search.jsp?r=19780012517 2018-05-06t13:01:13+00:00z · in conventional wind tunnels increase uith...

https://ntrs.nasa.gov/search.jsp?R=19780012517 2018-06-24T00:23:41+00:00Z

NASA Technical Memorandum 78635

Influence of Free-Stream Disturbances on Boundary-Layer Transition

William D. Harvey Lrrtrgley Remrrch C Y I I I C ~ Hlz~~rptotr. C 'irgitr icr

Nat~onal Aeronautics and Space Adm~nlstrat~on

Scientific and Technical Information Onice

Considerable exper imenta l evidence e x i s t s which shows that free-stream d i s - tu rbances ( t h e r a t i o o f root-mean-square p r e s s u r e f l u c t u a t i o n s to mean va lues ) i n convent ional wind t u n n e l s i n c r e a s e u i t h i n c r e a s i n g Mach .umber a t low super- s o n i c to moderate hypersonic speeds. I n a d d i t i o n t o l o c a l c o n d i t i o n s , t h e free- stream d i s t u r b a n c e l e v e l i n f l u e n c e s t r a n s i t i o n behavior on s imple test models. Based on t h i s obse rva t ion , e x i s t i n g n o i s e - t r a n s i t i o n d a t a obta ined i n t h e sane test f a c i l i t y have been c o r r e l a t e d f o r a l a r g e number o f r e f e r e n c e s h a r p cones and flat p l a t e s and are shorn t o c o l l a p s e along a s i n g l e curve. Th i s r e s u l t is a s i g n i f i c a n t improvement over previous a t t e m p t s t o c o r r e l a t e n o i s e - t r a n s i t i o n d a t a .

INTRODUCTION

The d i f f i c u l t y o f r e c o n c i l i n g l a r g e d i f f e r e n c e s i n t r a n s i t i o n measurements on simple bodies i n d i f f e r e n t wind t u n n e l s , as well as d i f f e r e n c e s i n tunne l measurements and f l i g h t measurements, is well-known. Many f a c t o r s in f luence t r a n - s i t i o n , and u n i t Reynolds number and wind-tunnel d i s tu rbance- leve l e f f e c t s a r e among t h e least c o n s i s t e n t and least understood. A e r o b a l l i s t i c ranges and t h e f l i g h t environment are e s s e n t i a l l y d i s t u r b a n c e free, and u n i t Reynolds number e f f e c t s that are somewhat d i f f e r e n t f r o m those observed i n wind t u n n e l s wi th i n h e r e n t free-stream d i s t u r b a n c e s have been found i n both c a s e s ( r e f s . 1 and 2 ) . A r e v e a l i n g eva lua t ion o f t h e many f a c t o r s (which can c o m b i ~ e i n complex ways) t h a t a f f e c t t r a n s i t i o n has been e x t e n s i v e l y desc r ibed i n r e f e r e n c e s 3 t o 5. Numerous r e s e a r c h e r s ( r e f s . 6 t o 13) have e s t a b l i s h e d t h a t t r a n s i t i o n i n super- s o n i c wind t u n n e l s is inf luenced by f ree-s t ream a c o u s t i c d i s t u r b a n c e s t h a t a r e r a d i a t e d f r o m t h e t u r b u l e n t boundary l a y e r s on t h e tunne l wa l l s . De ta i l ed amplitude and frequency composit ion o f t h e s e d i s t u r b a n c e s is g e n e r a l l y p e c u l i a r t o a p a r t i c u l a r tunne l ; t h e r e f o r e , Reynolds number s c a l i n g can be inf luenced t o v a r i - ous degrees . I n g e n e r a l , p r e s s u r e f l u c t u a t i o n s are t h e dominant f ree-s t ream d i s tu rbance mode t h a t a f f e c t s unsteady aerodynamic tests i n t r a n s o n i c wind tun- n e l s and boundary-layer s t a b i l i t y i n h igh supersonic /hypersonic wind t u n n e l s . Turbulent boundary l a y e r s r a d i a t e sound wi th i n c r e a s i n g e f f i c i e n c y a s t h e f ree - stream Mach number & is inc reased because r a d i a t i o n sources a c q u i r e r e l a t i v e superson ic phase v e l o c i t i e s , and t h e o v e r a l l i n t e n s i t y o f t h e sound is Reynolds number dependent. Reference 14 has shown t h a t , i n t h e JPL 20-Inch Supersonic Wind Tunnel, t h e r a d i a t e d i n t e n s i t y from t h e t u r b u l e n t boundary l a y e r s c a l e s wi th t h e square of t h e wa l l s h e a r stress and is about 1/5 t o 1/4 o f t h e s h e a r stress measured on t h e tunne l wall f o r superson ic flow ( 1 - 5 5 M 5 5 ) . If t h e p r e s s u r e f l u c t u a t i o n s a r e nondimensionalized by t h e mean s h e a r s t r e s s a t t h e wall, the Reynolds number e f f e c t d i sappears and t h e r a d i a t e d f i e l d s c a l e s wi th t h i s boundary-layer parameter ( r e f . 14 1.

Tt.,s paper is concerned u i t h a s i m p l i f i e d c o r r e l a t i o n o f wind-tunnel d i s - tu rbances with t r a n s i t i o n on simple models (cones and f l a t p l a t e s ) by us ing

free-stream d i s tu rbance l e v e l s a v a i l a b l e frum e i t h e r hot-wire anemometer d a t a o r a c o u s t i c p ressure t r ansducers t h a t are mounted f l u s h wi th t h e s u r f a c e o f t h e cones and f lat p l a t e s beneath laminar boundary l a y e r s . E a r l i e r a t t empts (refs. 6 t o 8, 15, and 16) t o relate t r a n s i t i o n Reynolds number with nondimen- s i o n a l free-stream pressure f l u c t u a t i o n s were unsuccessful i n c o l l a p s i n g r e s u l t s f o r s e v e r a l f a c i l i t i e s and l h c h numbers. References 13, 17, and 18 c o r r e l a t e d t r a n s i t i o n Reynolds numbers f o r a l a r g e number o f wind t u n n e l s a t 2 2 & 8 based on parameters such as local condi t ions a t t h e tunnel wa l l , tunnel s i z e , and rad ia ted sound. Reference 19 c o r r e l a t e d wind-tunnel r ~ i s e data with reason- able success over a range of Mach numbers and tunne l size: f o r supersonic / hypersonic flaw i n terms o f free-stream and l o c a l paramett*rs. Free-stream d i s - turbance l e v e l s and t r a n s i t i o n measurements i n a l a r g e number of wind t u n n e l s have r e c e n t l y been compared (ref. 201, and although no clear understanding o f t h e mechanisms i n f l u e n c i r g t h e t r a n s i t i o n process was revealed, no doubt remains t h a t t r a n s i t i o n is inf luenced by free-stream dis tu rbances over t h e e n t i r e speed range.

