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Z -JZ _- hJ --J This presentation is an overview of a joint NASA Lewls-McDonnell Aircraft Company Hot Gas Ingestion (HGI) test program in NASA Lewis' 9'x15' Low Speed Wind Tunnel (LSWT). This initial program is scheduled for testing in late 1986.
Advanced short takeoff/vertical landing (ASTOVL) aircraft capable of oper- ating from remote sites, damaged runways, figure l, aircraft carriers (figure 2) and small air capable ships are being pursued for deployment around the turn of the century. To achieve this goal, it is important that technologies critical to this unique class of aircraft be developed, ref. I. One of the ASTOVL concepts, the vectored thrust, has as its critical technology item, the potential of hot gas ingestion (which occurs during vertical flight operation while in ground effect) as a key development issue. Recognizing this need, NASA Lewis Powered Lift Section and McAir have defined a cooperative program for testing in the Lewis' g'xl5' LSWT.
(Rex) GROUND ATTACK / _ ' STOVL4oo It , ziJ STOL ; CTOL " _._'_"_ ,o oo In_!a[Ph.,ases Of Conflict _15 After Runway m ueivory a e Bombs/Hr .
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IOONM 200NM 15OHM 40QNM (AGARD CP 313) Distance To Target CO-B-I_ Figure 1.-STOVL improves Air Force operational effectiveness.
NAVAL CARRIER OPERATION
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TAKE-OFF,CATAPULT OPERATION CONVENTIONAL BATCH RELOCATE,REFUEL, REARM LAND, ARRESTING GEAR STOVL STOVL
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CONTINUOUS r/l/Ill/Ill OPERATION Sortie Rate LAND Figure 2_ -' STOVL improves Naval carrier operational effectiveness.
279-3) An artist's conceptual view of the vectored thrust concept (Model is shown in figure 3. The aircraft concept consists of: I. single engine; 2. bi-furicated inlet; 3. VTOGW 30,000 Ibs; 4. M Max. = ?.0; and two aft 5. Four nozzles - two forward NASL C - 8 4 - 4 5 4 0 Figure 3. - Model 2 7 9 - 3 w i t h improved L I D 8 and deflector.
The two f r o n t d e f l e c t o r nozzles w i l l be r e q u i r e d t o accommodate b u r n i n g o f the fan a i r f l o w . The two a f t d e f l e c t o r nozzles w i l l c o n t r o l t h e core a i r f l o w .
The concept may also have the f o l l o w i n g : 1. f r o n t f l o w d e f l e c t o r ;
2. sidewall d e f l e c t o r s (streaks)
The t e s t i n g o f t h i s vectored t h r u s t concept r e q u i r e s a unique model sup- p o r t system and modification t o t h e 9'x15' LSWT t e s t section.
The next f i g u r e ( 4 ) shows a schematic o f the 9.2% scaled Model 279-3 i n -
s t a l l e d i n t h e 9 ' x l 5 ' LSWT w i t h the unique model support system The model
support system provides f o u r degrees of freedom: V e r t i c a l movement, yaw, p i t c h , and r o l l c a p a b i l i t i e s . The v e r t i c a l movement range i s f o u r f e e t above t h e ground plane; yaw angle range i s + 1800; p i t c h angle range i s + 3O0, and the r o l l angle i s + 200 range. h o t h e r f e a t u r e shown i n f i g u r e 4 i s the: Ground plane which-has a s l i d i n g t r a p door.
3 4 4
PITCH DRIVE N LINE (MOVABLE) ROLL DRIVE SYSTEM- YAW DRIVE iiiiii MODEL H20 )T GAS LINES -MOVABLE SCREEN- TO EXHAUST FAN P BOTTOM OFTUNNEL Figure 4.-Schematic of model 279-3 and support system installed in the 9'X15' LSWT.
