AIA A-87-0035 Highlights of Unsteady Pressure Tests on a 14 Percent Supercritical Airfoil at High Reynolds Number, Transonic Condition Robert W. Hess, David A. Seidel, William B. Igoe, and Pierce L. Lawing NASA Langley Research Center VA

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AIM 25th Aerospace Sciences Meeting January 12-15, 19871Ren0, Nevada

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For permission to copy or republish, contact the American lnstttute of Aeronautics and Astronautics 9633 Broadway, N8w York, NY 10019

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

Robert W. Hess*, David A. Seidel**, W i l l i a m B. Igoe***, and P i e r c e L. Lawing****

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

Abs t rac t

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 n t h e Langley Research Center 0 . 3 4 Transonic Cryogenic Tunnel a t Reynolds numbers f r o m 6 x IO6 t o 35 x 106. 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 ampl i tudes from t.25 degrees t o ‘1.0 degrees a t f requencies f r o m 5 Hz t o KO Hz. The s p e c i a l requ i rements of t e s t i n g an unsteady pressure model i n a p ressu r i zed c ryogen ic tunne l a r e discussed. Se lec ted steady measured data a re p resented and a re compared with GRUMFOIL c a l c u l a t i o n s a t Reynolds number o f 6 x IO6 and 30 x 106. Exper imental unsteady r e s u l t s a t Reynolds numbers o f 6 x IO6 and 30 x IO6 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 of a t t a c k a t a Reynolds number of 30 x IO6 are a l s o examined.

C

f k

R V X

Nomenc 1 a t u r e

chord 1 i f t c o e f f i c i e n t o ressure c o e f f i c i e n t hodulus of osc i 1 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 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 v e l o c i t y , f t l s e c streamwise coo rd ina te measured from l e a d i n g edge, i n . peak o s c i l l a t i o n ampl i tude i n p i t c h , degrees, p o s i t i v e l e a d i n g edge up, deg.

steady o r mean dynamic ampl i tude i n p i t c h , p o s i t i v e l ead ing edge up, deg. micron, I 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 A e r o e l a s t i c i t y D i v i s i o n , Member A IAA.

**Research Engineer, Unsteady Aerodynamics Branch, Loads and A e r o e l a s t i c i t y D i v i s i o n , Member A I A A .

***Senior Research Engineer. Transonic Aerodynamics D i v i s i o n , Nat iona l Transon ic F a c i l i t y Operat ions Branch, Member A I A A . ****Senior Research Enai neer. T ranson ic - ~~~ .~ ~

Aerodynamics n i v i s i o n , Exper imental TechniqLer Rrancn, Senior Member A I M .

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Subsc r ip t s

C c o r r e c t e d va lue t t e s t measurement

I n t r o d u c t i o n

The advent of l a r g e c ryogen ic wind tunne ls a l l ows unsteady pressure measurements t o be made on models a t Reynolds numbers t y p i c a l l y exper ienced by h i g h performance a i r c r a f t , t hus e l i m i n a t i n g t h e need for a r t i f i c e s such as boundary- layer t r i p s t o s imu la te boundary l a y e r t r a n s i t i o n a t h i g h Reynolds number. New m i n i a t u r e t ransducers s p e c i f i c a l l y designed t o measure unsteady pressures i n a c ryogen ic environment make these measurements poss ib le . The study repo r ted i n t h i s paper was conducted i n t h e 0 . 3 4 Transonic Cryogenic Tunnel (0.3-m TCT) a t t h e NASA Langley Research Center. With i t s combined pressure, c ryogen ic temperature, and t r a n s o n i c speed c a p a b i l i t i e s t h e 0.3-m TCT can p rov ide f l i g h t equ iva len t a i r f o i l r e s u l t s f o r c u r r e n t a i r c r a f t . Th is tunne l was used i n t h e Advanced Technology A i r f o i l Tes t (ATAT) program i n ex tens i ve steady f low a i r f o i l s tud ies t h a t demonstrated t h e necess i t y f o r h i g h Reynolds number tes t ing . ’ The a i r f o i l used i n t h e present unsteady t e s t s i s a four teen-percent t h i c k , s u p e r c r i t i c a l a i r f o i l , des ignated Sc(2) - 0714, which was,developed a t t h e NASA Lanqey Research Center. The purpose of t h i s t e s t was t o o b t a i n unsteady t r a n s o n i c p ressure measurements f rom 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 t h e prev ious steady f l o w r e s u l t s . A secondary o b j e c t i v e of t h e t e s t was t h e development o f i ns t rumen ta t i on techn iques f o r measuring unsteady pressures a t c ryogen ic temperatures.

The two-dimensional model had a s i x i nch chord and an e i g h t i n c h span. The t e s t was concent ra ted a t a tunne l f reestream Mach number o f 0.72, which prev ious t e s t s i n d i c a t e d t o be t h e design Mach number. Reynolds number (based on a s i x i n c h chord) was va r ied f rom 6 Y I O 6 t o 35 x 106 and Mach number was va r ied a t two Reynolds numbers. The range o f t e s t f requenc ies 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 ampl i tudes which va r ied from 20.25 degrees t o 21.0 degrees. I n t h i s paper, se lec ted steady measured da ta are presented and a re compared w i t h GRUMFOIL ‘a l cu la t i ons a t Reynolds numbers of 6 x 106 and 30 x 106. Exper imental unsteady r e s u l t s a t Reynolds numbers of 6 x 106 and 30 x IO6 are examined f o r Reynolds number e f f e c t s . Measured unsteady r e s u l t s a t two mean angles of a t t a c k a t a Reynolds number o f 30 x 106 a re a l s o discussed.