This paper a l s o a t t empts t o extend earlier r e s u l t s by comparing t h e va r ia - t i o n s i n wall and stream d i s tu rbances over a wide range of Mach number and Rey- nolds number i n dind tunne l s of d i f f e r e n t s i z e s .

f frequency

H Mach number

P mean static pressure

q dynamic p ressure , 1 pH2 2

R u n i t Reynolds number

rmz roo t umn square

t temperature

U v e l o c i t y

a ang le of a t t a c k

T shear stress

Y r a t i o o f s p e c i f i c h e a t s

0 c cone ha1 f -angle

Subscr ipts :

a based on a c o u s t i c o r i g i n

edge c o n d i t i m s

local condi t ions

total condi t ions

t r a n s i t i o n

wall c o n d i t i m s

f r e e stream

f luc tua t ing values

MALL AND FREE-STRBAH PRESSURE FLOCiQAPIOIIS

The m e a s u r e m t d i f f i c u l t i e s and final data accuracy o f ind iv idua l refer- enceable r e s u l t s are r a r e l y reported. Therefore, da ta accuracy is d i f f i c u l t t o i den t i fy and is too numerous and eaapl icated to evaluate properly i n this paper. However, su f f i c i en t da ta are ava i l ab l e t o e s t a b l i s h t h e general l e v e l and t rend of wind-tunnel noise for a r a t h e r wide range o f test condi t ions, as shown i n f i gu re 1. (See t a b l e I f o r da t a source and refs. 2, 7 t o 9 , 13 t o 18, and 21 to 29.) The free-stream m pressure f l uc tua t ions have been divided by t h e mean static pressur- ( f i g . 1 ( a ) and by t h e dynamic pressure ( f i g . 1 ( b ) ) i n order t o compare t he va r i a t i on o f rad ia ted sound with and without free-stream Mach number effects. These Cata have been obtained by e i t h e r hot-wire anemumetry or pressure t ransducers Chat have been mounted f l u sh with t he surface o f sharp cones and f lat p l a t e s beneath laminar boundary l a y e r s (except for t h e electron-beam rms dens i ty r e s u l t s (discussed i n ref. 21 and shown i n f i g . 1 as cross-circles a t Eb, = 20) i n which mas dens i ty and pressure f l uc tua t ions have been assumed proport ional (ref. 1 '8 ) . that is, neg l ig ib l e v o r t i c i t y ) .

It has previously been shown t h a t t he acous t i c pressure (pv /p ) is not g r e a t l y changed a f t e r passing through weak obl ique shocks (refs. 15 and 19). Therefore, sur face acous t ic measurements a r e considered t o be an adequate represen ta t ion o f free-stream dis turbance i n t e n s i t y l eve l s . A l l values o f H, f o r helium i n f igure 1 (da ta source given i n t a b l e I) have been corrected t o equiv- a l e n t values of Mw f o r air by using the equation kfmvair = 1.291Ed,,~ which is based on (Y - 1 1 ~ 2 = Constant. (See ref. 30. ) The s o l i d symbols ?f ig . 1 ) are f o r data fmm t ransonic tunnels with e i t h e r perforated o r s l o t t e d walls. It

- has been experimentally shown t h a t a l p O D genera l ly decreases with increasing un i t Reynolds number f o r a given supersonic/hypersonic wind tunnel and increases with test Mach number when t h e wal l boundary l aye r is turbulen t ( r e f s . 6, 7, and 9). Therefore, the observed va r i a t i on i n dis turbance l eve l ( f i g . 1 ) f o r a given Pb, is mainly due t o a u n i t Reynolds number e f f e c t . It should be noted t h a t a l l the reference dis turbance da ta presented i n f i gu re 1 a r e p lo t ted aga ins t M, a t t h e sensor a x i a l loca t ion i n t h e i r respec t ive tunnel test sec t ion except f o r the unpublished data obtained i n t he Mach 5 p i l o t q u i e t tunnel a t t h e Langley Research Center and the published r e s u l t s ( r e f . 24) obtained i n the same tunnel. These l a t t e r r e s u l t s have been p lo t ted aga ins t

the upstream acous t i c o r i g i n k c h numbers (2.27 5 #a 5 5 ) t h a t correspond to severa l a x i a l probe loca t ions within t h e flow test rhombus by t r a c i n g along characteristic lines ca lcu la ted f o r the rapid-expansion Mach 5 p i l o t q u i e t tunne l at t h e -ley Research Center ( r e f . 24). These da t a demonstrate t h e i n t e r - estirtg r e s u l t t h a t dis turbances m decrease to very law l e v e l s i n t he upstream end o f the test rhombus o f t h i s nozzle (H, = 2.27) and can approach c lo se agree- slent with dis turbance l e v e l s from conventional tunnels that have more gradual expansions and with corresponding values of !&,. The dis turbance l e v e l i;r two d i f f e r e n t nozzle geosretries with = 2.3 ( f i g . 1) was measured and is reported i n re fe rence 29. A higher l e v e l uas obtained i n t h e two-dimensional half nozzle with one flat s i d e wall than i n a syumetric nozzle; t h i s f a c t suggested that favorable pressure g rad i en t s may be important i n reducing dis turbances. Rela- t i v e to the zero-pressure grad ien t case, the r a t i o o f rms wall pressure f luctua- tions t o dynamic pressure f l uc tua t ions is greater i n an adverse grad ien t than i n a favorable grad ien t . Further s tud i e s are required t o determine whether these new results are due t o t h e large favorable pressure g rad i en t s i n t h e rapid- expansion nozzle, to reduced energy a t low-frequency geometry, o r simply t o a reduction i n no ise l e v e l t h a t is caused by t h e decreasing boundary-layer thick- ness and Hach number as t he throat is approached.

Figure 1 ( a ) c l e a r l y shows that acous t ic dis turbances i n t ransonic tunne ls increase rap id ly with Hach number a t 0.2 < W, < 1.0. These r e s u l t s can be affected by diffUser design, tunnel-wall-hole resonance, flow through s l o t t e d walls, model support , and fan noise. It is genera l ly found t h a t acous t i c resonance effects i n t ransonic tunne ls with open s l o t s have c h a r a c t e r i s t i c frequencies less than about 200 Hz. The dis turbance l e v e l s f o r 1.5 5 H, S 2 i n supersonic tunne ls a r e considerably lower than f o r t ransonic tunne ls ( f i g . 1).

Barever, values of E/po0 aga in increase with increasing Mach number ( f i g . 1 (a)) although they have a tendency t o l e v e l o f f o r possibly decrease a t high hypersonic speeds (Ma 2 20 and a r e dominated by r ad i a t i on f r o m t he tu r - bulent boundary l aye r on t h e wall ( r e f s . 6 t o 13).

The t e n t a t i v e data f a i r i n g s i n f i gu re 1 are intended t o represent an aver- age trend without considering u n i t Reynolds number e f f e c t s . It should a l s o be noted that the re appears to be no s i g n i f i c a n t effect o f wall cooling on d i s tu r - bance l eve l , a t l e a s t f o r t h e l imi ted r e s u l t s shown.

Radiated pressure f l uc tua t ions divided by e i t h e r free-stream loca! s t a t i c pressure ( f i g . l ( a ) ) o r dynamic pressure ( f i g . 1 ( b ) ) e x h i b i t a wide s c a t t e r , mainly because o f u n i t Heynolds number e f f e c t s . The da ta and pred ic t ions ( r e f s . 16 and 31 presented i n f i gu re 2 (see t a b l e I1 f o r da ta source and r e f s . 6 t o 8, 13 t o 16, 21 t o 23, 25, 26, and 32 t o 42) include the effects o f u n i t Reynolds number (refs. 14 and 32) and Mach number by d iv id ing t h e rms pressure f l uc tua t ions measured e i t h e r a t the tunnel wall o r i n the free stream by t h e tes t - sec t ion l o c a l wall shear stress. Except f o r the electron-beam r e s u l t s o f reference 21, the free-stream data ( s o l i d symbols) have been obtained by e i t h e r hot-wire anemometry o r acous t ica l t ransducers on models beneatk laminar boundary l aye r s . The wall data (open symbols) have been obtained ei:her with acous t ic t ransducers beneath a curbulent boundary l aye r on t h e t u n n ~ l wal l o r under turbulent boundary l aye r s on f l a t p l a t e s ( r e f s . 40 and 41) i n the free

s t r e a m and were considered equivalent. A l l values of 4 f o r helium have beern corrected to values o f #, f o r air (ref. 30 1.