We have built in flexibility in this program. The aircraft inlet airflow is controlled independently of the nozzle airflow. The inlet airflow is con- trolled by a vacuum system and the nozzle airflow is supplied by a high pres sure-hot air system, with temperature ranging from ambient to lO00OF at the nozzles. The freestream velocity will vary from static to 65 kts.
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Iota Slots ---_
\
Ground Plane
4---Open for venting /
Open for venting-----_ Tunnel Floor Figure 5. - Modified 9'X15' Low Speed Wind Tunnel.
A cross-section of the 9'x15' LSWT is shown in figure 5. The 9'x15' LSWT has slotted sidewalls test section. The tunnel sidewalls will have an opening near the ground plane to allow the laterally-flowing hot gas from the exhaust nozzles to exit the test section.
THE PRIMARY OBJECTIVES oF THIS COOPERATIVE PROGRAM ARE TO INVESTIGATE TECHNIQUES WHICH WILL: O MINIMIZE/ELIMINATE HOT GAS REINGESTION DURING VERTICAL FLIGHT OPERATIONS WHILE IN GROUND EFFECTS.
O PERMIT PREDICTION OF OPERATING CHARACTERISTICS OF VECTORED THRUST CONCEPTS WITH FORWARD VELOCITY.
IN ADDITION, THE TEST PROGRAM WILL ESTABLISH A HIND TUNNEL HOT GAS INGESTION DATA BASE FOR: O BOTH NEAR/FAR FIELD INGESTION O FOUNTAIN FLOW EFFECTS, AND O GROUND VORTEX FLOW FIELD.
THE DATA BASE DEVELOPED SHALL BE APPLICABLE TO THE DEVELOPMENT OF UNIQUE ANALYTICAL CODES.
Figure 6.-Primary objectives.
The program objectives are shown in figure 6. In addition to the primary objectives, we shall establish a database in several needed areas, one of which is the ground-vortex-flow-field-jet interaction. The objective is to answer the question of what effect, if any, the boundary layer thickness has on the ground-vortex-flow-field-jet interaction.
The figures which follow are used to indicate the type of data parameters we will investigate. The trends shown on the figures are considered typical.
The results of the boundary-layer study, figure 7, will indicate the Forward extent of the ground vortex flow field-jet interaction due to the bo_ndary-I ayer thickness.
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qoa qN, x V- I Boundary-Layer Thickness X/dN, Figure 7.-Boundary-Layer thickness effect on the ground vortex flow.
Shown in figure 8 is a means of thickening the boundary-layer. Shown is a boundary-layer thickness configuration which consists of I/4" dia. rods in four rows. The rods would extend the width of the ground plane. Several configurations could be utilized; for example, 3" height rods or 6" height rods, to obtain several different boundary-layer heights.
0 0 o o o o 0 0 o o o o o o 0 0 0 0 0 o o 0 o o 0 o 0 0 0 0 undary-Layer Thickener Rods Configuration B 4_ Configuration A Figure 8.-Boundary-Layer Thickeners.
In addition to controlled thickening the boundary-layer, we also need to minimize the boundary-layer height due to axial distance. Figure 9 showns two methods of reducing the boundary-layer height. We have considered three tech- niques; a rotating belt was considered but was eliminated due to complexity and the temperature environment involved (lO00OF). As shown, another con- cept involves using a suction pump(s), which is located outside of the test section. The suction pump(s) would remove part or all of the boundary-layer.
The least concept involves lowering the front section of the ground plane.
This technique would relocate the initial boundary-layer growth point. Prior to the use of either of these concepts, we shall have established the extent of the ground vortex flow field on the ground plane.
Freestrsam Airflow ,,. I _mall suction pump Ground Plane / located outside of test section.
Freestream Airflow
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_'_-- Ground Plane Figure 9.-Boundary-layer removal configurations.
