fuselage) causes theoutboard end of the flaperon surface to rotate
upwards, towards the fuselage. This may be
accompanied by a roll transient. Adequate roll stickauthority should be available to counter the effect andmaintain control. With a flaperon surface separated
from one wing, landing should be made without
extending the TEF on the other wing (use alternate
landing gear extension, leave LG handle UP, andselect brakes CHAN 2). Hydraulic system pressures
should also be monitored. Upward rotation of aseparated-from-the-wing flaperon surface has caused
ISA movement that resulted in a leakage failure of
hydraulic system A.

STANDBY GAINSWhen operating on standby gains, aircraft response
is normal at low AOA. Because the LEF's are at zero
degrees (LG handle in UP and ALT FLAPS switch in
NORM) with a dual air data failure, buffet anddeparture susceptibility will be increased at higher
AOA (above 18 degrees). At flight conditions
higher/slower than the flxed gain conditions, aircraft
response is more sluggish requiring larger control
inputs for a given response. Landing the aircraft
should present no special problems.

ONE HYDRAULIC SYSTEM
Flight characteristics with one hydraulic systemshould be normal unless extremely large, rapid
control surface deflections are commanded. Under
these conditions, the hydraulic flow rate from the one
system may be inadequate which slows down control
surface movement rates and possibly causes sluggish
aircraft response.

SPEEDBRAKES
Speedbrakes may stick fully open or open asymmetrically. If a yawing moment is noted when the
speedbrakes are opened, close the speedbrakes. If the
speedbrakes fail to close, a significant increase in
drag results. Fully opened speedbrakes signiflcantly
reduce cruise range.

AIRCRAFT DAMAGE
Procedures for recovery of an aircraft with damage
will depend on the type and extent of the damage.
The following paragraphs provide general information based on analysis of past mishaps.

HORIZONTAL TAIL
Loss or partial loss of one horizon tal tail surface will
not result in an uncontrollable aircraft except for
certain combinations of flight conditions (machgreater than 0.80 and altitude less than 15,000 feet),aircraft loading (heavy CAT 111), and CG at or near theaft limit. Avoid abrupt maneuvering and maintainairspeed between 200 and 300 knots un til landing
approach. Place the STORES CONFIG switch to CAT
III if in CAT I. Some roll stick or roll trim may berequired to maintain wings level. Use a maximum of
11 degrees AOA during fmal approach. Aftertouchdown, lower nose to runway as soon as practical.

RUDDER AND VENTRAL FINS
Loss or partial loss of the rudder or a ventral fln will
not result in controllability problems unless theaircraft is above 1.5 mach. Ifloss or partial loss of therudder or a ventral fln occurs, avoid abrupt
maneuvering and reduce speed to subsonic, if
supersonic. Place the STORES CONFIG switch to
CAT III if in CAT I. During landing with loss or partial
loss of the rudder, lower the nose to the runway assoon as practical so that NWS can be used as requiredfor directional control.

LEADING EDGE FLAPS
Damage to the LEF's may result in locked LEF's. If
the LEF's have not automatically locked, they should
be manually locked. Refer to Leading Edge Flaps
Locked (Symmetric), this section for a discussion offlight characteristics.

WING
Loss of a portion of the wing produces a rolling motion
towards the damaged wing. The capability to stop thisroll depends on airspeed, g, and the amount of wing
surface lost. Higher airspeeds (above 250 knots) andlow g (less than 2) are essential for maximizing theamount of roll control authority available. If theaircraft is in a dive when a portion of the wing surface
is lost, apply roll stick force to stabilize the aircraft inroll before applying g to recover from the dive(altitude permitting). Use the lowest g level practicalto recover from the dive. As g is increased, additionalroll stick force is required to maintain wings level.
Depending on the amount of wing surface lost, theminimum speed to maintain adequate roll control
could be well above 200 knots.
Change 8

6-13

T.O. 1F- 16A-1

RADOME
Loss of the radome results in loss of two air data
sources (pitot probe) and two AOA sources (AOA
transmitters); however, loss of the radome does notnecessarily cause the aircraft to be immediately
uncontrollable. FOD to the engine and other damage
to the aircraft may also occur as the radome departs.
Loss of two air data sources will affect handling
qualities. If the two airspeed sources from the pitot
probe go to erroneously low values at the same time,
the remaining good source from the fuselage air dataprobe will be considered failed (Le. a single failure).
In this case the FLCS will not switch to standby
gains. Gain scheduling will be based on theerroneously low airspeed; thus, the aircraft will
become increasingly pitch sensitive as airspeed isincreased. If the two airspeed sources are lost in amanner that the FLCS recognizes as two failures,
then the FLCS reverts to standby gains. In standby
gains, the opposite problem occurs. With the landing
gear up at slow speeds, the aircraft will be sluggish
compared to handling qualities prior to the radomeloss. It may seem as though the aircraft isn't
responding to stick inputs if sufficient response time
is not allowed.
Loss of two AOA sources may not be detected as adual failure. The end result to the FLCS is a flXed
value of 13.6 degrees AOA (Le. no AOA limiter). The
more serious possibility associated with loss of both
AOA transmitters is exposed wiring. All four
branches of FLCS power go to each AOA transmitter.
If this wiring shorts against structure, FLCS powerin multiple branches can be degraded to the pointthat aircraft control is not possible.
If the radome is lost, attempt to attain 1 g level flightat 275 - 300 knots and place the landing gear handle
down. Airspeed and AOA indications won't be
accurate. Use a chase aircraft to help establish speed
and keep AOA below 12 degrees. If controllability
appears to be degrading (Le. shorting of AOA
transmitter wiring), eject before complete loss ofcontrol occurs. The engine control system may
default to a fixed mach number signal of 0.8; if thishappens, thrust in idle will be higher than normal.

DIVE RECOVERY
Refer to figure 6-1. Dive recovery capability is given
as altitude lost during pUllout and is a function of
pullout load factor, dive angle, true airspeed, andFLCS limiting. Plots to convert indicated airspeed ormach number into true airspeed are provided on thechart. Dive recovery during constant load factor
pullout may be on the AOA limiter prior to recovery,
6- 14

Change 8

under certain initial conditions, as airspeed isreduced. Dive recovery capability at constant loadfactor is nearly independent of store drag. The diverecovery chart is applicable to GW's between 20,00030,000 pounds. The dive recovery chart becomesincreasingly conservative for GW's less than 25,000
pounds and decreasingly conservative for GW's
greater than 25,000 pounds. Increase altitude lost
during full aft stick pUllout by 4 percent for each 1000
pounds in excess of 25,000 pounds GW if initial diveangle is ~ 45 degrees and initial airspeed is less than
500 knots.
For a constant g pullout, use the greater of constantg or limiter pullout altitude lost.
NOTE
The dive recovery chart is based on an
idle thrust, wings level, speedbrakes fully
open recovery. However, if airspeed is
below 350 knots, altitude loss is minimized by selecting/maintaining MIUAB
thrust and closing speedbrakes. If
airspeed is 350 knots or above, selecting!
maintaining idle thrust and opening
speedbrakes minimize altitude loss. In
either case, best dive recovery performance is obtained by making an ADI
referenced wings level pull.
SAMPLE PROBLEM:

•

GW

A.

KIAS

B.

Initial altitude= 25,000 pounds= 300= 15,000 feetC.

Temperature

= O°C

D.

KTAS

E.

Mach number

F.

Dive angle

G.

Category/KIAS= 381= 0.59= 60 degrees= CAT 1/300

H.

Altitude lost during
limiter pullout= 2130 feet

Altitude lost during
limiter pullout= 2130 feet

I.

J.

Maximum pullout loadfactor available= 5.8gK.

Pullout load factor

= 3g

L.

Altitude lost during a
3g pullout= 4150 feet10. 1F-16A- 1

SECTION VII
ADVERSE WEATHER OPERATIONTABLE OF CONTENTS

Page

Introduction .............................
Instrument Flight Procedures .............
Holding ...............................
Penetration. . . . . . . . . . . . . . . . . . . . . . . . . . . .
Instrument Pattern/Approaches. . . . . . . . .
Missed Approach . . . . . . . . . . . . . . . . . . . . . . .
Turbulence and Thunderstorms . . . . . . . . . . . .
Cold Weather Operation ..................
Before Entering Cockpit ................
Before Starting Engine .................
Starting Engine . . . . . . . . . . . . . . . . . . . . . . . .
After Engine Start .....................
Taxi.. . . . . . . . . . .. . . . . . . . . . . ... . .. . . . . . .
Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
In Flight ..............................
Landing in Icy or Wet Conditions ..........
Hot Weather and Desert Ground
Operation .............................
Volcanic Ash Operation ...................
Ground Operations . . . . . . . . . . . . . . . . . . . . .
In-Flight Operations ...................

7-1
7-1
7-2
7-2
7-2
7-2
7-2
7-5
7-5
7-5
7-5
7-5
7-6
7-6
7-6
7-6
7-7
7-7
7-7
7-8

INTRODUCTION
This section contains information and proceduresthat affect operation of the aircraft in adverse
weather and climatic conditions and which differ
from the normal procedures in Section II.

INSTRUMENT FLIGHT PROCEDURES
IWARNING,• The HUD should not be used as the solereference for instrument flight due tothe lack of adequate failure warning
but should be cross-checked with theprimarylbasic instruments.
• A delayed selection of ILSfrCN or
ILSINAV until the aircraft is nearly on
the ILS glide slope may cause the flight
director circle to be positioned incorrectly(full up or full down) for up to 90 seconds.
If the flight director circle is positioned
incorrectly when an ILS mode isselected, move the INSTR MODE knob
to TCN or NAV, then back to the desired
ILS mode. This action enables the flight
director to operate properly.
• ~ LN -93 RLG's may develop verticalvelocity error during extended climbs
and/or descents (over 3 minutes) at
vertical speeds of 2000-6000 fpm. Thiserror causes erroneous positioning of theFPM, an inaccurate RUD verticalvelocity indication, and degraded ILS
command steering glidepath information,
ground collision warning, LESS ~GCAS, LESS!MJ] DTS, FCR, and weapondelivery functions. This error may persist
for as long as 5 minutes after leveling off.

I

• ~ Certain RLG INU s may indicate afailure of the altitude and verticalvelocity data during supersonic flightor high rates of climb/descent. Thiscondition will blank the HUD FPM anddisable ground collision warning functions with no other failure indications
or advisories.

• It is possible for the displayed ADI
and/or HUD attitude to be in error withno ADI OFF or AUX warning flags inview and without an INS or HUD PFL.
Displayed RSI and/or HUD headings
may also be in error with no RSI OFF
or ADI AUX warning flags in view andwithout an INS or RUD PFL. Momentary warning flags may indicate
impending failure. To detect thesefailures and maintain proper flightorientation, basic and backup instruments shall be cross-checked.
• With DRIFT C/O selected and the FCC
off/inoperative, the FPM incorrectly
depicts the aircraft flightpath duringbanked flight.
• With certain failures of the INS with
ILS selected, a fixed aircraft reference
symbol is displayed at zero degrees
azimuth, 11 degrees below the boresight
symbol. This symbol is for ILS deviationreference in conjunction with thehorizontal and vertical deviation bars.
The aircraft reference symbol should
not be used for attitude reference.
Change 12

7-1

T.O. 1 F-16A-l

I WARNING I
Avoid touching the canopy transparency,
canopy frame or placing arms on the
canopy body positioning handles during
IMC. Contact may produce severe shock
as a result of static discharge.
NOTE• When rotating the HSI course settingfrom an eight-unit digit to a nine-unit
digit (e.g., 018 to 019), the tens digit
may rotate prematurely causing a
10-degree reading error (e.g., 029).
Cross-check counter setting withcourse arrow reading to insure propercourse setting.
• Interference effects from the scanningof the FCR antenna may degrade ILS
glide slope reception range and causeintermittent loss of signal indications.
The interference can be eliminated bystowing the FCR antenna (select STBY
or depress OVRD OSB). If FCR use isrequired, interference may be reduced
or eliminated by positioning the FCR
antenna tilt as high as practical,
preferably above the horizon.
HOLDING

Holding airspeed is 200-250 knots (maximum
endurance airspeed recommended).
PENETRATION

Penetrations are normally flown at 300 knots,
speedbrakes as required, and throttle at 70 percent
rpm minimum.
INSTRUMENT PATTERN/APPROACHES

Refer to figure 7-1 for a typical TACAN holding
pattern, penetration, and approach. Refer to figure
7-2 for a typical GCA. Instrument patterns arenormally flown at 200-250 knots clean. Approaching
final, lower the LG, slow the aircraft, and fly final
approach at 13 degrees AOA maximum.
MISSED APPROACH

Advance throttle as required, close speedbrakes, andretract LG after a positive climb is established. Adjust throttle to maintain between 200-250 knots.
Pitch transients resulting from LG and TEF changes
are mild and require minimum control compensation.
7 -2

Change 13

TURBULENCE AND THUNDERSTORMS
Avoid flight in turbulent air, hailstorms, andthunderstorms. There is a high probability of damage
to airframe and components from impact ice, hail, andlightning. Thunderstorm penetration airspeed is 300
knots or optimum cruise airspeed, whichever is lower.
At high airspeeds, personal discomfort and structuralstress are greater. At slower airspeeds, controllability
is reduced and inlet airflow distortion due toturbulence may cause compressor stall and/or engine
stagnation.
The GM mode of the radar can be used as an aid innavigation between or around storm cells. Refer toT.O. 1F-16A-34-1-1, M T.O. IF-16A-34-1-3 or ~T.O. 1F-16A-34-1-4 for expanded information. If
entry into adverse weather cannot be avoided, thefollowing procedures should be used:
1.

PROBE HEAT switch - Check PROBE HEAT.

2.

FLOOD CONSOLES knob - HIGH INT.

3.

ANTI ICE switch - ON.

4.

Airspeed - 300 knots or optimum cruise,
whichever is lower.
NOTE• Severe turbulence causes variations inairspeed and altitude. Do not change
throttle setting except for extreme
airspeed variations.
• An extremely loud screeching noise
may be heard in the headset while
flying in cirrus clouds or in the vicinityof thunderstorms. The noise may beeliminated by turning the UHF or VHF
radio off, by turning the volume(s)down, or by changing UHF antenna
positions.
• When flying in heavy rain, water tends
to be aerodynamically held on theforward portion of the canopy. At
higher airspeeds, this condition may
obscure visibility as much as 30
degrees back on each side of the
canopy. On final approach, the water isgenerally confined to the position of
the canopy immediately in front of the
HUD. It may be necessary to look out
the sides of the canopy to acquire therunw~y and to flare and land theaircraft.

IT.O. 1F-16A-l

TACAN Holding, Penetration, and Approach(Typical)

HOLDING (ALL ALTITUDES)MISSED APPROACH

AIRSPEED - 200-250 KNOTS• THROTTLE - AS REQUIRED• SPEEDBRAKES - CLOSE• LG -UP• AIRSPEED - ACCELERATETO 200-250 KNOTSFINALAPPROACH FIXTACAN APPROACH• LG -DN• THROTTLE - AS REQUIRED• AOA -13 DEGREES (MAX)• SPEEDBRAKES - OPEN

PENETRATION DESCENT• AIRSPEED - 300 KNOTS• SPEEDBRAKES - AS REQUIRED• THROTTLE - 70 PERCENT (MIN)• ALTIMETER -SET

NOTE:

I

A typical straight-in penetration and approach require approximately 400 pounds of fuel for a drag index of 0
and 600 pounds of fuel for a drag index of 200.

1F-16A-1-1192Ae

Figure 7-1.

Change 1

7-3

T.O. 1 F-16A-l

GCA (Typical)• LG - AS REQUIRED• SPEEDBRAKES - AS I---=======~~;===~~~~~~;:::-~~~s;:~~==-=-==:==s::::=====---REQUIRED~~ -:::::: ~ _~_ ~-• AOA -13 DEGREES_ -:z---~ ~-(MAX) (LG - ON)~~• LG - CHECK ON• SPEEDBRAKES - OPEN
• AOA - 13 DEGREES(MAX)• THROTILE- AS REQUIRED• SPEEDBRAKES - CLOSE• LG - UP• AIRSPEED - ACCELERATETO 200-250 KNOTS

NOTE:

I

A typical GCA pattern requires approximately 400 pounds of fuel for a drag index of 0
and 700 pounds of fuel for a drag index of 200.