With the exception o f data obtained on the flat p l a t e s (refs. 39 and 41 1, t he general trend f o r wall da ta (fig. 2) is to gradual ly increase i n l e v e l f o r 0.1 < Fi-< 10, with an indica t ion o f leve l ing off a t high speeds. The wall pre- d i c t i o n s i n reference 31 and the free-stream predic t ions i n reference 16 agree with the general trend o f the data that are mainly for ad iaba t i c condit ions ( t a b l e 11). For the l imi t ing values o f & = 0 and -, t he respec t ive normal-

ized dis turbances based on the present data fairings an @hW = 2.5 and 5.3. U t i l i z ing these l imi t ing values and an approach similar t o t h a t appl ied earlier (ref. 191, the following r e l a t i o n m y be used f o r the wall data and k/tt = 1.0 (shoun i n f i g . 2):

Except for the data o f reference 13 which seem excessively high f o r cur- r en t ly unknown reasons, t h e fme-stream data represented by t h e present f a i r i n g are considerably lower i n magnitude than the wall data at low speeds but increase with Hach nunher to an apparent peak at 8 < &, < 10 (fig. 2) . Surpris ingly,

above about M,,,, = 10, @hw decreases with increasing &,, perhaps due to the lower k/tt values f o r many o f these high Mach number r e s u l t s .

The r e s u l t s shown i n f i gu res 1 and 2 c l e a r l y i nd ica t e that inherent d i s tu r - bance l e v e l s i n most ground-based test facilities are much higher than t h e low l e v e l an t ic ipa ted f o r flight-sinnrlation and na tura l - t rans i t ion s t u d i e s (refs. 2, 43, and 44). However, if the da ta presented i n f i gu res 1 and 2 and reference 19 can be accepted as representa t ive o f the dis turbance l e v e l s i n conventional wind tmels, then estimates o f e i t h e r t h e wall o r free-stream pressure f luc tua t ions can be made f o r any wind tunnel. For example, t he free-stream pressure fluctua- t i o n s i n ava i lab le supersonic wind tunnels may be estimated from t h e following equation:

where R, is the free-stream u n i t Reynolds number and the constant of propor- t i o n a l i t y is presumably a function of tunnel geometry. Equation ( 2 ) is not appl icable over the subsonic and t ransonic speed range. In general , t he wind- tunnel free-stream rms pressure f l uc tua t ions divided by the mean s t a t i c pressure can be expected t o increase with increasing speed (1.5 5 Mw 5 10) a t the same u n i t Reynolds number up to M, 10 and then to decrease f o r b$+, 5 20. For &, less than 1.5, model support and diff 'user generated dis turbances can cause sound propagation upstream i n the test sec t ion r e su l t i ng i n t he r e l a t i v e l y high rnrs pressure f luc tua t ions i n t ransonic tunnels ( f i g . 1 ) . Presumably, choking the flow ahead o f the support and d i f f u s e r i n these tunnels would s i g n i f i c a n t l y reduce the stream disturbance l eve l .

CORRELATION OF TRANSITION REYNOLDS NUMBER WITH

FREE-STREAM DISTURBANCE LEVEL

Considerable t h e o r e t i c a l and experimental evidence e x i s t s which s u g g e s t s t h a t t h e t r a n s i t i o n p rocess on models is a func t ion o f both l o c a l c o n d i t i o n s and free-stream d i s tu rbance l e v e l . Thus, t r a n s i t i o n r e s u l t s f o r cones over t h e speed range are not g e n e r a l l y expected t o a g r e e wi th r e s u l t s f o r f l a t p l a t e s and c y l i n d e r s a t a = O0 (ref. 45). The e f f e c t o f l o c a l Mach number on t h e beginning o f t r a n s i t i o n was evaluated by comparing va lues o f R l , t r ob ta ined sn s h a r p cones , 2.870 5 8, 5 16.750, i n test f a c i l i t i e s o f s e v e r a l s i z e s f o r a near ly cons tan t va lue o f free-stream dis tu rbance l e v e l . These d a t a are shown

in f i g u r e 3 f o r t y p i c a l va lues of @/pa found i n fw ground test f a c i l i - t ies and are compared with t r a n s i t i o n d a t a on cones i n f l i g h t ( r e f . 46) and a e r o b a l l i s t i c ranges ( r e f . 1) where stream dis tu rbance l e v e l s are considered n e g l i g i b l e . The dashed l i n e s i n f i g u r e 3 are d a t a f a i r i n g s f o r n e a r l y c o n s t a n t

va lues of @/paD. A l l d a t a shown (Pig . 3 ) have u n i t Reynolds number e f f e c t s . For example, a b a r is used t o i n d i c a t e t h e spread i n R1 tr f o r t h e range d a t a f . 1 . The tunnel d a t a i n f i g u r e 3 c l e a r l y demonstrate a varying effect o f l o c a l Mach number on t r a n s i t i o n i n t h e low Mach number reg ion (M, < 5 ) f o r a given free-stream dis tu rbance l e v e l compared with a much s t r o n g e r e f f e c t i n t h e h igher Mach number region (5 5 M, 2 161, as shown i n r e f e r e n c e 9. It appears t h a t f o r e s s e n t i a l l y d i s tu rbance- f ree flows and smooth models, t r a n s i t i o n is mainly a func t ion of local Mach number and u n i t Reynolds number and R l , t r is expected t o i n c r e a s e wi th inc reas ing M,.

References 47 and 53 have shown t h a t t r a n s i t i o n d a t a on s h a r p cones can be s a t i s f a c t o r i l y c o r r e l a t e d f o r 3 5 Moo d 21 i n s e v e r a l t u n n e l s o f varying s i z e s based on a n o i s e - t r a n s i t i o n empi r ica l equat ion. Th is c o r r e l a t i o n f o r t h e end o f t r a n s i t i o n is a func t ion o f tunne l t e s t - s e c t i o n mean s k i n - f r i c t i o n c o e f f i - c i e n t , t u r b u l e n t boundary-layer displacement th ickness , t e s t - s e c t i o n c i rcumfer- ence, and re fe rence tunne l t e s t - s e c t i o n per imeter . The p resen t a t t empt t o correlate t r a n s i t i o n d a t a is o f course n o t completely new and stems from t h e b a s i c premise t h a t t h e s t a b i l i t y o f a laminar boundary l a y e r on a model is a f f e c t e d by i n p u t d i s tu rbance l e v e l and l o c a l dynamic p ressure . Although no d i r e c t account is taken o f t h e energy s p e c t r a , t h e r e is an i m p l i c i t i n c l u s i o n o f t h i s e f f e c t i n t h a t t h e s p e c t r a from t h e v a r i o u s f a c i l i t i e s a r e q u i t e similar. The fact t h a t t r a n s i t i o n was shown i n f i g u r e 3 t o be a s t r o n g func t ion o f both f ree - stream dis tu rbance l e v e l and M, a l s o l e n d s support t o t h e no i se parameter used here in . The in f luence o f free-stream d i s t u r b a n m l e v e l on t r a n s i t i o n f o r cones and f l a t p l a t e s is shown i n f i g u r e s 4 and 5 and may correspond t o mode 1 and mode 2 i n s t a b i l i t i e s , r e s p e c t i v e l y , i n accordancd with r e s u l t s from Mack's l i n e a r s t a b i l i t y theory (refs. 48 t o 50).

F igures 4 and 5 p resen t da ta f o r t h e beginrung o f t r a n s i t i o n as a func t ion o f e i t h e r free-stream turbu lence l e v e l o r rms p r e s s u r e f l u c t u a t i o n s d iv ided by local c o n d i t i o n s on sharp cones, c y l i n d e r s , and f l a t p l a t e s a t 3 = 00. The t r a n s i t i o n d a t a represen t a wide range o f experiments and c o n d i t i o n s ( s e e t a b l e I11 f w d a t a source and measurement techpique and refs. 1 , 6 , 7, 9 , 10 t o 13, 16, 18, 46, and 51 t o 61 1 where both R l , tr and f ree-s t ream d i s t u r b a n c e

l e v e l s were measured i n t h e same f a c i l i t y for a given test model and geometry. Based on p r o f i l e s o f t r a n s i t i o n f o r s e v e r a l u n i t Reynolds numbers o f t h e s e reference data, d a t a fo r t h e end of t r a n s i t i o n have been ad jus ted t o t h e b e g i ~ i n g o f t r a n s i t i o n by a cons tan t f a c t o r o f 1.55 f o r r e f e r e n c e s 11, 13, 17, and 18 and by a f a c t o r o f 1.33 f o r re fe rence 10; t h e s e data are included i n f i g u r e s 4 and 5.

