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NASA Technical Memorandum 89080

HIGHLIGHTS OF UNSTEADY PRESSURE TESTS ON A 14 PERCENT SUPERCRITICAL AIRFOIL AT HIGH REYNOLDS NUMBER) TRANSONIC CONDITION

ROBERT W, WILLIAM Be IGOEj AdD PIERCE Le LAWING

HESSj DAVID A * SEIDEL,

(NrSA-TU-89060) 616ELIGHIS CE UNSTEADY B87- 1 76 6 7 t w s s u m ~ T E S T S cn ~ r / i ~ P ~ R C E M SUPEBCRXTXCAL

C O Y D I T l C Y ( U S A ) 19 p C S C L O l A Uaclas 4.iIBFOlL AT U G H B E Y B G L D S t i U H B E U , TRANSONIC

(iJ/OZ Y3357

J AI4 U A RY 198 7

National Aeronautics and Space Administration

Langley Research Center Hampton. Virginia 23665

https://ntrs.nasa.gov/search.jsp?R=19870008234 2020-02-03T16:47:42+00:00Z

HIGHLIGHTS OF UNSTEADY PRESSURE TESTS ON A 14 PERCENT SUPERCRITICAL AIRFOIL AT HIGH REYNOLDS NUMBER, TRANSONIC CONDITION

Abstract

Robert W. Hess*, David A. Seidel**, W i l l i am B. Igoe***, and Pierce L. Lawing****

NASA Langley Research Center Hampton, V i r g i n i a 23665-5225

Steady and unsteady pressures were measured on a 2-0 s u p e r c r i t i c a l a i r f o i l i n t h e Langley Research Center 0 .34 Transonic Cry0 en ic Tunnel a t Reynolds numbers from 6 x 10% t o 35 x lo6. The a i r f o i l was o s c i l l a t e d i n p i t c h at amplitudes from k.25 degrees t o kl.0 degrees a t frequencies from 5 Hz t o 60 Hz. The special requirements o f t e s t i n g an unsteady pressure model i n a pressurized cryogenic tunnel are discussed. Selected steady measured data are presented and are compared w i t h GRUMFOIL c a l c u l a t i o n s a t Reynolds number o f 6 x 106 and 30 x lo6. Experimental unsteady r e s u l t s a t Reynolds numbers o f 6 x 106 and 30 x 106 are examined f o r Reynolds number ef fects . Measured unsteady r e s u l t s a t two mean angles o f a t tack a t a Reynolds number o f 30 x 106 are also examined.

Nomenclature

chord 1 i f t c o e f f i c i e n t pressure c o e f f i c i e n t modulus o f o s c i l l a t i n g pressure c o e f f i c i e n t frequency, Hz reduced frequency, based on semichord, ncf /v Mach number Pressures i n f l a w r e s t r i c t o r c a l i b r a t i o n , f i g u r e 5 Reynolds number based on chord ve loc i t y , f t / sec streamwise coordinate measured from leading edge, in. peak o s c i l l a t i o n ampli tude i n p i t c h , degrees, p o s i t i v e leading edge up, deg. steady o r mean dynamic amplitude i n p i t c h , p o s i t i v e leading edge up, deg. micron, 1 x 10-6 meters phase angle between o s c i l l a t i n g pressure and o s c i l l a t i n g wing p i t c h angle, deg.

*Senior Research Engineer, Unsteady Aerodynamics Branch, Loads and Aeroe las t i c i t y D iv i s ion , Member AIAA.

**Research Engineer, Unsteady Aerodynamics Branch, Loads and A e r o e l a s t i c i t y Div is ion, Member A IAA.

***Senior Research Engineer, Transonic Aerodynamics D iv i s ion , National Transonic F a c i l i t y Operations Branch, Member A IAA. ****Senior Research Engineer, Transonic Aerodynamics Div is ion, Experimental Techniques Branch, Senior Member A I A A .

Su b s c r i p t s

C corrected value t t e s t measurement

I n t roduct i on

The advent o f l a rge cryogenic wind tunnels al lows unsteady pressure measurements t o be made on models a t Reynolds numbers t y p i c a l l y experienced by h i gh performance a i r c r a f t , thus e l i m i n a t i n g t h e need f o r a r t i f i c e s such as boundary-layer t r i p s t o simulate boundary l a y e r t r a n s i t i o n a t h igh Reynolds number. New '

m in ia tu re transducers s p e c i f i c a l l y designed t o measure unsteady pressures i n a cryogenic environment make these measurements possible. The study repor ted i n t h i s paper was conducted i n the 0 . 3 4 Transonic Cryogenic Tunnel 10.34 TCT) a t the NASA Langley Research Center. With i t s combined pressure, cryogenic temperature, and t ransonic speed c a p a b i l i t i e s the 0 .34 TCT can prov ide f l i g h t equiva lent a i r f o i l r e s u l t s f o r current a i r c r a f t . This tunnel was used i n the Advanced Technology A i r f o i 1 Test (ATAT) program i n extensive steady f low a i r f o i l studies t h a t demonstrated the necessity f o r h igh Reynolds number test ing. The a i r f o i l used i n the present unsteady t e s t s i s a fourteen-percent t h i ck , s u p e r c r i t i c a l a i r f o i l , designated Sc(2)- 0714, which was developed a t t h e NASA Langey Research Center.3 The purpose o f t h i s t e s t was t o ob ta in unsteady t ransonic pressure measurements from an o s c i l l a t i n g s u p e r c r i t i c a l a i r f o i l over a wide range of Reynolds number t o supplement the previous steady f l ow resul ts . A secondary o b j e c t i v e o f t he t e s t was t h e development o f inst rumentat ion techniques f o r measuring unsteady pressures a t cryogenic temperatures.

The two-dimensional model had a s i x inch chord and an e i g h t inch span. The t e s t was concentrated a t a tunnel freestream Mach number o f 0.72, which previous t e s t s i nd i ca ted t o be t h e design Mach number. Reynolds number [based on a s i x inch chord) was var ied from 6 x 106 t o 35 x 106 and Mach number was var ied a t two Reynolds numbers. The range o f t e s t frequencies was from 5 Hz t o 60 Hz a t o s c i l l a t i n g p i t c h amplitudes which va r ied from k0.25 degrees t o 21.0 degrees. I n t h i s paper, se lected steady measured data are presented and are compared

o f 6 x 106 and 30 x lo6. Experimental unsteady r e s u l t s a t Reynolds numbers o f 6 x 106 and 30 x 106 are examined f o r Reynolds number e f fec ts . Measured unsteady r e s u l t s a t two mean angles o f a t tack a t a Reynolds number o f 30 x 106 are a l so discussed.