1F-18A-1-1193Ae

Figure 7-2.

7-4

Change 1

T.O. 1 F-16A-l

COLD WEATHER OPERATION

Engine operation under the following
conditions may result in engine damage due to icing:• Ambient temperatures between
20°F (-7°C) and 45°F (7°C) with
precipitation (rain, fog, sleet, orsnow).
• Dewpoint within 9°F (5°C) of
ambient temperatures between
25°F (-4°C) and 45°F (7°C).
• Ambient temperature below 45°F
(7°C) with standing water or a
mixture of water with ice or snow
within the immediate proximity ofthe engine inlet.
BEFORE ENTERING COCKPIT

All accumulated ice and snow must be removed fromthe aircraft before flight is attempted. Insure thatwater does not accumulate on control surfaces orother critical areas where refreezing may cause
damage or binding.

the FLCS battery test, external power must be
removed by placing the MAIN PWR switch to BATT.
Several attempts should be made to test the FLCS
batteries prior to aborting the aircraft as the current
flow in the batteries during the test may heat thebatteries sufficiently to produce the required power topass the FLCS battery test. Holding the FLCS PWRTEST switch in TEST for 5-10 seconds each time also
may improve the likelihood of passing the test.
STARTING ENGINE

If the aircraft is serviced with MIL-H-5606 hydraulicfluid and the aircraft has cold soaked for more than 1hour at temperatures below -40°F (-40°C), do notstart the JFS until ambient temperature increases toabove -40°F (-40°C) for at least 2 hours or until theengine bay is preheated. For temperatures above-40°F (-40°C), refer to JET FUEL STARTER
LIMITS, Section V
If the aircraft is serviced with MIL-H-83282
hydraulic fluid and the aircraft has cold soaked formore than 1 hour at temperatures below -20°F(-29°C), do not start the JFS until ambient
temperature increases above -20°F (-29°C) for atleast 2 hours or until the engine bay is preheated. For
temperatures above -20°F (-29°C), refer to JET
FUEL STARTER LIMITS, Section V
During cold start, oil pressure may be 100 psi for up
to 1 minute.
AFTER ENGINE START

Do not permit ice to be chipped or
scraped away.
BEFORE STARTING ENGINE

Extreme cold temperature may require cockpit
preheating to ease operation of rotary-type switches.
[(] The canopy may not latch on battery power alone.
Start the engine with the canopy closed as much as
possible.
If there is visible moisture and ambient temperature
is 45°F (7°C) or less, place the ANTI ICE switch to
ON. This reduces ice buildup on the engine front face,eliminates ice on the heated inlet strut, and reduces
the possibility of ice ingestion.
When the FLCS batteries have been cold soaked, they
may not produce sufficient power to pass FLCS
battery test. In this case, the use of external power
will operate the battery heaters and improve theperformance of the batteries. Prior to accomplishing

EPU fuel quantity can indicate as low as 90 percent
at temperature below 40°F (4°C).
For rapid cockpit warming, position the TEMP knob
to the desired MAN WARM range. Position the
RADAR to off and the DEFOG lever as required toclear fogging. Mter the cockpit reaches a comfortabletemperature, select a setting within the AUTO range.
If the engine was started with the canopy unlatched,wait approximately 10 minutes to warm the canopy
before fully closing it.
If probe icing is evident or suspected, turn the PROBEHEAT switch to PROBE HEAT at least 2 minutes
prior to accomplishing the FLCS self-test.

I WARNING I
If probe heat is or has been on, heat inprobes may be sufficient to cause
injury if touched.

7-5

T. O. 1F- 16A- 1
TAKEOFF

If the aircraft has cold soaked attemperatures below -20°F (-29°C),
repeated brake applications (25-30)
may be required before the brakes
work effectively.

If the aircraft has cold soaked at temperatures below-20°F (-29°C), LG retraction times may besignificantly increased. In addition, the nose geardoor may fail to close. If the nose gear door can bevisually confirmed as the only LG component that has
failed to retractJclose, then up to two extend/retract
cycles can be made in an attempt to achieve a normal
LG up condition. Observe LG extended or in transit
limitations, Section V.

TAXI

IN FLIGHT

To avoid brake icing, do not taxi in deep water, slush,or deep snow. When taxiing on ice or hard packed
snow, NWS may not be completely effective. Use a
combination of NWS and differential braking tomaintain directional control. Taxi at a safe speed
considering surface condition, GW, slope, and thrust.
If the aircraft has cold soaked at temperatures below-20°F (-29°C), the NWS may initially be sluggish,
but controllable.

Flight in areas of icing should be avoided wheneverpossible. If icing conditions are anticipated or cannot
be avoided, tum ANTI ICE switch to ON ~ andPROBE HEAT switch to PROBE HEAT. Frequently
check the aircraft leading edges for indication of ice
buildup. Make all throttle movements slower thannormal when in potential icing conditions to reduce
possibility of engine stalls and/or stagnation.
Consider diverting to an alternate field if required to
avoid icing conditions.

I WARNING I
Probe internal icing must be suspectedanytime the aircraft has been exposed
to near or below freezing conditions on
the ground. Internal icing may bedifficult to see and may remain present
even when current conditions do notappear conducive to ice formation.
Turn probe heat on at least 2 minutes
prior to takeoff anytime icing of probes
is possible.

IWARNING

1• If probe heat is inoperative, icing of the
AOA probes may cause loss of control
which may vary from mild pitch
oscillations to moderate pitchup or
severe pitchdown.

m• LESS
If a nonresettable PNEU flag
appears on the altimeter and the
CADC caution light is not illuminated,
the right AOA probe heater circuit
breaker may be open. Avoid areas ofknown or suspected icing conditions.

LANDING IN ICY OR WET CONDITIONS

If unable to control taxi speed or
direction, immediately shut down theengine.

NOTE

Mter cold soaking at temperatures
below O°F (-18°C), be alert for flat
MLG struts.
7-6

Change 8

Icy or wet runway conditions may pose severe
problems in directional control and braking effectiveness due to hydroplaning. Although possible, totalhydroplaning is not expected below 130 knots
groundspeed. Partial hydroplaning can occur tovarying degrees below 130 knots. Once hydroplaning
occurs, it can continue to speeds well below the onset
speed. Wheel spinup must occur to permit normal
antis kid braking. Hydroplaning can prevent wheel
spinup and can occur on runways which only appear
damp if heavy braking is applied at high speeds.
Hydroplaning tendency increases with water depth
and with smooth runway surfaces such as rubber
deposits or paint stripes.

T.O. 1 F-16A-l

I LESS 00 Approach and touchdown are the same as for(FE When stopping distance is critical, continuous

short field landing on dry runway. Immediately after
touchdown, make a deliberate effort to be sure brakes
are not applied while using the rudder. [N;J INQJ Deploy
the drag chute (if required) immediately after
touchdown. Use two-point aerodynamic braking until
approximately 80 knots; then fly the nosewheel to therunway. Maximum effective two-point aerodynamic
braking is achieved at 13 degrees AOA. After the
nosewheel is on the runway, open the speedbrakes
fully and maintain full aft stick for maximum
three-point aerodynamic braking and wheel braking
effectiveness. Test for braking effectiveness before
using full continuous braking by momentarily
depressing pedals, fully releasing pedals for at least
one-half second, and then depressing pedals again.
This technique gives the wheels a better opportunity
to spin up if hydroplaning conditions exist. If braking
effectiveness is not felt, continue to pump brakes as
speed decreases, making sure pedals are momentarily fully released between applications. Use continuous braking after braking effectiveness is felt. As
speed decreases, the antiskid system will increase
deceleration accordingly.

maximum wheel braking should be initiated in the
two-point attitude. Wheel braking effectiveness athigh speeds is very low compared to two-point
aerodynamic braking effectiveness. Low deceleration
at high speed may be mistakenly interpreted as abrake or anti-skid failure. When the wheel brakes
become effective, the nose will automatically lower.
Do not hesitate to lower the hook, if required.

I LESS (FE When stopping distance is critical, wheel
braking should be initiated in the two-point attitude
when below 100 knots. Wheel braking effectivenessat high speeds is very low compared to two-point
aerodynamic braking effectiveness. Low deceleration
at high speed may be mistakenly interpreted as abrake or anti-skid failure. If braking effectiveness oranti-skid cycling is not perceived, release brakes
momentarily and then reapply brakes. When thewheel brakes become effective, the nose willautomatically lower. Do not hesitate to lower thehook, if required.

I

LESS (FE If crosswinds preclude maintaining
two-point aerodynamic braking, test for braking
effectiveness as previously discussed before using fullcontinuous braking. Continue to pump the brakes
until braking effectiveness is felt and speed is below
100 knots.

(FE Approach and touchdown are the same as for ashort field landing on a dry runway. Immediately
after touchdown, make a deliberate effort to be surebrakes are not applied while using the rudder. Use
two-point aerodynamic braking until approximately
80 knots; then fly the nosewheel to the runway.
Maximum effective two-point aerodynamic braking isachieved at 13 degrees AOA. Mter the nosewheel is on
the runway, open the speedbrakes fully and maintainfull aft stick for maximum three-point aerodynamic
braking and wheel braking effectiveness. When
wheel brakes are used, they should be continuously
applied.

• LESS (FE Continuous wheel braking
above 100 knots is not recommended.
Hydroplaning may prevent spinup ofboth MLG wheels and wheel brakes
may become operative without antiskid
protection. Locked wheels and subsequent blown tires can occur.

• Rubber deposits on the last 2000 feet ofa wet runway make directional control adifficult problem even at very low speeds.
Braking should be started in sufficient
time to avoid excessive braking on thelast portion of the runway.

HOT WEATHER AND DESERT GROUNDOPERATION
Hot weather and desert ground operation requirethat added precautions be taken against damage
from dust, sand, and high temperatures. Particularattention should be given to those components andsystems (engine, fuel, oil, hydraulic, pitot-static, etc.)
which are susceptible to contamination, malfunction,
or damage from sand and dust. Inspect the pistons on
the LG and have them cleaned as required. Check theengine inlet duct for sand accumulation. Duringconditions of blowing sand and dust, the canopy
should be closed and sealed and all protective covers
installed when the aircraft is not in use.
In hot, humid conditions, fogging of the exterior
canopy surface after flight may reduce visibility to thepoint where the canopy must be opened prior totaxiing. Stow all equipment prior to opening the
canopy.

VOLCANIC ASH OPERATIONGROUND OPERATIONS

Modified ground operations on an airfield which has
experienced volcanic ash fallout are required evenafter cleanup is complete.
Change 8

7-7

T. O. 1F- 16A- 1

For preflight, carefully inspect the following areas forvolcanic dust:• ECS ram inlet ducts.
• Engine inlet.
• LG strut chrome surface.

IN-FLIGHT OPERATIONSFlight in a volcanic ash environment is extremely
hazardous. Airborne radar does not detect ash clouds
which visually are easily mistaken for normal clouds.
Some indications of volcanic ash cloud penetrationare:• Acrid odor.

• Pitot tube, air data probes, and AOA probes.

• Canopy opaquing.

• Static ports.

• Engine exhaust torching.

Mter engine start, keep ground operation time andthrust to a minimum and run air-conditioning at fullcold setting, if practical. Do not use anti-ice unlessrequired.
Taxi at a safe speed considering surface conditions,
GW, slope, and thrust. As volcanic ash, especiallywhen wet, reduces RCR, perform a rolling takeoff, if
possible.
Consider increasing the interval between takeoffs toallow clouds of suspended ash to clear.
Mter landing, consider shutting down the engine andhave aircraft towed to the parking area.
Use RCR of 18 for dry volcanic ash and 10 for wet
volcanic ash if actual RCR value is unknown.

7-8• Engine surges/malfunctions.
• Erroneous airspeed indications/fluctuations.
• FTIT rise.
• Saint Elmo's fire.
• Volcanic ash entering cockpit.
Upon detecting volcanic ash cloud penetration,
reduce thrust and maintain minimum thrust
required to sustain flight, exit the volcanic ash
environment, and land as soon as practical. Do not
use anti-ice unless required.
Engine operation above 80 percent rpm while
ingesting volcanic ash may cause buildup of melted
ash on turbine hardware and possible engine stalls.
If possible, keep rpm at 80 percent or less until clearof volcanic ash.

Change 8*U.5. ~ PRIN'l'IN:; OFFICE: 2001-660-002/42147•

T.O. 1F- 16A- 1

SECTION VIII
AIR REFUELING PROCEDURESTABLE OF CONTENTS

Page

Introduction ............................ .
General ................................. .
Control of TankerlReceiver Forces ...... .
Wingman Receiver Responsibilities ..... .
Airspeeds and Altitudes ............... .
Fuel Reserve Requirements ............ .
Weather .............................. .
Refueling Transfer Rate ............... .
Communications ...................... .
Oral Communications ............... .
Visual Signals ...................... .
Hot Armament Procedures .......... .
Lighting .............................. .
Receiver Director Lights KC-135 ..... .
Receiver Director Lights
KC-10IKDC-10 ................... .
Navigation and Position Reporting ..... .
Post Air Refueling .................. .

8-2
8-2
8-2
8-2
8-2
8-2
8-2
8-2
8-2
8-3
8-6
8-6
8-6
8-11
8-11
8-11
8-12

Flying Safety ............................ .
Boom Envelope Limits ................. .

8-12
8-12

Air Refueling Procedures ................. .
DeparturelEnroute Procedures ......... .
Cell Leader Responsibility ........... .
Buddy Departure ................... .
Taxi ............................. .
Line-Up ......................... .
Takeoff .......................... .
Aborts During Takeoff ............ .
Element Join-Up and Climb ....... .
Element Cruise Formation .......... .
Force Extension Procedures ......... .
VM C Procedures ................. .
IMC or Night Procedures ......... .
Post Refueling Procedures ........ .
Lost Wingman Procedures ........... .
Abort Procedures ................... .
Rendezvous Procedures ................ .
Altimeter Settings .................. .
Rendezvous Altitude Block .......... .
Track .............................. .
Tanker Rendezvous Equipment ...... .
Receiver Formation During
Rendezvous ...................... .
Early Arrival ....................... .
Tanker Identification ................ .

8-17
8-17
8-17
8-17
8-17
8-17
8-17
8-17
8-17
8-18
8-18
8-19
8-19
8-19
8-19
8-20
8-20
8-20
8-20
8-21
8-21
8-21
8-22
8-22

Point Parallel Rendezvous ........... . 8-22
Point Parallel Rendezvous With
Tanker Escort .................... . 8-23
Receiver Turn-On Rendezvous ....... . 8-24
Rendezvous Overrun ................ . 8-24
Enroute Rendezvous ................ . 8-24
Departure and Climb ............. . 8-24
Enroute Rendezvous Procedures ... . 8-24
Alternate Rendezvous Procedures .... . 8-26
Emission Option 3 .................. . 8-27
Missed Rendezvous Procedures ...... . 8-27
Special Air Refueling Procedures ....... . 8-27
Refueling Sequence ................. . 8-27
Fingertip Formation ........... , .. . 8-27
Echelon Formation (VMC Only) ... . 8-27
Fuel Management .................. . 8-28
Precontact ......................... . 8-28
Boom and Receptacle Procedures ..... . 8-28
Disconnect KC-135 ............... . 8-29
Disconnect KC-10IKDC-10 ........ . 8-29
Quick Flow Air Refueling
Procedures ....................... . 8-29
Toboggan .......................... . 8-31
Weather Abort Procedures ........... . 8-31
Afterburner Air Refueling ........... . 8-31
Separationtrermination Procedures .. . 8-31
Separation From a Single Tanker .. . 8-31
Cruise Cell Termination (VMC) .... . 8-31
Cruise Cell Termination (lMC) .... . 8-31
Cell Termination at Terminal
Approach Fix .................. . 8-31
Element Penetration ............. . 8-32
Receiver Radar Rejoin Procedures(IMC) ......................... . 8-32
Normal Air Refueling Procedures ....... . 8-34
Armament Safety Check ............ . 8-34
Precontact ......................... . 8-34
Contact ............................ . 8-34
Disconnect ......................... . 8-34
Post Air Refueling .................. . 8-34
Emergency Air Refueling Procedures ... . 8-35
Breakaway Procedures .............. . 8-35
System Malfunctions ................ . 8-35
Slipway Door Will Not Open ....... . 8-35
Slipway Door Will Not Close ....... . 8-35
Inoperative BoomlReceptacle
Latching ...................... . 8-36
KC-10/KDC-10 Boom FLCS Failure ... . 8-36
Brute Force Disconnect .............. . 8-36
Inadvertent Disconnect ........... . 8-36
Controlled Tension Disconnect ..... . 8-36
Change 13

8- 1

T.O. 1F- 16A- 1

INTRODUCTION
This section contains information and procedures for
F-16 air refueling with KC-135 and KC-I0IKDC-I0
aircraft. For basic flight crew air refueling procedures, refer to T.O. l-1C-I.

