THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 85-GT-184

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Copyright © 1985 by ASME

F-14 Re-Engining with The F110 Engine

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0. T. CASTELLS
General Electric Company
Evendale, Ohio

J. T. STRONG, JR.
Grumman Aerospace Corporation
Bethpage, New York

ABSTRACT TF30 powered F-14 performed excellently, realization

of the aircraft's full potential required the higher
The advantages of re-engining the F-14 aircraft thrust, new technology engine. Accordingly, the
with the FllO engine is presented. The areas of General Electric Company initiated a company
improvement and the engine development philosophy are sponsored program to develop an F-14 suitable engine
explained. A summary description of the pertinent derived from its successful FlOl-GE-100 turbofan that
engine design features of the FllO are presented. powers the Air Force B-1 aircraft. The FlOl
The flight test results on inlet/ �gine Qerivative £:ighter �ngine, now designated
.
compatability, afterburner operation, airplane FllO-GE-400, thus came into being. This engine,
performance, and developed through close Engineering effort between GE
maintainability/reliability/durability are and GAC, has been built to be compatible with the
interpreted. Finally, a description of the proposed original F-14B (the designation for the F-14 with the
version of the FllO engine for the F-14 is presented. larger engine) installation including the variable
geometry air inlet, aircraft accessories, and
THE OPPORTUNITY aircraft fuel system.

Conventional wisdom states that re-engining of The F-14A contains advanced aerodynamics (with
modern fighter aircraft is not possible. Various its swing wing) , structures (composite empennage) ,
reasons are given, such as: avionics (with its still unique Phoenix/AWG9
stand-off missile) but has an undersized, previous
- The speed of advancement in aerodynamics, generation engine. In addition, its inlet, fuel
structures, and avionics will render a fighter system, and the power plant accessories were all
design obsolete, such that insufficient payoff sized for a 35 percent larger engine. Therefore, it
will result from re-engining compared to a new wasn't a question of re-engining to improve an
design. airplane, but rather re-engining to realize an
airplane's original potential.
- The development time of a new engine is a
significant percentage of the total life of a H£ REQUIREMENT
fighter.
Early USN direction on this program emphasized
Technical and program difficulties of changing improvements in the following areas, in order of
interfaces, weight, center of gravity, priority, as shown by Table 1.
aircraft powerplant accessories requirements,
and nacelle structure all add time and cost, TABLE l
jeopardizing such an activity.
AREAS OF IMPROVEMENT
These arguments, however, were not germane to the
F-14. Uniquely, the F-14 was designed from the onset
to be configured with an advanced technology, 30, 000 o Operational Suitability
maximum thrust class engine, the P&W F401-PW-400 . o Reliability/Durability
Early aircraft models were to have been powered by o Maintainability
the available 20, 000 pound thrust class TF30-P-412, o Cost
the engine in the FB-lll, until the new engin� was o Performance
ready for production installation. Cancellation of o Weight
the F401 project aborted this plan. Although the

Presented at the Gas Turbine Conference and Exhibit

Houston, Texas - March 18·21, 1985
THE FllO DEVELOA-1ENT PH ILOSOPHY an equivalent of ten years of fleet operation. Such
testing will assure that the FllO will enter
production with the desired improvements in these key
Operational suitability for the F-14 encompassed
demonstrating: no throttle restriction operation areas.
anywhere in the aircraft maneuver envelope;
sufficient dry thrust to eliminate the need for All aspects of cost control in development
afterburner operation during catapult take-off; i
acquisition, and operations has been continual y
srrooth and continuous engine handling characteris tics addressed by the FllO. The developmental cost will
leading to reduced pilot workload; and finally be discussed here as it has an impact on how the
defining a suitable airstart envelope and airstart technical job was accomplished. Major cost savings
_