.*4+h GRUMFO:: ca lcu la t i ons a t Reynolds Embers

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Apparatus

Model

The Sc(2)-0714 a i r fo i l model is shown i n ' F i g . 1. I t was machined from an alloy (Vascomax-200) t h a t has superior dimensional s t ab i l i t y properties a t cryogenic conditions. A cavity machined i n the underside of the w i n g , Fig. 2, provided the space necessary t o mount the transducers. This cavity was closed by a cover plate on which some lower surface transducers were mounted. The wing was supported on one end by a c lose-f i t t ing tang fixed to a d r i v i n g plate w i t h machine screws; this end, seen on the l e f t i n F i g . 2 , was sealed with epoxy. The other end was supported by an integral shaf t which rotated i n a bushing i n the tunnel side wal l plate. A s l iding seal of f e l t was used to seal the gap between the end of the osc i l la t ing airfoi l and the fixed tunnel sidewall plate. The position of the supports was designed to locate the pitch a x i s a t t h i rty-fi ve percent chord.

Transducers

Forty-three unsteady pressure transducers were mounted internally in the model. Because of space constraints, forty of the transducers were mounted i n receptacles connected by a short length (nominally 0.75 inch) of tubing to the or i f ice . The remaining three transducers were mounted w i t h the transducer head less t h a n 0.1 inch below the surface of the wing. The or i f ices of these three transducers were paired with tube mounted transducers for comparison purposes. A series of t e s t s was conducted to examine the effects of or i f ice diameter, tube diameter, and tube length on the dynamic response of the system. A t atmospheric conditions there was no significantreduction of dynamic amplitude response or phase s h i f t of the t e s t configuration up t o 100 Hz.

The location of the transducers i s given schematically i n Figures 3 (a ) and 3(b) . The tube-mounted transducer or i f ices are located al ternately i n two rows 0.25 inches on e i ther side of the center line. On the top surface the o r i f i ce distribution of the twenty-five transducers resul ts in an or i f ice every 2% of chord to x/c = 0.1 and 4% chord to x/c of 0.70. The dis t r ibut ion of the 15 tube-mounted transducer or i f ices on the lower surface is 2% to an x/c of 0.1 and increases to 5% thereaf ter . The close-mounted transducers and reference or i f ices are located 0.5 inches from the center l ine.

The elements o f the transducer system are shown in Fig. 4. Since the different ia l pressure between the wing surface and the tunnel s t a t i c pressure could exceed the rated capabi 1 i ty of the transducer, the transducer was referenced t o a manifold which i n t u r n was vented to one of five reference or i f ices . A reference transducer measured the pressure different ia l between the manifold and the tunnel s t a t i c pressure.

The f inal configuration consisted of transducers w i t h a 10 psi range and with outputs of between 5 and 9 mv/psi. Each transducer was mounted i n a receptacle which i n t u r n was connected to the 0.015 inch diameter o r i f i ce by a 0.75 inch length of .030 inch i . d . tubing.

The connection between the manifold and the reference o r i f i ce was interrupted by a porous flow re s t r i c t e r which damped out the osc i l la t ing pressure from the s t a t i c reference o r i f i ce (replacing the long lengths of t u b i n g usually used for th i s purpose). The resul ts of a se r ies of cal ibrat ions made on different combinations of porous flow res t r ic tors and tube lengths are g iven i n Fig. 5. The flow res t r ic tors tes ted were c o n e r c i a1 ly avai 1 ab1 e sintered f i 1 t e r s composed of constant diameter par t ic les , w i t h diameters ranging from 1 0 ~ to 2511. Except for the resul ts for two 25p res t r ic tors i n se r ies (shown i n the curve labeled 1=.5, 2-25p) the data shown i n F i g . 5 are for s ingle flow res t r ic tors . The reduction i n unsteady pressures i s shown (Fig. 5) as the ra t io of the imposed osc i l la t ing input pressure amplitude, P , , to the output pressure amplitude, P,, as a function of frequency and for different combinations of f i l t e r s and tube length, 1. Also given i n the same figure is the time required, t, for the system t o reach equilibrium a f t e r the application of a s t a t i c pulse.

Two 25p f i l t e r s i n ser ies were selected for the wind tunnel t e s t . Because of lack of space i n the model the manifolds and flow re s t r i c t e r s were located outside the model during the t e s t and were connected to the model with approximately 6 inches of t u b i n g .

The transducers were t o be recovered a f t e r the t e s t and consequently could not be permanently bonded to the receptacle. A se r ies of t e s t s were conducted w i t h candidate mastics and dumy transducers a t 120 deg. K and 20 psi pressure different ia l t o determine which, i f any, would maintain a seal a f t e r repeated cycling. A mastic supplied by the transducer vendor was selected for the instal la t ion.

V r g e variations in temperature (120 deg. K to 320 deg. K) and stagnation pressure (1.4 atm. - 6 atm.) over the operating range of the C.3M-TCT resu l t in sidewall deformations tha t required special considerations i n the design of the osc i l la t ing drive system. Figure 6 shows the model and the drive system ins ta l la t ion in the t e s t section. The test section i s shown to the r i g h t with the test section cei l ing removed and can be ident i f ied by the two s lo t s on the f loor which run under the model. The model i s between the t e s t section sidewalls which i n t u r n are between the t u n n e l plenum spaces and f ina l ly , the tunnel pressure- shell or plenum walls. The c r i t i ca l elements of the system are ident i f ied in the schematic drawing i n F i g . 7. The hydraulic-rotary actuator required the maintenance of precise alignment for the duration of the t e s t . Since the 0.3-m TCT t e s t section f loa t s on a cable suspension system to accommodate thermal contraction a t the cold operating conditions,

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t h e a c t u a t o r and suppor t ing s t r u c t u r e were also supported by a system o f cables, b locks, and coun te r weights so t h a t they could m v e w i t h the t e s t sect ion.