Lower altitudes may be required for abnormallyhigh free air temperatures. The controlling agency
directing the mission will be responsible forobtaining enroute and air refueling altitude
clearance for training and operational missions.
FUEL RESERVE REQUIREMENTS

GENERAL
This section reflects Emission Option 2 procedures
unless noted within the text. Inflight situations andsound judgement may dictate discontinuing communications procedures outlined for Emission Option 2.
Both tanker and receiver crew must be thoroughlyfamiliar with all aspects of the refueling in order toadequately plan the mission. Planners will coordinate and crews will be thoroughly familiar withmission requirements as prescribed in the appropriate command directives.
The air refueling operation requires precise anddetailed planning to insure success. Each receiver unit
will maintain a file of T.O. l-1C-l along with thismanual.
CONTROL OF TANKER/RECEIVER FORCES

An airborne tanker force commander and alternate
commanders, as required by the mission, will be
designated for each air refueling area. During
operational missions, the tanker commander is incommand of the air refueling operation from theperiod subsequent to positive radio contact between
the tanker cell leader and the receiver leader during
rendezvous until the end of the refueling or
termination of route cell formation, as applicable. The
airborne tanker force commander will coordinate
with the receiver force commander to insure
successful mission completion.

For deployment operations, the last receiver in thecell will depart the penetration fix at the abort ordestination base with a minimum of 30 minutes offuel remaining, computed in accordance with AFI
11-202, Volume 3, as supplemented by command
guidance.
WEATHER

Weather minimums are prescribed by AFI 11-202,
Volume 3, as supplemented by command guidance.
Buddy departure minimums are 1500 feet and 3 NM
for day and 2500 feet and 3 NM for night takeoffs.
Rendezvous and air refueling will not be attempted
when inflight visibility is deemed insufficient for safe
air refueling operations. However, the aircraft may
close to the lock-on limits of the radar provided thatthe required altitude separation is maintained until
visual contact has been established. Without lock-on
capability, minimum visibility for rendezvous is 1 NM.
REFUELING TRANSFER RATE

The air refueling transfer rate averages 2,000 pounds
per minute with two tanker AIR pumps operatingwith a KC-135 and 3,000 pounds per minute with a
KC-I0IKDC-I0.
COMMUNICATIONSWINGMAN RECEIVER RESPONSIBILITIESTo assist the cell leader in insuring the safety andintegrity of the flight, the wingman receiver will:• Keep the leader in visual or electronic contact at all
times.
• Maintain briefed position at all times.
• Anticipate corrections/changes and plan accordingly.
• Monitor all aspects of formation operations andadvise the cell leader if an unsafe condition is noted.
AIRSPEEDS AND ALTITUDES

Cruise and air refueling KCAS is 310 knots at
30,000 feet unless specifically directed otherwise.
8-2

Change 13

Except during an emergency fuel situation, air refueling operations will not be
conducted when radio communication
capability is lost between tanker andreceivers. If radio communications are
lost, or unreadable between the boom
operator and receiver pilot, contacts will
not he attempted.
Emission Option 2 will be used as the normal
rendezvous and air refueling procedures. Emission
Option 2, 3, or 4 procedures do not preclude verbal
communications for safety of flight situations or toinsure mission success. Boom interphone should be
used when compatible.

T.o. 1F- 16A- 1
Communications procedures and plans for rendezvous and air refueling as outlined in pertinentcommand directives will apply. Deviations must be
specifically authorized by the appropriate command
headquarters.

Tanker(s) and/or receiver(s) will make an additional
radio call confirming level at the proper rendezvous
altitude if they are not at the proper rendezvous
altitude when the 15 minute prior to the rendezvous
control time call is made.

Unless directed otherwise, communication capability
between tankers and receivers will be maintained
during all normal rendezvous and air refueling
operations. Voice transmissions, however, will be held
to an absolute minimum during rendezvous and air
refueling to be in accordance with the Emission Option
being used.

NOTE

A "Talk-Through-The-Boom" interphone system isavailable after contact when air refueling with all
KC-135 tankers if hot mic is selected.
KC-I0IKDC-I0 aircraft are equipped with UHF, HF,

and VHF radios and a boom interphone system.
All crewmembers must be thoroughly familiar withall required oral, visual, and electronic means ofcommunications. Strict radio discipline must be
adhered to at all times. All calls will be prefaced withindividual call signs. Tankers will begin monitoring
designated frequencies and will have the Radar
Rendezvous Beacon operating at least 30 minutes
prior to the rendezvous control time. The AlA TACAN
will be tuned to the appropriate channel 15 minutes
prior to the rendezvous control time unless it isrequired for navigational purposes. Receivers willcall 15 minutes prior to the air refueling control time,advising the tanker(s) of call signs, any changes in
ETA (minutes early or late) or altitude, and hotarmament check (if required).

Tankers and receivers will include
altimeter setting with appropriate
altitude calls if other than 29.92 is used.
For example, "RENO 01, ONE TWO
THOUSAND FEET, ALTIMETER SETTING THREE ZERO ZERO FOUR, ON
TIME." If EMCON 3 or 4, altimetersetting must be prebriefed.
For all rendezvous and air refueling operations,tankers and receivers will normally use theirindividual flight call signs unless directed otherwise
in operational plans. When assured no other co-unit
formation will be in range of or using the frequency,
and/or a discrete tactical frequency has been assigned
to the formation, flight call signs may be abbreviatedfor clarity and brevity; for example, "RENO
FLIGHT. ... GO ECHELON" (acknowledge) "TWO""THREE."
Mandatory calls for the receivers are as follows:• Initial radio call 15 minutes prior to the rendezvous
control time.
• Notify the tanker when established on the proper
rendezvous altitude, if not at the proper rendezvous
altitude at the 15 minute prior to the ARCT call.
• Pre contact call (required by Flight Leader only).

NOTE• If tankers and receivers are in contact
with a common facility providing rendezvous assistance, radio contact between
the tanker and receivers may be delayed
to accomplish the rendezvous.
• During enroute rendezvous, all AIR
equipment operations, interplane
communications, and timing should bebased on the RZ time. For example, the
AlA TACAN should be tuned to theappropriate channel 15 minutes prior
to the RZIP unless it is required fornavigational purposes.
The tanker will advise the receiver(s) of their callsign, air refueling altitude and, if applicable, anychange in tanker timing that would affect therendezvous (in minutes early or late).

Oral Communications

NOTE• With the exception of the breakaway
calls, crewmembers may shorten individual flight call signs using only the
number. Example: "Tank II" would be"II."• Normally, the receiver leader willproceed to the pre contact position. Whenthe leader has completed refueling,
subsequent receivers will move from theobservation position as precoordinated.
The communications requirements should be established prior to the flight. For different EMCON levels,
refer to figure 8-1 for Emission Option Communications and figure 8-2 for Emission Option Emitter.

Change 138-3

10. 1F-16A- 1

Emission Option Communications
EMISSION OPTIONITEMACTION

1

2

3

4

*

**

1.

Radios Set 30 Minutes Prior to ARCT (if dual radio capable)

X

X

2.

15 Minute Call

X

X

3.

AlA TACAN Set 15 Minutes Prior to ARCT

X

X

4.

Beacon Positive Identification (if applicable)

X

5.

ADF Check (if applicable)

X

6.

Halfway Through Turn Call (Tanker)

X

7.

Mandatory Boom Operator Calls

I
8.

a. Pre contact Callb. Clear Receiver to Contact

X
X

c. Acknowledge Contact/Disconnect
d. Verbal Corrections

X

e. Advise Receiver(s) to Return to Precontact for checklist
or equipment considerations

***

X

X
X

Mandatory Receiver Calls Mter 15 Minute Calla. Visual Contact EstablishedlLost to Include Overrun
b. Pre contact Callc. When Contact or Disconnect is Maded. Verbally notify Boom Operator prior to Manuall
Emergency Boom Latching Procedures

X
X
X

X

X

X

X

9.

Post Air Refueling ReportX

10.

1 Mile Closure Call (Receiver)X* Radio silent. Use of other emitters is authorized unless prohibited by Supported Operations Plans.
** No emissions (UHF, VHF, TACAN, FCRJRadar, AIFFIIFF, exterior lighting, etc.) unless authorized by Air
Tasking Order, Rules of Engagement, Operations Plans, Safe Passage procedures, or other mission
directives.
*** Point parallel rendezvous only.
NOTE:
Variations may be indicated by, "EMCON 2. Item 7aJ8b COMM N/A." This would mean normal Emission Option
2 procedures, except the precontact call would be deleted.

Figure 8-1.

8-4

Change 13

T. O. 1F- 16A- 1

Emission Option Emitters
EMISSION OPTIONITEMEQUIPMENT

1

2

3

4

l.

FCRJRadarOn

On

As required

Off

2.

CARAJRadar AltimeterOn

On

As required

Off

3.

TACANIDMEOn

On

As required

Off

4.

AIFFIIFFOn

On

As required

Off

5.

UHFNHF RadioOn

On

MonitorMonitor

6.

Exterior LightingOn

On

As required

Off

NOTE:

Variations may be indicated by, "EMCON 3. Item 1 Emitters Off." This would mean normal Emission Option
3 procedures, except the FeR would be off.

Figure 8-2.

Change 13

8-5

T. O. 1F- 16A- 1

Normally, boom visual signals will be used
exclusively; however, if required or requested by thereceiver, the boom operator will begin communications when the receiver reaches approximately 50
feet from the contact position. Direction, if required,
will precede distance for receiver to move and will begiven until the receiver reaches the contact position.
Example: "Forward 50," "Up 4," "Back 2." Whencontact is established, the tanker will state, "(Tanker
call sign), contact."
For Emission Option 1 and 2, the boom operator will
make a precontact audio check with receiver(s) andthe receiver(s) will acknowledge. Example: Tanker
will say, "25/57." The receiver will reply, "25."
During receiver pilot demonstration of air refuelingenvelope limits, the boom operator will state boom
limit and give the boom position for the limit being
demonstrated in increments of 2.
When the tanker is required to use manual operation,
without disconnect capability, the boom operator willstate: "(Receiver call sign), the following contacts will
be made in tanker manual operation. Receiver air
refueling system will remain in normal and receiverpilot must initiate all disconnects. (Tanker call sign),
ready." Receiver pilot acknowledges by stating,"(Receiver call sign), ready."
Visual Signals

Refer to figure 8-3. Radio silent air refueling can be
conducted by use of visual signals provided thefollowing precautions and procedures are observed:• The method, time and place of rendezvous, andamount offuel to be transferred must be covered in
the briefing of each crew.
• The tanker will use the receiver director lights (redonly) to aid in positioning the receiver. A steady red
light indicates a large correction and a flashing red
light indicates a small correction in the direction
indicated.
• If the need for an emergency breakaway occurs
during radio silent air refueling, oral breakaway
procedures and the visual signal in figure 8-2 will be
used.
Hot Armament Procedures

Prior to rendezvous with the tanker for air refueling,receiver aircraft carrying forward firing ordnancewill conduct an Armament Safety Check inaccordance with NORMAL AIR REFUELING

8-6

Change 13

PROCEDURES, this section. When radio silence ismandatory, the receiver leader will conduct a visual
challenge with each member of his flight by pointing
his index finger straight forward and thumb upward(simulating a pistol). Each member of the flight willcomplete the safety check and respond to the leader
by raising his hand and showing a circle formed by his
index finger and thumb. To reduce the possibility of
inadvertent firing, receivers will not reposition any
electrical switches while behind a tanker unless thoseswitch changes are required for air refueling
operations or aircraft control.
LIGHTING

Refer to figure 8-4. While approaching the precontact
or contact position, the receiver pilot can adjust
exterior lighting as required by the boom operator.
NOTE

Visual contact for night air refueling canbe aided by requesting the tanker toflash his landing lights prior to andlor
during the tanker turn.
Single KC-135 tankers performing a rendezvous both
upper and lower strobes will display red.
NOTE

If the spare is used during the air
refueling, the appropriate color code will
be displayed until the receiver is in the
pre contact position. To further aid in
identification, tanker position lights will
be placed on BRIGHT and FLASHING
for numbers 1, 3, and 5. Position lights
for numbers 2 and 4 will be BRIGHT and
STEADY. Position lights will be set prior
to takeoff. After the receiver has
established visual contact and has closed
to 112 NM in trail, tankers will turn
position lights to STEADY and DIM andturn strobe light OFF. When any aircraft
will be flying visual wing formation on
the tanker, the tanker will also turn off
the strobe light. In this case, the last(outside) receiver aircraft with eachtanker will have anticollision lights ON.
When fighter receivers reach the observation position, tankers will turn underwing, underbody, and nacelle illuminating lights to DIM. Exterior lights willthen be adjusted as requested by thereceiver pilot.

10.lF-16A-l

Visual Signals
INDICATION
SIGNALBOOM AIR REFUELINGPROBE & DROGUEREFUELING

1. Boom in Trail

A. Extended 10 feet

*Ready for Contact

B. Fully extended

1. Tanker Manual Operation
without Tanker Disconnect
Capability

2. Acknowledge Receiver's
Manual Boom Latching Signal
C. Fully retracted

Offload Complete

2. Boom Stowed
A. Fully retracted

Tanker Air Refueling System
Inoperative

B. Extended 5 feet

System Malfunction, Tanker andReceiver Check Air Refueling
Systems

3. Flashing Receiver Director
Lights!Tanker Lower Rotating Beacon ON

BREAKAWAY

4. **Receiver Director Lights
Going OUT During Contact

Tanker Request for Disconnect,
Receiver return to Precontact
Position

5. Receiver Closing and Opening Receptacle Door when in
Pre contact Position

1. Manual Boom Latching

2. Acknowledge Tanker's
Manual Operation without
Tanker Disconnect Capability
Signal6. ***Steady Light from Receiver or Rocking of Wings

Emergency Fuel Shortage Exists

7. Flashing Light from Receiver
Cockpit Area

Initiate Toboggan Maneuver

Figure 8-3. (Sheet 1)
Change 13

8-7

T.O. 1F- 16A- 1

Visual Signals (Cont)
INDICATION
SIGNALBOOM AIR REFUELING
8. (a). (DAY): Same receiver returns to pre contact with receptacle door open. Pilot signals closed fist, thumb tomouth, plus hand signaling
number****(NIGHT): Same receiver
returns to precontact with receptacle door open, ready forcontact

PROBE & DROGUEREFUELING

Additional fuel required
EMCON 2-4

I(b). (DAY): Same receiver returns to precontact, ready forcontact. Pilot signals closed
fist plus hand signaling number****(NIGHT): Same receiver
returns to precontact, readyfor contact

Additional fuel required
EMCON 2-4

* Receiver(s) in the observation position will move to the precontact position in their briefed sequence only
after insuring that the boom is in the ready for contact position and the preceding receiver has cleared thetanker. The receiver will stabilize in the precontact position, then move to the contact position. The boom
operator will not give the ready for contact signal until the preceding receiver has cleared the tanker.
**The receiver(s) will advise the tanker of any pilot director light malfunctions/deficiencies.
*** If fuel shortage occurs at times other than scheduled air refueling, the receiver should be positioned so thesignal may be seen from the tanker cockpit.
*:1'** Additional fuel offioaded will be 5000 pounds for large receiver aircraft, or 2000 pounds for small receiveraircraft, on each subsequent contact.

