THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 84-GT-132

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Copyright © 1984 by ASME

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DEVELOPMENT OF THE F110-GE-100 ENGINE

Michael S. Coalson
Systems Engineering Manager
Aeronautical Systems rivision (ASD/YZYF)
Wright-Patterson AFB, Ohio

INTRODUCTION F-15 no doubt motivated GE to build the FIO1X, a pro-

totype derivative of the F101 engine. The engine was
The F110-GE-100 engine is now in the qualification thermodynamically similar to the F100 - low bypass
or "production verification" phase of development. ratio, high fan pressure ratio - and has the same
From a technical standpoint it represents the first core - already developed (and qualified) for the B-1
USAF fighter engine which has been developed with- bomber. Test results from the prototype fighter
out extreme emphasis on either high thrust or light engine or "derivative fighter engine" as it subse-
weight. Rather, since its inception the program quently became known, were sufficiently promising
emphasis has been on a balance of durability, oper- that the USAF and Navy jointly funded a demonstrator
ability and performance. For this reason alone, program for an engine which would fit both service
this paper should be interest to those whose busi- needs (i.e. F-15, F-16 and F-14. In this demon-
ness centers on fighter engine development. strator program, managed under the USAF's Engine
Model Derivative Program, or EMDP, durability test-
BACKGROUND ing, flight clearance testing and flight testing in
the F-16 and the F-14 was accomplished. In the
Two major USAF engine developments were launched in second phase, additional durability, component devel-
1970. The first was the F100, developed by Pratt & opment and flight testing in the F-16XL was accom-
Whitney Aircraft (P&WA) to power the F-15 air supe- plished. In the current program, Full Scale Develop-
riority fighter. The second was the F101 engine, ment to qualify the engine for production 4s being
developed by the General Electric Company (loser to accomplished. While these phases are important from
P&WA in the F-15 engine competition) for the B-1 a programming and budgeting standpoint, they are less
bomber engine. The engines were of different ther- important in a technical forum. Consequently, the
modynamic cycle and thrust requirements. The 1101 emphasis hereafter will be on the development accom-
being higher thrust, higher bypass ratio and higher plishments and the design features which have made
airflow than the F100. The F101 was (and is) a the accomplishments possible. Testing and demon-
physically bigger machine. strated accomplishments have been the byword of the
program. Durability testing in factory test cells
At the same time the USAF funded development of the and flight testing in F-16 and F-14 aircraft have
F100 engine, the Navy funded development of the F401 provided the demonstrated program accomplishments.
engine. The F401 and the F100 were to share the same
core (i.e. compressor, combustor and high pressure ENGINE DESCRIPTION
turbine), but mission requirements for the F-14 air-
craft dictated an engine with higher cycle pressure The engine is 182 inches long and has a maximum dia-
ratio and higher bypass ratio than the F100. About meter of 50 inches. The fan consists of three stages
half way through the engine development program the with variable flap inlet guide vanes. Nominal pres-
Navy decided to terminate funding for the F401 devel- sure ratio is 3.2 and nominal bypass ratio is 0.78.
opment. Consequently, all F-14 aircraft have been The fan is driven by a two stage, uncooled turbine
powered by the TF30 engine - an engine originally which is downstream of the single stage high pres-
viewed as interim until the F401 could be developed. sure turbine. The low spool is supported on the
three bearings - roller bearings at front and back
In 1977, President Carter cancelled production of the and a ball bearing behind the fan.
B-1 bomber and consequently that of the F101 engine.
This loss of business coupled with an awareness of The core of the engine, consisting of the compressor,
service problems with the engines in the F-14 and combustor and high pressure turbine is identical to
 E

that in the F101 engine. The augmentor is a scaled- A second feature is the high pressure turbine blade
down version of the F101 augmentor, modified for the pyrometer system. Unlike most other gas turbines
decreased bypass ratio in the F110. The number of which use thermocouples to measure gas temperature,
the fuel spraybars was doubled to better control the high pressure turbine blade metal temperature is
fuel distribution thereby enhancing combustion sta- measured. Blade temperature is monitored only for
bility and lightoff. The exhaust nozzle uses the exceedance - it is not governed as a scheduled value.
hinged-flap-cam-link design demonstrated on the F101
and used successfully on the 404. The nozzle actu- A third feature which contributes much to the super-
ators are the same as those used on the F101 except ior operability of this engine is the afterburner
that the stroke has been reduced. Frames and bear- light off detector (LOD). The LOD consists of an
ing supports use the same design approach as those ultra violet flame sensor which views the innermost