One mechanism for near field ingestion is the jet fountain. If the velo- city of the fountain or turbulence intensity is reduced, the effects of the near field ingestion will also be reduced. A means of reducing the fountain velocity and turbulence is to vary the front nozzles splay (laterally movement of the front nozzle) angle. It is anticipated that results will show a reduc- tion in both fountain jet velocity and turbulence intensity with increasing splay angle, as shown in figure lO.
Front Nozzles
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/ / / / / /
Front Nozzle Front Nozzle Splay Angle Velocity Turbulence Angle Ratio Intensity of Fountain Jet O" 0" 12' 12 e Distance from Centerline Figure l O.-Fountain turbulence and velocity.
Figure II illustrates the various configurations we will test to obtain the fountain flow characteristics. Shown is a schematic of four nozzle ar- rangements and the auxiliary inlets. The first configuration is Model 279-3 concept with both sets of nozzles flowing. The second configuration consists of only the front nozzles flowing. The third configuration consists of only the aft nozzles flowing. And the fourth configuration simulates a twin engine aircraft with one engine out. The jet temperature range is interchangable between the front and aft nozzles. These configurations will produce consid- erable information on the ground-vortex-boundary-layer-interaction.
Auxiliary Inlet
r-1 nn B
I Nozzles: [ 0 0 0 ,-- Forward ---, _-- Aft-_ [] D [] [] Figure 11.-Model deflected jet configurations.
Addressing the primary objectives of this joint NASA Lewis-McAir program, the major concern is hot gas ingestion in both the near and far field. In determining the effectiveness of the ingestion avoidance devices (IADs) for near field ingestion, the inlet temperature rise v.s. nozzle exhaust temper- ature will be plotted as shown in figure 12. Results from the configuration without IADs will be compared to a configuration with IADs. In general, a reduction should occur with ingestion avoidance devices.
_-- wlo IADs Inlet Temperature Rise
/
/ with IADs
/
/
/
Nozzles Exhaust Temperature Figure 12.-Effectiveness of ingestion avoidance devices in reducing inlet hot gas ingestion.
Figure 13.- Near field ingestion avoidance devices (lADs).
Figure 13 shows the three primary configurations: I. without IADs, 2. with IADs, option l: flow deflector and congitudinal streaks (2) instal Ied, 3. with IADs, option 2: flow deflector and congitudinal _treaks (2), aft fence and flow deflection sidewall (2) installed.
These configurations will be tested with the auxiliary inlets in the open and closed positions.
In addition to the near field, data applicable for determining the far field ingestion effect will also be obtained.
Pressure/temperature rakes are located on the ground plane (forward and aft of the model), figure 14. Also tufts will be located on the ground plane to give an indication of the far field airflow movement. The ground plane will contain static pressure and temperature taps. A thermo-vision system will be utilized to detect the most forward point of the hot gas at the vari- ous freestream speeds.
wlo IADs Configuration / Entrainment Separation caused by buoyancy and entrainment of Induced Inflow.
Figure 14. - Far field ingestion.
In the next several figures we shall briefly review some of the instru- mentation to be utilized during the test.
Figure 15 illustrates several of the rakes installed on the model. They are as follows: I. Nose boom rake which is used to measure the local freestream conditions.
2. Inlet plane undersurface rake which is used to measure the quali- ty of air entering the inlet region.
3. Fountain upwash rake will measure the upwash flow characteristics.
The rakes contain both total pressure and temperature measurements.
Nozzle hot air supply Inlet suction__\ i ; /'/ // Figure 15. - Model 279-3 external instrumentation rakes.
In addition to rake instrumentation, static pressure taps and high re- sponse thermocouples are located along the bottom and sides of the fuselage, as can be seen in figure 16. Using the fuselage instrumentations, we should have a good indication of the thermo-profiles along the fuselage.
Model'surface high response thermocoul ,i 0.062 in.
8urfacs_ X.._._
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detail 0,062 In.
Casing ._,,_it"iJ-,_ Thermocouple _" wires Figure 16. - Typical fuselage instrumentation.