Apparatus The f i n a l c o n f i g u r a t i o n cons is ted of t ransducers wi th a 10 p s i range and wi th ou tpu ts of between 5 and 9 mv lps i . Each t ransducer was

tl Model mounted i n a recep tac le which i n t u r n was connected t o t h e 0.015 i n c h diameter o r i f i c e by

The Sc(2)-0714 a i r f o i l model i s shown i n a 0.75 i nch l e n g t h of .030 inch i.d. tub ing . Fig. 1. It was machined from an a l l o y (Vascomax-200) t h a t has super io r d imensional The connect ioh between t h e man i fo ld and t h e s t a b i l i t y p r o p e r t i e s a t c ryogen ic cond i t i ons . R c a v i t y machined i n t h e unders ide o f t h e wing, Fig. 2, p rov ided t h e space necessary t o mount t h e t ransducers . Th is c a v i t y was c losed by a cover o l a t e on which some lower su r face t r a n s d u c k were mounted. The wing was supported on one end by a c l o s e - f i t t i n g tang f i x e d t o a d r i v i n g p l a t e wi th machine screws: t h i s end, seen on t h e l e f t i n Fig. 2, was sea led with epoxy. The o the r end was supported by an i n t e g r a l sha f t which r o t a t e d i n a bushing i n t h e tunne l s i d e w a l l p l a t e . A s l i d i n g seal o f f e l t was used t o seal t h e gap between t h e end of t h e o s c i l l a t i n g a i r f o i l and t h e f i x e d tunne l s idewa l l p la te . The p o s i t i o n of t h e supports was designed t o l o c a t e t h e p i t c h a x i s a t t h i r t y - f i v e percent chord.

Transducers

F o r t y - s i x unsteady pressure t ransducers were mounted i n t e r n a l l y i n t h e model. Because o f space c o n s t r a i n t s , f o r t y t h r e e o f t h e t ransducers were mounted i n receptac les connected by a s h o r t l e n g t h (nomina l l y 0.75 i n c h ) o f t ub ing t o t h e o r i f i c e . The remaining t h r e e t ransducers were mounted w i t h t h e t ransducer head l e s s than 0.1 inch below t h e su r face o f t h e wing. The o r i f i c e s o f these t h r e e t ransducers were pa i red w i t h tube mounted t ransducers f o r comparison purposes. A s e r i e s o f t e s t s was conducted t o examine t h e e f fec ts of o r i f i c e diameter, tube diameter, and tube l e n g t h on t h e dynamic response of t h e system. A t atmospher ic c o n d i t i o n s t h e r e was no s i g n i f i c a n t reduc t i on o f dynamic ampl i tude response o r phase s h i f t o f t h e t e s t c o n f i g u r a t i o n up t o 100 Hz.

The l o c a t i o n n f t h e t ransducers i s o iven I~ . ~ ~ ~ ~ . ~ . . .

schemat i ca l l y i n F igures 3 (a ) and 3 (b j . The tube-mounted t ransducer o r i f i c e s a r e l o c a t e d a l t e r n a t e l y i n two rows 0.25 inches on e i t h e r s i d e of t h e cen te r l i n e . On t h e t o p surface t h e o r i f i c e d i s t r i b u t i o n o f t h e t w e n t y - f i v e t ransducers r e s u l t s i n an o r i f i c e every 2% of chord t o x l c = 0.1 and 4% chord t o x l c o f 0.70. The d i s t r i b u t i o n o f t h e 15 tube-mounted t ransducer o r i f i c e s on t h e lower surface i s 2% t o an x/c of 0.1 and increases t o 5% t h e r e a f t e r . The c losemoun ted t ransducers and re fe rence o r i f i c e s a re l o c a t e d 0.5 inches from t h e cen te r l i n e .

The elements o f t h e t ransducer system a re shown i n Fig. 4. Since t h e d i f f e r e n t i a l p ressure between the wing surface and t h e tunne l s t a t i c p ressure cou ld exceed t h e r a t e d c a p a b i l i t y o f the t ransducer , t h e t ransducer was re fe renced t o a m a n i f o l d which i n t u r n was vented t o one o f f i v e re fe rence o r i f i c e s . A re fe rence t ransducer measured t h e pressure d i f f e r e n t i a l between t h e man i fo ld and t h e tunne l s t a t i c pressure.

re fe rence o r i f i c e was i n t e r r u p t e d by a porous f l o w r e s t r i c t e r which damped out the o s c i l l a t i n g pressure from t h e s t a t i c re fe rence o r i f i c e ( r e p l a c i n g t h e l o n g leng ths of t u b i n g u s u a l l y used f o r t h i s purpose). The r e s u l t s o f a s e r i e s of c a l i b r a t i o n s made on d i f f e r e n t combinat ions of porous f low r e s t r i c t o r s and tube leng ths a re g iven i n F ig . 5. The f low r e s t r i c t o r s t e s t e d were commercial ly a v a i l a b l e s i n t e r e d f i l t e r s composed of cons tan t d iameter p a r t i c l e s , w i t h diameters rang ing from IOU t o 25”. Except f o r t h e r e s u l t s f o r two 25u r e s t r i c t o r s i n s e r i e s (shown i n t h e curve l a b e l e d 1=.5, 2-2511) t h e da ta shown i n Fig. 5 a re f o r s i n g l e f l o w r e s t r i c to rs . The reduc t i on i n unsteady pressures i s shown (Fig. 5) as the r a t i o o f t h e imposed o s c i l l a t i n g i n p u t p ressure ampl i tude, P,, t o t h e ou tpu t p ressure ampl i tude, P,, as a f u n c t i o n of frequency and f o r d i f f e r e n t combinat ions o f f i l t e r s and tube leng th , 1. A l so g iven i n t h e same f i g u r e i s t h e t ime requ i red , t, f o r t h e system t o reach e q u i l i b r i u m a f t e r t h e a p p l i c a t i o n o f a s t a t i c pulse.