Figure 8-3. (Sheet 2)

8-8Change 13

T. o. 1F- 16A- 1

Exterior Lighting (KC-13SJ*A/R FLOOD LIGHTSSTROBE LIGHT* NACELLE ILLUMINATION LIGHT*NAVIGATION/POSITION
LIGHT

J

BOOM MARKER
LIGHTS* BOOM NOZZLE LIGHT*NAVIGATION/POSITION
LIGHT* NAVIGATION/POSITION
LIGHT

LANDING LIGHTTAXI LIGHTTAXI LIGHT
LANDING LIGHT* UNDERBODY ILLUMINATIONLIGHTTERRAIN LIGHT(RETRACTABLE)* NAVIGATION/POSITION
LIGHT

VIEW LOOKING UP

NAVIGATION/POSITION
LIGHT* DESIGNATES ADJUSTABLE LIGHTING
1 F-16X-1-8004X@l

Figure 8-4. (Sheet 1)

Change 13

8-9

T.O. 1 F-16A-1

Exterior Lighting (KC-l0IKDC-1.0J
FORMATION LIGHTSANTICOLLISION/RENDEZVOUS

LOGO

/

HOSE FLOODLIGHT

RECEIVER AIRPLANEFLOODLIGHTSWING TIP
ANTICOLLISIONFWD POSITION/
WING TIP ANTICOLLISION - -_--a.....:,--,-L--AFT POSITION

WING/ENGINE NACELLE SCANNING

HORIZONTALSTABILIZERUNDERSIDE LIGHTSAFT FUSELAGE-A
DIRECTOR
LIGHTS[]/NOTE:
THE INTENSITY OF ALL
EXTERIOR LIGHTING IS
ADJUSTABLE.
LEFT SIDE SHOWNRIGHT SIDE SAMEFORWARDFUSELAGEUNDERBODY
LIGHTSRECEIVERAIRPLANEFLOODLIGHTS

VIEW LOOKING UP

1 F-16X-1-8003X@

Figure 8-4. (Sheet 2)

8- 10

Change 13

T.O. 1 F-16A-l

KC-135 aircraft in cell formation will set their strobe
lights according to their position in the refueling cell.
Their lighting is as follows:• Tanker Number 1
Upper and Lower Lights - RED• Tanker Number 2Upper and Lower Lights - WHITE• Tanker Number 3Upper Lights - RED
Lower Lights - WHITE• Tanker Number 4Upper Lights - WHITE
Lower Lights - RED
Receiver Director Lights KC-135

Refer to figure 8-5. The receiver director lights do notgive true vertical and horizontal information. The up
and down lights change because of angularmovement of the boom and the fore-and-aft lights
change because of in-and-out movements of the boom.
The axis of the director lights system is inclined at a
30 degree angle to the fuselage. This angle causes aninteraction in both lights when a true vertical or
horizontal movement is made by the receiver. For
example, flying straight forward while in contact willcause the boom to compress and also increase its
angle with the tanker fuselage. The lights will showthat the aircraft is flying forward and down. If a true
up movement is made, the boom will both compress
and lessen its angle with the tanker fuselage and the
director lights will indicate that an up and forward
movement has been made. Small fore and aftcorrections can be made with little or no power change
by moving vertically.
Receiver director lights are on the bottom of thefuselage directly aft of the nose landing gear. They
consist of two rows of lights: the left row for elevationand the right row for telescoping. The elevation lights
consist of five colored panels with a green stripe,
green and red colors, and two illuminated letters, D
and U, for down and up respectively. The coloredpanels are illuminated by lights that are controlled byboom elevation during contact. There is an
illuminated white panel between each panel to serve
as a reference. The letters A, for aft, and F, forforward, augment the colored panels on the telescopeside. The receiver pilot director lights will remain
illuminated and follow boom movements in both thecontact made and disconnect conditions. There are nolights for azimuth position. A fluorescent yellow strip
on the bottom center of the tanker fuselage may be

used as centerline reference by the pilot. The
triangular-shaped panels are for elevation and therectangular-shaped panels are for forward and backmovement.
Receiver Director Lights KC-l O/KDC-l 0

Refer to figure 8-5. The receiver director lights consistof two rows of lights located forward of the wing root.
Relative elevation position is provided by the left row
and the right row provides telescoping position. The
elevation row contains one striped green, two amber
and two red triangular panels, and two white letters:
U at the forward end for UP, and D at the aft end for
DOWN. The colored panels and letters are dimly
illuminated by background lights. The telescoping
row contains one striped green, two amber, two red,and four white rectangular panels, and two whiteletters: A at the forward end for AFT, and F at the aftend for FORWARD. The colored panels are notbackground lighted; however, the letter at each end of
the row is dimly illuminated. Separation is provided
by the white panels. The pilot director lights areadjusted by the boom operator to the size air refuelingenvelope for each receiver and provide guidance
during contact.
To provide more response time, the appropriate panel
and letter are illuminated in anticipation of receiver
movement. The director lights provide commands
based on both receiver position and rate of movement.
Figure 8-5 shows the lights with no receiver motion.
With rapid motions of the receiver, the lights canshow a correction required even though the receiveris in the center of the envelope. The red panel andletter at the ends of each row can be illuminated by
the boom operator to aid the receiver in attaining thecontact position.
NAVIGATION AND POSITION REPORTING

When rendezvous is complete, tankers will be
responsible for all navigation, weather avoidance andposition reporting. The tanker pilot will, once eachhour, advise the receiver pilots of the cell geographic
position, heading, and distance and ETE to the next
checkpoint or destinations applicable. If the receiverpilot has not completed his onload upon reaching his
geographic air refueling abort point, the tanker pilot
will so advise.
NOTE

During refueling operations conductedwithin a pre-planned orbit pattern, thetanker is not required to provide theabove information unless requested by
the receiver leader.

Change 1 3

8- 11

T.O. 1F-16A- 1
Post Air RefuelingUpon completion of air refueling, the tanker will
normal1y climb to the top and the receiver(s) will
descend to the bottom of the air refueling block. The
receivers should maneuver to the prescribed
formation position while awaiting post air refuelingreport and further clearance. The tanker will give
post air refueling information to the receiver as
required. The receiver will advise the tanker of any
pilot director light malfunctions/deficiencies, e.g.,lights intermittent, inoperative, dim, dirty, etc. Upon
termination of air refueling, all exterior lights will be
as directed.

I WARNING I
Receiver will insure a safe clearance
from the tanker(s) as they proceed on
their assigned missions. Receiver(s)required to accelerate past the tanker(s)and climb on the refueling heading will
maneuver either left or right (aminimum of 1 NM) of track to preclude
climbing directly in front of thetanker(s) remaining receiver(s). Aircraft flying through departing receivers'
jet wash may experience damage to theaircraft and injury to personnel.

8- 12Change 13

FLYING SAFETY
BOOM ENVELOPE LIMITS

Refer to figure 8-6. The refueling envelope is limitedby the refueling receptacle location. As long as thereceiver is positioned within these limits, contact canbe maintained despite rolling, yawing, or pitching.

• Approaching boom limits at a relatively
high velocity can cause structural
damage as a result of an inability to
disconnect due to binding of the boom
nozzle.
• Due to the restricted refueling envelope,boom limit switch protection is notprovided in up elevation.
NOTE

The boom operator will disconnect at 25
degrees elevation (upper envelope limit).
In this position, a green DOWN arrow
will be illuminated.

10. lF-16A- l

Receiver Director Lights (KC-135J

'
--20,
0
---

0

--21.5 0
23.5 024.5 0- - - 26.0 0

t--

f'

JO. ooIF-l&o.- l ..(1162Xe

FIgu", 8-5. (Shed 1)
rhnnnA 1::t

A. 1 .1

T.O. 1F-16A-l

Receiver Director Lights (KC-10IKDC-10J

1 __
'

25"

_ _ 26_1°_28 4 "

1_.,....,- - 29.'-

C 1:IiD • C:=J • t=l •C:=J •

c:::J • -"'.I' [::>Figure EM. (Sheet 2)

8-14

Change 13

T. O. 1F- 16A- 1

Boom Envelope Limits (KC-135J

LEFT AZIMUTH LIMIT

6 FT EXTENDED
INNER LIMIT

~

t

20°
~",~
; - UP ELEVATION LIMIT

6 FT EXTENDED - - - - . /INNER LIMITDOWN ELEVATION LIMIT

18 FT EXTENDED
OUTER LIMIT

1F-16X-1-8002X@

Figure 8-6. (Sheet 1)

Change 13

8- 15

T. O. 1F- 16A- 1

Boom Envelope Limits (KC-10IKDC-10J
RIGHT AZIMUTH L1MIT~,--,-=--LEFT AZIMUTH L1MIT-

6 FT EXTENDED - - - - 'INNER LIMITDOWN ELEVATION LIMIT

21 FT EXTENDED
OUTER LIMIT

IF-16X-1-8001X@

Figure 8-6. (Sheet 2)

8- 16

Change 13

10. 1 F- 16A- 1

AIR REFUELING PROCEDURESNOTEDEPARTURE/ENROUTE PROCEDURES

On runways less than 300 feet wide, thereceiver will remain in the number 1
position until the tanker rolls.

Cell Leader Responsibility

The tanker cell leader or specified commander is
responsible for the command and control of theformation and the air refueling operation. The cell
leader or specified commander will coordinate withthe receiver force commander to insure successful
mission completion. Formation integrity and discipline begin with the formation briefing. The cell
leader must insure that all aspects of the mission areclarified and understood.
Normally, lead responsibilities pass from number one
to number two to number three; however, all pilots
must be prepared to assume full responsibility for theformation at any time. When it is necessary totransfer lead responsibilities, all airplanes in theaffected formation will be notified when the newleader assumes responsibility for the formation.
The tanker cell leader/specified commander musttake every feasible action to enhance the possibilityof completing air refueling.
Buddy Departure

A buddy departure is effected when the tanker(s) andreceiver(s) take off from the same base and visualcontact is maintained.

TAKEOFFI WARNING I
Wake turbulence generated by a preceding aircraft may create a hazard during
buddy takeoffs and join-ups.
The tanker will roll first, followed in 45 seconds by the
first receiver in his element. Each element will be
individually cleared for takeoff by the tower after thelast aircraft in the preceding element has passed theend of the runway. Takeoff interval may be varied
when weather, terrain, airfield conditions, or otherlocal considerations dictate.
ABORTS DURING TAKEOFF

An aborting aircraft will make an abort call on theprebriefed common frequency as soon as possible.
Frequency changes will not be made by tanker/receivers until all airplanes in the same element are
airborne.

TAXI

Mter engine start, receivers will check in with thetanker on the predetermined frequency. When readyto taxi, each tanker will call, "(Tanker Call Sign),
Taxiing." A distance of 300 feet will be maintained
between tankers and receivers. Tankers not scheduled to be used during the first air refueling will taxi
and takeoff first, followed by refueling elements
consisting of a tanker and his mated receivers for the
first refueling. Spare tankers will taxi and takeoff
last.
LINE-UPOn runways at least 300 feet wide, the KC-135 willline up on the downwind side of the runway. The
receivers will be positioned on the upwind side of the
runway, maintaining wing tip clearance.

ELEMENT JOIN-UP AND CLIMB

The climb speed schedule will be based on the firsttanker/receiver element climbing at 300 KCAS.
Individual tankers launching first will reduce climb
airspeed so that a 10 KCAS differential exists
between individual tankers and between the lastindividual tanker and the first refueling element.
Example:Tanker One . . . . . . . . . . . . . . . . . . . . . ..
Tanker Two . . . . . . . . . . . . . . . . . . . . . ..
Lead Element. . . . . . . . . . . . . . . . . . . ..
Second Element. ...... .... ... .....
Third Element .. . . . . . . . . . . . . . . . . ..
Fourth Element . . . . . . . . . . . . . . . . . ..

280 KCAS
290 KCAS
300 KCAS
310 KCAS
320 KCAS
330 KCAS

Change 13

8- 17

T.O.1F-16A-1

NOTE
For two and three element cells, tankers
without receivers are considered elements. The lead element will climb at
300 KCAS and the following speed
schedule will be utilized to expedite
closure: In a two element cell, the second
element will climb at 330 KCAS. In athree element cell, the second and thirdelem~nts will climb at 320 KCAS and
330 KCAS, respectively.

2.

The flight/element departing the tanker reforms and moves back from the right wing withoffset and establishes contact (Radar, TACAN,etc.) with the tanker. Maintain altitudedeconfliction from any joining receivers.

3.

The next flight/element may then begin to
move forward maintaining altitude.

4.

Once the flight/elements have horizontal
separation (established by Radar, TACAN,etc.), the next receivers can climb and thedeparting receivers can descend to briefed
altitudes.

5.

If a flight/element is to rejoin from the trail
position and the preceding flight/element
elects to stay in formation with the tanker, theflight/element in trail will close on the tanker
until they are established in a visual position
and then switch positions with the precedingflight.

If a ceiling is to be encounten~d prior to the completionof the join-up, the tanker will level-off with aminimum of 500 feet clearance below the cloud layer
and maintain assigned climb airspeed to provide
visual flight conditions during the element join-up.

The air refueling element (tankers and receivers)
join-up will be accomplished on the outside of the turnor on the tanker's left wing during a straight ahead
join-up. AU turns by the tanker will normally be 20
degrees of bank (Exception: See Lost Wingman
Procedures). Mter the receiver flight has stabilized inits proper forrnation position, the element leader will
inform the tanker pilot.
Element Cruise Formation

Receivers will fly route formation position on thetanker. (For extended VMC cell cruise formation
receivers may fly a spread formation.) Route andspread form ations will be as specified in theoperational procedures manuals. Spacing may be
reduced during IMC or night operations. When air
refueling is required, the {)ther receiver(s) will
aSt-'ume the observation position.
At night or in IMC, a maximum of 12 aircraft will beassigned to one tanker - a maximum of three aircraft
will fly on each wing of the tanker, and additional
aircraft/elements will fly in a 1-3 NM trail position, aminimum of 1000 feet below their tanker's altitude.
During tanker cell operations, only the last tanker
shou1d have receivers in trail. With nine or lessreceivers, only one element need be in trail (oneelement refueling while the other two elements are
s\vitching positions). With 10 to 12 receivers, twoelements will be in trail (one clement maintaining
trail, and the other two elements switchingpositions - and with four elements, one element
refueling). In this case, offset trail and additional
altitude separation will be used for deconfliction. The
following procedures will be used to change positionsof flights in trail:
1.

8- 18

The tanker should be In straight and level
flight.
Change 13

Force Extension Procedures

Force extension procedures are used when force
extensiontankersair
refueltankers(KC-10IKDC-I0IKC-135) escorting fighters during
deployment operations. Force Extension missions areoften complex and demanding to all aircrews,
especially in IMC. All facets of the mission, to include
the rendezvous, formation, air refueling, VMCIIMC
rejoin procedures, and cell break-up will be briefed
and clearly understood by all participants during
mission planning. Basic Point Parallel or Enroute
Rendezvous procedures apply, except as indicated.
Multi-ship tanker cells will fly 60 degree echelon, 2
N1f spacing unless otherwise directed. l\fi~sions thatencounter IMC conditions during air refueling may
increase air refueling echelon formation spacing from
2 miles to 3 miles. If a mid-mission rendezvous isplanned, the escon~ng- tanker will attempt to contact
the force extension tanker c!!1d pass the ETA to the
ARCP and updated weather information which may
effect air refueling. Inflight visibility will be thedetermining factor in utilizing Vl\1C versus IMC
procedures to conduct air refueling. For the purposeof these procedures, Vl'v1C is defined as visibility equal
to or greater than 2 NM. Il\1C is defined as visibility
less than 2 Nl\f.

