AIAA-84-1272 Improved Supersonic Performance Design for the F-16 Inlet Modified for the J-79 Engine L.G. Hunter and J.A. Cawthon General Dynamics, Fort Worth, Tx

AIAA/SAE/ASME 20th Joint Propulsion Conference

June 1 1-1 3, 1984/Cincinnati, Ohio

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 1633 Broadway, New York, NY 10019

IMPROVED SUPERSONlC PERFORMANCE DESlGN FOR

THE F-16 INLET MODIFIED FOR THE J-79 ENGINE

h u h G. Hunter' and J. A. Cawthon.'

General Dynamics Corp. Fort Worth, Texas

Abstract

The F-16 inlet modified for the J79-GE-119 engine, the integration of a new inlet incorporated into the F-l6A/B airframe, a new inlet bleed system, the design philosophy, development tests and performance of the inlet throughout the flight and maneuver envelope are reviewed. Model da ta a re included which show the performance of the inlet in terms of pressure recovery, distortion, and turbulence. An engine bypass system was developed for providing cooling air for the hot engine case and exhaust nozzle, to provide good ejector exhaust nozzle performance, and t o m a i n t a i n s t ab le inlet operation at high Mach numbers.

The F-16/79 inlet retains the same location, Capture area, cross-sectional shape and duc t length. The additions a re a longer compression ramp composed of a fixed and curved ramp surface for improved high - Mach number performance and a throat slot boundary-layer- control for reduced shock-interaction.

These improvements provide a 20% increase in pressure recovery and 60% decrease in spill drag a t Mach 2.0.

W Introduction

The F-16/79 aircraft is a tailored version of the operational F-16A/B aircraft. The F-16/79 demonstrator recently completed a rigorous flight development cer t i f ica t ion program w i t h USAF surveillance and participation. This contractor- funded development program integrated an improved version of the proven GE J79 engine into the F-16 aircraft which required a new inlet t o be incorporated into the basic F-16 structure and systems.

The F-16/79 (Figure 1) is a single engine tactical fighter with air-to-air combat capabilities. Advanced technologies are combined to make it one of the most maneuverable fighters ever built. The a i rc raf t is powered by a 17,820 pound-thrust J79-GE-119 turbojet engine. An airframe-mounted accessory drive gearbox powers a i rc raf t accessories, and a je t - fue l s t a r t e r provides self-starting of the engine. Recently in a quick- reaction s imula ted intercept mission, the F-16/79 demonstrated its record-breaking scramble capability by going from a cold start , with no ground power to 40,000 feet and twice the speed of sound in 6 minutes and 11 seconds. A major design milestone in the inletlengine integration of this airplane modification was to develop a new multiple ramp inlet with a sophisticated bleed system which would enable this airplane to operate and

v Engineering Specialist, Member AIAA ** Engineering Specialist, Sr. Copyright * 1984 by General Dynamics

Corporation. AU rights reserved.

easily accelerate t o M = 2.Ot. The purpose of this paper is to describe the design, validation and performance process through three wind tunnel tests and a flight test.

An Overview of the Inlet Desim

A cross section of the F-16/79 inlet and duct as installed in the development a i rc raf t is shown and compared w i t h the F-16 inlet and duct in Figure 2.

The new Inlet design utilizes the present F-16A/B inlet cowl and subsonic duct so that external nacelle lines are not changed aft of F.S. 160 in. (F.S. 160 is the inlet cowl location) and internal lines a re not changed aft of F.S. 188. The inlet capture area also remains unchanged a t 826.7 i d . Supersonic performance is considerably enhanced by the incorporation of a fixed geometry double compression ramp. The ramp is composed of a 6 degree initial wedge followed by an isentropic surface t h a t turns the local flow through an additional 6.67 degrees forward of the cowl lip for a total turn of 12.67 degrees. The new d o u b l e r a m p d e s i g n r e d u c e s t h e t h r o a t a r e a approximately 21% compared with the F-lGA/B. This approach accomplishes both design and performance goals. First, it provides an efficient means for matching the inlet throat size to the lower airflow requirement of the -119 engine. Second, it turns and compresses the flow in supersonic flight t o result in a 20 percent increase in total-pressure recovery and a 68 percent decrease in spillage drag a t Mach 2.0, compared with the basic F-16 inlet design (See Figure 3).