Two 25u f i l t e r s i n se r ies were se lec ted f o r t h e wind tunne l t e s t . Because o f l ack o f space i n the model t h e manifolds and f l o w r e C t r i C t w C . . . . . . . were loca ted ou ts ide t h e model du r ing t h e t e s t and were connected t o t h e model w i t h approx imate ly 6 inches o f tub ing . W

The t ransducers were t o be recovered a f t e r t h e t e s t and consequent ly cou ld no t be permanently bonded t o t h e receptac le . A s e r i e s o f t e s t s were conducted w i t h candidate mast ics and dummy t ransducers a t 120 deg. K and 20 p s i p ressure d i f f e r e n t i a l t o determine which, if any, would ma in ta in a seal a f t e r repeated cyc l i ng . A mast ic supp l i ed by t h e t ransducer vendor was se lec ted f o r t h e i n s t a l l a t i o n .

OriveT;;st,er arge v a r i a t i o n s i n temperature (120

deg. K t o 320 deg. K ) and s tagna t ion pressure (1.4 atm. - 6 atm.) over the opera t i ng range o f t h e 0.3M-TCT r e s u l t i n s idewa l l deformat ions t h a t r e q u i r e d spec ia l cons ide ra t i ons i n t h e design of the o s c i l l a t i n g d r i v e system. F igu re 6 shows t h e model and t h e d r i v e system i n s t a l l a t i o n i n t h e t e s t sec t ion . The t e s t s e c t i o n i s shown t o t h e r i g h t w i t h t h e t e s t s e c t i o n c e i l i n g removed and can be i d e n t i f i e d by t h e two s l o t s on t h e f l o o r which run under t h e model. The model i s between t h e t e s t sec t i on s idewa l l s which i n t u r n a re between t h e tunne l plenum spaces and f i n a l l y , the tunne l p ressure- s h e l l o r plenum wal ls . The c r i t i c a l elements of the system a re i d e n t i f i e d i n t h e schematic drawing i n Fig. 7. The h y d r a u l i c - r o t a r y a c t u a t o r r e q u i r e d t h e maintenance of p r e c i s e al ignment f o r t h e d u r a t i o n of t h e t e s t . Since t h e 0.3-11 TCT t e s t sec t i on f l o a t s on a cab le suspension system t o accommodate thermal c o n t r a c t i o n a t t h e c o l d ope ra t i ng cond i t i ons , W

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t h e a c t u a t o r and suppor t i ng s t r u c t u r e were a l s o supported by a system o f cables, b locks , and counter we igh ts so t h a t t hey cou ld move w i t h t h e t e s t sec t ion .

An i n s u l a t i n g spacer between t h e a c t u a t o r and t h e d r i v e s h a f t and two ho t a i r blowers were used t o reduce t h e c h i l l on t h e a c t u a t o r and system components ex te rna l t o t h e tunne l . A Te f lon (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 shaf 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 d i sk th rough a be l lows t h a t i s o l a t e d t h e sha 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 tunne l 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 t h e wing tang. The tang was ho l l ow t o p rov ide a p a t h f o r t ransducer cables and t u b i n g which went th rough a matching ho le i n t h e p l a t e and e x i t e d th rough t h e cab le p o r t s i n t h e d r i v e shaft . Th is end o f t h e wing was sealed w i t h epoxy and b o l t e d t o t h e r o t a t i n g disk.

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 shaf t and a bushing i n t h e s i d e w a l l p l a t e . This end o f t h e wing moved r e l a t i v e t o i t s mountino D la te and was sea led

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A Fast F o u r i e r con t iguous cyc les . Trans format ion average (FFT) was taken of t h e da ta t o c a l c u l a t e t h e harmonic components of t h e unsteady pressures. The da ta sample r a t e and number o f cyc les analyzed was se lec ted t o g i v e an accu ra te es t ima te of t h e f i r s t t h r e e fundamental harmonic componets. The harmonic p ressure c o e f f i c i e n t s a r e normal ized by t h e ampl i tude o f 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 p o s i t i o n .

S idewal l boundary l a y e r and ang le of a t t a c k c o r r e c t i o n s were a p p l i e d t o t h e measured steady pressure r e s u 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 are based on t h e theo ry o f Ref. 4 which i s used i n Ref. 5 w i t h measured values o f s idewa l l d isplacement 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 imenta l values i n t h i s paper. The ang le of a t tack c o r r e c t i o n s descr ibed i n Ref. 6 (sometimes r e f e r r e d t o as t h e "Barnwell -Davis- Moore" c o r r e c t i o n ) a d j u s t t h e theo ry o f Dav is - 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 . 3 4 TCT i s :

-ClC B(Itj)h a i =

w i t h f e l t t h a t matched t h e ' p r o f i l e of t h e wing. The wing sha f t was a t tached t o a r o t a r y a re : t ransducer by an i n s u l a t i n g shaf t . The ho l l ow wing s h a f t p rov ided a pa th f o r t h e remaining c = chord = 6 i n . i n s t r u m e n t a t i o n cables. The r o t a r y t ransducer was heated w i t h sur face heaters under thermosta t c o n t r o l and t h e e n t i r e assembly covered w i t h an i n s u l a t e d can.