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procedure. All of these efforts were addressed in were obtained by extensive use of existing assets and
simulated altitude and ground tests and were by utilizin9 the same engines for ground testing
.
confirmed during flight tests on F-14 tF7 in 1981. pr10r to flight tests. No new engines were used in
�
any of t e flight tests in the F-16A, F-14, and
Reliability/durability and maintainability F-16XL aircraft. The use of five engines in the
demonstrations required extensive ground testing program is shown in Figure 1. One engine has been
utilizing mission oriented cyclic testing. The utilized solely to maximize endurance time in the
limited duration flight test program provided an Accelerated Mission Tests. The other four have been
additional preliminary indication of the success of utiliz�d first in ground testing for system,
these efforts. A complete discussion of the mechanical, and performance testing, second in
reliability/durability and maintainability activities altitude functional and qualification testing at GE '
is presented in ref. 3. Suffice it here to state Arnold Engineering Development Center, and at the
that the FllO is successfully utilizing a severe '.
Naval Air Propulsion est Center, and finally, for
Accelerated Mission Test technique both to develop the three separate flight test programs. The F-14
and to qualify for production. The FllO has several �est aircraft selected was No. 7, an early model,
internal features incorporated for performance instrumented aircraft which had been utilized for the
retention. These include items like adding mass to F401 testing and had been in storage since 1974.
outer rings to thermally match the shrouds (or seals) Grumman Aerospace Corporation updated and modified
with the disks to maintain clearances during hot this aircraft for approximately 10 percent of its
rotor reburst. Some weight was added to assure original cost. Another example of a large cost
durability. Some features were proven by the savings was the proper selection of the aircraft
performance retention characteristics demonstrated in accessory gearbox. The current F-14A aircraft does
the Accelerated Mission Testing where the engine was not have an aircraft accessory gearbox as the TF30
still delivering 100 percent of the intermediate and has an integral gearbox with pads for the four
99 percent of max afterburning thrust at the end of aircraft accessories. The FllO did not have these

Summary of F110 Engine Usage

AEDC NAPC

Development/Flight Clearance
and
Three Separate Successful
Flight Tests

••
F-14

PLUS
5000 TAC Cycle Endurance Engine
I

L m •

FIGli1lli 1

2
pads so it was decided to utilize the Pratt & Whitney design. These features include a highly reliable
(P&W) F401 accessory gearbox and adapt the FllO power gear type main fuel pump, a backup control, dual
take-off to drive this box. P&W, at the request of independently powered ignition, automatic relight
NAVAIR, overhauled these gearboxes for much lower features, etc. Beyond these, the FllO includes two
cost than new ones. In summary, the total principal features which enhance operational
expenditures to fully qualify the FllO (scheduled for suitability; these are its integrated mixed flow
January 19 85) , including three separate flight tests, afterburner and its control mode.
will be less than the pi..blished expenditures on the
F401 program. The mixed flow afterburner consists of a
convoluted mechanical mixer which is integrated with
Performance, although fifth on the list, was not the afterburner flameholder assembly, the fuel

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stinted. The major effort was to improve the mission injectors, and pilot burner can. The augmentor
performance, i.e., range, intercept time, and loiter includes a light-off detector which insures that
time in fleet defense. In addition, single engine ignition has occurred before full core afterburning
rate of climb, acceleration time, and specific excess fuel flow is introduced. The core flow is fully
energy (Ps) were all improved significantly. These burned and serves to stabilize the fan flow prior to
are discussed in the flight test program results introducing fan afterburning fuel flow. This design
section. results in smooth continuous thrust modulation
without high pressure pulses at light-off or between
Weight, the last item, was important as it stages.
impacted both the empty weight of the aircraft and
the center of gravity. These aspects are discussed A second major feature is the basic control mode
in the final section of this paper. which directly maintains fan speed and fan operating
line. This mode produces the required thrust at all
Tl-E DESIGN FEATURES operating conditions as the fan speed determines the
engine airflow (which is proportional to thrust) and
The FllO is a two-spool, augmented turbofan also maintains the fan operating line which is
engine as illustrated in the cross-section of Figure proportional to tailpipe pressure and hence, thrust.
2. Some of its principal design features will be The control of the fan operating line is accomplished
discussed in this section. The FllO was designed by closed loop actuation of the exhaust nozzle to
from inception for both single engine and twin engine maintain the scheduled duct air velocity as measured
application, and, therefore, it incorporates many by a delta P/P parameter. This system maintains the
reliability features required for the single engine