Supersonic inlet performance is further enhanced by the use of boundary-layer control (BLC) provided by a transverse boundary-layer-bleed slot located on the ramp near the inlet throat (Figure 4). Low-energy boundary- layer air is removed through this slot and discharged through flush, aft-facing louvers located on the nacelle shoulders. This BLC system minimizes the degrading effect of shocklboundary-layer interaction, increases subsonic diffuser efficiency, and expands t h e stable operating range of the inlet a t supersonic Mach number. The bleed slot design also incorporates a differential- pressure-actuated door (see Figure 5) that closes the slot and prevents reverse flow during ground and low speed flight operations.

The inserted compression ramp is terminated a t F.S. 188 and from this point back to F.S. 335, F-16 internal duct lines are maintained. Between F.S. 335 and the engine inlet a t F.S. 361.5 a reducing transition section is inserted to match the smaller inlet diameter of the -119 engine. Figure 6 shows a comparison of flow area between t h e F-16 and the F-16/79 ducts. Also shown are enlarged views of the transition duct secondary-flow- system bypass valve mechanism as well as a cross-section of the bypass valve. The duct f low area decrease imposed by the transition section increases local flow velocity forward of the engine and along with the low-

_-

energy air removal through the bypass valve, results i n a decrease in pressure distortion and turbulence at the engine i n le t , t hus i m p r o v i n g i n s t a l l e d engine performance.

The bypass valve is the main secondary flow control device. The valve can be positioned either ful ly open or ful ly closed or at any intermediate position i n response to control signals. When the valve is open, a portion of the primary inlet airflow is bypassed. through the exposed annular slot into the nacelle. This bypass air i s drawn through the nacelle by the pumping action of the ejector nozzle. This secondary airflow provides engine Cooling air as well as enhanced gross thrust as the cool air mixes with the primary jet. At higher f l ight speeds (above 0.7 Mach), the valve pressure dif ferential is sufficiently high t o schedule the bypass valve in the open position which provides additional airflow at the high Mach numbers thus providing additional stability margin.

Preliminarv Inlet Desiun Considerations

Since the maximum inlet a i r f low (engine plus bypass) was aproximately 83% of the FlOO engine airflow, a corresponding in let throat reduction from 713 in2 (which is the F-l6A/B inlet throat area) t o 594 in2 was chosen as one of the primary test configurations. This configuration was designed with a 6 degree fixed first ramp and fixed second ramp w i t h a curved isentropic surface which turns the flow through an additional 5 degrees (see Figure 7). Along with this configuration are shown two additional designs each w i t h a Mach 2.0 shock system superimposed.

Configuration 1 was designed for operation at Mach 1.8 wi th minimum overspeed capability. Configurations 12 and 24 were designed for operation at Mach 2 with some overspeed capability. Thus, at Mach 2.0, the shock system on Configuration 1 focuses inside the cowl lip, while the other two longer ramp configurations focus the shock system outside the cowl lip. A l l configurations provided stable operation at Mach 2 but configurations 12 and 2 4 resulted in an increase in stability margin. Table 1 shows these 3 basic inlet configurations to be tested along w i t h other features such as variations in the BLC system, throat size, duct smoothing, vortex generators and sideplate variations. This test matrix was tested in the NASA/AMES 6' x 6' wind tunnel.

A t the conclusion of this f i rst wind tunnel test, i t was determined that Configuration 12 was the most promising design. Vortex generators, sideplate cutback and af t - ramp fill (smoothing t h e F.S. 188 i n l e t discontinuity) did not make a significant difference i n inlet performance and distortion and were discarded.

