F16C/D. FLIGHT MANUAL (2002) - page 57

T.O. GR1F16CJ11

Change 8A229

50

40

30

20

10

0

-10

-20

C

F

120

100

80

60

40

20

0

0 1 2 3 4 5

6

RUNW

A

Y

20 22 24 26 28 30 32 34 36 38 40 42 44

GW   1000 POUNDS

BRAKE ENERGY ABSORPTION   MILLION FOOT-POUNDS PER BRAKE

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

1

2

3

4

5

GR1F-16CJ-1-1-5026A37

Brake Energy Limits    Max Effort Braking

With Drag Chute

DATA BASIS ESTIMATED

ENGINE F100-PW-229

CONFIGURATION:

ALL DRAG INDEXES

SPEEDBRAKES   OPEN

TEF'S DOWN

EXCEEDING 24.5 MILLION
FOOT-POUNDS PER BRAKE
CUMULATIVE TOTAL ENERGY
MAY RESULT IN LOSS OF
BRAKING.

DEPLOYING DRAG CHUTE

AT AIRSPEEDS GREATER
THAN 170 KIAS MAY

RESULT IN LOSS OF THE
CHUTE CANOPY.

TEMPERA

TURE

DRAG CHUTE DEPLOYED

BELOW 170 KIAS

DRAG CHUTE RELEASED

PRIOR TO TAXI

ADD TAILWIND COMPONENT
OR SUBTRACT ONE-HALF

HEADWIND COMPONENT
FROM AIRSPEED WHEN
BRAKES ARE APPLIED.

FOR ABORTED TAKEOFF

AT AIRSPEED GREATER
THAN 100 KNOTS, ADD
2 MILLION FOOT-POUNDS

PER BRAKE IF BRAKES ARE

APPLIED SOONER THAN 4
SECONDS AFTER THROTTLE

IS RETARDED TO IDLE.

IF LANDING WITH ASYM-
METRICAL WING LOADING,

TAKE ACTION AS APPLIC-
ABLE FOR NEXT HIGHER

ENERGY ZONE TO ALLOW
FOR UNEQUAL BRAKE
ENERGY DISTRIBUTION.

46 48

24

NOTES:

CONDITIONS:

NORMAL IDLE THRUST

Figure

 A212.(Sheet 2)

T.O. GR1F16CJ11

A230Change 7

Brake Energy Limits   Max Effort Braking

ACTION TO BE TAKEN AS APPLICABLE TO THE AMOUNT OF BRAKE ENERGY ABSORBED

DANGER ZONE

1. REFER TO T.O. GR1F16CJ1, HOT BRAKES.

2. USE MODERATE BRAKING BELOW 25 KNOTS

GROUNDSPEED AND MAINTAIN FORWARD

MOTION.

3. THE SIDE AREA WITHIN 300 FEET OF THE MLG TIRES

SHOULD BE REGARDED AS UNSAFE FOR 45 MIN

UTES AFTER AIRCRAFT HAS STOPPED UNLESS THE

FUSIBLE PLUGS HAVE RELIEVED TIRE PRESSURE.

4. HYDRAULIC FLUID OR TIRE FIRE IS IMMINENT.

APPROACH MLG FROM FRONT OR REAR FOR FIRE

FIGHTING PURPOSES ONLY. APPLY EXTINGUISH

ING AGENT AS FOG OR FOAM DIRECTLY ON THE

WHEELS.

CAUTION ZONE

1. REFER TO T.O. GR1F16CJ1, HOT BRAKES.

2. THE SIDE AREA WITHIN 300 FEET OF THE MLG TIRES

SHOULD BE REGARDED AS UNSAFE FOR 45 MIN

UTES AFTER AIRCRAFT HAS STOPPED UNLESS THE

FUSIBLE PLUGS HAVE RELIEVED TIRE PRESSURE.

3. DO NOT ATTEMPT TAKEOFF UNTIL BRAKE HOUS

INGS, WHEEL RIMS, AND TIRES HAVE COOLED

ENOUGH TO PERMIT AT LEAST 15 SECONDS OF

CONTINUOUS BARE HAND CONTACT. THIS COOL

ING PERIOD IS TO PREVENT POSSIBLE TIRE FAILURE

DURING TAKEOFF OR IN FLIGHT.