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

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

d

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The system d i d not develop any problems d u r i n g t h e t e s t . The ang le of a t t a c k was checked v i s u a l l y aga ins t a s idewa l l s c r i b e mark a t t h e beg inn ing o f each days t e s t be fore t h e i n t r o d u c t i o n o f c ryogen ic 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 po r t s . The c o r r e l a t i o n o f t h e geometr ic f l o w angle of a t t a c k 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 du r 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 tunne l i n s t r u m e n t a t i o n were acqu i red us ing t h e 0.3M-TCT data a c q u i s i t i o n system. The model ang le of a t t a c k and pressure da ta were fed t o t h e sys tem's analog da ta 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 per second. S t a t i c da ta values a re 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 da ta were acqu i red us ing analog tape recorders . The i n s t r u m e n t a t i o n s i g n a l was a m p l i f i e d t o be a va lue of about one v o l t RMS. The model ang le of a t t a c k and pressures were taken d i r e c t l y from t h e 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 per second. TO o b t a i n ampl i tude and phase in fo rma t ion a t t h e frequency 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 pe r c y c l e of o s c i l l a t o r y mot ion f o r 64

j = aK/h K = 3.2 (semi e m p i r i c a l cons tan t ,

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

For C1 = 1.0

A; -1.73245 deg.

Resu l ts and D iscuss ion

The t e s t was designed t o e x p l o r e t h e e f fec ts of Reynolds number on unsteady pressures and t o generate a da ta base f o r v a l i d a t i n g unsteady-aerodynamic computer codes. The t e s t c o n d i t i o n s as de f ined by Mach number and Reynolds number a r e shown i n Fig. 8. Tes t p o i n t s were taken a t t h e design Mach number of 0.72 a t t e s t Reynolds numbers va ry ing from 6 x IO6 t o 30 x IO6. Mach number was v a r i e d a t two Reynolds numbers, 15 x I O 6 and 30 x lo6. A t o t a l o f 976 t e s t p o i n t s were taken. The p r imary data base was taken f o r p i t c h - o s c i l l a t i o n f requency between 5 Wz and 40 Hz a t ampl i tude of 20.25 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 da ta was i n hand, t h e p i t c h ampl i tude was increased t o 20.5 and t1.0 degree and t h e p i t c h f requency was a l s o inc reased t o 60 Hz a t t e s t cond i t i ons i n d i c a t e d by t h e s o l i d symbols.

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 of a t tack , ut, approx imate ly 2.5'. 2.0°, 1.5', and OO, and f o r two Reynolds

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numbers, 6 x 106 and 30 x 106, a re shown i n F ig . 9. The exper imental data a re shown as symbols and t h e 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 da ta have been c o r r e c t e d f o r s idewa l l e f f e c t s ( re fs . 4 and 5 ) and ang le of a t t a c k (ref . 6). Ca lcu la ted pressure d i s t r i b u t i o n s f rom t h e 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 (Fig. 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 l a w s o l v e r i n t e g r a t e d w i th a v iscous boundary l a y e r model. GRUMFOIL may be

en tered 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 t h e i n p u t data f o r t h e computed r e s u l t s which a re compared w i t h t h e co r rec ted exper imenta l values of Cp shown i n F ig . 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 h e tunne l t e s t cond i t i ons , t h e co r rec ted values, and t h e values r e s u l t i n g from t h e GRUMFOIL c a l c u l a t i o n s .

The comparisons between exper iment, shown as symbols, and c a l c u l a t i o n s , s o l i d l i n e s , i n Fig. 9 are very good. The shock moves a f t by approx imate ly 8% t o 10% of chord f o r a g iven va lue of at when Reynolds number i s inc reased from 6 x IO6 t o 30 x IO6. The GRUMFOIL code under -p red ic t s t h e p o s i t i o n o f t h e shock a t both Reynolds numbers by approx imate ly 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 severa l cases are shown i n F ig . 10 p l o t t e d aga ins t co r rec ted ang le of a t t a c k and aga ins t ang le of a t t a c k as computed by GRUMFDIL f o r i n p u t values o f Mach number and Cl f o r Reynolds number of 6 x IO6 and 30 x I O 6 . 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 t h e c o r r e c t i o n procedure of r e f . 6. Th i s t r e n d i s s i m i l a r t o t h e one shown i n re f . 8. I r r e s p e c t i v e of t h e ang le of a t t a c k c o r r e c t i o n s , an inc rease i n C 1 of approx imate ly 0.1 i s shown as Reynolds number i s inc reased from 6 x 106 t o 30 x IO6. This i nc rease r e s u l t s from t h e rearward mvement of t h e shock shown i n Fig. 9.

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 t h e unsteady pressure d i s t r i b u t i o n i s shown i n F ig . 11. Resu l t s are g iven i n terms of the modulus of t h e 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, 01, and t h e phase angle, +, between t h e o s c i l l a t i n g p i t c h ang le and t h e

unsteady pressure. Resu l ts a re shown f o r 4 i 1' and 2' and f o r R = 6 x IO6 and 30 x 106. Two o s c i l l a t i o n f requenc ies , 5 Hz and 20 Hz, are presented f o r a p i t c h ampl i tude o f 10.25 degrees. The upper surface pressure d i s t r i b u t i o n s a re shown i n Fig. I l ( a ) and t h e cor respond ing lower su r face pressures a re i n F ig . I l ( b ) .