During the NASA/AMES test i t was discovered that the bleed system was marginal at high Mach numbers and at the low Mach numbers a reverse f low was coming back through the bleed system causing f low separation in the throat slot region. Therefore, a second test was s e t up at the General Dynamics Engineering Test Laboratory (GD/ETL) to improve the bleed flow system and develop a set of bleed characteristic curves for the bleed louvers. In addition t o the above requirements, th is t e s t also served t o the determine static performance of the inlet. This test was designated the Static Model Test (Figure 8). it was during this test that the pressure di f ferent ia l throat slot door (Figure 5) was developed t o prevent the reverse f low i n the throat slot.

From the results of the NASA/AMES and GD/ETL tests, a final wind tunnel test was formulated to test Configuration 12 with a more elaborate bleed system

including a louver exhaust system (Figure 9). This configuration has 5 bleed cavities on the ramp which are bled outwardly through the louvers and centrally through the diverter chamber t o the 3 rear chambers. This f inal configuration was tested at the Rockwell 7' x 7' Trisonic Facility for the test conditions shown in Figure 10. After a comparison of results, which w i l l be discussed i n detail in the Wind Tunnel Results section, it was determined that a single throat slot bleed (Chamber 5 ) f lowing outward through the louvers and centrally through the diverter through chambers 6 and 7 would be adequate for the Mach 2.0 requirement. This was confirmed by the f l ight t e s t series which v e r i f i e d tha t th is f i n a l configuration was more than adequate to provide the necessary operating flow margins for Mach 2.0 flight.

L

Model DescriDtion and Instrumentation

A summary of the F-l6/79 inlet Test Program is shown in Figure 10. The first tes t was conducted in the NASA ARC 6' x 6' wind-tunnel using a 0.15 scale model consisting o f a forebody, boundary-layer diverter, p r i m a r y i n le t , subsonic duct, c o m p r e s s o r - f a c e instrumentation and remotely actuated throt t le plug for the primary inlet. The compressor face instrumentation consists o f 4 0 steady-state total pressure probes and 40 high-response probes where each pai r o f probes i s mounted joint ly in a single unit as shown in Figure 11. Throat to ta l pressure rakes are also shown in Figure 11 along with centerline static pressure locations (B.L. 3) on the upper ramp of the inlet. Other total and stat ic pressure instruments are located down the subsonic duct, along the diverter and in the ori f ice plate at the rear end of the diverter bleed chamber to measure bleed f low rates. The GD/FW ETL stat ic test used the same instrumentation that was used i n the NASA ARC wind tunnel test.

The Rockwell Tr isonic Wind Tunnel Test was L conducted in a blow-down faci l i ty wi th a 7' x 7' test section. The average run lasted between 2 0 and 40 seconds. Instrumentation was more l imi ted i n this faci l i ty and only one r ing of high-response compressor- face data were taken. The r ing selected for high response data acquisition was r ing 3 , where the data was processed specifically for input into the Melick code for instantaneous distortion levels. Figure 12 shows the test set-up at the Rockwell facility. A boundary-layer-bleed flow plug was used to control and measure the inlet bleed flow for a portion of the test and the louvers were used for the remainder of the test. A bypass plenum and flow plug were provided to simulate the engine bypass system.

Wind Tunnel Test Results

Figure 13 shows the major i n le t performance c h a r a c t e r i s t i c s of t h e F-16/79 basel ine i n l e t (Configuration 12 ramp and single slot bleed through bleed chambers 5, 6, and 7) along w i t h the corrected airflow as a function of freestream Mach number. The extended ramp and bleed system serve t o increase performance levels at the higher Mach numbers when compared to the normal shock inlet of the F-lGA/B. Note the bypass valve opens up at approximately Mach 0.6 and remains open at high speeds. The pressure recovery i s affected slightly a t M = 0.5 t o M = 0.8 by the opening of the bypass valve where the throat Mach number i S increased with the increased bypass airflow. In addition, the throat bleed closure door starts to open at approximately the same t ime the bypass opens, which has some effect on recovery. The above figure also shows the effect of sideslip on pressure distortion to remain w e l l below Nc = 0.12, which i s an upper distortion l imit. The parameter "Nc" i s defined in the JTS-GE-119 Engine Specification E2259. This parameter is a circumferential