F110 Crossection

Aug mentor

FIGCP.E 2

3
 THE FLIGH T TEST
stall margin during afterburner transients and is a
major factor in the no-throttle restriction operation
The F-14/FllO flight test program was conducted
of the FllO. A typical afterburner transient is
at the Grumman Aerospace Corporation test center at
shown in Figure 3.
Calverton, New York. The test consisted of 44
flights conducted in two phases, starting in July
1981 and continuing through March 19 82.
Afterburner Transient
During Phase I, cost considerations coupled with
Iner Pre-Open
Surge low emphasis on the low altitude, high speed flight
Margin
Light-Off
0.3
Ma:itAug
regime, precluded incorporation of certain
J_

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6.PIP
Schedule
modifications to aircraft 7 required for high " q"
t flight as foLlld necessary during aircraft structural
0.1 5°'a Surge Margin
Deer
Surge
Margin
flight testing. Prior to Phase II flight test, these
modifications were made, allowing operation in the
lower right hand corner of the envelope. The flight
925 test results have been reported in detail in
AB
Reference 1 and Reference 2, and only a summary and
interpretation of these results is included herein.
423

INLE T/ENGINE COMPATABILITY

Time

Flight tests were conducted throughout the flight

FIGURE 3 envelope, up to 50K feet altitude, as shown in Figure
The control system also eliminates the need for 6. Engine operability and inlet/engine compatability
engine trim which results in large savings in
maintenance time, fuel usage, and engine life
Flight Test Points
compared to the current system. As the engine ages, ---,
so ��ln � d� - �� ���--��� -� �
the fan speed and operating line are maintained and icat ed Air Speed (K nots )

the turbine temperature is increased. This is shown 50

in Figure 4. An optical pyrometer is used to Altitude Data Point
(1000 ft)
40
"No Trim" Feature
30

Int & Above Limit

�
20
Turbine
Fan Blade
Speed Temp 10
•New Engine

.4 .a 1.2 1.6 2.0 2.4

Inlet Temp Inlet Temp Mach No.

FIGURE 6
• New Engine Runs to Fan Speed ( Thrust)

• Fan Speed Maintained But Temp Increases as Engine were evaluated in over 1000 snap transients,
Matures
typically run on both engines simultaneously. The
• Once Temp. Limit is Reached Fuel Flow is Reduced snap throttle transients used were as follows:
FIGu1lli 4 starting from power for level flight down to idle,
idle to max (or military) thrust, and back to idle.
These tests were also successfully conducted during
directly measure the high pressure turbine
extreme aircraft maneuvers, as illustrated in Figure
temperature and once the limit is reached, the fuel
7. During all this testing, no engine stalls
flow is reduced. After the thrust drops below the
acceptable level, the engine is refurbished. A
schematic of the optical pyrometer operation is shown
in Figure 5. Inlet/Engine Compatibility

Optical Pyrometer - Limiter on 60

o Fixed Throttle
Blade Temperature ciMax Aug-ldle
50 • Aug Transients
Angle of
Attack 40
-aT
-Deg.
30

20
0 0
0
'• c 0
10
'· · 0
c
• ••

20 10 0 10 20

Angle of Sideslip - �T-Deg.