The shock wave, i d e n t i f i e d by t h e peak i n t h e unsteady pressures, moves a f t about 6% t o 10% chord as R i s increased from 6 x I O 6 t o 30 x 106 a t t h e same tunne l t e s t angles. The unsteady pressures, a t bo th Reynolds numbers, a re s i g n i f i c a n t l y g rea te r ahead of t h e shock a t

at = 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 t h e magnitude o f t h e unsteady pressures due t o a change i n frequency from 5 Hz t o 20 Hz.

For t h e t e s t c o n d i t i o n s shown, t h e pressures ahead of t h e shock are approx imate ly 180' ou t o f phase wi th t h e wing o s c i l l a t i o n . A t t h e shock wave t h e phase s h i f t s f rom -180' t o -360' t o be in-phase w i t h t h e w ing -p i t ch ing mot ion. A f t of t h e shock t h e phase ang le

remains a t -360' a t a t = 2 O , but i s more dependent on frequency a t t h e lower mean p i t c h

angle, at = lo, t end ing t o go back t o -18O0 a t 20 Hr and t o -360- a t 5 Hz.

The lower su r face pressures and phase angles are shown i n Fig. I l ( b ) . The pressures a re low and decrease from l e a d i n g t o t r a i l i n g edge; t h e phase angle i s c lose t o zero except a t t h e reg ion o f t h e lower sur face i n f l e c t i o n .

The e f f e c t of va ry ing t h e ampl i tude o f t h e 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 ig . 12. Pressure d i s t r i b u t i o n s are shown f o r R =

U

30 x 106 f o r two mean angles, at = 1' and 2* and f o r two f requenc ies , f = 40 and 60 HI. Data f o r p i t c h ampl i tudes of 0.25, 0.5, and 1.0 degree a re shown. I n most cases, i n t h i s and t h e f o l l o w i n g f i g u r e , t h e da ta p o i n t s a re n o t connected i n t h e neighborhood of t h e peak shock ampl i tude because t h e peak pressure i s n o t de f i ned by a f i n i t e number o f p ressure o r i f i c e s . The upper surface pressure d i s t r i b u t i o n , Fig. I 2 (a ) , shows a reduc t i on and broadening of t h e shock-generated peak ampl i tude as t h e p i t c h ampl i tude, a, i s inc reased from 0.25' t o 1" a t b o t h f requenc ies and mean angles. Note t h e 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 at = lo.

A secondary peak i n t h e magnitude o f t h e o s c i l l a t i n g pressure i s ev iden t immedia te ly

behind the shock a t i t = 2' and 1" which cou ld be a t t r i b u t e d t o f low seoara t ion and

v

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 t h e XTRANZLIOP~~ computer code p r e d i c t s t h i s secondary response, a l b e i t no t p r e c i s e l y a t t h e same chord l oca t i on . Ca lcu la t i ons w i t h GRUMFOIL shown i n F ig . 13 i l l u s t r a t e s t h a t a more probab le reason f o r t h e secondary response de r i ves from t h e

supersonic reg ions above t h e a i r f o i l . A t 5

1' the re i s a secondary supersonic reg ion behind t h e shock which i s engu l fed by t h e p r imary supersonic reg ion when t h e angle o f a t tack i s

inc reased t o i t = 2'. Tijdeman has noted t h a t t h e f l ow i n t h e supersonic reg ion p r i o r t o t h e fo rmat ion of a shock i s cha rac te r i zed by a s u b s t a n t i a l inc rease i n unsteady pressure generated by t h e movement o f wavelets.12

The upper sur face phase angle, 4 , shows changes i n t h e neighborhood of t h e shock and a f t p o r t i o n of t h e a i r f o i l w i t h p i t c h ampl i tude, F ig . 12(a).

The lower surface pressure ampl i tude and phase, m, are g iven i n F igu re 12(b). Both are

L

4

r e l a t i v e l y independent of p i t c h ampl i tude except i n t h e neighborhood of t h e i n f l e c t i o n o r cusp reg ion o f t h e a i r f o i l . Both pressure and phase decrease from l e a d i n g edge t o t r a i l i n g edge.

W The e f f e c t of va ry ing t h e o s c i l l a t i o n

f requency 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 a re shown f o r R = 30 x

l o 6 f o r mean angles i t = 1' and Z9 and f o r p i t c h ampl i tudes of 0.25O and 0.5*. Oata f o r f requencies o f 5, 15, 40, and 60 Hz a r e presented. I n general t h e excurs ion o f t h e shock on t h e upper surface, F ig . 14(a), i s reduced a t 60 Hz and again t h e second peak . is

g r e a t e r a t = 1' t han a t q = 2" and i s n o t a s t r o n g f u n c t i o n of frequency. As expected t h e phase ang le i s a f u n c t i o n of f requency 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 t h e p rev ious f i g u r e decreasing by approx imate ly 180° behind t h e shock. The lower surface pressures and phase angle, F ig . 14(b) again decrease f rom a maximum a t t h e l e a d i n g edge t o a minimum a t t h e t r a i l i n g edge and showing some dependence on frequency.