NORMAL ZONE

1. DO NOT SET PARKING BRAKE IF AIRCRAFT HAS

FLOWN IN THE PAST 2 HOURS OR IF CUMULATIVE

TOTAL ENERGY EXCEEDS 9 MILLION FOOT

POUNDS PER BRAKE. OTHER ACTION TO BE TAKEN

IS DETERMINED BY CUMULATIVE TOTAL ENERGY

AND AMBIENT TEMPERATURE, AS INDICATED BY

SAFE TIRE BEAD TEMPERATURE CHART.

2. IF CUMULATIVE TOTAL ENERGY EXCEEDS THE LIMIT

FOR SAFE TIRE BEAD TEMPERATURE:

A. DO NOT ATTEMPT TAKEOFF UNTIL BRAKE

HOUSINGS, WHEEL RIMS, AND TIRES HAVE

COOLED ENOUGH TO PERMIT AT LEAST 15

SECONDS OF CONTINUOUS BARE HAND

CONTACT. THIS COOLING PERIOD IS TO PRE

VENT POSSIBLE TIRE FAILURE DURING TAKE

OFF OR IN FLIGHT.

B. A COOLING PERIOD OF APPROXIMATELY 30

MINUTES PLUS AN ADDITIONAL 30 MINUTES

FOR EACH 1 MILLION FOOTPOUNDS BRAKE

ENERGY ABOVE THE LIMIT FOR SAFE TIRE BEAD

TEMPERATURE IS REQUIRED.

3. IF CUMULATIVE TOTAL ENERGY DOES NOT EXCEED

THE LIMIT FOR SAFE TIRE BEAD TEMPERATURE AND

THE AIRCRAFT HAS NOT FLOWN IN THE PAST 2

HOURS, PARKING BRAKE MAY BE SET AND NO

BRAKE COOLING IS REQUIRED PRIOR TO SUBSE

QUENT TAKEOFF. DETERMINE MAXIMUM BRAKE

APPLICATION SPEED IN CASE SUBSEQUENT TAKE

OFF IS ABORTED BY SUBTRACTING CUMULATIVE

TOTAL ENERGY FROM THE 23.5 MILLION FOOT

POUNDS.

Figure A212.(Sheet 3)

T.O. GR1F16CJ11

Change 7A231/(A232 blank)

1F-16X-1-1-1012A

-20

0

20

40

60

80

100

120

-30

-20

-10

0

10

20

30

40

50

AMBIENT TEMPERATURE    C

AMBIENT TEMPERATURE    F

CUMULA

TIVE TOT

AL ENERGY

(STOPPING ENERGY PLUS T

AXI ENERGY)

MILLION FOOT-POUNDS PER BRAKE

10

9

8

7

6

5

4

L

Safe Tire Bead Temperature     For Quick

Turnaround

COOLING

PERIOD

REQUIRED

NO COOLING

PERIOD

REQUIRED

Figure A213.

T.O. GR1F16CJ11

Change 7A31

PART 3 - CLIMB

TABLE OF CONTENTS

Page

Ground Operation Fuel

Consumption

A31

. . . . . . . . . . . . . . . . . . . . . . . . . . 

Climbout Fuel, Time, and

Distance

A31

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 

Cruise Ceilings and Optimum

Cruise Altitude

A31

. . . . . . . . . . . . . . . . . . . . . . . . 

MIL Climb

A32

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 

MAX AB Climb

A32

. . . . . . . . . . . . . . . . . . . . . . . . . . . 

LIST OF CHARTS

Figure

Page

MIL Climbout Fuel,

Time, and Distance

A31

A33

. . . . . . 

. . . . 

MAX AB Climbout Fuel,

Time, and Distance

A32

A34

. . . . . . 

. . . . 

Cruise Ceilings and

Optimum Cruise

Altitude

A33

A36

. . . . . . . . . . . . . . . . 

. . . . 

MIL Climb-Fuel

Consumed

A34

A37

. . . . . . . . . . . . . . 

. . . . 

MIL Climb-Distance

and Time

A34

A38

. . . . . . . . . . . . . . . 

. . . . 

MAX AB Climb-Fuel

Consumed

A35

A39

. . . . . . . . . . . . . . 

. . . . 

MAX AB Climb-Distance

and Time

A35

A310

. . . . . . . . . . . . . . . 