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on the F1O1. Like the F101, the F110 is a five-hear- "V" ring. When afterburning is selected the pilot
ing engine. As mentioned, the low-speed rotor is burner lights first and flame propagates around the
supported by three bearings; the high speed rotor by inner "V" ring. Only so called "min A/B flow" is
two, the aft of these being an intershaft bearing. allowed to flow by the control until the LOD con-
All five are identical to the bearings used on the firms that flame is in the inner "V" ring. Once the
F101. The lubrication system is similar to that inner ring is burning it is extremely unlikely that
used on the F101. Figure 1 is a pictorial of the a card A/B light off (at high fuel flows) will occur.
F110 engine. A hard A/B light-off would, of course, likely stall
the fan.
CONTROL
For single engine aircraft safety reasons the engine
The control approach represents a slight departure has a secondary mode of operation to safely accommo-
in mechanization from that on the F101 and utilized date failure of the electronic control. Afterburning
on the EMDP engine. In effect, what has been done is not available, but sufficiently high non augmented
is to make the electronic control "smarter" and re- thrust is available to assure safe take off if a
quire less of the hydromechanical control. Since failure occurs in this most critical flight regime.
the engine is being developed for a single engine To further enhance safety, air start procedures are
application, back-up features in the control system unchanged from normal to secondary and there are no
are included. For normal operation the electronic restrictions on throttle movement in the secondary
control accomplishes all control functions except mode.
for compressor variable vanescheduling. Figure 2
is a schematic of the control system operation. THERMODYNAMIC PERFORMANCE
Three aspects of the control system are particularly
noteworthy. The first is the control of fan operat- Engine SLS maximum thrust (uninstalled) ranges from
ing line through the control of fan duct mach number. approximately 25,000 lbf to 27,000 lbf depending on
Total and static pressure measurements are performed the airflow schedule. Intermediate thrust ranges
at the fan duct entrance and combined in the control from approximately 15,000 lbf to 16,000 lbf. Of
as A P/P where A P/P is the difference between total course the thrust varies throughout the flight en-
and static pressure divided by static pressure. velope depending on altitude, airspeed and air tem-
Schedules of A P/P vs engine inlet temperature, T2, perature. The basic cycle is one of moderate by-
are compared to the measured A P/P and the jet nozzle pass ratio ( — 0.8), high cycle pressure ratio
area is modulated to obtain the scheduled value. In (— 30) and high turbine inlet temperature 2500
('S-'

effect, then, the engine control is "self-trimming" °F).

in that periodic adjustment of control schedules to
maintain engine thrust is not required.