FIGURE 7
FIGLllli 5
 resulting from inlet distortion were experienced nor left hand corner the fuel controls were set to reduce
could stagnation stalls be induced. A total of 359 the maxirrum fuel/air ratio in a region labled
official, full data test conditions were "cutback region" shown in Figure 9. This was done to
accumulated. All tests were run with no throttle preclude any rich instability problems, since little
restrictions regardless of aircraft altitude, angle development testing in this region had been
of attack or airspeed. The test aircraft had completed. During each of the two phases, there were
complete dynamic and steady state pressure inlet rake two events of blowout-relight in this region. The
data available which allowed accumulation of engine immediately recovered without pilot action and
extensive distortion data. These data have been properly cycled to full afterburner. After the
analyzed to determine how much stall margin reduction flight test all these data were analyzed and some
the distortion caused. In addition, all other changes were introduced into the pilot burner to
destabilizing engine effects such as errors in improve the light-off capability. In addition,

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il variable geometry tracking, component deterioration, detailed tests were conducted at Arnold Engineering

IJ afterburner transient operation, and engine-to-engine

variation have been analytically included. These
Development Center (AEDC) on the afterburner metering
valve slew rates and another small change was

1· effects were compared to the available stall margin

at most critical conditions. These comparisons
identified. With these changes it was possible t o
significantly improve the light-off capability and to
(stability stacks) show positive stall margin eliminate completely the "cutback region" without any
remaining. These results provide high confidence further blowout relights occurring. These changes
that the FllO will operate successfully, fleet wide, have now been successfully confirmed in the ongoing
with no throttle restrictions. F-16XL flight test. Full afterburner capability is
available up to the extremes of the flight and
AFTERBURNER OPERATION maneuver envelope.

The afterburner envelope capability was fully AIRSTART ENVELOPE

investigated from 0.2M to 2.0M, with more than 915
afterburner light-off attempts. Even though a large Extensive evaluation of the airstart capability
nuntier of these tests were in the upper left hand was conducted. Three types of airstarts (assisted,
corner, (the most difficult area for afterburner spooldown, and windmill) were tested to establish
operation) successful lights were obtained in over 98 limits in a total of 83 attempts. The assisted
percent of the attempts with the no-lights confined airstarts were excellent at all altitudes up to 31J<
to the high altitude, low speed region. Outside that with only one failure to start which was associated
region, the light-off reliability was almost with an improperly positioned metering valve. These
perfect. Some of the no-lights were successful on an results are shown in Figure 10. This problem in the
immediate recycling attempt at the same flight
condition. When no- lights occurred, the engine Cross Bleed Airstarts
properly remained at intermediate power. The areas
\
Predicted F110 Air Start
Legend
where the light attempts were made in Phase I and oSuccessful Capa y
•Unsuccessful
Phase II are shown in Figures 8 and 9. In the upper
30

oooe
Engine Transients F110 Flight Envelope Std Day

Including
10 �------�
Standard Da ___ _ / =-i
After-Burner
60
No.7/-;;
'\,.
F·14A/C
1 Failure to Start -

I
Associated with Metering
Valve Positioned to Mmimum

Operation 50 I • 10
I
All Successful Starts
Produced Low ( <775°C)(930°C) Limit

.A
Turbine Blade Temps

14 • :
•• •
I
40

'•··-.

I//l'.-J• ;' I
; I �.�1�--'� 0.4---:0�.5��o� .6--o�7
�.3-_.,
o .2--o . o.9
30 ; . ,---o�.a,---::'
Alt.
10·' Ft I • • •• • . ; True Mach Number, Mr
/
.
, . .. .
20 .

rn
: :.
-

'"-
. ..