Another major region of concern is the inlet. Vie need to know what effects inlet temperature rise has on the fan face Mach number rise. Figure 17 shows what might be a typical plot of inlet temperature rise vs inlet fan face Mach number. That is, the inlet temperature rise reaches a plateau at some fan face Mach number. This particular curve is a function of the model height above the ground plane.
X Inlet "-. f(h) Temperature Rise Inlet Fan Face Mech Number ,Figure 17. - Inlet fan face temperature rise., Typical model inlet and nozzle instrumentations are shown in figure 18.
The nozzles contain total pressure and temperatures probes. The engine fan face rake will also contain 32 total pressures and temperature measurements.
To determine the severity of the hot gas ingestion, the inlet temperature rise and contour maps will be obtained utilizing the fan face rake. A typical contour map of a fan inlet instantaneous temperature profile is shown in figure Ig.
Nozzle Engine Face Rake Porous Plate-_ _Pt Pt___-i TIC _--Stabilizer Ring _,_g .......
_xrThermocoup le ._. _ _ _ A i _Total Pressure \\ __ ) ) Probe, Pt _ "_ / _Static Pressure/_ \\ _ Pt'-/ _ Section A.A Figure 18. - Typical model inlet and nozzle instrumentations.
0 o zT0° t -@ Figure 19. - Contour map of the fan inlet temperature profiles.
A t the conclusion o f the 9'x15' LSWT t e s t , we w i l l have pressure/tempera-
t u r e contour maps a t the f a n face f o r various freestream v e l o c i t i e s and model a t t i t u d e s . Rut what we would l i k e t o a s c e r t a i n as an end i t e m is t h e e f f e c t t h e h o t gas i n g e s t i o n has on t h e actual engine.
W e a n t i c i p a t e , as a f o l l o w - o n program, u s i n g b o t h t h e pressure and temper- a t u r e d i s t o r t i o n p r o f i l e s from t h e 9'x15' LSWT program and implement these i n t o a f u l l scale engine program. T h i s f u l l scale t e s t i n g would e s t a b l i s h t h e c h a r a c t e r i s t i c s o f t h e engine s e n t i ti v i ty due t o t h e temperature, pressure and
a combination o f temperature-pressure d i s t o r t i o n . A t NASA Lewis Research
Center, we have an a l t i t u d e t e s t chamber (PSL) i n which w e do f u l l s i z e engine t e s t i n g . F i g u r e 20 shows a view of t h e A l t i t u d e l e s t Chamber w i t h a TF-34 engine i n s t a l l e d . This f a c i l i t y ' s a l t i t u d e s i m u l a t i o n range from sea l e v e l t o l O O K feet.
F i g u r e 20. - T F - 3 4 engine i n s t a l l e d in the A l t i t u d e T e s t C h a m b e r .
i l l u s t r a t e s t h e e x t e n t o f t y p i c a l engine instrumentation. The
n consisted o f steady-state and dynamic t o t a l pressures, s t a t i c
pressures, and t o t a l temperature measurements. 7-ransient t o t a l temperature
and h i gh-response pressure data are a1 so recordes.
I 35 6
m2.5 Front Labyrinth ,---Rotating / f-2.6 Ikhead_ sea!7 / screen i-2 2.47, I / j-2. 7 , 2C F3 ,)a // r I / i L Two rows of 2D 3 L 2R Station 1 static taps at // Hydrogen / 550 and 2350 burner -/ 0o 10 o 0 o 0° 200300 3250 II
o
315o\ _"_ /450 o31° \/ ! ',,F,'_ 30o \2\ i _/ \ 27o°_I_,,o,0,,.o( "--_---)o,o_,,_I- _0o \ /_ ; "w"\i2o o / i / \13°° 250°I_ _25o_"/_LLL-_>,oo" O 2o2 _o,/___5o I_5O 180-_In° /_ LAU _00 200Gl_0U 1800 ....