T. O. 1 F- 16A- 1
VMC PROCEDURESFighters will join on the force extension tanker whencleared by the escorting tanker pilot. After all fighters
have joined on the force extension tanker, theescorting tanker will be cleared for refueling.
Fighters should fly a loose wing formation andremain with their force extension tankers in the eventa breakaway occurs.
IMC OR NIGHT PROCEDURES

Air-to-air radar equipped fighters, when cleared by
the escorting tanker pilot, will fly 1 112 to 2 NM trail(6 o'clock) position, 2000 feet below their assigned
force extension tanker. Non-air-to-air radar equippedfighters will comply with the VMC procedures.
POST REFUELING PROCEDURES

Once all air refuelings are complete, the escortingtanker(s) will descend 1000 feet, offset slightly to theright and then move to a position 1 NM in front of theforce extension tanker(s). Once the escortingtanker(s) are stabilized 1000 feet below and forwardof the force extension tanker, the escorting tanker willassume lead for the formation after a positive verbal
lead change. The escorting tanker will then clear thefighters forward to rejoin.
Non-air-to-air radar equipped fighters will reJOIn
visually with their respective escorting tanker.
Air-to-air radar equipped fighters will rejoin using
radar guidance if required. If IMC prevails and poor
visibility precludes visual rejoins, force extension
tanker(s) may momentarily reduce separation to 1/2
NM and 500 feet vertical separation to facilitate therejoin.
Once all fighters have rejoined on their respective
escorting tanker, the force extension tanker(s) willdepart the stream from the rear of the formation.
If the formation reaches the End NR point and visualrejoins are not possible, force extension tankers willcontinue along the receivers' route of flight until
visual rejoins are possible, fuel permitting. If theforce extension tanker(s) reach BINGO FUEL at, or
after, the End AIR point and the fighters have notrejoined with the escorting tanker(s), the entire
formation will abort to a suitable alternate airfield.

The tanker radar should be used for position
monitoring throughout the maneuver. It is the force
extension tanker leader's responsibility to inform theentire formation of current heading and airspeed
until relieved of that responsibility by a lead change.
The force extension tanker/cell will reform at the topof the air refueling block. Once the fighters haverejoined on their respective escorting tankers, theescorting tanker/cell will reform at the bottom of theblock. Cell separation will be accomplished by theforce extension or escorting tanker cell increasing/decreasing airspeed as determined by mission requirements and/or the pre-mission brief. Force extension
or escorting tankers will not make any climbing ordescending turns to depart the stream until thetanker cells are identified visually or by radar, arewell clear, and verbal coordination is made between
tanker cell leaders. All aircrews must clear
aggressively and be cognizant of potential converging
headings or conflicts.
When simultaneous refueling of fighters andescorting tankers is required, 310 KCAS
(KC-10/KDC-10) or 295 KIAS (KC-135) normally will
be used as refueling airspeed for the formation. The
lead force extension tanker will determine air
refueling airspeed based on aircraft type, altitude,
weight, weapons load, etc.
The above procedures are the standard for force
extension tankers refueling escort tankers. Any
deviations to these procedures are not authorized
unless coordinated between all tankers (escortingand force extension), receivers, and mission commander. If a pre-departure briefing is not conducted
due to geographically separated departure locations,the escorting tanker will coordinate changes inflightto the force extension tanker upon initial contact,prior to the air refueling rendezvous.
Lost Wingman Procedures

In the event a receiver aircraft becomes lost during
refueling operations or during buddy cruise, thefollowing procedures will apply:1.

If a tanker or receiver detects weather (by radar,
reduced visibility, etc.) which may require a
receiver to go lost wingman, the tanker shouldmake any required navigational corrections,then fly straight and level. Any subsequent
turns required should be made with ten degrees
of bank maximum and called over the radio (i.e.,"Tanker 21 is rolling into a ten degrees rightturn.").

2.

The receiver flying on the tanker's wing will
simultaneously inform the leader and turnaway using 15 degrees of bank for 15 seconds,
then resume heading.

3.

The receiver flying on the wing of the above
receiver (second in the echelon) will simultaneously inform the leader and turn away using 30
degrees of bank for 30 seconds, then resume
heading.
Change 13

8- 19

T.O. 1F-16A- 1

4.

5.

The receiver flying on the wing of the above
receiver (third in the echelon) will simultaneously inform the leader and tum away using
45 degrees of bank for 30 seconds, then resume
heading.
The receiver who is flying in the precontact orcontact position will simultaneously reduce
airspeed 10 knots and descend 500 feet below
refueling altitude. Hold refueling heading andafter 30 seconds resume normal airspeed.

6.

If the above procedures are started while
tanker(s) are in a turn, request tanker(s) rollwings-level and call roll-out heading.

7.

Notify flight leader or tanker commander of thesituation.

8.

Attempt rejoin only after receiving clearance
from the tanker and when within radar or VMC
capability.

Abort Procedures

If a receiver aborts during an air refueling mission,

the receiver leader will determine the course of action
to be taken.
RENDEZVOUS PROCEDURES
NOTE

Initial visual contact between thereceiver and tanker may be enhanced,
inflight weather conditions permitting,if the tanker jettisons fuel to increase its
visual signature. This procedure may beinitiatedJrequested by the tanker, receiver, or the ground agency controlling
the rendezvous. It should only be used if
a receiver low fuel state or other similar
circumstances require the rendezvous
be expedited.
The type of rendezvous will be dictated by mission
requirements, weather conditions, etc. If weather
conditions for join-up at cruise altitude are VMC, anenroute rendezvous may be used. When tanker andreceivers operate as separate flights, the Point
Parallel rendezvous will be primary.
After radio contact has been established, bothtankers and receivers should be tuned to the same
NAVAID, if possible, to improve rendezvous
capability.
8-20

Change 13

NOTE• For receiver's directed rendezvous, thereceiver's airborne radar equipment (ifavailable) will be the primary means toaccomplish a rendezvous. AlA TACAN
will be used as a backup.
• Ground radar assistance should be
utilized to the maximum for all
rendezvous. The tanker commander
will monitor ground control radio
frequency and confirm receiver range/
bearing information to insure positive
identification during the rendezvous.
Ground radar assistance will be terminated when the receiver reports visualcontact with the tanker.
Altimeter Settings

Unless otherwise directed, an altimeter setting of
29.92 inches Hg will be used for air refueling
operations at or above transition altitude or whenover water and operating in accordance with ICAO
procedures. For all other air refueling operations, thebriefed altimeter setting will be used.
Rendezvous Altitude Block

Four consecutive altitudes shall be requested by thetanker for rendezvous and refueling. When four
altitudes are available, the rendezvous will beeffected with the tanker at the second altitude andthe fighter(s) at the third. For example, when therefueling altitudes are FL290, 300, 310, and 320, thetanker will be at FL310 and the fighter(s) at FL300
for rendezvous, thus providing 1000 feet above thetanker and 1000 feet below the fighter(s). Whentankers are in cell, they will stack up from the second
altitude from the top of the block, FL310 in thisexample.
When only three altitudes are available, the tanker
shall be at the top altitude with the fighter(s) at the
middle altitude, providing 1000 feet below thefighter(s).
When only three altitudes are available and therefueling involves tank.er cell formation, the highest
tanker within the cell should be at the top of the block.
To accomplish this, the tanker leader shall place
himself at an altitude that will permit the highest
tanker in his cell to be at the top of the block. For
example, when there are two tankers and theavailable block is FL270-290, the tanker leader would
be at FL285 with number two tanker at FL290. The
fighters would rendezvous at FL275 (1000 feet below
the lowest tanker).

T.O.1F-16A-1

The above procedures do not apply when aircraft areoperating on an Altitude Reservation (ALTRAV) or
when clearance has been granted for aircraft tooperate as an enroute cell. In these cases, the altitude
block will provide airspace necessary to accommodate
the type of formations being used (standard or
nonstandard), with at least 1000 feet between thehighest receiver and lowest tanker during rendezvous, and at least 1000 feet below the air refuelingformation once the rendezvous is complete.
Track

Receivers shall pass over the ARIP, if applicable, andmake good the planned inbound track to the ARCP. If
a deviation is required because of weather, etc.,
receivers will not attempt rendezvous or proceed to
the ARCP until the deviation has been approved by
air traffic control and coordinated with the tanker. If
radio contact between the tankers and receivers is notestablished prior to the ARCT, the tankers will beover the ARCP at the ARCT.
Tanker Rendezvous Equipment

Tanker rendezvous equipment consists of thefollowing:• KC-135• AlA TACAN - DME only

• Radar beacon - AN/APN-69 (all aircraft) and
AN/APN-134 (some aircraft)• Automatic Direction Finder - AN/ARA-25• KC-10IKDC-10• AlA TACAN - Range and Bearing• Radar beacon - APX-78 (Two Pulse, Variable
Width)• UHFIDF
• INS
NOTE

Radar beacon contact will be shortrange due to radar polarization incompatibility.

Receiver Formation During Rendezvous

Formation procedures after level-off or from the ARIP
until join-up with the tankers will be as follows:
NOTE

Formation lead changes and join-ups
will normally be completed prior to
departure from the ARIP. Should suchmaneuvers be required subsequent to
departure and prior to join-up on thetanker(s), the rendezvous will not be
continued unless the flight leader ispositive of his position in relation to
the tanker(s) and the published AIR
track.
• Day VMC (Visibility five miles or greater). Flights
will be in trail, offset to the right of the precedingflight. When all aircraft are in visual contact withthe tanker(s), each aircraft/flight will join with his
respective tanker as briefed.
• IMC or Night. The receiver formation will be formed
into flight(s) of four aircraft in close or routeformation. Succeeding flights will be in a likeformation, positioned 1500 feet to the rear of the firstflight or 2-3 NM radar trail, depending on weather
conditions. When the tankers are established on the
on-course track, the receiver leader will position his
flight 3 NM to the rear of the last tanker. When alltankers in the cell are in positive radar contact,
receivers will climb to 1000 feet below base altitude.
If visual contact has been established, theflight/element leader will initiate join-up on the lasttanker. When the appropriate wingmen have visualcontact and are within 1 NM of their tanker, theleader will drop them off and proceed to the nexttanker. The receiver leader will continue as above
until all wingmen are on their appropriate tanker,
then join the lead tanker. If not in visual contact at
3 NM, the receiver flight/element leader will clear
the last receiver flight to join on the last tanker. The
last receiver flight will then turn to the right and joinon the last tanker while climbing to 500 feet below
base altitude, then maintain this altitude until
visual contact with the tanker is established. The
first receiver flight will join on the number 2 tanker,
maintaining 1000 feet below base altitude until
visual contact is established. Each receiver leader
will offset his tanker target 15 degrees to the left andclose at approximately 50 knots above the tanker
KCAS. Aircraft with operable airborne radar
equipment will close no closer than 1500 feet.
Receivers losing radar lock between 1 NM andminimum range will ensure 1000 feet altitude
separation is maintained and discontinue rendezvous attempts until adequate range separation (1NM) is achieved or radar lock-on is regained. Range
Change 13

8-21

T.O. 1F- 16A-1

closure limitation for non-radar equipped receiver(s)or receiver(s) without radar lock-on is 1 NM. Whenvisual contact with the tanker has been established,
the receiver element will form in the precontact
position, and the receiver flight leader will turn left,then right back to track heading and join on the leadtanker in the cell using the same procedures. The
flight leader's wingmen, after flight separation, will
echelon to the right. If visual contact with thetankers is not established, receivers will maintain 15
degrees offset, applicable altitude, and minimum
slant range until cell termination procedures areaccomplished.

If the radar beacon is to be used topositively identify the tanker, thereceiver should attempt to identify thetanker rendezvous beacon as soon as
possible so that an alternate method of
identification may be used if unable toreceive the beacon. Only the tanker cell
leader or his designated alternate, if
applicable, will operate the rendezvous
beacon.
Point Parallel Rendezvous

Early Arrival

Once a join-up is initiated, and it is necessary for thejoining receivers to hold while waiting for a precedingflight to complete their operations, the joining flightwill join in a position 600-800 feet out (laterally) fromthe receivers in the observation position ifin VMC; or
fly offset laterally 1-3 NM in trail if in IMC, andmaintain 1000 feet below tanker base altitude. The
decision on which side to join will be based on the
direction of orbit of the tanker, departure intentionsof the receiver flight refueling, and the presence ofadditional holding receiver flights.

IWARNING

1

Joining flights of receivers should not
close astern of a tanker that is
conducting refueling operations withother receivers. The wake turbulence
generated by these aircraft during
departure/changing of positions, if
encountered, can result in loss ofaircraft control.
In the event the receiver(s) arrives ahead of thetanker at the ARIP or ARCP point, the receiver will
orbit at an altitude that ensures at least 1000 feetseparation between tanker and receiver or any
elements of tanker and receiver cells. If receivers holdat the ARCP, they will normally enter a left handholding pattern using 2 minute legs at 1000 feetbelow air refueling altitude.
Tanker Identification

Tanker identification is critical in congested refuelingairspace. Available aids used in any combination
should be used to confirm tanker location/identification prior to and during the rendezvous. These aids
include ground radar, tanker/receiver radar, INS, N A
TACAN, UHFIDF steers, common ground TACAN
stations, and radar beaconslIFF/SIF interrogation
systems for receivers so equipped.
8-22

NOTE

Change 13

The point parallel rendezvous is used to effect a rapid
join-up between the tanker and receiver withminimum receiver maneuvering. The tanker andreceiver approach on reciprocal headings offset, leftor right, a distance equal to the tanker turn diameter.
At a predetermined turn range, the tanker executes
a turn to the receiver heading to roll out
approximately 1-3 NM ahead of the receiver.
Normally, the tanker aircrew has responsibility forthe overall refueling operation and rendezvous andestablishes the offset and turn point. When tanker
systems are degraded, the situation dictates, or for
training, the Tactical Air Controller or receiver may
be responsible for the execution of the rendezvous.
Specific rendezvous responsibilities will be inaccordance with AFI 11-214. Receivers will monitor
the air refueling frequency and attempt to establish
contact as soon as possible. If both tankers andreceivers are on a common radio frequency to obtainground radar rendezvous assistance, the change to
air refueling frequency may be delayed until positive
radar/visual contact is established. If under radar
control, obtain bearing and distance to the tanker
prior to changing to air refueling frequency.
A successful point parallel rendezvous requires thetanker to maintain the proper offset and the receiverto fly the specified rendezvous track from the ARIP to
the ARCP. Emission Option 2 will be the normal
rendezvous and air refueling procedure. The receiverwill call 15 minutes prior to the ARCT and relay callsign, ETA (minutes early or late), and altitude. The
tanker will then confirm his call sign, air refueling
altitude and timing (minutes early or late) if it willaffect the rendezvous. If either the tanker or thereceiver is not on appropriate rendezvous altitude,
another radio call will be made when the proper
rendezvous altitudes are established. Receivers
departing the ARIP will maintain cruise speed at1000 feet below tanker base altitude until positive
radar or visual contact is made from in trail with theappropriate tanker(s). The receivers will proceed

T.O. 1F- 16A- 1

from the ARIP to the ARCP using all navigational
aids available to maintain the centerline. When it isdetermined the receiver is at or inside the ARIP, thetanker will turn to, or continue on, the reciprocal of
the receiver's inbound track and will establish andmaintain the proper offset until reaching the planned
turn range. However, receivers will not deviate fromthe ARIP/ARCP centerline unless directed to do so by
the tanker.
The tanker will adjust to appropriate air refuelingspeed when rolled out toward the ARCP The tanker
INSIDNS will be the primary means of maintaining
the offset and the AlA TACAN will be primary forrange information. To provide AlA TACAN ranging,
the tanker and the receiver (1 airplane per 0011) willtune the assigned AlA TACAN channels 15 minutes
prior to the ARCT. The receiver will set thenumerically lower AlA TACAN Y-channel and thetanker will set the higher V-channel. AlA TACAN
should be left in AlA until the receiver reaches
precontact.

NOTE• The above slant range/bearing chart
applies only to the KC-135 at approximately 28,000 feet. It is based on 9.5
NM lateral offset with the tanker
using 30 degrees of bank during theturn to refueling heading. The
KC-IO/KDC-10 normally uses 25 degrees of bank or less during the turn torefueling heading and approximately
an 11.5 NM offset. Normally, the turnrange with the KC-10/KDC-10 will be
2-6 NM greater than the turn range
with the KC-135.
• When more than one tanker isinvolved, it may be necessary to add 1
NM to the turn range for each
additional tanker in the formation toensure receiver leader is 3 NM in trailof the appropriate tanker at roll-out.