. . . . 

Data needed to plan for MIL and MAX AB climbs is

contained in this part, including climbout from take

off, climb to cruise, and ceiling altitudes. Refer to Part

8 for information regarding combat ceiling and climb

to combat ceiling.

GROUND OPERATION FUEL

CONSUMPTION

Idle fuel flow is approximately 20 pounds per minute.

An average of 25 pounds per minute is used for

ground operation.

CLIMBOUT FUEL, TIME, AND

DISTANCE

Figures A31 and A32 contain data describing fuel,

time, and distance from end of runway brake release

to climb airspeed. Effects of temperature, GW, alti

tude, and drag index are shown. A constant throttle

position (MIL or MAX AB) from brake release to MIL

or MAX AB climb speed is used. After takeoff, a

constant pitch attitude of 12 degrees is held until

2500 feet AGL. A level acceleration to climb speed is

then made. In some cases, climb airspeed will be

reached prior to gaining 2500 feet AGL. This tech

nique was developed for performance calculations

only and not as an operational procedure.

REFER TO FIGURES A31 AND A32.

Enter chart with temperature (A), proceed horizon

tally to altitude (B), and then down to intersect GW

(C). From there, proceed horizontally to drag base

line and parallel nearest guideline to drag index (D).

Finally, proceed horizontally to read fuel consumed

(E), time (F), and distance (G).

SAMPLE PROBLEM (MIL, FIGURE A31).

A. Temperature

= 40

°

C

B. Altitude

= 2000 feet

C. GW

= 33,000 pounds

D. Drag index

= 150

E. Fuel consumed

= 385 pounds

F. Time

= 2.1 minutes

G. Distance

= 9.0 nm

CRUISE CEILINGS AND OPTIMUM

CRUISE ALTITUDE

MIL cruise ceiling, MIL service ceiling, and optimum

cruise altitude, are shown in figure A33. All data is

based on use of optimum cruise mach number. A

correction factor to adjust MIL cruise ceiling to ser

vice ceiling is given on the chart.

REFER TO FIGURE A33.

Enter upper portion of the chart with GW (A), proceed

upward to drag index (B), and then proceed to the left

to read cruise altitude (C). Enter lower portion of chart

with GW (A), proceed upward to drag index (B), and

then proceed to the left to read MIL cruise ceiling (D).

T.O. GR1F16CJ11

A32Change 7

SAMPLE PROBLEM.

A. GW

= 33,000 pounds

B. Drag index

= 230

C. Optimum cruise

altitude

= 33,110 feet

D. MIL cruise ceiling

(300 fpm)

= 34,180 feet

E. MIL service ceiling

(100 fpm)

= 34,850 feet

MIL CLIMB

Figure A34 contains MIL climb data. Fuel consumed

data is shown in sheet 1 and time and distance data

is shown in sheet 2. The data is for climbs starting at

sea level, but performance data for climbs from any

altitude to a higher altitude may also be determined.

The climb schedules are defined by airspeed/mach

number. Climb at the scheduled airspeed to the

scheduled mach number, then maintain the mach

number to the desired altitude. When starting a climb

at an altitude above the airspeed/mach number tran

sition point, climb at the scheduled mach number.

The schedules are tabulated as a function of drag in

dex and are selected to maintain maximum fuel effi

ciency while still providing near maximum rate of

climb. To obtain data for climb to cruise ceiling, use

the dashed cruise ceiling drag index lines.

REFER TO FIGURE A34.

Enter sheet 1 of chart at initial GW (A), proceed hori

zontally to final altitude (B), vertically to drag index

(C), and horizontally to air deviation temperature

(D). Continue horizontally to read fuel consumed (E).

If initial altitude is above sea level, reenter chart at

initial GW (A), proceed horizontally to initial altitude

(F), and continue as above to read fuel used (G). The

difference between fuel consumed to final altitude

and fuel used to initial altitude is the fuel used to

climb from initial to final altitude. Climb time and

distance are found in a similar manner from sheet 2.

SAMPLE PROBLEM.

MIL climb to optimum cruise altitude.