An i n s u l a t i n g spacer between t h e ac tua to r and t h e d r i v e s h a f t and two ho t a i r b lowers were used t o reduce t h e c h i l l on t h e ac tua to r and system components ex te rna l t o t h e tunnel . A T e f l o n (reg. t r a d e mark) bushing and pressure seal were t h e remaining f i x e d support p o i n t s f o r t h e ho l l ow aluminum d r i v e shaf t . The s h a f t was a t tached t o t h e r o t a t i n g s idewa l l d r i v e disk through a bel lows t h a t i s o l a t e d t h e s h a f t from t h e r e l a t i v e i n - l i n e movement of t h e tunnel s idewa l l . The r o t a t i n g d r i v e d i sk was Tef lon coated on i t s c i r c u m f e r e n t i a l bear ing surfaces and had a rec tangu la r s l o t t o accommodate the wing tang. The tang was ho l l ow t o prov ide a pa th f o r t ransducer cables and t u b i n g which went through a matching ho le i n the p l a t e and e x i t e d through t h e cab le p o r t s i n t h e d r i v e shaft . Th i s end o f t h e wing was sealed wi th epoxy and b o l t e d t o the r o t a t i n g d isk .

The o the r end o f t h e wing was supported by t h e i n t e g r a l ho l l ow wing s h a f t and a bushing i n t h e s idewa l l p la te . Th is end o f t h e wing moved r e l a t i v e t o i t s mounting p l a t e and was sealed w i th f e l t t h a t matched t h e p r o f i l e o f the wing. The wing sha f t was at tached t o a r o t a r y t ransducer by an i n s u l a t i n g sha f t . The hol low wing s h a f t prov ided a pa th f o r t h e remaining i ns t rumen ta t i on cables. The r o t a r y t ransducer was heated w i t h sur face heaters under thermostat c o n t r o l and t h e e n t i r e assembly covered w i th an i n s u l a t e d can.

The system d i d no t develop any problems d i i r i ng t h e tes t . The angie o f a t t a c k was checked v i s u a l l y against a s idewa l l s c r i b e mark a t t h e beginning o f each days t e s t be fo re the i n t r o d u c t i o n o f cryogenic n i t r o g e n caused ex tens i ve f r o s t t o be formed on t h e view por ts . The c o r r e l a t i o n o f t h e geometric f l ow angle o f a t tack between t h e s c r i b e mark a t t h e t r a i l i n g edge and t h e i ns t rumen ta t i on d i d not vary dur ing t h e t e s t .

Data A c q u i s i t i o n and Reduct ion

S t a t i c da ta f rom t h e model and tunnel i ns t rumen ta t i on were acqui red us ing t h e 0.3M-TCT data a c q u i s i t i o n system. The model angle of a t tack and pressure data were f e d t o the system's analog data a c q u i s i t i o n channels. The system has 192 channels which a re f i l t e r e d w i t h a 10-Hz low-pass f i l t e r and then d i g i t i z e d a t 20 samples pe r second. S t a t i c data values are acqu i red by averaging t h e d i g i t i z e d values over a one second i n t e r v a l .

Dynamic model data were acqui red using analog tape recorders. The ins t rumen ta t i on s i g n a l was a m p l i f i e d t o be a va lue o f about one w!t RYS. The mcde! :ngle o f att:ck and pressures were taken d i r e c t l y from the a m p l i f i e r s and recorded on two 28-channel analog tapes opera t i ng a t 15 inches pe r second. To o b t a i n ampl i tude and phase in fo rma t ion a t the f requency o f o s c i l l a t i o n and t h e lowest harmonics, t h e data was d i g i t i z e d a t 32 samples p e r c y c l e o f o s c i l l a t o r y mot ion f o r 64

cont iguous cycles. A Fast Four ie r Transformat ion average (FFT) was taken o f t h e data t o c a l c u l a t e t h e harmonic components o f t h e unsteady pressures. The data sample r a t e and number o f cyc les analyzed was se lec ted t o g i v e an accurate est imate o f t h e f i r s t t h r e e fundamental harmonic componets. The harmonic pressure c o e f f i c i e n t s a r e normal ized by t h e ampl i tude of t h e harmonic wing mot ion i n degrees. A l l phase angles were r e l a t i v e t o t h e wing pos i t i on .

S idewal l boundary l a y e r and angle of a t t a c k c o r r e c t i o n s were app l i ed t o t h e measured steady pressure resu l t s . The s idewa l l boundary l a y e r c o r r e c t i o n s a re based on t h e theory of Ref. 4 which i s used i n Ref. 5 with measured values o f s idewa l l displacement and momentum th i ckness t o compi le t h e t a b l e s which were used t o c o r r e c t t h e exper imental values i n t h i s paper. The angle of a t tack co r rec t i ons descr ibed i n Ref. 6 (sometimes r e f e r r e d t o as the "Barnwell -Davis- Moore" c o r r e c t i o n ) ad jus t t h e theory o f Davis- Moore w i t h exper imental data. The w a l l induced downwash immediately over t h e model f o r t h e 0.34 TCT i s :

-c. C

The parameters necessary t o make t h e c o r r e c t i o n are:

c = chord = 6 i n . h = tunne l semi-height = 12 in. a = s l o t spacing = 4 in . 6 = w i d t h o f s l o t = 0.2 i n .

j = aK/h K = 3.2 (semi emp i r i ca l constant,

f u n c t i o n of 6 and a )

For C1 = 1.0

A: = -1.73245 deg.

Resul ts and D i s c u r n

The t e s t was designed t o exp lo re t h e e f f e c t s o f Reynolds number on unsteady pressures and t o generate a data base f o r v a l i d a t i n g unsteady-aerodynamic computer codes. The t e s t cond i t i ons as def ined by Mach number and Reynolds number a r e shown i n Fig. 8. Test p o i n t s were taken a t t he design Mach number o f 0.72 a t t e s t Reynolds numbers va ry ing from 6 x 106 t o 30 x lo6. Mach number was va r ied a t two Reynolds numbers, 15 x lo6 and 30 x 106. A t o t a l o f 976 t e s t p o i n t s were taken. The pr imary data base was taken f o r p i t c h - o s c i l l a t i o n frequency between 5 Hz and 40 Hz a t ampl i tude of t0.25 degree as i n d i c a t e d by t h e open and c losed symbols. Once t h i s data was i n hand, t h e p i t c h ampl i tude was increased t o 20.5 and kl.0 degree and t h e p i t c h f requency was a l s o increased t o 60 Hz a t t e s t c o n d i t i o n s i n d i c a t e d by the so!id symbn!~.