NOTERadio silence will be broken if thetanker or receiver determines thateither the tanker or receiver will exceed
ATe protected airspace while maneuvering to attain the offset.
The tanker will note the receiver's distance when
halfway through the turn back to the ARCP. This isthe best time to determine if an overrun condition
exists and the best time for visual sighting. If an
overrun condition exists, appropriate action should
be taken. Tactical Air Controllers or receivers who
have established radar contact and positively
identified the tanker will offset the tanker in order toestablish a 26 degree relative bearing at 21 NM
range. The receiver/weapons controller will offset thetanker as follows:
APPROXIMATESLANT RANGE(NM)BEARING(DEGREES)

100
80
60
50
40
30
25
21
20
13

5
6
8
10
13
18
21
26
27
45

At the turn range, the tankers will turn to therefueling track and adjust to refueling formation. The
receiver will assume responsibility for closing on thetanker at the turn range. The last tanker in a cell willturn the radar beacon to operate, single code, onroll-out to refueling heading at the completion of theturn. The receivers will normally be 3 NM in trail ofthe tanker(s). When rolled out toward the ARCP, thetanker will adjust to the appropriate air refuelingspeed.
Point Parallel Rendezvous With Tanker Escort

Refer to figure 8-7. The receivers will join Dn theescorting tanker in the briefed sequence. The orbitingand escorting tankers are responsible for effecting therendezvous.
An air refueling anchor is a left-hand racetrackpattern with legs separated by 20 NM and normallya minimum leg length of 50 NM. Tankers will adjustfrom enroute cell formation to the air refuelingformation of 20 degrees right echelon, 1 NM
nose-to-nose separation, stacked up at 500 footintervals during the final turn to the air refuelingtrack.
When the receivers are inbound, the rendezvous will
be directed by the Tactical Air Controller or thereceiver leader. The Tactical Air Controller or thereceiver leader (as appropriate) will determine thetype rendezvous to be made. The tanker will adjust to
Change 13

8-23

T.O. 1F- 16A- 1

refueling airspeed when directed by receiver leader.
Receivers will rendezvous 1000 feet below refuelingbase altitude until visual contact is established. In
the event Tactical Air Control radar is not available
to control anchor refueling operations, the following
alternate procedure will be used:
The tanker will establish a normal point parallel
rendezvous at the anchor point. Receiver flights willproceed to an anchor ARIP a minimum of 70 NM
upstream from the anchor point. Receivers willrendezvous 1000 feet below the refueling basealtitude until visual contact is established. Normal
point parallel rendezvous orbit (fighter) procedures
will be used for the rendezvous.

NOTE

Unless directed otherwise by the tanker,
the receiver flight will accomplish a 360
degree left turn at the upstream end of
the anchor pattern to enable the tanker
to turn toward the receiver flight for
rendezvous.
After the receiver flight is joined up, the anchor
pattern will be used for refueling. If cleared by thetanker commander, subsequent receiver flights may
depart the ARIP prior to the previous receiver flights
departing the anchor area when the receivers havethe capability to assure safe aircraft separation andto join on the tanker using receiver turn-on
rendezvous procedures.
Ensure at least 1000 feet separation is provided
between each joining flight, and between the highest
flight and the lowest refueling element until visualcontact is established. Use of secondary frequency isrecommended. To preclude conflict with receivers
clearing the tanker or during a breakaway, ensure
fighters maintain adequate in-trail spacing from therefueling formation.

NOTE

If subsequent flights are cleared todepart the ARIP and air refueling is
not complete prior to join-up, a fighter
turn on rendezvous should be made. If
this is not possible, these fightersshould hold at the ARIP until air
refueling is complete.
8-24

Change 13

Receiver Turn-On Rendezvous

Receiver turn-on rendezvous will be conducted inaccordance with the procedures established incommand guidance. Receivers will maintain required
vertical separation until visual contact is made withthe tanker(s).
Rendezvous Overrun

In the event of an overrun by fighters, the receiverCs)
will pass 1000 feet below the tanker to insure positive
vertical separation. The receiver will decelerate to
290 KCAS and maintain air refueling heading. The
tanker will accelerate to 355 KIAS (350 KCAS) or
Mach 0.90, whichever is lower, and maintain air
refueling heading. When the tanker is in positive
visual contact ahead of the receiver, the tanker will
decelerate to air refueling airspeed and normal
closure procedures will be employed to establish
contact.
Enroute Rendezvous

DEPARTURE AND CLIMB

The receiver departure time will be adjusted to place
him at altitude in trail of the tanker.
The tanker will level off, on course, at the
programmed cruise altitude and establish 260 KCAS
to permit receiver overtake. The receiver will level off,on course, 1000 feet below the tanker's base altitude
and establish a closing airspeed. All available
rendezvous aids will be used to effect tanker/receiverclosure until visual contact is made. Receivers willcall "visual" when visual contact is established withthe tanker. Tankers will accelerate to enroute or
refueling airspeed at the direction of the receiver.
ENROUTE RENDEZVOUS PROCEDURES

An enroute rendezvous may be used when thetanker(s) and receiver(s) fly individual flight plans to
a common rendezvous point (RZ), where join-up isaccomplished, and continue enroute cell formation to
the ARCP.

These procedures may provide an orbit delay ortiming triangle enroute to the ARCP. It is not
appropriate to accomplish a point parallel rendezvous at the RZ because the length of the orbit legs
cannot be extended. Tankers will depart the RZ tomake good the ARCT or the receiver's ETA to the
ARCP.
Either tanker(s) or receiver(s) may be scheduled toarrive at the RZ first, orbit if necessary, and thendepart at a preplanned time.

T.o. 1F- 16A- 1

Racetrack Patterns (Typical)

Anchor Pattern

INBOUND COURSEANCHOR POINT

Anchor Pattern (Alternate Procedure)~-------------~~--------------~----~ARIP

ANCHOR POINT

RECEIVER)HOLDING(
POINT

Enroute Rendezvous (Typical)
I

I

ARCP

100 NM PT

ARIP

I

I
J

/'

"

50 NM

RZ
M----r
6'01:.,'-----. ____------'__--+________+--___1 _40__
N_
'-;T

ARCP

100 NM PT

_~'

\)"

/'

TIMING TRIANGLES\ ............
\........

'0'

,

,

1 F-16X-1-8006X@

Figure 8-7.
Change 13

8-25

T.O. 1 F-16A-1

IWARNING

1

When close interval stream operations
are being conducted, do not use orbit
delays to control timing.
The RZ will be located a minimum of 50 NM prior to
the ARIP/SD. Tracks from the ARIP/SD may beestablished from any direction and need notnecessarily be an extension of the air refueling track.
If orbit delays are required, they will be accomplishedby orbiting at the RZ point along an extension of the
track from the RZ to the ARIP/SD. Orbit in a
racetrack pattern using 30 degree banked turns anda maximum of 15 NM straight legs (unlessoperational directives specify longer straight legs).

NOTE

For peacetime training missions, theARIP or the ARCP may be designatedas the RZ. In these cases, cells will
echelon and start descent at the baserefueling altitude as soon as practicalafter rendezvous completion.
If prebriefed, tanker(s) and receiver(s) may adjust to

air refueling airspeed and begin air refueling after
passing the RZ. Once departing the RZ/ARIP, thetanker(s) should fly centerline. The receiver is the
maneuvering aircraft. If the tanker is behind thereceiver, the tanker should accelerate and passslightly off the left wing of the receiver.
Alternate Rendezvous Procedures

Tanker(s) and receiver(s) will join-up at the RZ bycontrolling timing so they arrive at the RZ at the sametime. Timing to the RZ may be adjusted using
differential airspeeds, orbit delays or timing
triangles. If a planned orbit delay is used, receiver(s)and tanker(s) may accomplish join-up in the orbit.
Assigned altitudes at the RZ will provide at least 1000
feet separation between affected airplanes (highest
tanker and lowest receiver), with the receivers
normally at the highest altitude. If the receiver(s)planned level off altitude is within 30 minutes flying
time from the ARIP, the receiver(s) may level off
below the tanker and maintain an altitude which
provides a minimum of 1000 feet vertical separation
between the highest receiver(s) and the lowest
tanker(s).
Communications will be in accordance with specified
emission option. If radio contact between the airplanes
has not been established prior to the rendezvous
control time, or the adjusted rendezvous control time,
airplanes will maintain altitude and depart the RZ tocross the ARCP at the ARCT. Delays at the ARCP will
use normal orbit procedures unless otherwise directed.
If there is minimal separation between following
aircraft or cells using the same track, orbits at the
ARCP will require close coordination and a thoroughcrew briefing to ensure altitude separation.
When the aircraft or cells pass the ARIP/SD, thetanker(s) and receiver(s) will echelon and thereceiver(s) will begin descent to the base air refueling
altitude. Receiver(s) will descend to be at the basealtitude 80 NM prior to the ARCP. TankerCs) will
maintain published buddy cruise KCAS and adjust to
air refueling airspeed crossing the ARCP.

8-26

Change 13

Tanker and receiver crews must be prepared at all
times to accomplish the rendezvous using whatever
resources are available. When rendezvous equipment
is degraded, tankers and receivers will fly the sameprofiles as described in previous paragraphs. The
following are some suggested alternate rendezvous
procedures which should be used in any combination
to ensure a successful rendezvous:
NOTE

Initial visual contact between thereceiver and tanker may be enhanced,
inflight weather conditions permitting,if the tanker jettisons fuel to increase its
visual signature. This procedure may beinitiated/requested by the tanker, receiver, or the ground agency controlling
the rendezvous. It should only be used if
a receiver low fuel state or other similar
circumstances require the rendezvous
be expedited. If required, the tanker will
dump fuel in 500 - 1000 pound increments until positive visual contact canbe maintained.
• RadarlRendezvous Beacons. The receiver/tanker
beacons may be used for range and offset
information with suitably equipped airplanes.
Depending on equipment capability, one airplane
should maintain the planned outbound or inboundtrack while the other airplane maneuvers toestablish the planned offset. The tanker will clearly
establish which airplane will be maneuvering.

10. 1F- 16A-l• Common Ground Station. If AlA TACAN is notavailable, switching to a common ground TACAN/
VORTAC station for range information may benecessary. The final turn to refueling track is madewhen the DME difference equals proper turn range.
• UHFIDF. For DF steers, receivers will be requested
to use the MIC switch without talking. The receiverwill transmit on the air refueling frequency
approximately 10 seconds out of every 20 second
period, ending each transmission with the receiver's call sign. When the receiver position shows
proper turn range bearing (No Wind) from thetanker heading, the tanker will turn to the refuelingtrack. Notify the receiver when the turn is started.
At the receiver's request, the tanker will transmit
a homing signal.
• ETA. When adequate navigational check points areavailable, the tanker may adjust final orbit pattern
to arrive over the ARCP on the air refueling headingat the receiver(s) ETA to the ARCP.
• Ground Assistance. Ground radar facilities may be
used for vector and separation advisories. FAA
ground radar assistance will be used to the
maximum when conducting rendezvous withsignificantly degraded equipment to ensure a
successful rendezvous.
Emission Option 3• Inbound courses to the RZ/ARIP• Radio silent termination time in the event of a
missed rendezvous
Missed Rendezvous Procedures

If contact is not established at the RZ/ARIP, thetanker will arrive at the ARCP at the ARCT. This
procedure begins when either aircraft arrives at the
ARCP and does not have visual contact with theother. In this case, a left hand orbit should be enteredand orbit controlled so as to be over the ARCP atintervals of every eight minutes (ARCT plus 8, plus
16, etc.). While in the orbit, every attempt should bemade to establish visual contact with the otheraircraft. The length of the delay and decision as tohow long to continue radio silence should bedetermined during mission planning/development
prior to flight.
SPECIAL AIR REFUELING PROCEDURES

The tanker boom is controlled by the boom operator
while the fuel transfer (pressure, flow, quantity, etc.)
is normally controlled by the tanker crew from thepilots' compartment. In IMC, when visibility is suchthat Lost Wingman Procedures may be necessary,receiver formations and the refueling sequence will
be structured so that no more than three aircraft areon each wing of the tanker.

The elimination of the 15 minute prior calls increase
the element of risk, and the following guidelines
should enhance safety considerations:Refueling Sequence• Normally accomplish when clear of clouds.

Normally, the leader will proceed to the precontact
position. Number 2 will proceed to the lead element's
observation position. The second element will proceed
to an observation position on the tanker's opposite
wing. Each subsequent receiver will visually clear
and move from the observation position to theprecontact position. The refueling sequence will be
designated by the receiver leader. Each receiver, after
refueling is -completed, will rejoin to an outside wing
position of his original element. When all receivers
have completed refueling, the receiver force will
rejoin to the left or right, as briefed, and slightly below
the tanker.

• If unable to remain clear of clouds, tanker(s) andreceiver(s) will immediately confirm altitudes.

• The receiver and tanker inbound courses to the
RZ/ARIP must be separated by a minimum of 30
degrees.
• The receiver and tanker inbound legs to the
RZ/ARIP must be a minimum of 40 NM in length.
This type rendezvous should be an enrouterendezvous at the ARIP with both aircraft using thesame RZ time. The receiver should rendezvous 1000
feet below the tanker. An ETE from the ARIP to the
ARCP should be planned which permits an airspeed
which falls in the middle of the aircraft speed
performance envelope. It is essential that crews/planners coordinate certain items during mission
planning/development. Minimum items include:• Rendezvous altitudes• RZ time and ARCTFINGERTIP FORMATIONECHELON FORMATION (VMC ONLy)

Normally, the leader will proceed to the precontact
position. Number 2 will proceed to the observation
position with the remainder of the flight. Refuelingsequence will be as directed by the receiver leader.
Each receiver will visually clear and move from theobservation position to the precontact position. The
receivers, after refueling is completed, will rejoin inechelon formation on the tanker's opposite wing.
Change 13

8-27

T.O. 1 F-16A-l
Fuel Management

The ENG FEED switch should be in the NORM or
proper position, and the fuel distribution will bechecked within flight manual tolerances on the fuelquantity indicator prior to contact with the tanker.
The fuel system operation is automatic (fuel being
distributed to internal and external tanks simultaneously).

NOTE• Upwash and downwash effects may
occur, drawing the aircraft together.
Low pressure areas created by anoverrunning receiver flying under thetanker will affect static ports, causing
possible erroneous airspeed and altitude indications to both aircraft. The
tanker autopilot altitude hold function
may sense the low pressure as aclimbing indication and initiate adescent into the lower aircraft.
Boom and Receptacle Procedures

NOTE• If a partial fuel load is onloaded, a fuelspread in excess of flight manual limits
should be anticipated.

• [ID Disconnect from the boom may occur

before all tanks are full if the externalfuel tank configuration consists of only
a centerline fuel tank. Such a disconnect
typically occurs when refueling with an
initial internal fuel load of 4000 pounds
or more and the centerline tank empty.
At disconnect, the aircraft total fuel may
be up to 1600 pounds less than full, withmany occurrences resulting in approximately 1000 pounds less than full.

When cleared, move forward to the contact position
and the boom operator will make contact. The
receiver may request assistance from the boom
operator in obtaining and maintaining position.
From the pre contact position, the receiver moves
slowly with a 2-3 knot closure until reaching thecontact position. When closing on the boom, constantcross reference between the boom and the tanker
fuselage will alleviate any tendency to "chase"variations of boom trail position due to turbulence.
When stabilized in the contact position, maintain thisposition. The boom operator will then make thecontact.

Precontact

All precontact air refueling checks will be completed in
the observation position or prior to reaching 1 NM in
trail, except for final exterior light adjustment. Afterthe receiver has stabilized in the precontact position,
the receiver will move to the contact position.

I WARNING I• The receiver will stabilize in theprecontact position and attain a zerorate of closure. If the receiver fails to
attain a stabilized position, or it
becomes apparent that a closure overrun will occur, a breakaway will be
initiated. Failure to initiate a breakaway under closure overrun conditions
can result in a midair collision.
8-28

For night operations, prior to closing
for contact with the tanker, coordinate
with the boom operator on exterior
lighting to avoid impairing nightvision .

Change 13• If the receiver director lights fail to
illuminate when contact is established,
the receiver pilot will inform the boom
operator if he wishes to continue
refueling operations. If refueling iscontinued, verbal corrections from theboom operator may be requested.