-,

distortion indicator which has limiting value of 0.12 for the 5-79 engine. For values of Nc less than 0.12 (which is a very conservative steady-state value) the distortion levels may be regarded as stall free. Flight test data indicate that the Nc l i m i t is a very conservative number and that engine stalls are not likely to be encountered until an Nc value of a t least 0.20 to 0.25 is encountered (see Figure 14).

J

At high speeds both the NASAIAMES and Rockwell test data are combined and shown in Figure 15 where pressure recovery is slightly higher for the Rockwell test and the stable airflow margin is significantly increased. Two effects are responsible for this: (1) increased bleed in the Rockwell test and (2) higher Reynolds number in the Rockwell tunnel (1.5 x 106 vs 9 x 106). The pressure recovery differentials a re more substantial at higher angles of attack. Distortion is also decreased in the Rockwell t e s t as a result of less boundary-layer interaction due to the higher Reynolds number and greater bleed. A further comparison of t h e variations in t h e bleed system is shown in Figure 16 where performance with the porous ramp and throat slot are compared w i t h the throat slot alone. Perfomance levels are not significantly changed, however with the porous ramp the inlet is stable over a much larger airflow range, especially at Mach 1.8 and 2.0.

The subcritical s tab le operating range of the F-16/19 inlet a t cruise flight attitude is shown in Figure 17. A t Mach 1.6 and above, the least stable operating point (indicated by a vertical line on the left-hand side of the pressure recovery characteristic) was defined during the wind-tunnel test. A t duct corrected airflows less than those indicated by this limit, boundary-layer on the compression ramp will separate, the shock system wil l become unstable and begin t o oscillate and the inlet wil l be operating in an unstable mode commonly referred to as "inlet buzz". In this mode of unstable operation, pressure recovery a t the engine inlet wil l decrease and distortion and turbulence will increase, possibly resulting in compressor stall.

A s ment ioned a b o v e , t h e i n l e t o p e r a t i n g characteristics shown in Figure 17 apply only for cruise attitude flight. During certain aircraft maneuvers where a pushover or sideslip is imposed, the m i n i m u m stable operating airflow will be more than that shown. Limited data on the effect of maneuver were acquired during the wind-tunnel tests and are shown i n Figure 18.

During low speed flight operation, the throat static pressure is much lower than the ambient s ta t ic pressure thus inducing a reverse flow in the BLC system. At Mach 0.6, Figure 19 shows the effects of the reversed flow on inlet performance. Smoke tests performed a t these same conditions show strong vortex flows in the corner regions of the throat slot which will degrade the performance and increase distortion. To eliminate this problem the throat slot door was added t o the design.

Figure 20 shows the ramp and duct static pressures just inside the inlet at M = 2.0 as a function of duct airflow, where airflows greater than 190 lblsec represent supercritical flows. Also shown a re the inlet throat region total pressure rake measurements. The flow is well-behaved in the inlet throat region for the sub- critical flows. For the supercritical flow conditions, some separation is observed on the upper ramp surfaces

-which is indicated by the decreased t o t a l pressure measured by t h e station 187 throat rake.

Flight test data a t M = 2.0 (Figure 21) shows the inlet ramp pressure distribution along with the Rockwell

wind tunnel data. The flight test ramp data is in good agreement w i t h the tunnel data. Using th i s data as a reference, t ime varying da ta were taken during a pressure anomoly (intermittent buzz condition) where ramp pressures a re shown as a function of time. The shock movement is correlated with the PS3 compressor pressure in the engine, the PLA, RPM and EGT for this time sequence. The pressure in the inlet bleed cavity during this t ime sequence also increases as the shock moves out on the ramp which increases the bleed through the louvers and tends to stabilize the inlet flow.