Steady Pressures

Steady pressure d i s t r i b u t i o n s f o r f o u r

angles o f a t tack , at, approx imate ly 2.5", 2.0", 1.5", and O", and f o r two Reynolds

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numbers, 6 x 106 and 30 x lo6, are shown i n F i g . 9. The experimental data are shown as symbols and the c a l c u l a t i o n s as s o l i d l i n e s . The pressure data have been co r rec ted f o r s idewa l l e f f e c t s ( re f s . 4 and 51 and angle o f a t t a c k ( r e f . 6). Ca lcu la ted pressure d i s t r i b u t i o n s f rom the f u l l p o t e n t i a l GRUMFOIL computer code ( r e f . 7 ) are a l s o compared (F ig. 9). The GRUMFOIL code c o n s i s t s o f a f u l l p o t e n t i a l equat ion f low so l ve r i n t e g r a t e d w i t h a v iscous boundary layer model. GRUMFOIL may be

entered by spec i f y ing e i t h e r a o r C1 . The c o r r e c t e d values o f Mach number and C1 were used as the i n p u t data f o r the computed r e s u l t s which are compared w i t h the co r rec ted exper imental values o f Cp shown i n F i g . 9.

Below each f i g u r e a re l i s t e d M, a, and C1 values f o r t he tunnel t e s t cond i t i ons , t he co r rec ted values, and the values r e s u l t i n g f rom t h e GRUMFOIL ca l cu la t i ons .

The comparisons between experiment, shown as symbols, and ca l cu la t i ons , s o l i d l i n e s , i n F i g . 9 are very good. The shock moves a f t by approximately 8% t o 10% o f chord f o r a g iven

va lue o f i t when Reynold2 number i s increased f rom 6 x 106 t o 30 x 10 . The GRUMFOIL code under -p red ic t s the p o s i t i o n o f the shock a t both Reynolds numbers by approximately 2-3% o f chord even though C1 i s matched.

L i f t c o e f f i c i e n t s f o r several cases are shown i n F i g . 10 p l o t t e d aga ins t co r rec ted angle o f a t tack and against angle o f a t tack as computed by GRUMFOIL f o r i n p u t values o f Ma h

and 30 x lo6. The angles c a l c u l a t e d from GRUMFOIL a re c o n s i s t e n t l y l a r g e r than those determined from the c o r r e c t i o n procedure o f r e f . 6. T h i s t rend i s s i m i l a r t o the one shown i n r e f . 8. I r r e s p e c t i v e o f the angle o f a t tack c o r r e c t i o n s , an increase i n C 1 o f approx imate ly 0.1 i s shown as Reynolds number i s increased from 6 x lo6 t o 30 x lo6. Th is i nc rease r e s u l t s f r o m the rearward movement o f t he shock shown i n F ig . 9.

number and C1 f o r Reynolds number o f 6 x 10 E

Unsteady Pressures

The e f f e c t s o f Reynolds number and frequency o f o s c i l l a t i o n upon the unsteady pressure d i s t r i b u t i o n i s shown i n F i g . 11. Resul ts are g iven i n terms o f the modulus o f the unsteady pressure c o e f f i c i e n t normal ized by t h e o s c i l l a t i n g p i t c h angle, a, and the phase angle, Q, between the unsteady pressure and t h e o s c i l l a t i n g wing pos i t i on . Resul ts are shown

f o r a t = 1" and 2" and f o r R = 6 x 106 and 30 x 106. Two o s c i l l a t i o n f requencies, 5 Hz and 20 Hz, are presented f o r a p i t c h ampl i tude o f k0.25 degrees. The upper sur face pressure d i s t r i b u t i o n s a re shown i n F ig . l l ( a ) and the corresponding lower sur face pressures a re i n F i g . l l ( b ) .

The shock wave, i d e n t i f i e d by the peak i n t h e unsteady pressures, moves a f t about 8% t o 10% chord as R i s increased from 6 x l o 6 t o 30 x lo6 a t t h e same tunnel t e s t angles. The unsteady pressures, a t both Reynolds numbers, a re s i g n i f i c a n t l y greater ahead o f the shock a t

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is = 1 degree than a t 2 degrees. There i s no s i g n i f i c a n t d i f f e rence i n the magnitude o f t he unsteady pressures due t o a change i n frequency from 5 Hz t o 20 Hz.

F o r the t e s t c o n d i t i o n s shown, t h e pressures ahead o f t he shock are approximately 180" o u t o f phase w i t h the wing o s c i l l a t i o n . Imned ia te l y a f t e r the shock wave the phase angle a t t he pressure ab rup t l y changes from -180" t o approx imate ly 0" t o be in-phase w i t h the wing p i t c h i n g mot ions. A f t o f t he shock wave the

phase angle remains a t 0" a t a t = 2", b u t i s more dependent on frequency a t t he lower mean

p i t c h angle, a t = lo, tending t o go back t o -180" a t 20 Hz and t o 0" a t 5 Hz.

The lower sur face pressures and phase angles are shown i n F ig . l l ( b ) . The pressures are low and decrease from lead ing t o t r a i l i n g edge; the phase angle i s c lose t o zero except a t the reg ion o f t he lower sur face i n f l e c t i o n .

The e f f e c t o f va ry ing the ampl i tude o f t he p i t c h o s c i l l a t i o n a t M = 0.72 i s shown i n F i g . 12. Pressure d i s t r i b u t i o n s are shown f o r R =

30 x l o 6 f o r two mean angles, ht = 1" and 2" and f o r b o f requencies, f = 40 and 60 Hz. Data f o r p i t c h ampl i tudes o f 0.25, 0.5, and 1.0 degree are shown. I n m s t cases, i n t h i s and the f o l l o w i n g f i g u r e , the data p o i n t s are n o t connected i n the neighborhood o f the peak shock ampl i tude because the peak pressure i s n o t de f i ned by a f i n i t e number o f pressure o r i f i c e s . The upper sur face pressure d i s t r i b u t i o n , F ig . 12(a) , shows a reduc t i on and broadening o f t he shock-generated peak amp1 i tude as the p i t c h ampl i tude, a, i s increased from 0.25" t o 1" a t bo th f requencies and mean angles. Note the s u b s t a n t i a l change i n mean shock p o s i t i o n due t o p i t c h ampl i tude a t f = 60 Hz and U t = 1".