Boundary Layer S t a t e

The unsteady pressure t ransducers used i n t h i s t e s t a l s o enabled measurements t o be ob ta ined which a r e o f i n t e r e s t rega rd ing t h e s t a t e of t h e boundary l aye r . The t ime h i s t o r i e s of t h e pressures a t f i v e t ransducer l o c a t i o n s taken when t h e a i r f o i l was locked a t f i x e d angle of a t t a c k a re shown i n F ig . 15. The data shown i n t h i s f i g u r e a re a l l a t a ga in o f 10 bu t t h e t ransducer s e n s i t i v i t y , g iven w i t h each t race ,

W has no t been app l i ed t o pu t the t ime h i s t o r i e s i n eng inee r ing u n i t s . The t ime h i s t o r i e s taken

a t two f i x e d angles o f a t t a c k (4 = 0' and 2 O )

a t R = 35 x 106 a re shown i n Fig. 15(a) and F ig . 15(b) r e s p e c t i v e l y . The steady pressure d i s t r i b u t i o n s a re shown a t t h e r i g h t o f each f i gu re . The s o l i d p o i n t s on t h e pressure d i s t r i b u t i o n mark t h e l o c a t i o n (x/c = .14, .28, .46, .62, .75) of t h e f i v e t ransducers . I n

F ig . 15(a) (4 = OD) t h e 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 of a t u r b u l e n t boundary

l a y e r . However i n Fig. 15(b) (& = 20) t h e pressure i s qu iescent a t x/c of 0.14 and 0.28 i n comparison w i t h t h e t ransducer responses a t

q = 0". A t an x/c o f 0.46 t h e e f f e c t of shock movement i s observed. A t an x/c o f 0.62 t h e shock movement i s s t i l l observed and tu rbu lence i s apparent. A t x/c o f 0.75 t h e

s i g n a l i s comparable t o t h a t a t = OD. The most obvious d i f f e r e n c e between t h e cond i t i ons a t t h e two angles of a t tack i s t h e presence o f a shock and t h e s l i g h t l y more favorab le p ressure

g rad ien t a t at = 2'. The fo rgo ing d i scuss ion

suggests t h a t a t & = 2', t h e boundary l a y e r i s laminar ahead of t h e shock and t h a t t h e p o s s i b i l i t y e x i s t s t h a t l ong runs o f l am ina r f l o w may have e x i s t e d i n t e r m i t t e n t l y du r ing t h e t e s t .

Conclusions

Steady and unsteady pressures a t a 14 pe rcen t s u p e r c r i t i c a l a i r f o i l a t t r a n s o n i c Mach numbers have been measured a t Reynolds numbers from 6 x 106 t o 35 x 106. I ns t rumen ta t i on techn iques were developed t o measure unsteady pressures i n a c ryogen ic tunne l a t f l i g h t Reynolds numbers. Exper imental steady data, c o r r e c t e d f o r w a l l e f fec ts , show ve ry good agreement w i t h c a l c u l a t i o n s 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 bo th show a shock p o s i t i o n t h a t i s dependent on Reynolds number. For a super- c r i t i c a l p ressure d i s t r i b u t i o n a t a chord Reynolds number o f 35 x 106, laminar boundary l a y e r f low was observed over a s i g n i f i c a n t percentage o f t h e a i r f o i l chord.

Acknowledgement

The au thors wish t o acknowledge t h e ass i s tance o f Clyde Gumbert of t h e T h e o r e t i c a l Aerodynamics Branch, NASA Langley Research Center, i n t h e GRUMFOIL c a l c u l a t i o n s .

References

'Ray, E. J.; Ladson, C. L.; Adcock, J. 8.; Lawing, P. L.; and H a l l , R. M.: "Review o f Design and Opera t iona l C h a r a c t e r i s t i c s o f t h e 0.3M Transonic Cryogenic Tunnel." NASA TM 80123 1979.

"A Review of Reynolds Number S tud ies Conducted i n t h e Langley O.3M-Transonic Cryogenic Tunnel ." A I A A l A S M E 3rd J o i n t Thermophysics, F l u i d s , Plasma and Heat T rans fe r Conference, S t . Lou is , MO, June 7-11, 1982. A I A A Paper 82-0941.

'Harr is, C. D.: "Aerodynamic C h a r a c t e r i s t i c s o f a 14-Percent Th ick NASA S u p e r c r i t i c a l A i r f o i l Designed f o r a Normal-Force C o e f f i c i e n t o f 0.7.'' NASA TM X-72712, J u l y 1975.

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

'Jenkins, R . V.; and Adock, J. R.: "Tables f o r C o r r e c t i n g A i r f o i l Data Obtained i n t h e Langley 0.3-Meter Transonic Cryogenic Tunnel f o r S idewal l Boundary Layer E f fec ts . " NASA TM 87723 June 1986.

'Barnwell, R . W.: "Design and Performance Eva lua t i on of S l o t t e d Wal ls f o r Two-Dimensional Wind Tunnels." NASA TM 78648, February 1978.

'Mead, H. R.; and Melnik, R. E.: "GRUMFOIL: A Computer Code f o r t h e Viscous Transn ic Flow Over A i r f o i l s . " NASA CR 3806, October 1985.

'Gumbert, C. R . ; and Newman, P. A.: " V a l i d a t i o n of a Wall I n t e r f e r e n c e AssessmentICorrect ion Procedure f o r A i r f o i l Tes ts i n t h e O.3M-Transonic Cryogenic Tunnel." A I A A 2nd App l ied Aerodynamics Conference, August 21-23$ 1984.

Mundel l , A. R . G.; and Mabey, n. G.: "Pressure F l u c t u a t i o n s Caused by Transon ic ShocklBoundary-Layer I n t e r a c t i o n . " Aeronaut ica l Journa l , August/September 1986.

'Ray, E. J.:

5

"Whitlow, W., Jr.: "XTRRNZL: A Program f o r S o l v i n g t h e General Frequency,l Unsteady Transonic S m a l l D is turbance Equation, NASA TM 8572{1 November 1983.