3 STAGE FAN 9 STAGE COMPRESSOR REDUCED ENVELOPE

3.2 PR 9.5 PR LIGHT OFF DETECTOR
-
(SLIMLINE)
270 PPS MAX

DE ICE
- E NO TRIM" ELECTRONIC GEAR PUMP OPTICAL PYROMETER
SYSTEM CONTROL SYSTEM

HYDROMECHANICAL
BACK-UP

Figure 1 Pictorial of F11O-GE-100 Engine

 it Y ^^
wig wv ioat _T2

x - W[
N` ^NI1^ 1Cli',!
uri-v — rs^

GttU ,if

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I w1- - r51

r^u [^u'.iu ca,u

';°cof ,'1[akY
u Inc.
:xl

SECO;JD4Fy^ 41.
a,

Figure 2. Schematic of F110 Control

than does Titamuim; hence the higher thermal conduc-

Thrust retention is a very desirable attribute in an tivity assures higher heat transfer rate than would
engine. The F110 has design features which promise otherwise be possible. In a drive to reduce engine
to virtually eliminate early engine removals due to weight, without sacrificing engine durability,
low performance. The first feature is a large amount weight was taken out of nearly every component and
of temperature margin (i.e. the difference between system of the engine. Exceptions were the core of
"red-line" turbine temperature and the specification the engine - which is required to maintain commonal-
turbine temperature). Projections are that new pro- ity with the F101 engine. In fact, weight was act-
duction engines will be shipped with nearly 100°F ually added in the high pressure turbine disks to
temperature margin throughout the flight envelope. achieve design life under Engine Structural Inte-
These projections are based on development engine grity Program (ENSIP) criteria.
tests. Furthermore, considerable effort was expend-
ed by the engine designers to maintain thermal match ENGINE STRUCTURAL INTEGRITY PROGRAM (ENSIP)
between the turbine rotor and stator during tran-
sients to assure rub free operation. The engine ENSIP (Reference 2) is a general approach for assur-
break in procedure is used to, in effect, accomp- ing the structural integrity of USAF engines. It is
lish final machining of clearances between the blade an orderly "checklist" of analyses, component test
tips and shrouds. Thereafter the stiff shroud sup- and engine tests in a development program which
port and thermally matched design minimize dete- serve to define production quality control and life
rioration due to opened clearances. management requirements. The objective of ENSIP is
"to reduce the probability of unanticipated struc-
DESIGN REFINEMENTS FOR QUALIFICATION tural problems by providing analystical and/or ex-
perimental identification of potential problems and
Certain refinements to the basic design, incorpora- to develop methods to correct these problems before
tion of additional features and changes in approach they cause field delays and significant costs." The
in engine subsystem areas were accomplished in re- F110 ENSIP builds upon the F101 ENSIP because of the
fining the design for engine qualification. common core. Essentially all of the structural de-
sign verification is being performed to satisfy ENSIP.
One of these was engine basic external dimensions.
While the engine which was flight tested in the F-16 TEST EXPERIENCE OF TIlE F110 ENGINE
and the F-14 fit reasonably well in the airplanes it
would not fit the F-15. Interference of approxi- Several papers by both government and industry engi-
mately '2" occurred at a major bulkhead. Thus to meet neers have described the evolution and application
the requirement for a common engine design for the of Accelerated Mission Tests (AMT) to the development
F-15 and F-16 the external cases were reduced in of aeronautical gas turbine engines (Reference 3).
diameter about 1z" and the controls and accessories The objective is to simulate real operational usage
were repackaged. Further making the fit a challenge to which the engine will he subjected and then to
was the requirement for anti-icing the engine front impose that usage on the engine, within constraints
frame and bullet nose. The EMDP design engines had of economics and knowledge of those aspects of usage
no anti-icing system. Space has to be available for which are most damaging to an engine. Throttle ex-
routing 5th stage compressor bleed air forward to cursions from low power to (or from) high power in-
the fan frame. To increase the effectiveness of the duce low cycle fatigure damage on not section com-
anti-ice air the inlet guide vanes' trailing edge ponents and some (e.g. HPT disks) rotating cold sec-
flaps were made of Aluminum. Interestingly, Aluminum tion components. Also continuous operation at high
has an order of magnitude higher thermal conductivity power is relatively more damaging than operation at
 U

low power because stress rupture life is consumed,

oxidation rate increases and erosion is accelerated. 60
Thus from a first order standpoint throttle movements indicated Air Speed (Knots)
over wide power excursions and operation at high
power must be simulated in an AMT. Other operation 50 Ir..s • •
is relatively less damaging and can either be ignor- • e 6p0
ed or addressed in other ways. 40 • • • •
0 •
From its initial testing in 1979 the F110 was tested 0 •
• r .- • • 8p0
to an ANT which was derived from usage surveys on F- •