F i re 4 shows t h e v a r i a t i o n of R l , t r with free-stream dis tu rbance l e v e l f o r 0 rMm : 4 where n o i s e - t r a n s i t i o n dominance could be inf luenced by mcde 1 i n s t a b i l i t i e s ( r e f s . 48 t o 50) . T r a n s i t i o n R 1 tr i n wind t u n n e l s i n c r e a s e s wi th decreas ing d i s t u r b a n c e l e v e l as expected bu t approaches d i f f e r - i n g l e v e l s o f near ly cons tan t R l , t r a t low d i s tu rbance l e v e l s ( l e s s than 103) due t o c h a r a c t e r i s t i c f requenc ies and wave o r i e n t a t i o n i n t h e d i s tu rbance spec- trum o f t h e re fe rence d a t a , d i f f e r e n t wall temperature, and d i f f e r e n t Mach num-

ber . A t t h e lowest value of = 2 x t r a n s i t i o n - p o i n t behavior on wind-tunnel models t ends t o aCi~r i%~h-the lower l e v e l s i n d i c a t e d f o r cones i n free f l i g h t ( r e f . 46) and a e r b b a l l i s t i c range cones ( r e f . 1 ) t h a t have presuma- b l y n e g l i g i b l e in f luence due t o d i s tu rbance l e v e l .

Var ia t ions i n R l , t r f o r t h e f l i g h t ( r e f . 46) and range ( r e f . 1) d a t a are i n d i c a t i v e o f u n i t Reynolds number and Mach number effects and are not p l o t t e d a t a p a r t i c u l a r d i s tu rbance l e v e l but merely s e r v e a s a comparison f o r t h e tcn- n e l data. Actual d i s tu rbance l e v e l s r e p r e s e n t a t i v e ol' t h e range and f l i g h t data are expected t o be a t least 1 o r 2 o r d e r s o f magnitude lower than i n d i c a t e d i n f i g u r e 4. The t rend o f t h e t r a n s i t i o n r e s u l t s shown i n f i g u r e 4 (an i n c r e a s e i n R l , t r with a decrease i n d i s tu rbance l e v e l ) is i n agreement wi th t h a t expected f r o s Mack's s t a b i l i t y theory ( r e f s . 48 t o 50) f o r first mode i n s t a b i l i t y a t 0 2 M, 5 4 i n c r e a s i R l , tr with i n c r e a s i n g Mach number f o r low d i s t u r - bance l e v e l ) . The d a t a o f f i g u r e 4 do not i n d i c a t e a s i g n i f i c a n t e f f e c t o f model geometry on t h e n o i s e - t r a n s i t i o n r e s u l t s , but i n d i c a t e on ly t h e i n f l u e n c e o f t h e free-stream d i s tu rbance l e v e l d ivided by edge cond i t ions and t h e inher- e n t c h a r a c t e r i s t i c energy s p e c t r a of a given re fe rence wind tunne l t h a t a r e i m p l i c i t l y included.

Figure 5 shows t h e d i f f e r e n t v a r i a t i o n s o f R 1 tr with a nondimensionalized free-stream d i s tu rbance l e v e l f o r s h a r p cones and f i a t p l a t e s a t supersonic / hypersonic speeds ( 4 <= MoD 5 23) where t h e r e s u l t s o f r e f e r e n c e s 48 t o 50 may be i n d i c a t i v e o f the t rend. It is discussed i n re fe rence 4 and shown i n r e f e r - ence 45 t h a t t r a n s i t i o n Reynolds numbers on cones and c y l i n d e r s may d i f f e r by a varying f a c t o r f o r 3 < Mz < 10 a t about t h e same u n i t Reynolds number. For Me - 3 , t h e e s t a b l i s h e d f a c t o r v a r i e s from 3 t o 1.6; f o r 5 2 M, 2 7, it v a r i e s f r o m 1.4 t o 1.1; and f o r Me = 8, no d e f i n i t e conclusion was reached as t o t h e cause of these observed d i f f e r e n c e s . The r e s u l t s i n f i g u r e 4 ( 0 2 MoD =< 4) show no l a r g e d i f f e r e n c e s i n R l , t r between axisymmetric and p lanar bodies. T r a n s i t i o n va lues shown i n f i g u r e 5 (4 2 M, 2 23) over a range o f u n i t Rey- nolds numbers f o r cones and f l a t p l a t e s d i f f e r i n l e v e l by about t h e same fac- t o r s as might be expected from t h e r e s u l t s o f r e f e r e n c e 45. Boundary-layer s t a b i l i t y o r R l , t r i n c r e a s e s wi th decreas ing d i s tu rbance l e v e l ( f i g . 5 ) i n a t rend similar t o t h a t f o r Ma0 < 4 ( f i g . 4 ) and similar models, but i t reaches s i g n i f i c a n t l y higher values f o r Me > 10. I n p a r t i c u l a r , t h e sharp-cone r e s u l t s ( f i g . 5) a t t h e lowest d i s tu rbance l e v e l s tend t o approach t h e R l , t r va lues

obtained for f l i g h t data at % = 15. This trend i n t he wind-tunnel t r a n s i t i o n da ta is not too su rp r i s ing s ince references 49 and 50 show t h a t , f o r mode 2 ins ta - b i l i t i e s and constant Reynolds number, the maximum s p a t i a l amplif icat ion rate changes by about t he same order o f magnitude as do the t r a n s i t i o n r e s u l t s i n f i " i g u r e 5 over the Mach number range 4 5 5 20. A comparison of the experimentally observed, most unstable frequencies (refs. 51 and 52) with t h e second- mode theo re t i ca l r e s u l t s (refs. 49 and 50) f o r sharp cones a t hypersonic speeds has served t o i d e n t i e t he est imation of t r a n s i t i o n t rends a t high speeds and, to a c e r t a i n ex ten t , the periodic waves.

It should be reca l led that t h e rad ia ted free-stream pressure f l uc tua t ions

P, p, increase with increasing % and reach very high l e v e l s (fig. l(a)) . P/ The highest values of R l , t r shown i n f i g u r e 5 were obtained on models i n wind tunnels and correspond to high Mach numbers (M, = 20), bu t a l s o t o high values of absolu te dis turbance leve l . This fact provokes quest ions concerning t h e effects o f absolute dis turbance l e v e l on t r a n s i t i o n ; both dis turbance l e v e l and t r a n s i t i o n are known t o vary with u n i t Reynolds number and from tunnel t o tunnel. However, when R l , t r is cor re la ted with the free-stream disturbance leve: divided by l o c a l dynamic pressure, t h e da t a tend t o co l lapse and agree with t h e individual data f a i r i n g s (fig. 5) t o within about 20 percent f o r cones and f la t p la tes . Accounting f o r both free-stream dis turbance l e v e l and l o c a l condi t ions is c e r t a i n l y not the only funct ion con t ro l l i ng t r a n s i t i o n but does serve as a s ign i f i can t improvement over previous at tempts ( r e f s . 7 and 16) t o c o r r e l a t e ao ise- t rans i t ion data f o r a la rge number of tunnels. It may be that t h e a b i l i t y of large-amplitude ex te rna l dis turbances t o induce t r a n s i t i o n is reduced a t l o c a l condi t ions t h a t correspond t o high Mach number flow. I n a l l probabi l i ty , the amplified c r i t i c a l frequencies are more important than dis turbance l eve l . The r e s u l t s i n f i gu res 4 and 5 ind ica t e that t r a n s i t i o n mea~urements on models could more c lose ly simulate those i n flight i f ground test f a c i l i t i e s were prop- e r l y designed t o reduce flow noise tc approximately those l e v e l s and scales expected i n f l i g h t ( r e f . 43). It would be highly des i r ab l e t o be a b l e t o vary both disturbance l e v e l and energy spec t ra i n such quiescent tunnels. An e f f o r t has been made toward the development of a Mach 5 p i l o t q u i e t tunnel a t t he Langley Research Center (refs. 2 and 24) that can be used t o inves t iga te the inf luence of free-stream disturbance l e v e l on t r a n s i t i o n .