"User 's Manual f o r XTRANZL (Vers ion 1.2): A Program f o r So lv ing t h e General -Frequen:y Unsteady Transonic Smal l -Disturbance Equation, NASA TM 87737, J u l y 1986

"Tijdeman, H.: ' ' I n v e s t i g a t i o n 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.

Se ide l . D. A.; and Bat ina, J. T.:

A-08.

A - 1 0 0

A-12.

Driving A-14.

A-18. end

A-lea A-20.

CP,;:z1

F ig . 1 Ex te rna l view of model.

0b-09

.b-11

8 6 - 1 3

.6-16

88-97

.6-19

06-27

1

F ig . 2 I n t e r n a l c o n f i g u r a t i o n OF model. Inches

2 - 0

0- 0-

.l -

.2 -

.3 -

.4 - X I E

3- . 5 -

.6 -

.7 -

.a -

.9 - e- 1.0-

1 -

2 -

4 -

5 -

0- 0

* I C

Drlvlnp BDd

upper 8Urf.sCe lnSlt"mcnlallo"

a OrifiCD e close mtd. Iran b RBI. Orifice

1%

I a D - 3 3 *D-34

e D - 3 6 .D-35

.0-37

c-47+ e6-23

*E-25 A-240

( a ) Upper surface

F ig. 3a Transducer l o c a t i o n s .

0

1 E __ G-O?!. .F-44

E-280 1 eD-30

I eF-48 1 OD-31

E-32. 1

1 €48. 1 * F - 4 5

eD-40 E-41.

( b ) Lower sur face

F ig. 3b concluded.

. . ., .. Fig. 4 Elements of t he t ransducer system.

6

I\

P21P1

-'.Lo 0 Frequency, Hr '

insulated rotary transdUCQr housing

Heated rotary transducer

bearing surfaces

PresSUrQ seal

insulating spacers

F ig . 7 Schematic drawing o f model i n s t a l l a t i o n . ._ Fig. 5 Dynamic f low-restr ictor- tube c a l i b r a t i o n resu l ts .

9 Frequency and amplitude 0 Frequency

e e

e D e . e o D O Mach

number .7

.6 I I

0 10 20 40 4 0 to6

Re

Fig. 8 Mach number and Reynolds number t e s t conditions.

F ig . 6 Model i n s t a l l a t i o n .

. . ..._,

7

o Upper surface -2 .o r 0 Lower surface

CP

R = 6.035 X

0

106

1.2 0 .2 .4 .6 .8 1.0

X l C

o Upper surface -2.Or 0 Lower surface

CP

R =30.0 X 106

0

1.2 0 .2 .4 .6 .8 1.0

X I 0

Tunnel Cor rec ted Grumfoi 1 Grumfoi 1 M 0.719 0.705 0.705

cl 2.51 0.756 1.399

Tunnel Cor rec ted

2.504 0. 844 M 0.720 0.701 0.701

C1 0.9581 0.9753 0.9837 1.385 C1 1.0123 1.0256 1.0336

4 0 Upper surface 0 Upper surface 0 Lower surface 0 Lower surface

R-30.0 X lo6

-2.0

R =6.01 X l o 6

CP

0

1.2 0 .2 .4 .6 .8 1.0

x l c X l C

Tunnel Cor rec ted Grumfo i l Tunnel Cor rec ted Grumfoi 1 M 0.72 0. 701 0.701 M 0.721 0. 705 0.705 -

5 1.997 0.393 0.910 C, 0.926 0.939 0.9453

2.002 0.525 0.961 C 1 0.8523 0.8676 0.8757

F ig . 9 Comparisons o f steady t e s t r e s u l t s w i t h c a l c u l a t e d r e s u l t s a t a tunne l Mach number o f 0.72.

W

o Upper surface e Lower surface

R=6.035 X IO6

CP

0

1.2 0 .2 .4 .6 .8 1.0

x l c

Tunnel Cor rec ted Grumfoi l M 0.719 0.701 0.701 0 1.495 0.201 0.493 C, 0.7467 0.7601 0.7680

o Upper surface 0 Lower surface

R = 6.02 X IO6 d

CP

0

x l c

Tunnel Corrected Grumfoi 1 M .721 0.705 0.705 0 -0.005 -1.036 -0.715 c1 n.5951 0.6034 0.6099

0 Upper surface 0 Lower surface

R = 29.98 X lo6

CP

0

1.2- 0 .2 .4 .6 .8 1.0

% I C

Tunnel Corrected Grumfoi l M 0.718 0.705 0.705 0 . 1.501 0,051 0.552 C1 .R373 0.849 0.8555

0 Upper surface 0 Lower surface

R =30.05 x lo6

CP

0

1.2- 0 .2 .4 .6 .8 1.0

% I C

Tunnel Cor rec ted Grumfoi 1 M 0.720 0.701 0.701 0 0.004 -0.92 -0.398 C1 0.5288 0.5383 0.5484

F i g . 9 concluded.

9

W

1.0-

.Q

.8

. 7

.6

.5

1.1

1 .o

.Q

.8

CI

.7

.6

.*

.4

-

-

-

-

-

M - 0.72, R = 6 X 106

" r

I I I I I -1 0 1 2 3 -

a. deQ

M E 0.72, R = 30 X l o 6

I

/

Method

+ Barnwell 0 Orurnfoil

I I I I I 1 2 3 .41

- 2 -1 0- a. deg

Fig. 10 Comparison of l i f t coef f ic ient versus corrected angle-of-attack.

1 . Hr k 1 .o 0 5.0 ,0116

0 20.0 ,0466

.6 Close rntd. trans.