A. Initial GW

= 33,000 pounds

B. Final altitude

= 33,110 feet (cruise

altitude for 33,000

pounds, drag index =

230)

C. Drag index

= 230

D. Air temperature devi

ation

= +10

°

C

E. Fuel consumed to final

altitude

= 1353 pounds

F. Initial altitude

= 2500 feet

G. Fuel consumed to ini

tial altitude

= 72 pounds

Fuel consumed to

climb 1353-72

= 1281 pounds

Note that the cruise altitude used above is based on

an initial climb GW of 33,000 pounds. This cruise alti

tude should be adjusted to account for the fuel con

sumed during climb and the climb fuel recomputed.

Initial GW is unchanged.

A. Initial GW

= 33,000 pounds

H. Revised final altitude

= 33,950 feet based

on endofclimb

weight (33,000-

1281 = 31,719 from

figure A33)

C. Drag index

= 230

D. Air temperature

deviation

= +10

°

C

I. Fuel consumed to

final altitude

= 1419 pounds

F. Initial altitude

= 2500 feet

G. Fuel consumed to

initial altitude

= 72 pounds

Fuel consumed in

climb

= 1419 - 72 = 1347

pounds

Distance in climb

=  89 - 2 = 87 nm

Time in climb

= 11.4 - 0.4 = 11.0

minutes

Climb speed (for drag

index = 230)

= 356 KIAS/0.80 mach

MAX AB CLIMB

Figure A35, sheets 1 and 2, presents MAX AB climb

data. The climb schedule given on sheet 1 results in

minimum timetoclimb to altitude at subsonic

speeds.

REFER TO FIGURE A35.

Refer to instructions under MIL CLIMB, above.

T.O. GR1F16CJ11

Change 7A33

30

20

10

0

0

100

200

300

G

DIST

ANCE   NM

DRAG INDEX

1F-16CJ-1-1-1027A

DATA BASIS FLIGHT TEST

CONFIGURATION:

CONDITIONS:

ENGINE F100-PW-229

MIL Climbout Fuel, Time, and Distance

DATA INCLUDES GROUND RUN FUEL,

LG/TEF'S RETRACTED 8 SECONDS

AFTER TAKEOFF

CLIMBOUT AT 12 DEGREES PITCH
ATTITUDE TO 2500 FEET AGL; THEN
ACCELERATE TO MIL CLIMB
AIRSPEED

TIME AND DISTANCE FROM END OF

RUNWAY BRAKE RELEASE TO CLIMB

AIRSPEED

400

NOTE: GW = GW AT BRAKE RELEASE.

6

4

2

0

TIME   MINUTES

F

8

6

4

2

0

FUEL CONSUMED   100 POUNDS

BASELINE

E

60

40

20

0

-20

A

TEMPERA

TURE    C

NO WIND

BASELINE

BASELINE

B

C

C

D

D

D

C

Figure A31.

T.O. GR1F16CJ11

A34Change 7

1F-16CJ-1-1-1028A

15

10

5

0

DIST

ANCE   NM

0

100

200

300

BASELINE

DRAG INDEX

DATA BASIS FLIGHT TEST

CONFIGURATION:

CONDITIONS:

ENGINE F100-PW-229

MAX AB Climbout Fuel, Time, and Distance

CLIMBOUT AT 12 DEGREES PITCH

ATTITUDE TO 2500 FEET AGL; THEN
ACCELERATE AT MAX AB TO MIL

CLIMB AIRSPEED

NO WIND

DATA INCLUDES GROUND RUN FUEL,

LG/TEF'S RETRACTED 8 SECONDS

AFTER TAKEOFF

TIME AND DISTANCE FROM END OF

RUNWAY BRAKE RELEASE TO CLIMB

AIRSPEED

3

2

1

0

15

10

5

0

FUEL CONSUMED   100 POUNDS

20

400

NOTE: GW = GW AT BRAKE RELEASE.