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15 and F-16 aircraft. In fact, the AMT run was the
same one run on the F100 engine which currently
powers these two aircraft. This approach provided
not only a realistic assessment of the probable ser-
•
vice exposure, but also a direct comparison to the 0 e •
F100 engine which was undergoing the exact same test • •
cycle in qualifying engineering changes in CIP. 0^,

.4 .6 .8. 1.2 1.6 2.0 2.4

The AMT is primarily a hot section test. It does not
Mach Number
provide such a good test of components which may be
subjected to other primary failure modes - such as
operation at critical speeds, high mach number, par- Iigure 3. F110/F-16A Operability Demonstrations
ticular attitudes or those which are random in nat-
ure. In May of 1981 the F110 demonstrated durabil-
ity comparable to the F100 (except) hot section hard- make the engine stall. When the engine failed to
ware was in very good shape. By June of 1982 the stall, confidence was gained in the engine as con-
durability demonstration was more than twice that of tractor hyperbole was replaced by test data.
the F100 and by December of 1982, with a test comple-
tion at 5004 TAC cycles the engine had demonstrated A/B lights in the Upper Left Hand Corner (ULHC) of
nearly three times the hot section durability as new the aircraft flight envelope illustrate the effec-
FI00 engines then being bought by the USAF. tiveness of the light off detector. During tests
to evaluate optimum pilot burner flow there were
AMTS are always behind. Typically they have not been some A/B "no-lights" - i.e. the initial flow was
anticipatory of service usage. In late 1982 a new either too rich or too lean to sustain combustion.
AFT was being defined for F-15 and F-16 operation The light off detector did its job - it refused to
based primarily on F-16 usage since it is more severe. allow large amounts of fuel to flow into the A/B
That AMT compared to the previous one added several spraybars in those cases. Without a LOD, hard A/B
thousand "small throttle excursions", increased high lights and possible fan stalls would have occurred.
power operating time, increased assumed sortie length,
and added operation at Mach 1.6. and above. Mechanical Systems were shown to operate satisfac-
torily. There was a rather intensive "learning ex-
By May of 1983 another development engine completed perience" very early in the program when General
2000 TACs to this refined AMT and its hardware con- Dynamics and General Electric engineers had to
dition provided additional confidence in the inte- scramble to understand the dynamics in the fuel
grity of the engine design. line when the A/B fuel pump was turned on and off.
An aluminum elbow failed during flight and a dead
Many design features contribute to producing hot sec- stick landing was made. However a change in A/B
tion durability. Obviously blades and vanes have to pump on-off slew rate and installation of an accu-
he extensively and uniformly cooled, but component mulator in the line solved the problem and testing
efficiencies must also he high and combustor pattern resumed without further incident.
or profile factor must be low. To illustrate, 1% in
°
compressor efficiency is worth 20 F in turbine inlet F-16 system performance cannot, of course, be dis-
temperature at intermediate power for an F110 engine cussed in a forum such as this. Suffice it to say
cycle. Nominal compressor and high pressure turbine that changes, as compared to the F100 powered F-16,
efficiencies are 85% and 87% respectively and have were not judged to be striking.
deteriorated only approximately 2 in over 4000 TAC
,
cycles of durability testing. Furthermore th com- Compatibility test results are shown in Figure 4
bustor pattern factor is approximately 0.2-0.3 and which describes the range of angle of attack and
it too deteriorates at a low rate. side slip to which the airplane was flown. The
implication, of course, is that inlet distortion is
F-16 FLIGHT TEST more severe at the extremities of the maneuver en-
velope. There was no inlet instrumentation. The
An F101 DFE engine first powered an F-16 in flight engine was operated "throttled back" below its max-
on 19 December 1980. During the next five months, imum airflow capbility throughout the flight test.
58 flights for 75 hours were accomplished at the Air Its nominal capability is 270 pps, but it was oper-
Force Flight Test Center (AFFTC). The objectives of ated to 254 pps. There were no inlet changes and
the test were to evaluate (1) operability (2) mech- 254 was arrived at analytically based on an engine
anical systems performance (3) F-16 performance in stability audit using F-16 wind tunnel model data.
simulated combat and (4) compatibility. Figure 3 In point of fact, at static operation (M=0) the flow
tells the operability story very well. The engine was further cut back to 248 pps. To demonstrate net
was operated throughout the flight envelope with positive stall margin it was statically operated to
throttle excursions in combinations calculated to 254 during initial checkout.