• Attempts to affect a contact during lossof any air refueling lighting that
results in less than desired illumination will be at the discretion of theboom operator.
To maintain proper contact elevation and boom
extension, refer to the director lights located on thebottom of the fuselage of the tanker (See figure 8-5).
While in contact position, there is freedom in all threeaxes as depicted in figure 8-6.

10. 1 F- 16A- 1

If, for any reason, fuel is not transferring or istransferring at less than normal rate, the receiver pilot
will disconnect and monitor the aerial refueling statusindicator. The bottom lamp (DISC) lights amber whena disconnect has been accomplished. The system willautomatically reset to ready and the top lamp (RDY)
relights blue after a 3-second delay. A second contact
may then be accomplished. If this does not resolve theproblem, the pilot may then disconnect, confirm
disconnect with the boom operator, and recycle thesystem by closing and opening the slipway door using
the AIR REFUEL switch.
DISCONNECT KC- 135

In the event of failure to obtain a contact and after
each disconnect, the receiver will move aft andstabilize in a position in trail of the boom or inprecontact position and await clearance from theboom operator to return to the contact position.

• Remain stabilized in the contact
position until visually confirming a
disconnect has been made. This will
prevent damage to the boom and/or
receptacle through a brute force
disconnect.
• Brute force disconnects can occur
unintentionally as the result of rapidly
exceeding boom limits or failure of thereceptacle toggles to release when adisconnect is initiated.
DISCONNECT KC-1 O/KDC-1 0

The KC-lOIKDC-lO aerial refueling boom is controlled by a digital fly-by-wire system. Certain failure
conditions of this system may-cause one or more axesof the boom control system to become inoperative.
Should this occur, the boom operator may not be able
to maneuver the boom to avoid striking the receiver
airplane. In this situation, the boom operator will
issue instruction to direct the receiver to a position
where a safe disconnect can be effected.

I WARNING I• When notified that a KC-lO/KDC-lO
boom flight control system failure has
occurred, do not initiate a disconnect
unless directed by the boom operator.

• Follow the boom operator's instruction
explicitly. To reduce the probability ofboom strike after disconnect, it may benecessary to remain in a stabilized
position to allow for aerodynamic
fairing of the boom control surfaces.
Another feature of the KC-lOIKDC-lO is the
Independent Disconnect System. This system allows
the boom operator to obtain a disconnect even when
the receiver's toggles remain in the latched position.
This system should be used in lieu of a brute force
disconnect.
Quick Flow Air Refueling Procedures

Refer to figure 8-8. Fighter type receivers may use
Quick Flow procedures to expedite air refueling
operations. Quick Flow allows receivers to minimize
refueling time with maximum fuel transfer. Quick
Flow may be used during day or night operations, in
VMC conditions only. If it appears that the flight may
encounter adverse weather conditions, standard IMC
procedures will be used. Coordination between
tanker(s) and receivers prior to initiation of Quick
Flow procedures is required. Air tasking guidance,
direct communication with the tanker unit, or adding
the term "Quick Flow" to the initial radio call will
satisfy coordination requirements. Tanker lead is the
final authority for Quick Flow operations. Rightechelon formation is normally used for Quick Flow;however, variations are authorized with flight leadcoordination and tanker lead approval.
Normally, the receiver flight will join on the tanker
with the flight lead moving to the precontact position.
Remaining aircraft will proceed to the rightobservation position. Once the flight lead commences
refueling, the second aircraft in the air refuelingsequence will move to the On-Deck position. The
On-Deck position is normally flown as a routeformation with approximately 10' spacing. When theflight lead completes refueling, that aircraft moves toan observation position on the tanker's left wing. The
second receiver moves from the On-Deck position to
the precontact and contact position. With three ormore receivers, the third receiver moves to the
On-Deck position. The right to left flow continues
until all fighters have refueled. When the air
refueling operation is complete, the flight may departthe tanker or, if additional refueling is required,remain in echelon formation on the tanker's left wing
and reverse the Quick Flow procedures, with a left toright flow. The second receiver will assume a left
On-Deck position and Quick Flow will continue inorder. Additional receivers arriving prior to the firstflight completing refueling operations will remain intrail position until they are cleared by the tanker tothe observation and/or precontact position.
Change 13

8-29

T.O. 1 F-16A-1

Quick Flow Air Refueling~~~~~~==~~~~~

1F-16X-1-8007X@

Figure 8-8.

8-30

Change 13

T. O. 1F- 16A-1
In the event of a breakaway, the On-Deck receiver
follows the receiver that was on the boom. Any
receivers on the wing will remain with the tanker. In
the event a breakaway is initiated while a receiver istransitioning from the observation position to the
On-Deck position, that receiver will follow thereceiver that was on the boom.
Toboggan

When altitude and atmospheric conditions result inthrust requirements that exceed the receiver's
available thrust, a toboggan will be necessary. The
toboggan technique is a coordinated effort between
the tanker pilot and the receiver pilot in which
refueling is accomplished in a slight descent, allowing
the receiver to perform the refueling with available
thrust.
The receiver pilot must signal or call on boom
interphone that a toboggan maneuver will be
required before reaching full military power.
The tanker pilot will very gently reduce power andinitiate a rate of descent of approximately 300 FPM
while maintaining the air refueling airspeed
throughout the toboggan maneuver.

SEPARATION FROM A SINGLE TANKER

The tanker and receiver leader will coordinate on themethod of separation. Normally, after the receiver
flight has reformed, they will clear the tanker bydescending or as directed by the controlling agency.
The tanker will advise and receive clearance from thereceiver leader before changing altitude or heading.
Receivers will maintain a safe clearance from thetanker as they proceed on their assigned mission.
CRUISE CELL TERMINATION (VMC)

When cleared by the receiver leader, elements will
join on their respective flight leaders. The receiverforce will then reform to the left and slightly below
the lead tanker. After receiving clearance from thetanker leader and the appropriate controlling
agency, the receivers will proceed on their assigned
missions, maintaining safe clearance from thetanker formation.
CRUISE CELL TERMINATION (IMC)

Mterburner Air Refueling is not recommended.

Refer to figure 8-9. Ten minutes prior to reaching thecell termination point (if the point is other thandestination approach fix), the receivers will reform inleft echelon on the left wing of the tanker(s). Uponreaching the cell termination point, the tanker(s) will
climb straight ahead 3000 feet and then turn to thedesired track, maintaining cell formation. Receivers
will maintain heading, altitude, and airspeed for 3
minutes. At this time, if flight formation rejoin is
impractical, number 1 receiver element will descend
1000 feet below base altitude, number 2 receiver
element will descend 500 feet to the base altitude,number 3 receiver element will maintain altitude(which will be 1000 feet above the base altitude), andnumber 4 receiver element will climb 500 feet to an
altitude which is 2000 feet above base altitude.
Receiver elements will then proceed with their
mission independently.

Separation/Termination Procedures

CELL TERMINATION AT TERMINAL APPROACH FIX

Weather Abort Procedures

Receivers must take every feasible action to enhance
the possibility of completing air refueling. Suchactions include altitude and course deviations
necessary to avoid severe weather. Deviations, whenrequired, must be made judiciously. When thereceiver leader determines that weather conditions
are such as to make formation refueling hazardous,
he may abort the cell. When the cell is to be aborted,
the receiver leader will instruct the tanker leader toclear refueling track. Normal end refueling procedures will apply.
Afterburner Air Refueling

Following completion of air refueling, the receiver(s)will maneuver to the prescribed formation position,
obtain tanker post air refueling report, and return tothe primary refueling frequency (if applicable). Afterthe receivers have reformed, the tanker leader will
provide the receiver leader with present position in
relation to the planned completion point. Additional
information will be provided if requested; i.e.,weather information, nearest abort bases, etc. The
receiver leader will request the no wind heading anddistance to the next checkpoint unless he has apositive fix from which to navigate.

Due to the many possible combinations of tanker/receiver formations, terminal destination weather, andterminal airfield penetration facilities, it is impractical to designate one optimum method for penetration
at the destination. The following methods should beapplied as applicable:
1.

From the final air refueling point, tankers andreceivers can be scheduled at their individual
optimum airspeeds to provide spacing for the
penetration.
Change 138-31

10. 1F- 16A- 1

2.

After the receivers have a positive TACAN
lock-on, they will normally depart the tankers
and proceed to destination as directed by theappropriate controlling agency.

3.

When available, approach control should be
used with enroute descents to obtain aircraft
separation.

4.

The element (one tanker/two receivers) may
penetrate as a unit. Weather minimums forthis type approach are 2500 feet and 3 NM.
NOTE

A low fuel altitude will be designated2000 feet below base altitude for
immediate descent of receivers withlow fuel or an emergency condition.
Receiver altitude changes will becoordinated by the receiver cell leader
with air traffic control. All aircraft ofthe cell will note individual altimeter
errors at the cruise altitude with 29.92
inches Hg set on the altimeter and fly
their assigned altitudes after cell
separation with these errors applied.

8-32

Change 13

ELEMENT PENETRATION
If conditions exist which necessitate a more
expeditious recovery (fuel shortage, emergency, etc.),
a basic cell penetration may be made. Penetration
airspeed and descent rate will be coordinatedbetween the tanker and receiver leader. When VFR,
the receivers will break off and enter initial for a VFR
landing.

RECEIVER RADAR REJOIN PROCEDURES (IMC)
If receiver radar rejoin is desired at the completion of
the cell termination, the following procedure will be
initiated:
1.

Each receiver will maintain his respective
altitude.

2.

The receiver formation leader will maintainheading and each of the following receiver
elements will simultaneously turn left 15
degrees on the formation leader's command.

3.

Numbers 2, 3, and 4 elements will maintainthis heading for 1, 2, and 3 minutes
respectively, and then resume the original
heading.

4.

The flight will then rejoin on radar using theprocedures for radar join-up with tankers.

T.O. 1F- 16A- 1

Cell Termination Procedure~;-,---------------~::::.::.::.::.--------'---,.:,------'-.. - - - -- - - - - - - - - - - ;,.-----------,~:;Z - - - ' ---~ - - - - - - - - - - - - - - '-------~~-

-

, ,

c;--

,------------, '-'-- ---- -,.:,----------- -------'"1-,-,,.-----------" !.

'-------\.~----->----,-------,------- -----------\ -~::=~J' - - ,30,000- - - - - - - - 29,000

,

-.r~

',NO.3,28,000

',NO,

tuw

LL

---~

2a j~~---50-rfO~FT--{----

I

w

o

:J

W - - - - - ~-::::_-J' _-----'tL-...__ - - - - - 27,000 «~I-

BASE ALTITU DE

\ NO.

~

1
1000 FTj----------'------~~::.::_,:J' - - - - - - - - - - - - - - - - - - - - - - 26,000-c.;-_-=_----;' , LOW FUEL,,-"ALTITUDE- 25,000

ALTITUDES LISTED ARE FOR SAMPLE ONLY USING 27,000 FT AS A BASE ALTITUDE.

1F-16X-1-8005X@

Figure 8-9.

Change 138-33

10. 1 F- 16A- 1
NORMAL AIR REFUELING PROCEDURES
Armament Safety CheckIf making an outer limit disconnect,high separation rates should be
avoided to prevent damage to the boomor receptacle.

Prior to closing within lethal range of the tankers,
complete the following checks:
1.

MASTER ARM switch - OFF or SIMULATE.

2.

LASER ARM switch - OFF.

3.

SMS - Confirm ordnance safe.

4.

CHAFF/FLARE switches (4) - OFF.

2.

Remain stabilized in the contact
position until positive visual confirmation of boom separation is confirmed by
the boom operator.

Precontact

Prior to air refueling, the following checks will becompleted:
1.

TACAN - As required.

2.

Emitters (ECMlRadarIRDR ALT) - As required (Quiet/SilentlSTBY/OFF).

3.

HOT MIC CIPHER switch - HOT MIC.

4.

Exterior lights (Night) - DIM, STEADY.

5.

ANTI COLLISION switch (Night) - OFF.

6.

AIR REFUEL switch - OPEN.

7.

AR status indicator light - RDY.

AR status indicator light - DISC.

Post Air Refueling1.

AIR REFUEL switch - CLOSE.

Failure to close the air refueling door
will result in the FLCS remaining intakeoff and landing gains, the roll rate
restricted to a fixed value, and thefail ure of external fuel to transfer.

Contact

2.

AR status indicator lights (3) - Off.

1.

AR status indicator light - ARINWS.

3.

Fuel quantity - Check.

NOTE

4.

MASTER ARM switch - As required.

Once contact is made, boom interphone communications can be established with the boom operator if HOT
MIC is selected. Volume is controlled
by the intercom volume control. Theboom interphone capability is provided on all KC-10IKDC-10 and
KC-135 tankers.

5.

SMS - As required.

6.

CHAFF/FLARE switches (4) -As required.

7.

ECM - As required.

8.

TACAN - As required.

9.

Radar - As required.

10.

RDR ALT - As required.

11.

LASER ARM switch - As required.

12.

Exterior lights .- As required.

2.

Fuel transfer - :Monitor.

Disconnect

1.

8-34

NR DISC button - Depress momentarily,then release.

Change 13

'"""""""",~''"'''''''-'-'-'-'-'"'"'"'"EMERGENCY AIR REFUELING PROCEDURES.
Breakaway Procedures

Relative position of both airplanes must be closely
monitored by all crew members during all phases of
air refueling. When either a tanker or receivercrewmember determines that an abnormal condition
exists which requires an immediate separation of the
airplanes, that crewmember will transmit the
breakaway call on air refueling frequency. Abnormal
conditions include excessive rate of closure, closure
overrun, and engine failure. The receiver does nothave to be in the contact position to call a breakaway.
For all breakaways, transmit the tanker's call signand the word "breakaway" three times (Example:"Chevy 2, breakaway, breakaway, breakaway") andsimultaneously take the following actions:• Actuate disconnect switches as applicable.
• Retard throttle and establish a definite rate ofdescent, using speed break if necessary.
• If possible, drop aft of tanker until entire tanker isin sight and monitor flight instruments.
The tanker pilot will increase power to obtain forward
separation. Unless lateral separation cannot be
assured, the tanker will accelerate in level flight andwill not climb. The lower rotating beacon will beturned on, the pilot director lights will be flashed, andthe RadarlRendezvous Beacon will be turned tooperate, if appropriate. When the receiver is wellclear, the breakaway may bt~ terminated. The receiverpilot will be notified of and will acknowledge any
reduction in power by the tanker to resume air
refueling speed. If a climb is required, the tanker pilot
will disengage the autopilot and climb straightahead. If in a turn, the tanker will maintain theestablished bank angle until the receiver is well clear.

10. 1F-16A-l~~• With certain gross weights and aircraft configurations, the tanker rate ofacceleration on a breakaway may
exceed the rate of acceleration for thereceiver aircraft in the observation~~~

syste::~:::nctions~

When any system malfunction or condition exists
which could jeopardize safety, air refueling will not be
accomplished except during fuel emergencies or whencontinuance of fueling is dictated by operational
necessity. At any time fuel siphoning is noticed, fueltransfer will be stopped and the receiver notified. The
requirement to continue fuel transfer will be at thediscretion of the receiver pilot.
NOTE

A small amount of fuel spray from thenozzle and receptacle during fueltransfer does not require fuel transfer tobe terminated. The receiver pilot should
be notified if this condition exists andthe air refueling operations will be
continued or discontinued at his discretion.
SLiPWAY DOOR WILL NOT OPEN

~
~
~
~
~~

No back-up system is provided to open or close theslipway door if hydraulic system B fails.

.."1IIIIIIIISLiPWAY DOOR WILL NOT CLOSE~

If the slipway door will not close, perform thefollowing:

NOTE• If a breakaway is called prior to any
receiver reaching the observation
position, the entire receiver flight will
execute the breakaway procedure. If a
breakaway is called after receiver(s)have reached the observation position,
only the receiver in the contact orprecontact position will execute thebreakaway procedure. The receiver(s)in the observation position will maintain formation on the tanker.

~

1. AR switch - CLOSE.
Normal FLCS gains and tank pressures will
be regained.
NOTE

The RDY, ARJNWS, and DISC lights
will not indicate normally. The NWS
light will not illuminate when nosewheel steering is engaged.