.1
I •1g •Elevated Angle
of Attack
FIGURE 10
-e
:.• .
/ Full Adverse
Rudder
hydromechanical control was identified and corrected
early in the program and no further problems
0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4
occurred. The spooldowns were successful in most
FIGURE 8 True Mach Number, Mr
cases when attempted with the engine stabilized at
power for level flight as shown in Figure 1 1. Some
F·14/ F110 Phase II Program
Spool Down Air Starts
After-Burner Lighting Tests
50
35
oSuccenful Legend
c)
•No-Light F-14 oSuccesstul
40 30 "'F-16 •Unsuccessful

25
30
Altitude 10·' Ft
I
Altitude 20 I
10' Ft I
0 0 00
20 I
15 I
I
I
10 NOTE: Some Open Symbols 10 I
Denote Multiple Events I F-14A (No. 7)
5 L.-vmin

0.4 0.8 1.6 2.0 0.2 0.3 0.4 0.5 0.6 0.7 0.B 0.9

True Mach.Number. Mr
Mach Number
FIGURE 9 FIGURE 11

5
hot starts resulted when attempted from intermediate
power at high speed. Additional work on the engine
controls was done after the flight test, and
subsequently tests at AEDC have shown that airstart Max Power Acceleration and
reliability in that region has now been improved. Climb Times
The windmill tests were all successful but required 60,000
high aircraft speeds ( · 420 Knots) to develop • Indicated Times

sufficient rpm to effect a light. The test results 50.000

112 67 sec ______ll� Are an Average
tmprovement ol 30%
and planned production changes insure that the FllO Over Current Engine
40,000
will have a large envelope where reliable airstarts 125
61
Altitude 30.000
of all types can be performed. 1
Feet

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20.000
PERFOR--1ANCE 1
62
10.000

The FllO flight test performance has been

converted to operational fleet defense mission (being 0.4 0.8 1.2 1.6 2.0 2.4

used at that time) and these results are shown in Mach Number

Figures 12 and 13. These improvements of 62 percent

FIGURE 11
F-14A Deck-Launched Intercept
Mission Radius Comparison
•M=1.5
•Missiles Carried All the Way
• Weapon load: (4) Phoenix+
300 (2) Sparrows+ (2) Sidewinders+
(2) 280 Gallon Drop Tanks

�
F110
Intercept
200 Specific Excess Energy Improved
Radius, n. mi. 100/150 fps in Combat Arena (Max Power)
� • Dry Ps Improved 77%; M.7/10K
Current
100 Engine 60,000
• Combat Weight
(55,000 lbs)
50,000 4 Sparrows
lg
Altitude 40,000

FIGURE 12 Feet
30,000

F-14A Cap Mission Tradeoff 20,000

g
• Ps=Th�=1��;a Velocity
Notes:
10,000
2.8 •Combat At T = D (Partial A/B) at • From Ref. 1
M=1.35
2.4 •Weapon Load: (4) Phoenix+
0.4 0.8 1.2 1.6 2.0 2.4
(2) Sparrows+ (2) Sidewinders+
(2) 280 Gallon Drop Tanks Mach Number
2.0

1.6
Time On
FIGCRE 15
Station,
Hr 12

.8

.4
Engine

o������-'--'
0 100 200 300 400 500

Combat Radius, n. mi.

FIGGRE 13
increase in deck launched intercept radius or 34%
increase in time-on-station at a corrbat radius of 150
F-14A - Single Engine Rate of Climb
nautical miles will significantly enhance the
effectiveness of the weapons system.
(4) Phoenix+ (2) Sparrows+ (2) Sidewinders+ (2) 280 Gallon Drop Tanks

While the principal objective of performance -Current • Sea Level Std Day
--- F110 • Flaps and Gear Extended
improvements was the enhancement of the fleet defense
missions, the flight test also demonstrated large Takeoff Power Settings

gains in most of the other aircraft performance Max AIB Power

2 Max TOGW
categories. The improvements in acceleration, clirrb, Rate of Climb,
74,349 lb
1000 ft/min
and specific excess energy (Ps at lg) are shown in
Figures 14 and 15. In addition, data were taken on
,........
. ... ........... ...,
....
Intermediate
..........
. ....
single engine rate-of-clirrb in order to verify the .............
ability to operate from carriers without the need o f
......
.. ....
.. ......
54 58 62 66 70 74 78
the afterburner. The data substantiate that the
Gross Weight. 1000 ·1b
F-14/FllO corrbination will be able to take off under
a single engine condition with positive rate of clirrb
up to its maxirrum take-off gross weight as shown in
FIGURE 16
Figure 16.