CONCLUDING REMARKS

A s ign i f i can t quant i ty of ava i lab le da ta has been compiled and an attempt has been made to evaluate the e f f e c t s of wind-tunnel dis turbance l e v e l s on boundary-layer s t a b i l i t y and t r a n s i t i o n on t e s t models.

I n general , wind-tunnel root-mean-square disturbance l e v e l s divided by free-stream s t a t i c pressure can be expected t o increase with increas ing f ree- stream Mach number ( 1.5 6 % <= 10) a t the same u n i t Reynolds number up t o low hypersonic Mach numbers and can be expected t o decrease a t very high Mach numbers (M ? 20 ) .

The s t a b i l i t y of a laminar boundary l aye r and the subsequent onset of t ran- s i t i o n f o r simple axisymmetric and planar surfaces a r e influenced both by free-

ORIGINAL PAGE 1s OF POOR Q U A L ~

stream disturbance l eve l and loca l condi t ions f o r a wide range of test condi- t i ons . The beginning of t r a n s i t i o n on these su r f aces has been shown t o co l l apse along a s i n g l e curve when cor re la ted w i t h the nondimensional free-stream d i s tu r - bance l e v e l based on l o c a l dynamic pressure. The general t rend f o r t r a n s i t i o n t h a t is expected from l i n e a r s tabi l . i ty theory a t r e l a t i v e l y low ampl i f ica t ion and hypersonic speeds supports t he se r e s u l t s . The present r e s u l t s are considered a s i g n i f i c a n t improvement over previous at tempts t o c o r r e l a t e noise- t r a n s i t i o n r e s u l t s . Apparently, f o r s u f f i c i e n t l y low dis turbance l e v e l s i n tunnels , t r a n s i t i o n r e s u l t s t h a t approach corresponding measured values i n f l i g h t without dis turbance effects can be obtained on simple shapes. On t h e o ther hand, very high noise l e v e l s i n tunnels probably dominate t r a n s i t i o n f o r any given shape.

Langley Research Center National Aerow u tics and Space Administration Hampton, VA 23665 February 22, 1978

REFERENCES

1. Potter, J. Leith: Boundary-Layer Transition on Supersonic Cones in an Aero- ballastic Range. AIAA J., vol. 13, no. 3, Mar. 1975, pp. 270-277.

2. Beckwith, I. E.: Development of a High Reynolds Number Quiet Tunnel for Transition Research. AIAA J., vol. 13, rio. 3, Mar. 1975, pp. 300-306.

3. Reshotko, Eli: Stability Theory as a Guide to the Evaluation of Transition Data. AIAA J., vol. 7, no. 6, June 1969, pp. 1086-1091.

4. Morkovin, Mark V.: Critical Evaluation of Transition From Laminar to Tur- bulent Shear Layers With Emphasis on Hypersonically Traveling Bodies. AFFDL-TR-68-149, U.S. Air Force, Mar. 1969. (Available from DDC as AD 686 178.)

5. Morkovin, Mark V.: Instability Transition to Turbulence and Predictability. IIT Paper presented at the Fluid Dynamics Panel Symposium (Lyngby, Denmark), May 2-4, 1977.

6. Stainback, P. C. ; Fischer, M. C. ; and Wagner, R. D. : Effects of Wind-Tunnel Dtsturbances on Hypersonic Boundary-Layer Transition. Pts. I and 11. AIAA Paper No. 72- 181 , Jan. 1972.

7. Stainback, P. Calvin: Hypersonic Boundary-Layer Transition in the Presence of Wind-Tunnel Noise. AIAA J., vol. 9, no. 12, Dec. 1977, pp. 2475-2476.

8. Stainback, P. Calvin; Wagner, Richard D.; Owen, F. Kevin; and Horstman, Clifford C.: Experimental Studies of Hypersonic Boundary-Layer Transition and Effects of Wind-Tunnel Disturbances. NASA TN D-7453, 1974.

9. Fischer , M. C. ; and Wagner, R. D. : Transition and Hot-Wire Measurements in Hypersonic Helium Flow. AIAA J., vol. 10, no. 10, Oct. 1972, pp. 1326-1332.

10. Laufer, John: Factors Affecting Transition Reynolds Numbers on Models in Supersonic Wind Tunnels. J. Aeronaut. Sci., vol. 21, no. 7, July 1954, PP. 497-498.

11. Laufer, John; and Marte, Jack E.: Results and a Critical Discussion of Transition-Reynolds-Number Measurements on Insulated Cones and Flat Plates in Supersonic Wind Tunnels. Report No. 20-96 (Contract No. DA-04-495-0rd 18), Jet Propul. Lab. , California Inst. Technol. , Nov. 30, 1955.

12. Wagner, R. D., Jr.; Maddalon, D. V.; and Weinstein, L. M.: Influence of Measured Freestream Disturbances on Hypersonic Boundary-Layer Transition. AIAA J., vol. 8, no. 9, Sept. 1970, pp. 1664-1670.

13. Pate, S. R.; and Schueler, C. J.: Radiated Aerodynamic Noise Effects on Boundary-Layer Transition in Supersonic and Hypersonic Wind Tunnels. AIAA J., vol. 7, no. 3, Mar. 1969, pp. 450-457.

14. Laufer, John: Some S t a t i s t i c a l P r o p e r t i e s o f the Pressure F i e l d Radiated by a Turbulent Boundary Layer. Phys. F l u i d s , vo l . 7 , no. 8 , Aug. 1964, pp. 1191-1197.

15. Laderman, A . J.: Review of Wind-Tunnel Frees t ream Pressure F l u c t u a t i o n s . A I A A J . , vo l . 15, no. 4 , Apr. 1977, pp. 605-608.

16. Bergstrom, E. R . ; and Raghunathan, S.: Aerodynamic Noise and Boundary- Layer T r a n s i t i o n Measurements i n Supersonic T e s t F a c i l i t i e s . A I A A J., vo l . 10, no. 11, Nov. 1972, p ~ . 1531-1532.

17. Dougherty, N. S . , Jr.: Cor re la t ion o f T r a n s i t i o n Reynolds Number With Aero- dynamic Noise Levels i n a Wind Tunnel a t Mach Numbers 2.0-3.0. A I A A J . , vo l . l j , no. 12, Dec. 1975, pp. 1670-1671.

18. Dougherty, N. S . , Jr.: Prepared Comment on t h e Cone T r a n s i t i o n Reynolds Number Data C o r r e l a t i o n Study. Flight/Ground Test ing F a c i l i t i e s Corre la- t i o n , AGARD-CP-187, 1976, pp. 3A-1 - 3A-7.

19. Sta inback, P. C. ; and Rainey, R . A . : Corre la t ion of Freestream Pressure Dis turbances i n Supersonic Wind Tunnels. A I A A J., vo l . 14, no. 2 , Feb. 1976, pp. 286-288.

20. Whit f ie ld , Jack D . ; and Dougherty, N . Sam, Jr.: A Purvey o f T r a n s i t i o n Research a t AEDC. Laminar-Turbulent T r a n s i t i o n , AGb3D-CPP-224, 1977, pp. 25-1 - 25-20.

21. Harvey, William D. ; and Hunter, W i l l i a m W. , Jr. : Experimental Study o f a Free Turbtllent Shear Flow a t Mach 19 With Electron-Beam and Conventiorial Probes. NASA TN D-798 1, 1975.