A secondary peak i n the magnitude o f the o s c i l l a t i n g pressure i s ev iden t immediately

behind the shock a t a t = 2" and 1" which cou ld be a t t r i b u t e d t o f l ow separat ion and reattachment as discussed i n Ref. 9. However, an i n v i s c i d c a l c u l a t i o n us ing the X T R A N Z L ~ O S ~ ~ computer code p r e d i c t s t h i s secondary response, a l b e i t n o t p r e c i s e l y a t the same chord l o c a t i o n . Ca lcu la t i ons w i t h GRUMFOIL shown i n F i g . 13 i l l u s t r a t e s t h a t a more probable reason f o r the secondary response der ives f rom the

supersonic reg ions above the a i r f o i l . A t at = 1" t h e r e i s a secondary supersonic reg ion behind the shock which i s engu l fed by the pr imary supersonic reg ion when the angle o f a t tack i s

increased t o i t = 2". Tijdeman12 and o the rs have noted t h a t t he f l o w i n the supersonic reg ion p r i o r t o the format ion o f a shock i s cha rac te r i zed by a s u b s t a n t i a l increase i n unsteady pressure.

The upper su r face phase angle, Q, shows changes i n the neighborhood o f the shock and a f t p o r t i o n o f the a i r f o i l w i t h p i t c h ampl i tude, F ig . 1 2 ( a l .

The lower sur face pressure ampl i tude and qhase, $, are given i n F igu re 12(b) . Both are

-

4

r e l a t i v e l y independent o f p i t c h amp1 i tude except i n the neighborhood o f the i n f l e c t i o n o r cusp reg ion o f the a i r f o i l . Both pressure and phase decrease from leading edge t o t r a i 1 i ng edge.

The e f f e c t o f vary ing the o s c i l l a t i o n frequency a t M = 0.72 i s shown i n F ig . 14. Pressure d i s t r i b u t i o n s are shown f o r R = 30 x l o6 f o r mean angles i t = 1' and 2" and f o r p i t c h amplitudes o f 0.25" and 0.5'. Data f o r frequencies o f 5, 15, 40, and 60 Hz are presented. I n general the excursion o f the shock on the upper surface, Fig. 14(a), i s reduced a t 60 Hz and again the second peak i s g rea te r a t a t = 1' than a t at = 2" and i s n o t a s t rong func t i on o f frequency. As expected t h e phase angle i s a f u n c t i o n o f frequency showing s i m i l a r c h a r a c t e r i s t i c s as shown i n the prev ious f i g u r e decreasing t o approximately 0' behind the shock. The lower surface pressures and phase angle, F ig . 14(b) again decrease from a maximum a t the leading edge t o a minimum a t t he t r a i l i n g edge and showing some dependence on frequency.

- -

Boundary Layer S ta te

The unsteady pressure transducers used i n t h i s t e s t a l so enabled measurements t o be nhtai ned which are o f i n t e r e s t regard ing the s t a t e o f the boundary l aye r . The t ime h i s t o r i e s o f the pressures a t f i v e transducer locat ions taken when the a i r f o i l was locked a t f i x e d angle of a t tack are shown i n F ig . 15. The data shown i n t h i s f i g u r e are a l l a t a gain o f 10 b u t the transducer s e n s i t i v i t y , given w i t h each trace, has no t been app l i ed t o p u t the t ime h i s t o r i e s i n engineering un i t s . The t i m e h i s t o r i e s taken a t two f i x e d angles o f a t tack ((It = 0" and 2') a t R = 35 x 106 are shown i n F ig . 15(a) and F ig . 15Eb) respect ive ly . The steady pressure d i s t r i b u t i o n s are shown a t the r i g h t o f each f igure. The s o l i d p o i n t s on t he pressure d i s t r i b u t i o n mark the l o c a t i o n (x/c .14, .28, .46, .62, .75) o f the f i v e transducers. I n Fig. 15(a) (it = 0') t he t ime h i s t o r i e s have t h e c h a r a c t e r i s t i c s o f a t u r b u l e n t boundary l aye r . However i n Fig. 15(b) ( i t = 2') the pressure i s quiescent a t x/c o f 0.14 and 0.28 i n comparison w i t h the transducer responses a t a t = 0'. A t an x/c o f 0.46 the e f f e c t o f shock movement i s observed. A t an x/c o f 0.62 the shock movement i s s t i l l observed and turbulence i s apparent. A t x/c o f 0.75 the s igna l i s comparable t o t h a t a t i t = 0'. The most obvious d i f f e rence between the condi t ions a t the two angles o f a t tack i s the presence o f a shock and the s l i g h t l y more favorable pressure g rad ien t a t i t = 2'. The t ime h i s t o r i e s i n d i c a t e t h a t laminar flow was present a t a t = 2" and t h a t t r a n s i t i o n t o turbulence was between an x/c o f 0.28 and 0.46 corresponding to t r a n s i t i o n Reynolds numbers between 9.8 x l o6 and 16.1 x lo6. The p o s s i b i l i t y e x i s t s t h a t l o n g runs of laminar f l ow ex i s ted i n t e r m i t t e n t l y dur ing the tes ts .

-

-

Conclusions

Steady and unsteady pressures on a 14 percent supercr i t i c a l a i r f o i l a t t ransonic Mach numbers have been measured a t Reynolds numbers from 6 x l o 6 t o 35 x lo6. Inst rumentat ion techniques were developed t o measure unsteady pressures i n a cryogenic tunnel a t f l i g h t Reynolds numbers. Experimental steady data, corrected f o r w a l l e f fec ts , show very good agreement w i t h ca l cu la t i ons from a f u l l p o t e n t i a l computer code w i t h an i n t e r a c t e d boundary l aye r . The steady and unsteady pressures both show a shock p o s i t i o n t h a t i s dependent on Reynolds number. Fo r a super- c r i t i c a l pressure d i s t r i b u t i o n a t a chord Reynolds number o f 35 x lo6, laminar boundary l a y e r f l ow was observed over a s i g n i f i c a n t percentage o f the a i r f o i l chord.

Acknowledgement

The authors wish t o acknowledge the assistance o f Clyde Gumbert o f the Theore t i ca l Aerodynamics Branch, NASA Langley Research Center, i n the GRUMFOIL ca l cu la t i ons .

References

'Ray, E. J.; Ladsm, C. L.; Adcock, d . 8.; Lawing, P. L.; and H a l l , R. M.: "Review o f Design and Operational Charac te r i s t i cs o f t he 0.3M Transonic Cryogenic Tunnel .I1 NASA TM 80123, 1979.