Station 1 Station IB Station 2 3470 0o 120 0o 3350 ?o 270 323°_ ,450 3060 . 530
2700 _ oo
,_ ', ._" ,N.5_ o o ! '_ "-'oJ '118 225o 143 ° 225° 1920 180o 180o 2070 1800 Station 2. 4 Stations 2.5, 2.6 and 2. l Station 2C 34003550 0° x Steady-state total temperature o Steady-state total _-_r 295 ° pressure iI o o Steady-state static 90 o pressure Yaw pressure probe Boundary layer yaw probe / B Stages6 to 1:3 Stages 1 to 5 C Stages 1 to 13 2250 / \ 19_ 160° Stations 2D to 2R Station 3 Figure 21. - Instrumentation layout for the TF-34 turbofan engine.
(Stations viewed looking upstream.
See symbols for description of station locations.)
Figure 22. - Pressure distortion generator with rotatable screen
as s em bly .
Pressure and temperature distortions can be imposed on the f u l l size en-
gi ne by using pressure and temperature d i storti on generators. In1 e t pressure
distortion (pressure lower t h a n average) i s generated using one o f three
screen configurations, figure 22. The pressure distortion circumferential
extent o f a 180"can be varied by a rotatable screen assembly which i s mounted
upstream o f the engine inlet,
The gaseous-hydrogen-fueled b u r z c r device, f i g u r e 23, i s used t o produce t h e
tirne-dependent temcerature d i s t o - t i o n and i s i n s t a l l e d upstream o f t h e engine i n l e t bellmouth. The burner has the c a p a b i l i t y o f being r o t a t e d + 300 from t h e c e n t e r p o s i t i o n and i s divided i n t o f o u r i n d i v i d u a l l y controTled quad- rants. A i r passing through t h e burner i s heated i n selected 900 sectors.
Each s e c t o r has t h e f o l l o w i n g : 6 swirl-can p i l o t burners, i g n i t i o n source f o r hydrogen.
1.
2. 6 annular g u t t e r s supported by 1 r a d i a l g u t t e r , 3. 6 c i r c u l a r - t u b e manifolds ( 1 i n s i d e each annular g u t t e r ) w i t h m a l l holes f o r hydrogen i n j e c t i o n .
H i gh-response valves c o u l d Se energized i n any d e s i r e d combination t o produce t h e temperature d i s t o r t i on.
Figure 23. - T o t a l temperature distortion g e n e r a t o r w i t h a
g a s e o u s - h y d r o g e n - f u e l e d b u r n e r .
By using the above distortion devices, we can arrive at the distortion sensitivity parameters for the engine inlet as shown in figure 24. The engine stall line is temperature and/or pressure distortion sensitive. This is ulti- mately the type of information you need to know about the model-inlet-engine characteristics.
Solid symbols - stall Open symbols - nonstall Stall line Q Pressure __ Estimated stall line_. //j_._ distortion a_/St all line amplitude Temperature distortion amplitude Figure 24. - Distortion sensitivity at the engine inlet.
In conclusion: I. We shall obtain data which will permit prediction of operating characteristics of vectored thrust concepts with forward velocity.
2. We shall minimize/eliminate hot gas ingestion during vertical flight operations while in ground effects.
3. We shall establish a data base for near and far field ingestion, fountain flow effects, and ground vortex flow field- jet interaction.
4. We shall also obtain _istortion results which can be utilized for full size engine testing in the altitude test chamber facility.
_ N 5. We shall obtain the extent of ground effects on the vectored thrust ASTOVL concept.
6. It is important to develop analytical codes which will predict the overall effects of hot gas ingestion.
REFERENCE I. Kuhn, R. E. and Eshleman, J., "Ground Effects on V/STOL and STOL Aircraft - A Survey," NASA TM 86825, Nov. 1985.