I
 Li

commented upon favorablN: - the unrestricted throttle

and the spool down/JFS airstarts. The engine lights
off quickly and quickly (approx 24 sec) accelerates
to idle. Afterburner operational restrictions in
the ULHC in place on the F100 engine are not requir-
ed. Throttle transients from idle-max and mil-max
were performed In the ULHC to the limit of the air-
30 ! craft without engine stall or flameout.
Y y
U
rC)
^I The engine also performed well in the Reliability and
20 I

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Maintainability (R&M) areas. Flight line (0 level)
o I ® I maintenance was recorded and the engine demonstrated
I I substantially lower MMH/FH than the F100 engine.
rn ^' 10 •
C d Maintainability even in this tight installation was
Q e I! demonstrated with various components ("LRU")removals
! I and the high basic engine reliability meant no engine
caused removals were required throughout the program.
•20 •10 0 10 20
Angle Of Sideslip 1T ~ Degrees TWO IMPORTANT ELEMENTS OF THE FSD PROGRAM

ENGINE MONITORING SYSTEM

Figure 4. F110/F-16A Compatibility Demonstration,
Automatic engine monitoring has been around for a
F-14 FLIGHT TEST long time. As far back as the early 60's there were
a couple of F4's and F105's, flying with engine moni-
Subsequent to the F-16 flight test, two F1OIDFE en- toring systems. Those systems did not go into pro-
gines were used in the F-14 flight test program duction. Although the C-5 was developed with its
which was conducted in late 1981 and early 1982. MADAR system and B-1 (beginning of the 70's) develop-
The objectives of this flight test were simply to ed with its CITS neither the F-15 or F-16 were deve-
evaluate F-14 system performance and engine operabil- loped with engine monitoring systems. It is now USAF
ity. Again it would be inappropriate to discuss policy that all new engine developments will have en-
system performance in an forum such as this - perhaps gine monitoring systems. This new policy is reflect-
doubly so for an Air Force representative describing ed in development of the F109 (NGT), F108 (KC135),
Navy aircraft performance. However, some inferences F1O1 (B1), and F110 (F-15, F-16) and F100-220 (F-15,
can be drown from Table 1 which shows some gross F-16). Development of an Engine Monitoring System
TF30 characterics as compared to the F101DFE. Fur- is an FSD contractual requirement. The F11O Engine
thermore, the FIOlDFE can be oeprated to full 270 Monitoring System (EMS) is being developed as a
pps air flow capability in the F-14. Recall that maintenance aid, diagnostic tool, usage tracking and
the F-14 was intended, initially, to have the higher performance trending aid. Generally speaking an EMS
thrust, higher airflow F401. is the collection of everything (people, training,
T.O.s, computers, plans, etc.) required to accomp-
lish the EMS function. Only a limited description
is provided here and is focused more on the engine
and aircraft devices themselves.
F1OIDFE TF30-PW-414A A so called EMS Processor or EMSP is to be engine
mounted. This unit receives data from the AFT con-
MAX THRUST (SFC) 24750 (2.08) 20900 (2.78)
trol, which already receives most of the inputs of
INT THRUST (SFC) 14870 (0.67) 12350 (0.69) use to an EMS, and converts it from analog to digit-
al form. The EMSP then provides data to the Engine
AIR FLOW (PPS) 254 (nom) 240 Monitoring System Computer, or EMSC, which is air-
BYPASS RATIO 0.78 0.91 craft mounted. The EMSC does all computation to
flag faults, accomplish performance trending, track
WEIGHT (POUNDS) 3830 4176 life limited components and note limit exceedance.
Since the EMSC may not be mounted in a routinely
accessible location, it provides "No GO" input to a
remote panel located, for example, in a wheel well.
TABLE I. F1OIDFE and TF30 Performance The Data Display and Transfer Unit, or DDTU, is used
Comparison to "milk" information from the EMSC and feed it to
ground data processing facilities. Figure 5 is a
schematic of the F110 EMS.
F16XL FLIGHT TESTING