"A Review o f Reynolds Number Studies Conducted i,n the Langley 0.3M-Transonic Cryogenic Tunnel. AIAA/ASME 3rd J o i n t Thermophysics, F lu ids , Plasma and Heat Transfer Conference, S t . Louis, MO, June 7-11, 1982. A I A A Paper 82-0941.

'Harris, C. 0.: "Aerodynamic Charac te r i s t i cs o f a 14-Percent Thick NASA Supercri ti ca l A i rf o i 1 Desi gned f o r a Normal-Force C o e f f i c i e n t o f 0.7." NASA TM

'Sewall, W. G.: "The E f f e c t s o f Sidewall Boundary Layer i n Two-Dimensional Subsonic and Transonic Wind Tunnels." A I A A Journal, Vol. 20, No. 9, September 1982, pp. 1253-1256.

'Jenkins, R. V.; and Adock, J. 6.: "Tables f o r Cor rec t i ng A i r f o i l Data Obtained i n the Langley 0.3-Meter Transonic Cryogen?c Tunnel f o r Sidewall Boundary Layer E f fec ts . NASA TM 87723 June 1986.

'Barnwe1 1, R. W. : "Design and Performance Evaluat ion o f S l o t t e d Wall s f o r Two-Dimensional 'vi4 nd Tunnel 5."

'Mead, H. R.; and Melnik, R. E.: "GRUMFOIL: A Computer Code f o r the Viscous Transnic Flow Over A i r f o i l s . " NASA CR 3806, October 1985.

'Gumbert, C. R.; and Newman, P. A.: "Va l i da t i on o f a Wall I n t e r f e r e n c e AssessmentKorrection Procedure f o r A i r f o i 1 Tests i n the 0.3M-Transonic Cryogenic Tunnel ." A I M 2nd Appl ied Aerodynamics Conference, August

'Mundell, A. R. G.; and Mabey, D. G.: "Pressure F1 u c t u a t i ons Caused by Transonic Shock/Boundary-Layer In te rac t i on . " Aeronautical Journal, August/September 1986.

LRay, E. J.:

X-72712, J u l y 1975.

KASA TM 78648, February i978.

21-23 1984.

5

"Whitlow, W., Jr.: "XTRAN2L: A Program f o r Solv ing the General Frequency Unsteady Transonic Small Disturbance Equation," NASA TM 85721 November 1983.

"User's Manual f o r XTRAN2L (Version 1.2): A Program f o r Solv ing the General -Frequency Unsteady Transonic Small -Disturbance Equation," NASA TM 87737, Ju l y 1986.

"Ti jdeman, H.: " I n vest i gat 1 on o f t h e Transonic Flow Around O s c i l l a t i n g A i r f o i l s , " NLR TR 77090 U, 1977.

'Seidel, D. A.; and Batina, J. T.:

Fig. 1 External view of model.

Fig. 2 In te rna l conf igurat ion o f model.

Fig. 4 Elements o f the transducer system.

6

0- 0 -

.1 -

.2

.3

.4 - Inches 3 - .5 -

.8 -

.7

.8 -

.e

1 - - -

2 -

X I C

4 - -

5 - -

8- 1.0-

Upper surface instrumentation

0 Orifice 8 Close mtd. tran. e .Ref. orifice

Z - 0

A-02.

A - 0 8 0

A - 0 8 0

A-100

A-120

Driving ~ - 1 4 .

A-180

A-18.

A-200

C-428

C-48e

end

C-A;321 c-47

A-24.

A-28.

08-07

0 8-00

08-11

08-13

08-15

.8-17

08-19

08-21

0 8-23

08-25

1.25

(a) upper sur face

Fig. 3a Transducer locat ions.

Lower Surface instrumentation

0 Orifice Close mtd. tran.

e Ref. orifice

0 - 0 -

.1

.2 - .3

.4

- 1 -

- 2 -

- xlc

Inches 3 - 5 -

.8 -

.7

.8 -

.o

4 - -

5 - -

6- 1.0-

G-01 E-280

E-32 0

Drivina end

E-38.

E-300

E-410

1 .25 4-

(b) Lower sur face

Fig. 3b concluded.

1 ,F-44

06-27 0d-29 0d-30

00-31 eF-48

0 D-33 0 D-34

D-35 0-36

00-37

8 F-45

eF-49 0 D-40

& .25

ApprOX. delay time

0.5 15

7 Flow restrlctors t ,-Transducer

P2lP1

- 1 0 -

I . I . I . I . 1 0 i o 20 3 0 40 50

Frequency, Hz

F i 9. 5 Dynami c f 1 ow-rest r i c t o r -tube c a l i b r a t i on resul ts .

Fig. 6 Model i n s t a l l a t i o n

Insulated rotary transducer housing \

Heated rotary transducer Bushing ~.

L

Teflon coated bearing surfaces

Plenum wall Cable sort

Pressure seal

Insulating spacers

coupling H Fig. 7 Schematic drawing o f model i n s t a l l a t i o n .

0 Frequency and amplitude 0 Frequency

e e e o e

Mach 1 0 number .7 e

0

I I I 1 0 20 30 4 0 106

.6

0

Fig. 8 Mach number and Reynolds number t e s t condi t i ons .

0 Upper surface 0 Lower surface

0 Upper surface -2.0 0 Lower surface

R-6 .035 X lo6 R = 30.0 X 106

cP

0

1.2 0 .2 .4 .6 .8 1.0

x ic x i c

Corrected Grumf o i 1 Tunnel M 0.720 0.701 n. 701 - a 2.504 0.844 C1 0.9581 0.9753

1.385 0.9837

Tunnel Corrected Grurnf o i 1 M 0.719 0.705 0.705

2.51 0.756 1.399 C1 1.0123 1.0256 1.0336

0 Upper surface 0 Lower surface

R =30.0 X lo8

0 Upper surface 0 Lower surface

R =6.01 X lo6

0 .2 .4 .6 .8 1.0 x i c x i c

Corrected Grumfoi 1 Tunnel 0.72 n. 71-11 n. 701 -

a 2.002 0.525 c1 0.8523 0.8676

n. 961 0.8757

Grurnf o i 1 M 0.721 n. 705 0.705

c1 n.926 0.939 0.9453

Tunnel Corrected

2; 1.997 0.393 0.910

F i g . 9 Comparisons of steady t e s t resul ts with calculated resul ts a t a tunnel Mach number o f 0.72.