BASELINE

BASELINE

60

40

20

0

-20

TEMPERA

TURE    C

TIME   MINUTES

Figure A32.(Sheet 1)

T.O. GR1F16CJ11

Change 7A35

1F-16CJ-1-1-1029A

DATA BASIS FLIGHT TEST

CONFIGURATION:

CONDITIONS:

ENGINE F100-PW-229

MAX AB Climbout Fuel, Time, and Distance

CLIMBOUT AT 12 DEGREES PITCH

ATTITUDE TO 2500 FEET AGL; THEN

ACCELERATE AT MAX AB TO MAX

AB CLIMB AIRSPEED

NO WIND

DATA INCLUDES GROUND RUN FUEL,

LG/TEF'S RETRACTED 8 SECONDS

AFTER TAKEOFF

TIME AND DISTANCE FROM END OF

RUNWAY BRAKE RELEASE TO CLIMB

AIRSPEED

15

10

5

0

0

100

200

300

400

15

10

5

0

FUEL CONSUMED   100 POUNDS

20

DIST

ANCE   NM

NOTE: GW = GW AT BRAKE RELEASE.

DRAG INDEX

20

BASELINE

TEMPERA

TURE    C

60

40

20

0

-20

2

1

0

TIME   MINUTES

BASELINE

3

25

BASELINE

Figure A32.(Sheet 2)

T.O. GR1F16CJ11

A36Change 7

Cruise Ceilings and Optimum Cruise Altitude

DATA BASIS FLIGHT TEST

STANDARD DAY
OPTIMUM CRUISE MACH NUMBER

CONDITIONS:

ENGINE F100-PW-229

20

24

28

32

36

40

44

GW   1000 POUNDS

OPTIMUM CRUISE ALTITUDE

CRUISE CEILING (300 FPM)

MIL CRUISE CEILING   1000 FEET

OPTIMUM CRUISE AL

TITUDE   1000 FEET

20

24

28

32

36

40

44

GW   1000 POUNDS

20

30

40

50

60

20

30

40

50

60

MACH

0.84

0.84

0.84

0.82

0

50

100

150

1F-16CJ-1-1-1030A

C

D

0.80

0.80

0.75

200

250

300

DRAG INDEX

OPTIMUM CRUISE

FOR MIL SERVICE CEILING

(100 FPM) INCREASE

ALTITUDE 670 FEET.

NOTE:

NOTE: FOR ALL DRAG INDEXES

+20 C OR -20 C OFF

STANDARD TEMPERATURE

EFFECTS ARE NEGLIGIBLE.

NOTE: FOR ALL DRAG INDEXES

+20 C OR -20 C OFF

STANDARD TEMPERATURE

EFFECTS ARE NEGLIGIBLE.

48

48

B

A

B

A

0.71

0.68

350

400

Figure A33.

T.O. GR1F16CJ11

Change 7A37

1F-16CJ-1-1-1031A

1600

1200

800

400

0

20

10

0

FUEL CONSUMED   POUNDS

AIR TEMP BASELINE

100

150

200

300

400

400/0.84

385/0.82

365/0.80

340/0.75

320/0.68

INITIAL GW   1000 POUNDS

SL

CRUISE CEILING

(300 FPM CLIMB

POTENTIAL)

A

I

E

G

200

600

1000

1400

MIL Climb    Fuel Consumed

DATA BASIS FLIGHT TEST

STANDARD DAY

CONDITIONS:

ENGINE F100-PW-229

1800

CLIMB SCHEDULE

DRAG INDEX KIAS/MACH

0

50

445/0.84

425/0.84

2000

40

32

24

20

28

36

44

48

D

C

H

B

F

AIR TEMP DEVIATION

FROM STANDARD DAY ( C)

Figure A34.(Sheet 1)

T.O. GR1F16CJ11

A38Change 7

1F-16CJ-1-1-1032A

20

10

0

0

5

10

15

AIR TEMP BASELINE

0

20

40

60

80

100

AIR TEMP BASELINE

20

SL

A

TIME   MINUTES

DIST

ANCE   NM

INITIAL GW   1000 POUNDS

DATA BASIS FLIGHT TEST

STANDARD DAY

CONDITIONS:

ENGINE F100-PW-229

MIL Climb    Distance and Time

100

150

200

300

400

400/0.84

385/0.82

365/0.80

340/0.75

320/0.68

CLIMB SCHEDULE

DRAG INDEX KIAS/MACH

0

50

445/0.84

425/0.84

24

28

32

36

40

44

48

120

F

H

C

C

CRUISE CEILING

(300 FPM CLIMB

POTENTIAL)

AIR TEMP DEVIATION

FROM STANDARD DAY ( C)

Figure A34.(Sheet 2)