Change 13

8-35

~

~

~

~~
~

~
~~
~

~"I'I'~~"'I"""""'~.,.

10. 1 F-16A-l.,. ::E::::~:::i::~:::l:f~:~t::::er have

~

"~

~
~~

~~

failed, and an actual fuel shortage emergency aboard
the receiver airplane exists, fuel can be transferred bymaintaining boom/receptacle contact using a slight
extend pressure on the boom telescope lever. Unusualand varying trim changes may be required of bothtanker and receiver airplanes.

Brute Force Disconnect

There are two types of brute force disconnects:
Inadvertent and Controlled Tension.

NOTE
Enter any brute force disconnect as adiscrepancy in the AFTO Form 781.
The entry will specify which type ofbrute force disconnect occurred.

~ If a fuel shortage emergency requires:
Boom operator - Inform of the need toaccomplish manual boom/receptacle pres-1.

surerefut:ARNING

~
~
~
~

I

The receiver pilot must inform thetanker he is ready to receive fuel andcoordinate the disconnect cycle for theconclusion of refueling.

Prior to attempting this method oftransferring fuel, the boom operator
will brief the receiver pilot andthoroughly coordinate the procedures
to be used. Both tanker and receivercrews will monitor the refueling withextreme caution.

~ KC-l O/KDC-l 0 Boom FLCS Failure~...

Do not disconnect until cleared by boom operator.

~~~

I WARNING I~~INADVERTENT DISCONNECT
An inadvertent brute force disconnect is defined as
any unplanned disconnect which is the result of oneof the following:• The receiver aircraft moving rapidly to the aft limit,
causing mechanical tanker/receiver separation.
• Boom pullout occurs at 38 degrees elevation or
below.

Following an inadvertent brute force
disconnect, air refueling will be terminated except during fuel emergencies
or when continuation of air refueling is
dictated by operational necessity.
CONTROLLED TENSION DISCONNECT

A controlled tension brute force disconnect is defined
as an intentional, coordinated disconnect occurring
above 38 degrees elevation, accomplished by gradualmovement of the receiver aircraft to the aft limit, andending with a smooth tension boom pullout.
Coordination between the receiver pilot and the boom
operator is required to ensure as smooth a disconnect
as possible.
1. Slide out boom with gradual power reduction.
2.

• When notified that a KC-IO/KDC-IO
boom flight control system failure has
occurred, do not initiate a disconnect
unless directed by the boom operator.
• Follow the boom operator's instruction
explicitly. To reduce the probability ofboom strike after disconnect, it may benecessary to remain in a stabilized
position to allow for aerodynamic
fairing of the boom control surfaces.

8-36

Change 13

When at full boom extension, tension
disconnect will occur with slight power
reduction.

A controlled tension brute force disconnect will be accomplished only as a lastresort, after all other normal andemergency methods of disconnect havefailed.

~""""""""'~~~~SSS
S• Fly stabilized at contact altitude until
certain the nozzle is clear of thereceptacle and slipway.

• The receiver pilot must not jerk theboom out with rapid thrust change
toward idle or by using speedbrakes; todo so may cause serious structural
damage. Gradual power reduction will
suffice to effect a disconnect.

• Air refueling for the receiver which
required controlled tension disconnect
will be terminated except during fuelemergencies or when continuation of
air refueling is dictated by operational
necessity.

Change 13

8-37/(8-38 blank)

~
~~

~
~
~
~
~
~
~
~
~
~S
~

T. o. 1 F- 16A- 1

GLOSSARY
STANDARD AND NONSTANDARD ABBREVIATIONS

A
AlA
AAM
AB
ac,AC
AlCGW
ACM
ACMI

ADC
ADG
ADI
AFI
AFTO
AGL
AGM
AIFF

I

•
I

•

AIM
AIS
AJ
AL
ALOW
ALT
ALTRAVAMAMMO
AMRAAM

ANT
AOA
AP
AR, AIR
ARCP
ARCT
ARI
ARIP
ARMT
ATT
AUX
AVTR

Air to Air
Air-to-Air Missile
Mterburner
Alternating Current
Aircraft Gross Weight
Air Combat Maneuvering
Air Combat Maneuvering Instrumentation
Air Data Converter
Accessory Drive Gearbox
Attitude Director Indicator
Air Force Instruction
Air Force Technical Order
Above Ground Level
Air-to-Ground Missile
Advanced Identification, Friend or
Foe
Air Intercept Missile
Aircraft Instrumentation System
Antijamming
MtJLeft
Automatic Low Altitude Warning
Altitude or Altimeter or Alternate
Altitude Reservation
Amplitude Modulation
Ammunition
Advanced Medium Range Air-to-Air
MissileAntennaAngle of Attack
Autopilot
Air Refueling
Air Refueling Control Point
Air Refueling Control Time
Aileron -Rudder Interconnect
Air Refueling Initial Point
Armament
Attitude
Auxiliary
Airborne Videotape Recorder
B

BAKBARO
BATT
BDU
BIT
BL
BLU
BSU
BUC

Arresting Cable Prefix (e.g., BAK-9)
Barometric
Battery
Bomb Dummy Unit
Built-In Test or Binary Digit
Buttock Line
Bomb Live Unit
Bomb Stabilizing Unit
Backup Fuel Control

C

CADC
CARA
CBU
CCIP
CCRP
CCW
CDI
CENC
CG
CHAN
CIU
CIVV'sCONFIG
CONT
CO2
CRIU
CRS
CRV
CSD
CTVS
CW
CWI

Central Air Data Computer
Combined Altitude Radar Altimeter
Cluster Bomb Unit
Continuously Computed Impact
Point
Continuously Computed Release
Point
Counterclockwise
Course Deviation Indicator
Convergent Exhaust Nozzle Control
Center of Gravity
ChannelCentral Interface Unit
Compressor Inlet Variable Vanes
Configuration
Control
Carbon Dioxide
Conventional Remote Interface Unit
CourseCanadian Rocket Vehicle
Constant-Speed Drive
Cockpit Television Sensor
ClockwiseContinuous Wave IlluminatorDdBA
dc, DC
DEEC
DESIG
DF
DI
DIA
DIFF
DIS
DISC
DME
DN
DNS
DSGRS
DTOS
DTU
DTS
DVAL
DWAT

Adjusted (human ear response)
Decibels
Direct Current
Digital Electronic Engine Control
DesignatelReturn to Search
Direction Finding
Drag Index
Diameter
Differential
DisableDisconnect
Distance Measuring Equipment
Down
Doppler Navigation System
DesignatelReturn to Search
Dive Toss
Data Transfer Unit
Digital Terrain System
D-Value
Descent Warning Mter Takeoff
Change 13

Glossary 1••

T.O. 1F-16A- 1

E
EAS
ECA
ECM
ECP
ECS
EDU
EEC
EED
ELECT

•

ELEV
EMCON
EMER
ENG
EO
EPAF
EPU
EQUIP
EST
EXT

Equivalent Airspeed
Electronic Component AssemblyElectronic Countermeasures
Engineering Change Proposal
Environmental Control System
Engine Diagnostic Unit
Electronic Engine Control
Electroexplosive Device
Electronic (primary altimeter operating mode)
Elevation
Emission Control Option
Emergency
EngineElectro-Optical
European Participating Air Force
Emergency Power Unit
Equipment
Estimate
ExternalF•

FAA
FALTACK
FC
FCC
FCNP
FCR
FCS
FF
FFAR
FFP
FLCC
FLCP
FLCS
FM
FO
FOD
FORM
fpm, FPM
FR
ft, FT
FTIT
FWD

Federal Aviation Administration
Fault Acknowledge
Flight Control
Fire Control ComputerFire Control/N avigation Panel
Fire Control Radar
Flight Control System
Fuel Flow
Folding Fin Aircraft Rocket
Fuel Flow Proportioner
Flight Control Computer
Flight Control Panel
Flight Control System
Frequency Modulation
Foldout
Foreign Object Damage
Formation
Feet per MinutefFlightpath Marker
ForwardJRight
Feet
Fan Turbine Inlet Temperature
ForwardGND
GP
GPS
GRDCUSGS
GW

H
HDG
HDGSEL
HF
Hg
HSI
HUD
HYD
HYDRAZNHz

Glossary 2

Force of Gravity
GallonGuided Bomb Unit
Ground Controlled Approach
Ground Collision Avoidance System
Generator
Ground Map

Change 13

HeadingHeading SelectHigh Frequency
Mercury
Horizontal Situation Indicator
Head -Up Display
HydraulicHydrazine
Hertz
I

Indicated Airspeed
In Accordance With
International Civil Aviation Organization
Identification
ID
Identification, Friend or Foe
IFF
Instrument Landing System
ILS
Instrument Meteorological CondiIMC
tions
Instrument Mode Select Panel
IMSP
in., IN.
InchesIncrease
INC
Indicator
IND
Inoperative
INOP
Inertial Navigation Set (or System)
INS
INST, INSTR Instrument
Intensity or Internal or Interval
INT
Inertial Navigation Unit
INU
Identification of PositionlIP
Integrated Servoactuator
ISA
lAS
lAW
ICAO

J
JETT/JTSN
JFS
JOAP

G
g,G
gal, GAL
GBU
GCA
GCAS
GEN
GM

GroundGroup
Global Positioning System
Gulf Range Drone Control Upgrade
SystemGlide Slope
Gross Weight

JettisonJet Fuel Starter
Joint Oil Analysis Program
K

K
KCAS
KEAS
KIAS
KT{S)
KTAS
KVA

Thousand (e.g., 40K = 40,000)
Knots Calibrated Airspeed
Knots Equivalent Airspeed
Knots Indicated Airspeed
Knot(s)
Knots True Airspeed
Kilovolt AmpereI10. 1 F- 16A-1

L
L
LADD
LAU
lb, LB
LBIHR
LBIMIN
LCO
LCOS
LD
LE
LEF'S
LG
LMLG
LOC
LOD
LOX
LPU
LRU
LTS
LWD

Left
Low Angle Drogue Delivery
Launcher Armament Unit
Pound(s)Pounds per Hour
Pounds per Minute
Limit Cycle Oscillation
Lead Computing Optical Sight
Load or Low Drag
Leading Edge
Leading Edge Flaps
Landing Gear
Left Main Landing Gear
Localizer
Light-Off Detector
Liquid Oxygen
Life Preserver Unit
Line Replaceable U ni t
LightsLeft Wing Down

M
M
MAAS
MAC
MAJCOM
MAL
MAL&IND
MAN.
MAU
MAX
MAXAB
mb
MECH
MEM
MFC
MFL
MHz
MIC
MIL
MIN
MK
MLG
mm
MPBA
MPO
MRKBCNmsec
MSL
MUXBUS
MWOD

•

.Mach
Mobile Aircraft Arrestment System
Mean Aerodynamic Chord
Major Command
Malfunction
Malfunction and Indicator
ManualMiscellaneous Armament Unit
Maximum
Maximum MterburnerMillibar
Mechanical
MemoryMain Fuel Control
Maintenance Fault List
Megahertz
Microphone
Military
Minute or MinimumMark (Equivalent of Model)
Main Landing Gear
Millimeter
Multiple Practice Bomb Adapter
MANUAL PITCH OverrideMarker Beacon
Milliseconds
Missile or Mean Sea Level
Multiplex Bus
Multiple Word of Day

N
NA
NAM
NAVAID
NLG
NM,nmNo., NO.
NORMNOZPOS
NVG
NWS

Not Applicable
Nautical Air Miles
Navigation Aid
Nose Landing Gear
Nautical Miles
NumberNormalNozzle Position
Night Vision Goggles
N osew heel Steering

•
•

0

OAT
OCP
OHEAT
OP
OPT
OVRD
OXY

Outside Air Temperature
Organic Change Proposal
Overheat
Operational or Optimum
Optional
Override
Oxygen
P

PBG
PIDS
PFL
PLD
PMG
PNEU
PNL
pph, PPH
PRE
PRESS.
PRI
PSA
psi, PSI
PTO
PWR

Pressure Breathing for g
Pylon Integrated Dispense System
Pilot Fault List
Personnel Lowering Device
Permanent Magnet Generator
Pneumatic (secondary altimeter
operating mode)
Panel
Pounds per Hour
PresetPressure, Pressurization
Primary
Pneumatic Sensor Assembly
Pounds per Square Inch
Power Takeoff (shaft from engine
gearbox to ADG)
Power
Q

QTY

Quantity
R

RAD
RCFI
RCR
RCVV
RDR
RDY
REL
REO
RER
RETSRCHRadio (e.g., RAD 1 or RAD 2)
Radio ChannellFrequency Indicator
Runway Condition Reading
Rear Compressor Variable Vanes
Radar
Ready
ReleaseRadarlElectro-Optical
Radial Error Rate
Return to Search

Change 13

Glossary 3

T. O. 1F- 1 6A- 1
RIT
RMLG
RNDS
RNG
rpm, RPM
RS
RSVR
RT
RV
RWD
RZ
RZIP

I

Reduced Idle Thrust
Right Main Landing Gear
Rounds (gun)Ranging
Revolutions per Minute
Return to Search
Reservoir
RetardedReceive Variable
Right Wing Down
Rendezvous
Rendezvous Initial Point
S

SAl
SCP
SD
SEC
SEL
SEAWARS
SFO
SIF
SL
SMS
SNSR
SPDBRKSPL
SQSTA
STAPAC
STBY
STD
SUU
SV
SW
SYM
SYS

•

Standby Attitude IndicatorStores Control Panel
Start Descent Point
Secondary Engine Control
SelectSeawater Activated Release System
Simulated Flameout Landing
Selective Identification Feature
Sea Level
Stores Management System
SensorSpeedbrakeSound Pressure Level
Squelch or Square
StationStabilization Package
StandbyStandard
Suspension Utility Unit
Secure Voice
SwitchSymmetrical
System
T

TACAN
TAS
TCN
TCTO
TEF's
TEMP

Glossary 4

Tactical Air Navigation
True Airspeed
TACAN
Time Compliance Technical Order
Trailing Edge Flaps
Temperature

Change 13

TER
TEU
TGM
TGT
THEO
TISL
T.O.
TOD
TR, TIR
TRV
TT
TV
TVS
TWS

Triple Ejector Rack
Trailing Edge Up
Training Guided Missile
TargetTheoryTarget Identification Set, Laser
Takeoff
Time of Day
TransmitlReceive
TravelTotal Temperature
Television
Television Sensor
Threat Warning System
U

UFC
UHF
UNK
UWARS

Unified Fuel Control
Ultra High Frequency
Unknown
Universal Water Activated Release
System
V

VAC
VDC
VHF
VIP
VMC
VMS
VOL
VVI

Volts acVolts dc
Very High Frequency
Visual Initial Point
Visual Meteorological Conditions
Voice Message System
VolumeVertical Velocity Indicator
W

WI
WB
W/O

WOD
WOW
WPN
WPNREL
wt,WT

With
WidebandWithout
Word of Day
Weight on Wheels
WeaponWeapon(s) Release
WeightYY

Yaw

10. 1 F- 16A- 1

TERMS/SYMBOLSPo

Density at Sea Level, Standard Day()

Density Ratio~

Angle of Sideslip

And

Fe

Elevator Stick Force (lb)<i>Roll Rate (Deg/Sec)

Fa

Aileron Stick Force (lb)

1jJ

Yaw Rate (Deg/Sec)

Fp

Rudder Pedal Force (lb)Ay

Lateral Acceleration (FtJSec2)

N2

eEngine Compressor RPM or Nitrogen

Pitch Rate (Deg/Sec)P

PressureaAngle of Attack (Deg)

T

Ambient Air TemperaturePo

Pressure at Sea Level, Standard
Day

W

Specific Weight (LblFt 3)

PtTotal Pressure

Vc

Calibrated Airspeed

Ps

Static Pressure or Specific Energy
Rate

Vt

True Airspeed~H

Delta Change Altitude (Ft)

If

Altitude Rate (FPS)An

Normal Acceleration (G's)

aoSpeed of Sound at Sea Level, Standard Day&

qc

Impact Pressure (Pt - Ps)

Change 13

Glossary 5/(Glossary 6 blank)