22. Harvey, William D.; Bushnel l , Dennis M . ; and Beckwith, Ivan E. : Fluctu- atlng P r o p e r t i e s o f Turbulent boundary Layers f o r Mach Numbers up t o 9 . NASA TN D-5496, 1969.

23. Kemp, Joseph H . ; and Owen, F. Kevin: Experimental Study o f Nozzle Wall Boundary Layers a t Mach Numbers 20 t o 47. NASA TN D-6965, 1972.

24. Beckwith, I . E.; Anders, J. B.; S ta inback , P. C . ; Harvey, W. D . ; and Srokowski, A. J.: Progress i n t h e Development o f a Mach 5 Q u i e t Tucnel. Laminar-Turbulent T r a n s i t i o n , AGARD-CPP-224, 1977, pp. 28-1 - 28-14.

25. Ladermar, A . J .; and Demetriadea, A . : Mean and F l u c t - ~ a t i n g Flow Measure- ments i n t h e Hypersonic Boundary Layer Over a Cooled Wall. J. F lu id Mech., v o l . 63, p t . 1, Mar. 18, 1974, pp. 121-144.

26. Donaldson, J. C. ; and Wallace, J . P. : Flow F luc tua t ion Measurements a t Mach Number 4 i n t h e Test S e c t i m o f t h e 12-Inch Supersonic Tunnel. AEDC-TR-71-143, U.S. Air Force, A u g . 1971. (Ava i l ab le from DDC a s A D 728 630. )

27. !bu@erty, It. S., Jr.; and S t e i n l e , Frank U., Jr.: Trans i t ion Reynolds NUB- bsr Corparisoas i n Several hj3r P ~ O I I ~ C ¶ ' u~e l s . AIAA Paper No. 74627, Ju ly !974.

28. Dodds, J u l a B. : Jr . ; and Hanley, Richard D. : Evaluation o f Transonic and Supersonic Wind-Tunnel Background Noise and Effects o f Surface Pressure Fluctuat ion &mu.rements. A U A Paper No. 72-1004, Sept. 1972.

29. u, 8.; and h v i g l i o , J.: Sane Heasurements i n the Trans i t iona l Super- sonic Wake of a Transverse Cirzulap Cylinder Uith h p h a s i s on t h e Ef fec t af External Noise. Iaminar-Tu-bulent Trans i t ion , AGARD-CPP-224, 1977, pp. 15-1 - 15-11.

30. Hatting, Fred b .; Cham, Dean R.; Nyholm, Jack R.; and Thoeras, Andrew G.: Turbulent Sb-rl Fr i c t i on a t High Mach Numbers and Reynolds Numbers i n A i r and Helim. M A TR R-82, 1961.

31. L i l ley , G. H.: Uall Pressure Fluctuat ions Under Turbulent Boundary Layers a: Subsonic and Supersonic Speeds. CoA Note 140, College of Aetlonaut., Craufield (England), Mar. 1963.

32. Kistler, A. L.; and Chen, U. S.: The Fluc t lu t i ng Pressure F i e ld i n a Super- m u i c Turbulent Boundary Layer. J. Fluid Plech., vol. 16, p t . 1, Hay 1963, pp. 41-64.

33- Bull , tl. K. : Proper t ies of t h e Fluctuat ing Mall Pressure Field o f a Turbu- l e n t Boundary Layer. AGARD Rep. 455, Apr. 1963.

34. Y i l h r t h , U. 21.; and Roos, F. Y.: Resolution and S t ruc tu re of t h e Mall Pressure F ie ld Beneath a Turbulent Boundary Layer. J. Fluid Nech., vol. 22, p t . 1, Hay 1965, pp. 81-94.

35. Harrison, Hark: Pressum Fluc tua t ions on the Wall Adjacent t o a T u r b ~ l e n t Boundary Layer. Rep. 1260, David Taylor Hodel Basin, Navy Dept . , Dec. 1958.

36. Fischer, M. C.; Haddalon, D. V.; Meinstein, L. N.; and Wagner, R. D., Jr.: Boundary-Layer P i t o t and Hot-Mire Surveys a t Q 20. AIAA Paper 70746, June-July : 970.

37. Felcher, Pe te r H.: Predict ions of Boundary-Layer-Turbulence Spectra and Correlat ions f o r Supersonic F l igh t . Se ~ o n g r k In t e rna t iona l dlAcoustique, vol. I ~ , h i e 1 E. CoPasins, ed., Congr. I n t . Acoustique, 1965, pp. [~54] 1-4.

38. Lowson, M. V. : Predict ion of Boundary Layer Pressure Fluctuat ions. AFFDL-TR-67-167, U.S. Air Force, Apr. 1968. (Available from DDC a s AD 832 715.)

39. Speaker, b. V. ; and Ailman, C. N. : S p e c t r ~ and Space-Time Correlat ions of the Fluctuat ing Pressures a t a Wall Beneath a Supersonic Turbulent Bound- a ry Layer Perturbed by Steps and Shock Waves. NASA CR-486, 1966.

40. Raman, K. R . : Surface Pressure F l u c t u a t i o n s i n Hypersonic Turbulent Bound- a r y Layers. NASA CR-2386, 1974.

41. Chen, W i l l i a m S . : Measurements of Aerodynamic Noise on a F l a t P l a t e i n Supersonic Flow. Aeronaut. Eng. Thes i s , C a l i f a r n i a I n s t . Technol., 1961.

42. Maest re l lo , L.; Monteith. J. H., Manning, J. C. ; and Smitn. D. L.: Measured Response o f a Complex S t r u c t u r e t o Supf?rsonic Turbulent Boundziry Layers. AIAA Paper No. 76-83, Jan. 1976.

43. MacCready, Paul B., Jr.: Turbulence Measurements by S a i l p l a n e . J. Geophys. Res., vol . 67, na. 3, Mar. 1962, pp. 1041-1050.

44. Pfenninger, Werner; and Reed, Ver l in D.: Laminar-Flow Research and k p e r i - ments. Astronaut. & Aeronaut., J u l y 1966, pp. 44-50.

45. P o t t e r , J. Le i th ; and Whi t f i e ld , Jack D.: Boundary-Layer T r a n s i t i o n Under Hypersonic Condit ions. AEDC-TR-65-99, U.S. A i r Force, Hay 1965.

46. Wright, Robert L.; and Zoby, Ernes t V.: F l i g h t Boundary Layer T r a n s i t i o n Hea~urements on a Slender Cone at Kach 20. A I A A Paper 77-719, June 1977.

47. Pa te , S. R.: Comparison o f NASA Helium Tunnel T r a n s i t i o n Data Yith Noise- T r a n s i t i o n Cor re la t ion . AIAA J., vo l . 12, no. 11, Nov. 1974, p. 1615.

48. Mack, L e s l i e M.: Linear S t a b i l i t y Theory and t h e Problem o f Supersonic Boundary-Layer Trans i t ion . A I A A J. , v o l . 13, no. 3, Mar. 1975, pp. 278-289.

49. Mack, L e s l i e M.: A Numerical Method f o r th,? P-diction o f High-Speed Boundary-Layer T r a n s i t i o n Using Linear Theory. Aerodynamic Analysis Requiring Advanced Computers - P a r t I , SASA SP-347, 1975, pp. 101-;20.

50. Mack, L e s l i e M.: On t h e Appl icat ion o f Linear S t a b i l i t y Theory t o t h e Problem o f Supersonic Boundary-Layer Trans i t ion . A I A A Paper No. 74-134, Jan. -Feb. 1974.

51. ~ e m e t r i a d e s , Anthony: Hypersonic Viscous Flow Over a Slender Cone. P a r t 111: Laminar I n s t a b i l i t y and T r a n s i t i o n . A I A A Paper No. 74-535, June 1974.