6
RELIABILITY/DURABILITY/MAINTAINABILITY
Configuration F or USAF/USN
The flight test was of 70 hours duration, USAF Engines
producing a total of 140 very demanding engine flight
hours. The short duration provided only a minimal
check on reliability/durability/maintainability,
however, the results were excellent with no need to
utilize a spare engine and only a few minor engine
maintenance events. This record (no spare engines
utilized) has continued from the original 1980 flight

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test in the F-16A through the F-14 and the still
on-going F-16XL flight tests. A total of 366 engine
flight hours and 253 flights accumulated through 27
March 1983. In addition, in the FllO-F -14 flight
Dillerences:
test, not a single day was lost due to engine •Added Tailpipe Length • Repackaged C & A
•A I C Acces. Gearbox • Mount Locations
maintenance or repair. I n fact, no ground test cell
was even set up at Calverton, and none was needed.
The no-trim feature of the engine was demonstrated by FIGURE 18
replacement of the electronic control in about two
hours without a sl..b sequent need to retrim the The FllO re-engining of the F-14 has been shown
engine. The engine was just run up at the end of the by ground and flight testing to be an attractive,
runway immediately prior to take-off. This will be improved, cost-effective, and affordable approach.
an important advantage in carrier operation, The final determination of whether the "conventional
eliminating the need to place the aircraft on the wisdom" was correct in the case of the F -14 has now
trim pad. been made, and the USN will proceed to improve the
F-14.
During the testing, 16 different pilots,
including GAC, USN, and other Government
representatives flew in the aircraft, and they were
unanimous in their enthusiasm on the FllO's
operability, handling response, and performance.
GAC's chief test pilot, the auther of Reference 1,
concluded that the FllO significantly enhances the
operational and fighter capabilities of the F-14
airplane. His conclusions are included here as
Figure 17.

Chief Test Pilot Conclusions:

(Reference 1)

• No Throttle Restrictions, Regardless of

Altitude, Airspeed or Angle of Attack

• Significantly Improved Airplane Performance

- Climb, Both Military and Maximum Thrust
- Accelerations, Both Military and Maximum Thrust

- Turning Performance

- Single Engine Rate of Climb

• No Engine Trim
- Manpower!Time/Fuel Savings

FIGURE 17

THE FLEET INTRODUCTION

The USN is currently planning to upgrade the F-14

aircraft, and the FllO engine is a major element of
this upgrade.

Some changes from the flight test engines are

planned for production versions. These changes
include a longer tailpipe to closely match the TF30
center of gravity because the FllO is shorter than
the TF30. A new integral engine mounted aircraft
accessory gearbox will also be supplied. These
changes are shown in Figure 18. The basic engine up
to the tailpipe would be identical to the US AF
version. With these changes, the total propulsion
package would have approximately the same center of
gravity and weight as the current system.

7
REFERENCES

1. Sewell, c., Chief Test Pilot, Grumman Aerospace

Corporation, Calverton, NY. F-14/FlOl DFE Flight
Test, July 1982. Unpublished presentation at
" Fly Navy West, " San Diego, CA, and at The
Society of Experimental Test Pilots Symposium,
Septerrber 1982.

2. Strong, J. T., Jr., FlOl OFE Flight Test

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E valuation in the F-14 Aircraft, Septerrber 1982,
included in - Flight Testing Technology: A State
of the Art Preview - Society of Flig,t Test
Engineers 13th Annual Symposium, NYC.

3. Castells, O. T., Accelerated Mission Testing of

FllO Engine, June 1983, AIAA-83-1235, 19th Joint
Propulsion Conference, Seattle, Washington.