E n ~ n e - Inlet Compatability C i rcumfe ren t i a l t o t a l p ressure d i s to r t ion is

measured by the distortion parameter, Ne. The steady- s ta te limit value for Nc is 0.12, where Figure 14 shows the F-16/79 distortion to be well below this l imit .

A historical record of the Nc pressure distortion measured in flight tests on three different aircraft using the 579 engine is also shown in Figure 14. Again, the F-16/79 data show a wide margin of safety.

Eight dynamic total pressure measurements were also recorded at the engine face during t h e wind tunnel tests. These data were used with the Melick methodology t o predict m a x i m u m instantaneous t o t a l pressure distortion. This prediction is compared w i t h steady-state resuits and shown across the ranges of test Mach number and maneuver attitude in Figure 22. it is noted that the very conservative specification Ne s tea ty-s ta te limit value is exceeded by the instantaneous distortion only at high subsonic Mach number during severe sideslip conditions.

The results of the engine-inlet compatibility have been well-demonstrated in a comprehensive flight test program during which there have been no incidents of engine stall.

Concludina Remarks

A fixed double ramp inlet with a throat slot bleed system has been developed for the F-16/79 derivative of the F-IGA/B. Three wind tunnel tests were required to optimize the inlet design and performance. In the subsonic and transonic Mach number range, the F-16/79 performance distortion are essentially unchanged from the F-16AIB levels. However, at Mach 2.0, performance is increased 2 0 % and spill drag decreased 60% over the F-lGA/B levels.

Various ramp/throat bleed configurations were tested where the final configuration chosen was a single transverse throat bleed slot with a bleed closure door. A bypass valve opens a t approximately M = 0.6 and remains open at high speeds which stabilizes the inlet.

However, should higher Mach numbers than 2.0 be required, provision for additional ramp bleed cavities have been incorporated into the current airplane design, and can be activated should that option be exercised.

The results of the engine-inlet compatibility have been well-demonstrated in a comprehensive flight test program during which there have been no incidents of engine stall.

..;

.+.

Figure 1 F-16!79 Fixed Ramp Inlet Compared With F-16A/B Normal Shock Inlet

F I B INLET RAMP TE INTERSECTS F.16 OUCT AT FS 188

PRESENT F.16 COWL AND OUCT LINES BELOW WL 70.9

Figure 2 Model E-14/79 Compression Ramps

3(a) INLET PRESSURE RECOVERY COMPARISON

J79-GE-119 ENGINE COMBAT PLUS

ALT = 30.000 FT 0070

SPILLAGE DRAG

COEFFICIENT

C D SPILL

/ /

0 0 4 8 1 2 1 6

MACH NUMBER. Mo

3(b) INLETSPILLAGE ORAG COMPARISON

Figure 3 F-16 and F-16/79 Performance Comparisons

(Incommg Ai r )

Figure 4 Internal View of the Model F-16/79 Inlet

CENTER DOOR HINGE 7 ELC BLEED SLOT

DAMPERJ

Figure 5 Inietdlot Boundrry Layer Bleed Closure DoodDamper

Figure 7 Mach 2.0 Shock System for Candidate Inl& Designs

u

VALVE OPENIXB

W Figure 6 F-16/79 Inlet Duct Area Distribution and Transition Duct and Bypass Valve Arrangement

Table I F -16 ,79 INLET DESIGN CHARACTERISTICS

4

- CONFIG.