52. Kendall , J. M.: Wind Tunnel Experiments Rela t ing t o Supersonic and Hyper- s o n i c Boundary-Layer T r a n s i t i o n . A I A A J., vo l . 13, no. 3, Mar. 1975, pp. 290-299-

53. Pa te , S. R. : Measurements and C o r r e l a t i o n s of T r a n s i t i o n Reynolds Numbers on Sharp Slender Cones a t High Speeds. A i A A J . , vo l . 9 , no. 6 , June 1971, pp. 1082-1090.

54. Schubauer, G. B.; and Skramstad, H. K . : Laminar-Boundary-Layer O s c i l l a t i o n s and Trans i t ion on a F l a t P l a t e . NACA Rep. 909, 1948.

55. Uells, C. S i n c l a i r , J r . : E f f ec t s of F r e e s t m a Turbulence on Boundary-Layer Transi t ion. AIM J., vol. 5, no. 1 , Jan. 1967, pp. 172-174.

56. Boltz, Frederidr U. ; Kenyon, George C. ; and Allen, Clyde 0. : The Boundary- Layer Trans i t ion Cha rac t e r i s t i c s o f Two Bodies o f Revolution, a F l a t P l a t e , and an Unswept Uit* i n a Low-Turbulence Wind Tunnel. NASA TN D-309, 1 960.

57. Dryden, Hugh L.: A i r Flcw i n t he Boundary Layer N e a r a P la te . NACA Rep. 562, 1936.

9. Hall, A. A. ; and Hislop, G. S. : Experiments on t h e Transiticm o f t he Lami- nar Boundary Layer on a F l a t P la te . R. 6 #. No. 1843, B r i t i s h A.R.C., 1938.

59. Spar.&ler, J. G. ; and Uells, C. S., Jr. : Ef fec t s o f Freestream Disturebances on Boundary-Layer Transi t ion. A I A A J., vol. 6, no. 3, Mar. 1968, pp. 543-545.

60. Norkovin, Mark V.: On Trans i t ion Experiments at Moderate Supersonic Speeds. J. Aeronaut. Scl . , vol. 24, no. 7 , Ju ly 1957, pp. 480-486.

61. Coles, Donald: Measurements o f Turbulent F r i c t i o n on a Smooth F l a t P l a t e i n Supersonic Flow. J. Aeronaut. Sci., vol. 21, no. 7, July 1954, pp. 4333%.

ORIGINAL PAGE IS or POOR Q U ~

DEFINITIONS OF FACILITY ABBREVIATIONS USED IN TABLES I TO I11

ARC 11' TUT

ARC 14' TUT

ARC n-50

ARC 9 x 7' SUT

ARC 3.5' H W

ARC 12' PUT

AUT(4T)

CAI, 48" HST

CSUT

GDFTA

GDFTB

GDFTD

IMST(S7i or ( S 8 )

JPL 20" SKT

JPL 21n HKT

LaRC 4 ' SPT

LaRC HeT

LaRC HRNT

LaRC HNT

LaRC 6

LaRC 16'TT

Ames 11-Foot Transonic Wind Tunnel

Ames 14-Foot Transonic Uind Tunnel

Ames M-50 helium tunnel

Ames 9- by 7-Foot Supersonic Wind Tunnel

Ames 3.5-Foot Hypersonic Uind Tunnel

Ames 12-Foot Pressure Wind Tunnel

AEDC Aerodynamic Wind Tunnel (4T)

Calspan 48-inch leg of the hypersonic shock tunnel (form- erly Cornell Aeronautical Laboratory 48-inch hypersonic shock tunnel)

The Johns Hopkins University continuous supersonic wind tunne 1

AEDC von K Q d n Gas Dynamics Facility Tunnel A

AEDC von Karkn Gas Dynamics Facility Tunnel B

AEDC von K a h n Gas Dynamics Facility Tunnel D

low turbulence wind tunnel of the Institut de Mbcanique Stati.tique de la Turbulence (asymmetric nozzle or sym- metric nozzle)

JPL 20-Inch Supersonic Wisd Tunnel

JPL 21-Inch Hypersonic Wind Tunnel

Langley 4-foot supersonic pressure tunnel

1.52-meter (60-inch) diamcter test section of the Lang- ley high Reynolds number helium tunnel

Mach 6 high Reynolds number tunnel at the Langley Research Center

Langley hypersonic nitrogen tunnel

Mach 5 pilot quiet tunnel at the Langley Research Center

Langley 16-foot transonic tunnel

LaRC 2OW

LgRC 22'

LaRC M ( 1 ) or ( 2 )

LaRC VD

LW BI,C

L r n

m 41' rn 2

PwT( 16s)

U G E K

Langley 20-inch Mach 6 tunnel

22-inch aerodynamic leg of Langley hypersonic helium tunnel f a c i l i t y

Langley Unitary Plan wind tunnel ( t e s t section 1 or 2)

Langley Hach 8 variable-densi t y hypersonic tunnel

L'F11 Research Center Boundary Layer Channel

Laughborough University of Technology Gun 'Punnel

MBS 4:-foot wind turmel 2

AEDC ~ m p u i s i o a Wind Tunnel ( 16s)

AEDC Range It

TABLE 1.- DATA SOURCE FOR FIOURE 1

-- - -

-- -. 3 F ~ ~ i l i t y Source per meter per inch -- - -- -- -- . - -- -~ .. - - . . . .- -. . -. - - -

0.09 x 106 JPL 20" SWT Laufer ( re f . 14)

. - . ~. .34

----- - 0.47 1.4 ARC 3.5 HUT Stainbaok, Wagner, Owen,

11.46 -29 1 and Horrtamn ( re f . 8 ) -~ - . .- - .- - -. - ., -- -- ~ .. -. -

0.39 1.4 :.62 x 106 0.092 x lo6 CAL 48* HST Harvey, Bushnell, and 5.00 .I27 Beokwith ( r e f . 22) 6.02 ,153 6.97 ,177

- .- --a --- 1.66 x lo6 0.0422 x lo6

32.8 ,833

0.574 1.67 4.22 x 106 ---~--- -,.. - - .

0.107 x 106 --.- .-

Hot wire -0.85 1.4 15.0 x 106 0.381 x 106 La1.C M5 Bookwith, Anders, Stalnback. Harvey, and Srokowskl ( r e f . 24)

-----

-0.85 1.4 ( 1 1 . 7 t o 5 1 . 2 ) ~ 1 0 6 ( 0 . 2 9 7 t o 1 . 3 ) ~ 106 La6CM5 Unpublished (21.2 t o 50.5) (0.538 t o 1.28) (19.7 to 48.0) (0.50 t o 1.22)

0.267 x '106 LUTOT Bargatrun and Raghumthan ,306 ( r e f . 16) .366

--.- ,413

1 .O 1.67 7.09 x lo6 0.180 x lo6 LaRC 22n Fisoher and Wagner ,430 ( r e f . 9) .780

aConverted rma pitot to rms a ta t io pressure. b~onverted rma density to ram s t a t i c pressure. OM, a t acoustic origin.