T.O. GR1F16CJ11

Change 7A39

1F-16CJ-1-1-1033A

20

10

0

2800

2400

2000

1600

1200

800

400

0

AIR TEMPERATURE BASELINE

FUEL CONSUMED   POUNDS

24

32

40

INITIAL GW   1000 POUNDS

COMBAT CEILING

(500 FPM CLIMB

POTENTIAL)

SL

20

28

36

3200

DATA BASIS FLIGHT TEST

STANDARD DAY

CONDITIONS:

ENGINE F100-PW-229

MAX AB Climb    Fuel Consumed

100

150

200

300

400

545/0.90

545/0.88

530/0.88

515/0.85

495/0.82

CLIMB SCHEDULE

DRAG INDEX KIAS/MACH

0

50

590/0.91

570/0.90

48

44

AIR TEMP DEVIATION

FROM STANDARD DAY ( C)

Figure A35.(Sheet 1)

T.O. GR1F16CJ11

A310Change 7

1F-16CJ-1-1-1034A

DATA BASIS FLIGHT TEST

STANDARD DAY

CONDITIONS:

ENGINE F100-PW-229

MAX AB Climb    Distance and Time

20

10

0

AIR TEMP DEVIATION

FROM STANDARD DAY ( C)

6

4

2

0

AIR TEMP BASELINE

TIME   MINUTES

8

40

32

24

INITIAL GW   1000 POUNDS

SL

COMBAT CEILING

(500 FPM CLIMB

POTENTIAL)

36

28

20

30

20

10

0

AIR TEMP BASELINE

DIST

ANCE   NM

40

50

60

100

150

200

300

400

CLIMB SCHEDULE

DRAG INDEX KIAS/MACH

0

50

545/0.90

545/0.88

530/0.88

515/0.85

495/0.82

590/0.91

570/0.90

44

48

Figure A35.(Sheet 2)

T.O. GR1F16CJ11

Change 7A41

PART 4-CRUISE

TABLE OF CONTENTS

Page

Cruise Data

A41

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 

Optimum Mach/Optimum

Altitude Cruise

A41

. . . . . . . . . . . . . . . . . . . . . . . . 

Fuel Flow Conversion

A42

. . . . . . . . . . . . . . . . . . . . . 

Subsonic Cruise Tables

A42

. . . . . . . . . . . . . . . . . . . . 

Diversion Decision

A43

. . . . . . . . . . . . . . . . . . . . . . . . 

Best Cruise Altitude for

Short Range Mission

A44

. . . . . . . . . . . . . . . . . . . 

LIST OF CHARTS

Figure

Page

Optimum Cruise

A41

A45

. . . . . . . . . . 

. . . . 

Fuel Flow Conversion

A42

A46

. . . . . . 

. . . . 

Subsonic Cruise-Effects

of Bank Angle

A43

A47

. . . . . . . . . . 

. . . . 

Subsonic Cruise-

Sea Level

A43

A48

. . . . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

4000 Feet

A43

A413

. . . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

8000 Feet

A43

A418

. . . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

12,000 Feet

A43

A423

. . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

16,000 Feet

A43

A428

. . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

20,000 Feet

A43

A433

. . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

25,000 Feet

A43

A438

. . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

30,000 Feet

A43

A443

. . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

35,000 Feet

A43

A448

. . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

40,000 Feet

A43

A453

. . . . . . . . . . . . . 

. . . . 

Subsonic Cruise-

45,000 Feet

A43

A456

. . . . . . . . . . . . . 

. . . . 

Diversion Decision-

Divert

A44

A458

. . . . . . . . . . . . . . . . . . 

. . . . 

Diversion Decision-

Loiter

A44

A460

. . . . . . . . . . . . . . . . . . 

. . . . 

Best Cruise Altitude for

Short Range Mission-

Penetration Descent

A45

A462

. . . . . 

. . . . 

Best Cruise Altitude for

Short Range Mission-

Maximum Range 

Descent

A45

A463

. . . . . . . . . . . . . . . . 

. . . . 