NO. - 1

2

14

12

15

21

22

23

24 -

~

RAMP L.E. FUS. STA

(1N.I

134.3

134.3

134.3

131.3

131.3

131.3

131.3

131.3

130

THROAT SIZE LIN21

594

594

594

562.4

562.4

562.4

562.4

562.4

563

TYPE BLC

THROAT SLOT

THROATSLOT

THROAT SLOT &POROUS PLATE

THROATSLOT

THROATSLOT

THROATSLOT

THROAT SLOT

NONE. TAPE0 SLOT

THROATSLOT

SIDEPLATE CONFIG.

SOLID

SLOTTED

SOLI0

SOLID

SOLI0

SOLID

SOLI0

SOLI0

SOLlD

SUBSONIC RAMP

BASIC

BASIC

BASIC

BASIC

BASIC

FILLEO

FILLEO

FILLED

BASIC

- VG'I

- NONE

NONE

NONE

NONE

YES

NONE

YES

YES

NONE

F i p r e 8 0.15-Scale F-16179 Inlet Model in the General Dynamics Static Test Facility

BLEEO CHAMBERS

OlVlOER PLATE

A

B L 0 t -POROUSPLATE

BLEED CHAMBER ITYP)

8.L. BLEEO EXIT LOUVERS

BLEEO CHAMBER I N FALSE OIVERTER

SECT A 4

Figure 9 Rockwell Inlet Model Detail

A 0 15-SCALECOMPOSITE WIND.TUNNEL MODEL

FULL-SCALE SLC SYSTEM MODEL

IWO WINO-TUNNEL ENTRIES

iNLETlENGlNE COMPATIBILITY CONFIGURATION DEVELOPMENT

NOV - JAN 1979.60

NASA A R C 6 I 6 WT

REiFT I 5 10'

0 I5.SCALECOMPOSlTE INLET MODEL WIF-16 COMPRESSOR FACE

THROAT-SLOT 8 2ST POROUS- RAMP eLc

MACHRANGE 0 6 7 0 2 0

*LYRANGE -5'TOm'

.ORANGE O ' T O ? I '

STATIC MODEL TEST . MARCH 1980 GOIFW ETL

f 0 I S ~ S C A I E MOOEL

* COMPOSITE CONFIG 6 ISOLATE0 INLET WIBELL-

THROAT~SLOT 8 !STPOROUS- OAMP 8LC

. Mlil 1980 GD'FWETL

It2-SECTIONOF FULL~SCALE 8LC SYSTEM

. JUNE 1960 ROCKWELL TRISONIC 7 r 7 W

REiFT ~ 9. 10 . 0 15.SCALECOMPOSlTE INLETMOOEL W/F-16,79 COMPRESSOR FACE . THROAT-SLOT 6 2ND POROUS- RAMP eLc . MACH RANGE 0 TO 2 0 LYRANGE . 5 * TO 20-

0 RANGE 0'10 10'

Figure 10 Summary of Model 161'79 Test Program

INSTRUMENTED SUBSONIC DUCT I

REMOVABLE FOREEODY 40 PROBES LOCATED IN

EOUAL AREAS HIGH-RESPONS PROBE d? , r7

STEADY-STA

. .' -..-.._

DUCT RAKES PROBE FULLY INSTRUMENTED (Steady-Stale and High-Response) SIMULATED COMPRESSOR FACE

*GD/FW ETL STATIC TEST

STEADY-STATE

41 STATICS

72 TOTALS

HIGH-RESPONSE

40 TOTALS (3 C.F.

48 STATICS

72 TOTALS

HIGH-RESPONSE

8 TOTALS @I C.F, I :::z 1 I ::K: Figure 11 Inlet Wind Tunnel Model and Primary Pressure Instrumentation

CUTAWAY VIEW //-?T-..