8

0 Upper surface 0 Lower surface

R = 6.035 X 10'

1 . 2 r 0 .2 .4 .6 .8 1.0

X I C

Tunnel Corrected Grumf o i 1 M 0.719 0.701 0.701 h 1.495 n.2ni 0.493 C! 0.7467 0.7601 n. 7680

0 Upper surface 0 Lower surface

R -29.98 X lo6

CP

0

1 . 2 c 0 .2 .4 .6 .8 1.0

X I C

Tunnel Corrected Grumfoi 1 M 0.718 0.705 n. 705 a 1.501 n.051 0.552 C1 .8373 0.849 0.8555

-2.0 t o Upper surface 0 Lower surface

R = 6.02 X lo6

0 Upper surface 0 Lower surface

R = 3 0 . 0 5 x 10'

CP CP

0 0

0 1.2

0 1.2 X I C X I C

Tunnel Corrected Grumfoi 1 M 0.720 0.701 n. 7ni

C, 0.5288 n. 5383 n. 5484 h 0.004 -0.92 -0.398

Tunnel Corrected Grumf o i 1 M .721 0.705 0.705

-0.005 -1 036 -0.715 c1 n.5951 n. 6034 n. 6099

Fig. 9 concluded.

9

M = 0.72, R = 6 X l o 6 1 * 1 r 1.0

.9

.8

.7

.6

.5

-

-

-

-

-

-

Method

,-O Barnwell --- 0 Grumfoil

.5 c .4 I 1 I I I I

- 2 - 1 0 1 2 3

a, deg

Method

---O Barnwell - - -0 Grumfoil

Fig . 10 Comparison o f l i f t c o e f f i c i e n t versus c o r r e c t e d a n g l e - o f - a t t a c k .

f, HZ k 1 .o 0 5.0 . O l l 6 .

-+ 20.0 .0466

.8 Close mtd. trans.

.6 R - 6.64 X 106

ICp I /deg .4

.2

0

2oo r 120

40

-40

-120

- 2 0 0

@, deg

-280 t -360 I I I I I

0 .2 .4 .6 .8 1.0 xlc zt = 2.06', a = 0.59'. C

f. Hz k

0 5.0 .Or16 0 20.0 .0465

Close mtd. trans. ICplldeg .4

.2

0

40 r -40

- 1 2 0 @, deg

-200

-280

-360 l - 2 2 - 2 0 .2 .4 .6 .8 1.0

xlc Ot 1.036', ic = -0.561'

( a ) Upper surface

lCpl ldeg

1, Hz k 0 5.0 .0153 0 20.0 .0618

0 Close mtd. trans.

.8

.6

.4

.2

0

40

-40

- 1 2 0

- 2 0 0 4, deg

-280 t -360

0 .2 .4 .6 .8 1.0 x lc

it = 2.05', ic = 0.446'

.8 f, Hz k

0 5.0 .0154 .6 0 20.0 .061Q

Close mtd. trans. ICplldeg .4

R=30.03 X lo6

.2

0

40 r -40

-120

4, deg - 2 0 0

-360 -280 0 L .2 .4 .6 .8 1.0

x l c

at = 1.05', ic = -0.255'

Fig. 11 Unsteady pressure test results a t a tunnel Mach number o f 0.72 and a t a = 20.25'.

11

f, Hz k o 5.0 .Oil6

-0- 20.0 .0466

0 Close mtd. trans.

ICplldeg .2

0

- 1 0 0 1 o .2 .4 .6 .8 1.0 X I C

'Jt = 2.06', 'J = 0.59' C

f , Hz k 0 5.0 .0153 4 20.0 .0618

Close mtd. trans.

.2 R = 30.05 X 106

ICpI/deg

0

loor

0 .2 -4 .6 .8 1.0 X I C

it = 2.05', ic = 0.446'

f, Hz k f , Hz k

o 5.0 .0116 0 5.0 .0154 4 20.0 .0465 20.0 .0610

0 Close mtd. trans. Close mtd. trans. .4 -4

R = 30.0 X lo6 lCplldeg .2

0 0

ICplldeg .2 R = 6.04 X lo6

100

0, deg 0 0, deg loo+- 0

-100 -100 0 .2 -4 .6 .8 1.0

X I C

it 1.05", ic = -0.255' zt = 1.036', ic -0.561"

(b) Lower surface

Fig. 11 concluded.

12

-320 c -360 1 I I I J

0 .2 .4 .6 .a 1.0 x l c

f = M H z

lCpl ldeg

.0

.e

ICpl ldeg .4

.2

0

-320 -360 I I I I J

.4 .6 .a 1.0 0 .2 X I C

f = " t

I a. de0 at, deg I

.25 2.05

.5 2.08

O r

-320 1 --Re(! I I I

.4 .6 .a 1.0 0 .2 x I C

f = 6 0 H z

-380 .2 .4 .6 .a 1.0

X l C

f = 60 Hz

Fig. 12 Variation of Cpl/deg and 0 a t M = 0.72 and R = 30 x lo6 with pitch amplitu b e.

13

.2 r

loo I a. d e g 6,. d e g

.25 1.02 1.07 ----- .5

-- 1.0 1.03

-100 I I I I I I 0 .2 .4 .6 .8 1 .o

x I C

f = 4 0 H z

. 2 r

lo0c a, d a g at. d e g

.25 1.06

.5 1.05 1 .o 0.97 --

-1001 I I 1 I I 0 .2 . 4 . 6 .B 1 .o

x l c

f = 60 Hz

I C p l l d e g 0 ~

l o o r 0. d e g

-100 1 1 1 1 1 J 0 .2 .4 .6 .8 1 .o

x / c

f = 4 0 H z

I C p l l d e g 0 2 100

0, d e g 0

-100

f = 60 Hz

( b ) Lower surface

Fig. 12 concluded.

a. d e g 6,. d e g

. 25 2.05

.5 2.08 -- 1.0 2.05 -----

1 1 1 1 1 .2 .4 .6 .B 1 .o

X l C

Fig. 13 Sonic regions a t M = 0.72 calculated by GRUMFOIL code.