CRUISE DATA

The cruise data enables the mission planner to calcu

late the optimum cruise performance available for

most combinations of flight conditions. It is apparent

from the Subsonic Cruise tables that the GW's for

some of the drag indexes are not realistic. This data

is only used to give good results when extrapolation

of the data is required. One conversion chart is sup

plied as an aid in computing TAS and fuel flow. The

following considerations will assist in selecting the

combination of flight conditions and techniques

which will result in required mission performance:

D

Optimum Mach/Optimum Altitude:A cruise

climb technique is required. Mach remains constant

throughout the cruise while altitude increases as

fuel is consumed. Changes in optimum mach due to

changes in GW are insignificant. Use of this

technique will result in maximum attainable range

(including maximum aircraft capability).

D

Optimum Mach/Constant Altitude:Mach number

is decreased as fuel is consumed. This cruise

technique will yield maximum cruise performance

at a given altitude.

D

Constant Mach/Constant Altitude:Cruise at a

given mach and altitude condition. This technique

does not usually provide maximum performance

but is often used due to time and flightpath

constraints.

OPTIMUM MACH/OPTIMUM

ALTITUDE CRUISE

Detailed optimum cruiseclimb performance data is

given in figure A41. Cruise data is shown as specific

range and optimum altitude versus GW for lines of

drag index. Optimum cruise mach numbers are tabu

lated on each drag index line. Fuel flow may be com

puted from figure A42.

REFER TO FIGURE A41.

Enter chart with cruise GW (A) and proceed vertically

upward to drag index (B) in both the lower and upper

portions of the chart. In the lower portion of the chart,

proceed to the left from (B) to read specific range (C).

In the upper portion of the chart, proceed to the left

from (B) to read optimum cruise altitude (D). Opti

mum cruise mach number is obtained from the mach

numbers indicated on the drag index lines in the

lower portion of the chart.

T.O. GR1F16CJ11

A42Change 7

SAMPLE PROBLEM.

A. GW

= 33,000 pounds

B. Drag index

= 230

C. Specific range

= 0.105 nm per

pound of fuel

D. Optimum cruise alti

tude

= 33,110 feet

D

Optimum cruise

mach

= 0.80

FUEL FLOW CONVERSION

The Fuel Flow Conversion chart, figure A42, is used

to convert specific range and speed into fuel flow.

REFER TO FIGURE A42.

To convert specific range into fuel flow, enter chart

with mach number (A), proceed to the right to temper

ature (standard day temperature is shown on figure

NO TAG) (B), and then proceed upward, reading

KTAS at (C). Continue to specific range line (D), and

finally, proceed to the left to read fuel flow (E).

SAMPLE PROBLEM.

A. Optimum cruise alti

tude

= 33,110 feet

D

Optimum cruise

mach

= 0.80

B. Temperature

= -56.5

°

C (Standard day

temperature is shown)

C. KTAS

= 458

D. Specific range

= 0.105 nm per pound of

fuel

E. Fuel flow

= 4174 pounds per hour

SUBSONIC CRUISE TABLES

The Subsonic Cruise tables, figure A43, present dry

thrust fuel flow data for a range of constant cruise

altitudes (sea level45,000 feet), true airspeeds (180

690 knots), GW's (20,00048,000 pounds), and drag

indexes (0400). True airspeeds and fuel flows for

maximum range/endurance cruise at constant alti

tude and drag index are presented for a range of GW's.

If Vmin (minimum true airspeed based on MIL) is

greater than 180 knots, then Vmin and the fuel flow

at Vmin are shown. Vmax (maximum true airspeed

based on MIL) and the fuel flow at Vmax are also

shown. Temperature effect factors are presented for

20

_

C ambient temperature deviation from stan

dard. Cruise KTAS are presented in increments of 30

KTAS. The fuel flows are shown in PPH; therefore,

the distance flown and fuel consumed at some speci

fied cruise time may be quickly evaluated.

REFER TO FIGURE A43.

To find fuel flow for cruise at a constant true airspeed

and altitude, enter the table with appropriate drag

index, KTAS, and GW. Then read the fuel flow in

PPH. To find fuel flow and KTAS at constant altitude

cruise for Vmin, Vmax, maximum endurance, or max

imum range, enter the table with drag index and GW.

Then read standard fuel flow and KTAS for the speci

fied condition. Temperature effect factors are found

on the right side of each chart. Multiply standard day

fuel flows by their respective factor to get final fuel

flows for  20

_

C deviation from standard. To compute

fuel flows for other temperatures, simply ratio the dif

ference between standard day values and  20

_

C val

ues for the particular temperature deviation.