BYPASS PLENUM

TUNNEL CEILING

PRIMARY FLOW PLUG

E.L. BLEED FLOW P

TUNNE5FLOOR

Figure I 2 Rockwell Trisonic Wind Tunnel Test Setup

PRESSURE RECOVERY B l P L I O CLOSED

I ,I ,. 0 3 ' I'

0 4 3 1.2 1 6 2 0

CORRECTED AIRFLOW 200

180

W,% 7

160

140

4 3 1 2 1 6 20 M.3

0 120

Figure 13 F-16/79 Inlet Performance Characteristics

0 4

03

PRESSURE DISTORTION

NC 0 2

PREDICTED INSTANTANEOUS-

07

STEADY-STATE -

0

F-4C (SOLID -STALL, RA5-C F - 1 0 4 A i S

I

05 10 1 5

MACH NUMBER Mo

Figure 14 Inlet Pressure Distortion Comparison

0 M-2.0 A. -5 E 0 F 161179 0 M.2.O A- -2 8.0 0 M.2.0 A- 1 8.0 TEST NO 314 b M.2.O A. IO 8.0 OATE 11ilP-IIII(l V M-2.0 A. 20 0-0 CONFIG 12

S O L I O S Y M B O L S R F P R E S E N T R O C K W E L L O A T A OPfN SYMBOLS REPRESENT NASAIAMES OATA

NASIAMES WT TEST

1.0

El c

: t 0.9 > = w > D u : 0.5

E .A M Y

0.1 100.0 150.0 200.0

WCZ. LBISEC 0.06

e. c e

t

5 : 0.02

0.01

I Y

rn = c

0.00 100.0 150.0 200.0

WC2, LBISEC

El I

> m x

: 10.0 > m

Y > a 2

(E

a m I

D m

> 5.0

0.0 100.0 150.0 200.0

wc2. LBISEC - 0.1 - c &

e 4

e a

: 0.2 z c

El c

s

E 0.1

a I

0.0 100.0 150.0 200.0

WC2, LBISEC

STALL

r UNCERTAINT

4- STALL FREE MODEL

SPEC LIMIT

Figure 15 Comparison of Rockwell and NASAIAMES Data at M, = 2.0 Show the Effect of Angle of Attack

WZ 081 091 OD1 021 W l t8'

98'

88'

ob'

x

ts

96'

86'

0 0 1 lNlOd lV3111M3

0 'd

d

- Z I

ooz 081 091 o w OZL 001 w

98'

88' '11 D v) rn

0 6 ' "

rn 5

w r n n

96' <

86'

DO 1

h

h

MACH NUMBER

Figure 18 Effect of Aircraft Maneuver on F-16/79 Inlet Stability Limit

w c Q U C I L B I S I C

Figure 19 Effects of Closed vs Open Throat Slot at M = 0.6

0.8

PtIPt,

0.4

0.01 I I I I I 1 130.0 140.0 150.0 160.0 110.0 180.0 190.0

STA 166 8 1 3.0 I

8.0

Pt/Pl,

4.0

0.0 0.0 1.0 8.0

DIST. FROMSUAF. - IN.

.OWER STA 166 8L 18.0

RAKE

8.0

1 .o

0.0 0.0 4.0 6.0 0151. F R O M SURF. -IN.

M = 2.0

CONF. 1 2 0 = 1.0; 0

RUN WCZ - - 0 - 218172 0 - 218178 A - 218182 0 - 218190 v - 218200 0 - 218219

Figure 20 Duct Static and Total Pressures

PSI IPSIAI

PLA

RPM

EGT 1%

'W TIME, SEC FUSELAGE STATION. IN

Figure 21 Flight Test Data Showing Effect o f Pressure Anomoly on Ramp Static Pressure, PS3. PLA. RPM and EGT

.*%I ,.,- ,,,..c-.

D..e.10 .~sl^ .l~" ..-. ,,.-.*

-l(- I .>I I m. :I*. I I " I I I ., I I .. / I ,. I I ,.I I I* I .- -.lj- i * I , I * . I I * a i 8 " . , ! .. ~ ! . I I , * / I ..I -Is1 1 I . d I. I # * . I ! * . / I .. I i .. I1 (. I1 . I I1 I. I Figure 22 Comparison of Steady-State and Predicted Maximum Instantaneous Distortion at the Engine Inlet