Due to the successful F-16 flight test in 1981 a high

level USAF decision was made to evaluate the F110 SUPPORT EQUIPMENT
(F101 DFE) in the F-16XL. F-16XL flight testing One of the aspects of the F110 development making it
with the F1O1DFE started in late October 1982. The attractive is the F11O adaption to F100 and F101 sup-
F-16XL is a derivative of the basic F-16 designed port equipment. Field level (i.e. base) support
for improved weapons/fuel carriage and increased equipment can in many cases be adapted directly from
range/payload/penetration speed. Two aircraft were what is already available for the F100. Figure 6
built by General Dynamics and tested at the AFFTC. illustrates the adaptation to F100 work stands and
Very limited propulsion dedicated testing was done, trailers. Note the only change required is that to
however. Two areas of engine operation were again the links which serve to support the engine. At the

5
 E

TO FLIGHTLINE AND RASE DATA

PROCESSING FACILITIES
ADDITIONAL AND GEMS/DEPOT I ADDITIONA
SENSORS I SEN SOBS
I ^ I
AUG
FAN
105 ENGINE RS232C
DISFLAY AND •______---1+
TA
_
RS232C
ORIN:3
INE AUGMENTOR I
CONTROL PROC
SYSTEM
SR^ -1NIT
TRANSFER --'S+
^R
'l ESSEN
TEM

ENGINE NO ! ., ENGINE NO 2
PS232C

MI L- STD-15539 ,I
J

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EMS FAULT _
`-^ GINEBATTERY LOW
EMDC RTU CODE M0N IT ORP413 ,I F.EMOTE
SYSTEM 0^ I PANEL
L CC"A 7:1'ER

PILOT OPTION CONTROL

(tea ^) REMOTE STATUS (FIG)
^„° ---- j READY STATUS (FIE)

AIRCRAFT
SYSTEMS

Figure 5. F110 Engine Monitoring System Block Diagram

BRACKET (RIGHT) I)

7 \/l
- 1fC^

LUG
VIEW OF
RIGHT SIDE SUPPORT
ESSOR INTERMEDIATE
,Doi

II A^Q

BRACKET

Figure 6. F110 Work Stand Adaptation

6
depot level, the existence of the F101 tools, proce-
dures and test benches is also being exploited for
the F110. Since the cores of the engines are identi-
cal and since the control approaches (i.e. electronic
supervision, hydromechanical back up, etc.) are iden-
tical and control implementation very nearly so, much
support equipment is either directly usable or re-
quires only slight adaptation.

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CLOSURE

The F110 engine design and program have been reviewed.

Qualification of the F110 design is planned for com-
pletion in September 1984. The engine is designed
for application to the F-15 and F-16 aircraft. The
engine is part of the F101 family of engines and
shares the same core with the F101 bomber engine.
Past program accomplishments provide confidence that
the F11O engine will provide the USAF substantially
improved durability, reliability, operability and
life cycle cost.

REFERENCES

1. Coalson, Michael S, ASME paper 82-GT-183 "Status

Report of the USAF's Engine Model Derivative Porgram".

2. Tiffany, C.F., and Cowie, W.D., ASME paper 78-WA/-

GT-13 "Progress on the ENSIP Approach to Improved
Structual Integrity in Gas Turbines/An Overview."

3. Ogg, V.S., and Taylor, W. R., AIAA paper 77-992

"Accelerated Mission Testing of Gas Turbine Engines."