R = 6.64 X IO6 .6

lCplldeg .4

.2

0

-100

-200

0. deg -3001 \ 0000

-400

-500

-600 0 .2 .4 .6 .6 1.0

% I C

4 = 2.06", & = 0.59'

f. Hz k 4

.6 0 5.0 ,0116 + 20.0 ,0465

.4 Close mtd. trans.

I Cp I l d e g R-6.04 X lo6 .2

0

-100 -I 0, d e g -200

-300

-400 1.1.I.I.I.I 0 .2 .4 .6 .8 1.0

X l C

& = LOW, & = -0.561'

( a ) Upper surface

1 , Hr k

0 5.0 ,0153 -20.0 ,0616

Close mtd. trans.

lCplldeg .4 R =30.05 X IO6

.6

.6

.2

0

-100

0. de0 -200 -3001, -400 I , I , ,L1 0 .2 .4 .6 .6 1.0

X l C

6 = 2.05'. & = 0.446'

f , Hz k .6 0 5.0 ,0154

Q 20.0 ,0619

e Close mtd. trans. .6

l C p l l d e g .4 R =30.03 X IO6

.2

0

-100

0. d e g -200

-300

X l C & i 1.05", & -0.255"

Fig. 11. Unsteady pressure tes t results a t a tunnel Mach number of 0.72 and a t a = i0.25".

11

f, Hz k 0 5.0 .0116

-0- 20.0 ,0466

.4r Close mtd. trans.

1 , Hz k o 5.0 ,0153 e 20.0 ,0616 L,

Close mtd. trans.

.2 R = 30.05 X lo6 lCplldeg .2 R =6.64 X lo6 lCplldeg

0 0

100 100

0, deg 0 0. deg 0

-100 . I " < " ' > ' 0 .2 .4 .6 .8 1.0

X l C

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

X l C

4 = 2.05'. & = 0.446' 4 i 2.06', = 0.59'

f, Hz k

0 5.0 .0116 -3 20.0 ,0465

Close mtd. trans.

lCplldeg .2 R =6.04 X lo6

0

100

0, deg 0

-100 o .2 .4 .6 .8 1.0 x l c

& = 1.036", = -0.561'

o 5.0 .0154 -c- 20.0 .0619

.4 Close mtd. trans.

ICpl/deg .2

0

100

0, deg 0

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

x l c

&, = 1.05O, & = -0.255'

W

( b ) Lower surface

F i g . 11 concluded.

C'

W

-1ooc

$ . d e s -201% ,

-300

-400 0 .2 .4 -6 .8 1.0

0.

"I- .6

a. deg zt, dag

- .25 2.07 ----- .6 2.06 --1.0 2.07 I .:IL j,L - .25 zt. 2.07 deg

----- .6 2.06 --1.0 2.07

.4

.2 , ,J ,,".

0

- 1 0 0 ~

. I C f = 40 Hz

-400 I I I I 1

f = 40 HZ

0 .2 .4 .6 .e 1.0

X l C

m. ' ieg -200

V, deQ -200 -300

-400 .* .a .6 .8 I . 0

X l C

-300

-400 f = 60 Hz 0 .2 .4 .6 .6 1.0

X l C f = 60 Hz

F i g . 12 V a r i a t i o n of C deg and rn a t M = 0.72 and R = 30 x 106 w i t h p i t c h a m p l i t u e . PI '

13

.2 r

loor 01, deg zt, deg

- .25 1.02 ----- .5 1.07 -- 1.0 1.03

-1001 L , 0 .2 .4 .6 .8 1.0

X l C

f = 40 Hz

loor - .25 1.06 ---_- .5 1.05 --1.0 0.97

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

X I C

f = 60 HZ

.2 r

loor a. de0 XI, deg - .25 2.07

.5 2.06 1.0 2.07

---_- --

\--.----- m ' d e g -100

ICplIdeY '2i.----__ 0

0 .2 .4 .6 .8 1.0

X I C

f = 40 Hz

0 . de9 zi, dB9

- .25 2.05 .5 2.08

1.0 2.05 ----- -_

( b ) Lower su r face

Fig. 12 concluded.

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

X I C

f = 60 Hz

( b ) = 2"

Fig . 13 Sonic reg ions a t M = 0.72 c a l c u l a t e d by GRUMFOIL code.

18.

v

ICpl ldeg

T O 1

I C P I l d e g

m. d e 4

4

1 . H2 k c d 6 8 4

I .2 .4 .e .8 1.0

- 4 o o L 4 0

X I (

a i .50 - 5 0 0 0 . 2 .4 .6 .a 1.0

(L i .50

( a ) Upper surface

- 1. Hr k atdeg

- 5 ,0154 1.05 ----- 15 ,0461 1.04 -- 40 ,123 1.02 60 ,1845 1.06 I0 l ldes .2

L $ $ -- --- .... 0

loor

-1001 I I I I 0 .2 .4 .e .8 1.0

x / e a = 2 5

l0 l ldeg

- f . Hr k at. deg

- 5 ,0154 2.08 15 ,0461 2.08 4 0 ,123 2.06 60 ,1845 2.05

W ----- -_ - --

ICplIdeg 0

'O0I-

- I. H Z k a,deg I. Hz k 51, deg

,0154 1.06 - 5 ,0154 2.15 15 ,0461 2.09 __- 40 ,123 2.07

_---_ -_--_ ,185 1.05

l c p l l d e g 0

loor l 0 O Y

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

X I C X I C

01 = . 5 a = .5

( b ) Lower surface

F i g . 14 concluded.

W

16

L 1

r

c

A