14

.6

!Cp!!deg .4

.2

I 1 1. Hz k 6 , d e g

0

-3201 , I . , , I -360

0 .2 .4 .6 .0 1.0 x l c

a = .25'

1. Hz k 8. deg

5 ,0154 1.06 .8 -

ICpl/deg .4 -

0

0

-80

-180

- 2 4 0

-320 1 I , I I I

-380 0 .2 .4 .8 .8 1.0

X l C

.8

.e

ICplldeg .4

.2

1, HZ k 81, deg

5 ,0154 2.15 15 ,0481 2.00 4 0 ,123 2.07

! --- 80 ,1845 2.07

----- --

0

-80

6 , de9 -180

-240

-320 -380

a = .25'

I I I I I .2 .4 .8 .8 1.0

X l C

1. Hz k 6,. deg - 5 ,0154 2.076 ----- 15 .048 2.081 4 0 .123 2.058 80 ,185 2.075

.8 - --

ICplldeg .4 - -- -

.2 -

0

0

-80

6 , deg -160

-240

-320 1 -360 I I I I I I

0 .2 .4 .8 .8 1.0 x i c

( a ) Upper surface

F ig . 14 Variat ion of C /deg and $ a t M = 0.72 and R = 30 x 106 with frequency I PI

15

1. Hz k at. deg - 5 .0154 1.05 ----- 15 .0461 1.04 -- 4 0 . I 2 3 1.02 IQlldeg .2 --- 60 .1645 1.06

loor

-1001 I I I I J 0 .2 .4 .6 .8 1 .o

x l c

a = . 2 5 O

10 I ldeg

1 . Hz k zit, deg - 5 .0154 1.06 ---- - 15 .0461 1.08 -- 4 0 .123 1.07 --- 60 .185 1.05

loor 0, deg a

/---- I

-100 0 .2 .4 .6 .8 1 .o

x lc

a . 5 O

.2

ICplldeg 0

100

0, deg 0

- 1 0 0

f, Hz k it, deg - 5 .0154 2.08 ----- 15 .0461 2.08 -- 4 0 . I 2 3 2.06 - -- 60 .1845 2.05

I I I I I .2 .4 .6 .8 1 .o

x l c

a = . 2 5 O

1. Hz k a,, deg

5 .0154 2.15 ----- 15 .0461 2.09 -- 4 0 .123 2.07 --- 6 0 .1845 2.05

ICplldeg 0

loor l -e-

I

-1001 I I I I I 0 . 2 .4 .6 .8 1 .o

x lc

a = . 5 O

(b) Lower surface

Fig . 14 concluded.

16

~-

. _ - I .

- - 0 2 4 6 6 10 1 2 14

Tome. rec

ORIGINAL PAGE IS w. FX)OR QUALITY

-2.0 - 1 . 5 L

0 r,

0 .l .2 .3 .4 . 5 . 6 . I .8.91.0 a l c

A. a l c __ 0.14. Senr i 6.45 mvlPSl _ _

6 . x l c 0.26. Sens: ~ 7.23 ~ mvlPSl

I

E. ,xc -0-75. Sens 6-85 mv1PSI - I

. , L... ,z .. ,4 ~ ~ . ~. > 0 .6 .8 1.0 1.2 1.4

Tome. rec

-2.0r

( a ) Gt = 00.

1 5 L L I 1 L U 0 . 1 .2.3 .4 .5 . 6 . 7 .6 3 1 . 0

a l c

Fig. 15 Time h i s t o r i e s a t f i v e chord s t a t i o n s for u = 0, M = 0.72, and R = 35 x 106.

17

Standard Bibliographic Page

___ -- _____ _ _ _ ~ . __

6 19 Sei i irity Clrssif (of this report)

.. Report No. 12. Government Accession No.

-~ -7-

--__ 1 A05 I

18 -

NASA TM-89080, Cor rec ted C0p.y I L. Title dnd Subtitle

H i g h l i g h t s o f Unsteady Pressure Tests on a 14 Percent S u p e r c r i t i c a l A i r f o i l a t High Reynolds Number, T ranson ic C o n d i t i o n

Rober t W. Hess, David A. S e i d e l , W i l l i a m B. Igoe, and P i e r c e L. Lawing

_ _ - ~ _ _ ._ 7 . Author(s)

__ __ _ _ _ - ~ _ _ _ _ _ _ _ ). I’erforniing Organization Name and Address

NASA Lang ley Research Center Harnpton, V i r g i n i a 23665-5225

-~ - - -____ 12. Sporisoring Agency Name and Address

N a t i o n a l Aeronaut ics and Space A d m i n i s t r a t i o n Washington, DC 20546

1.5. Supplementary No T h i s paper wiy$ be p resen ted a t t h e A I A A 2 5 t h Aerospace Nevada, January 12-15, 1987, as A I A A Paper No. 87-0035.

~

3. Recipient’s Catalog No.

5. Report Date

January 1987 6. Performing Organization Code

505-63-21-01 8. Performing Organization Report No.

10. Work Unit No.

___.____ ~- 11. Contract or Grant No.

13. Type of Report and Period Covered

Techn ica l Memorandum 14. Sponsoring Agency Code

Sciences Meet ing, Reno,

16. Abstract

Steady and unsteady pressures were measured on a 2-D s u p e r c r i t i c a l a i r f o i l i r t h e Langley Research Center 0.3-m T ranson ic Cryogenic Tunnel a t Reynolds numbers f rom 6 x l o 6 t o 35 x l o 6 . f rom k.25 degrees t o k1.0 degrees a t f r e q u e n c i e s f rom 5 Hz t o 60 H z . s p e c i a l requi rements o f t e s t i n g an unsteady p r e s s u r e model i n a p r e s s u r i z e d c r y o g e n i c t u n n e l a re d iscussed. a r e coinpared w i t h GRUMFOIL c a l c u l a t i o n s a t Reynolds number o f 6 x l o 6 ?nd 30 x 106. Exper imenta l unsteady r e s u l t s a t Reynolds numbers o f 6 x l o 6 and 30 x 106 a r e examined f o r Reynolds number e f f e c t s . two mean angles o f a t t a c k a t a Reynolds number o f 30 x l o 6 a r e a l s o examined.

The a i r f o i l was o s c i l l a t e d i n p i t c h a t amp l i t udes The

Se lec ted s t e a d y measured da ta a r e p resen ted and

Measured unsteady r e s u l t s a t

17. Key Words (Suggcsted by Authors(s)) I 18. Distribution Statement

Reynolds number Steady and Unsteady Pressure S u p e r c r i t i c a l Wing Mach Number

U n c l a s s i f i e d - U n l i m i t e d S u b j e c t Category - 02

For sale by the National Technical Information Service, Springfield, Virginia 22161 N A S A Langley Form 6 3 ( . lune 1985)

- _________