True airspeeds for Vmin, Vmax, maximum endur

ance, and maximum range are affected by ambient

temperature and correction factors for these air

speeds are presented on the right side of the chart.

These factors are used to correct KTAS in the same

manner as described for fuel flows. If the factors are

greater than 1, final fuel flow and KTAS increases. If

the factors are less than 1, final fuel flow and KTAS

decreases.

SAMPLE PROBLEM.

A. Altitude

= 30,000 feet

B. Drag index

= 0

C. GW

= 20,000 pounds

D. KTAS

= 360

E. Standard day ambi

ent temperature

= -44

°

C

F. Ambient temperature = -34

°

C

G. Temperature devi

ation

= 10

°

C hot

T.O. GR1F16CJ11

Change 7A43

Find fuel consumed and time required to cruise at 360

knots for 180 nm:

H. Standard day fuel flow= 1809 PPH

I. Temperature effect

fuel flow factor at

+20

°

C and 360 KTAS = 1.070

J. Fuel flow at +20

°

C

hot is 1.070

1809

= 1936 PPH

K. Fuel flow for ambient

temperature of -34

°

C

is 1809 + (127

10

20

) = 1873 PPH

L. Time to travel 180

nm at 360 KTAS is

180

360

= 0.5  hour

M. Fuel consumed in

0.5 hour of cruise

at 360 KTAS is

1873

0.5

= 937 pounds

Find fuel consumed and air distance traveled for

maximum range cruise for 30 minutes:

N. Standard day maxi

mum range airspeed = 433 KTAS

O. Standard day fuel

flow

= 2091 PPH

P. Temperature effect

factor for KTAS at

+20

°

C

= 1.043

Q. Temperature effect

fuel flow factor at

+20

°

C

= 1.046

R. KTAS at 20

°

C hot is

1.043

433

= 452

S. Fuel flow at 20

°

C hot

is 1.046

2091

= 2187 PPH

T. KTAS for ambient

temperature of -34

°

C

is 433 + (19

10

20

)

= 443

U. Fuel flow for ambient

temperature of -34

°

C

is 2091 + (96

10

20

)

= 2139 PPH

V. Air distance traveled

in 30 minutes at 443
KTAS is 443

60

30

= 222 nm

W. Fuel consumed in

30 minutes of cruise

at 443 KTAS
is 2139

60

30

= 1070 pounds

To find the fuel flow and KTAS for maximum endur

ance cruise, use the method outlined above for maxi

mum range cruise; then loiter time = fuel consumed/

fuel flow.

REFER TO FIGURE A43.

If an average bank angle of 30 degrees were used in

the above problem, an effective GW of 23,100 pounds

would have to be used to enter the chart. Find the

effective GW by entering the lower right plot with GW

(A), follow the guidelines to bank angle (B), and read

effective GW (C).

SAMPLE PROBLEM.

A. GW

= 20,000 pounds

B. Bank angle

= 30 degrees

C. Effective GW

= 23,100 pounds

DIVERSION DECISION

The Diversion Decision-Divert and Diversion Deci

sion-Loiter, figure A44, contain range and time

available data to be used in deciding whether to divert

to another base or wait (loiter) until the runway is

reopened. Data is given for fuel quantities up to 2000

pounds and for initial altitudes from sea level to

40,000 feet. Range and time available by staying at

initial altitude or by climbing to optimum altitude are

given. Range and time for climbs to optimum altitude,

cruise or loiter, and descents to sea level are included

in all data as applicable. No reserve fuel is included

in the data.

SAMPLE PROBLEM.

Assume that you have arrived over base at 5000 feet

MSL with only 600 pounds of fuel remaining and are

informed that the runway is closed due to an accident.

Twenty to 30 minutes is required to open the runway.

Can you wait (loiter) for the runway to open, or should

you divert to a base only 50 nm away? Checking figure

A44 yields the following information:

Range Available at 5000 feet MSL = 66 nm (cruise

at mach = 0.46, begin enroute descent 16 nm from

destination with 104 pounds fuel used in descent).

Range Available Using Optimum Altitude = 88 nm

(MIL climb at 423 KIAS or optimum altitude mach

number, whichever is less, to 30,000 feet, cruise at

mach 0.72, and begin enroute descent 72 nm from

destination with 306 pounds fuel used in descent).