TABLE I.- Continued

- ~... -

Type of

I coustic, 100 cone

--

2 l~coustic, 2 lo0 cone 2.5 1 2.5

Acoustic , 100 cone

IZLiG- 1 100 cone

Acoustic, flat plate

-- . - . - - - -. - . - --

%,

per inch

15.75 x lo6 0.4 x 106 1.6

1 .O 13.39 x lo6 0.1, x 106 9.06 I .23

1 1 23.62 .60 19.29 .49 48.0'3 1.22

1 .O 4.92 x 106 0.125 x 106 3.94 .10 4.92 .I25 1 .?.76 .07 3.94 .10 1.97 .05

to 0.25) x 1 (6.57 to 13.11) (0.167 to 0.333) (6.57 to 13.11) (0.167 to 0.333) (6.57 to 13.11) (0.167 to 0.333) (9.84 to 13.11) (0.25 to 0.333) (6.57 to 13.11) (0.167 to 0.333)

(0.167 to 0.25) --- -

1 .O 0.216 x lo6 13.11 ,333

. . -

Facility

- . - . - - . - - - . - LaRC HRNT

---- LaRC 22*

AWT(4T) (porous walls)

-- ARC 14' m (slots)

. - - - . . -

ARC 11" TWI

--

Source I -

Beckwith (ref. 2) t I

Dougherty (ref. 17) 1

Dougherty and Steinle 1 (ref. 27)

Dougherty and Steinle (ref. 27)

.- - . -- - - - Dodds and Hanley (ref. 28)

TABLE 11.- DATA SOURCE FOR FIGURE 2

Symbol

w 4

v

-i .---

Y 1-

Location of b/tt y Facili ty I I I 1 I

Hot wire Free stream ---- 1.4 LUTGT

I Hot wire Free stream 1.0

I

Acoustic, Free stream 160 cone ,

Hot wire 1 ~ r e e stream [ 1 .0 I I .4 1 GDFTD

Acoustic, Free stream 0.62 ' 1 .4 LaRC HRUT 100 cone

Acoustic, Free stream 0.62 1 .4 LaRC 20" 100 cone

100 cone

I Acoustic, Free stream 0.47 1 .4 ARC 3.5 HUT

Acoust?.~, Free stream 0.40 1.4 LaRC VD 160 cone

--- ;

Hot wire Free stream 0.75 1 . 4 GDFTB

Hot wire F r ~ e stream 1 .4 ARC 3.5 HUT I Hot wire Free stream 0.40 1.4 JPL 21" HWT -ii

Source

Bergstrom and Raghunathan ( re f . 16)

----- Stainback, Fischer, and

Wagner ( re f . 6 ) ~ .--.

Stainback, Fischer, and Wagner ( r e f . 6 )

-- Donaldson and Wallace

( re f . 26)

Stainback ( re f . 7)

-

Stainback ( r e f . 7 )

Stainback, Wagner, Owen, and Horstman ( re f . 8)

Statnback ( re f . 7)

Laderman ( re f . 15)

Stainback, Wagner, Owen, and Horstman (ref . 8)

Laderman and Demetriades ( re f . 251

Stainback, Fischer, and Wagner (ref . 6)

TABLE 11.- Concluded

~ +. - - -- 4 - t--- +- -+ - -

-t 1.0 1-3, 0.75 1 Acoustic Wall , ; 1 . 4 1 ----------- IWlllrnarth and Roos ( r e f . 3 - 4

Symbol

= O . 15 1.0 i 1 . 4 ----------- Harrison ( r e f . 35)

~ - -. .. . . . - - - - . - -- . -- . -- - - --

Mm Faci l i ty Source

-- -- - - - . - 1.6 t o 5 1 Hot wire Free stream 1 .O

-- - -- - ~-

- - ~ I ~ c o u s t i c j wail 0.3, 0.5 1 . O

-'--- - -

LaRC 22" Fischer, Madealon, Weinstein,

I ! I and Wagner ( r e f . 1 36) ~ -- . - ~~. ... -

I m ( . 5 t o 5 1 Acoustic t Wall ~ - - ~ ~ ~ ~ t ; ~ ~ - ~ l s t l e r and Chen ( r e f . 321 7 1 0.95 t o 1.9 1 Acoustic

0 1 0.35 t o 0.78 I Acoustic

a,.l 8.24 t o 8.89 Acoustic

Wall 37) ------------ i ~ l r Flight [Belcher ( r e f . -. - - - - - - - . - - - -

1

wall I 1 .0 11.4 I ----------- l~owson ( r e f . 38) I Wall, f ree 0.09 t o 0.18 1.4 CAL 48" HST Harvey, Bushnell, and

stream I / I Bectuith ( r e f . 22)

/ 6 1 18.72, 19.76 1 Electron beam

Flat p la te ------------ Flat p la te 1 .O

- Free stream 0.574 -- Free stream 0.18

Wall i - - I .o

----------- GDFTA -

ARC M-50

LaRC HNT

- -- -

Chen ( r e f . 41,

Pate and Schueler ( r e f . 13) -- - - . - - - - . . - - . ~ - . - - -. -

Kemp and Owen ( r e f . 23 1 ---- .- A

Harvey and

Maestrello, Monteith, Manning, and Smith ( r e f . 42) I

TABLE 111 .- DATA SOURCE FOR FIGURES 3, 4 , AND 5

Swlrce Symbol R l , tr measurement

- Free-stream noise

measurement

A

a 0

0 0

0

v

0 A

-b D - 0 I

0 0

Mm

Heat t ransfer

Heat t ransfer

Heat t ransfer

Heat t ransfer

Heat t ransfer

Heat t ransfer

Heat transfer

Acoustic

Heat t ransfer

Heat transfer

Heat t ransfer

Heat t ransfer

Acoustic

Acoustic

Photo

Heat transfer

Acoustic

-21

-18

8

7.32

2.2, 4.75

2.5, 4.5, 1.8

1.3 t o 3.7

1.6, 2

6

8

6

3 , 5

1.6, 2, 2.86

2.9, 3.5, 4.6

2 . 3 , 5

2.5, 2.75

0.6, 1.0, 1.3

Y

Hot wire

Hot wire

Hot wire

Hot wire

Hot wire

Hot wire

Hot wire

lrcoust i c

Acoustic

Acoustic

Aooust l c

Acoust i c

Acoustic

Acoustic

Acoustic

Acoustic

Acoust Ic

1.67

1.67

1.4

1.4

1.4

1.4

1.4

1.4

1.4

1.4

1.4

1.4

1.4

1.4

1.4

1.4

1.4

t w / t t

- (Fischer and Wagner

( re f . 9)

Pischer and Wagner ( rdf . 9)

Dernetriades ( r e f . 51)

Stainback, Fischer, and Wagner ( r e f . 6) -

Kendall ( r e f . 52)

Laufer ( r e f . 10)

Laufer and Marte ( re f . 11 )

Oougherty ( re f . 1 8 )


Stainback ( r e f . 7 )


Pate i:.ef. 53)

Dougherty ( r e f . 18)

Dougherty ( re f . 18)

Potter ( r e f . 1 )



1.0

. - 1.0

0.75

0.47

1.0

1.0

1 .O

-1.0

0.62

0.40

0.62

-1.0

-1.0

Mode 1 Faci l i ty

2.870, 160, sharp cone

1.870, 50, 100, sha:p cone

5O sharp cone

16O sharp cone

lo0 sharp cone

lo0 sharp cone

5O sharp cone

100sharpcone

100 sharp cone

160 sharp cone

16.750 sharp cone

5O sharp cone

lo0 sharp cone

IARC 22"

------ LaRC HeT

GDFTB

ARC 3.5 HWT

JPL 20" SWT

JPL 20" SWT

JPL 20" SWT

LaRC4'SPT

LaRC HRNT -

LaRC VD

LaRC 20"

GDFTA

LaRC 4' SPT

LaRC 4 ' SPT

RangeK

PWT(16S)

AWT(4T)

-1.0

0.52,0.19

-1.0

-1 .O

lo0 sharp cone

bO, 1O0sharp cone

lo0 sharp cone

lo0 sharp cone

TABLE 111.- Concluded

- -- t w / t t Model F a c i l i t y R l , t r Free-stream noise

measurement measurement - -

Source

0.46 t o 1.0 lo0 sharp cone F l igh t test Heat t r a n s f e r I I Wright and Zoby I ( ~ e e n t r y PI I ( r e f . 46)

-- Tunnel wall Hot wire

Schubauer and Skrars tad ( r e f . 54)

Wells ( r e f . 55)

Sharp f l a t

-

1 Sharp f l a t

I p l a t e

Sharp f l a t T p l a t e

ARC 12' PWT Heat t r a n s f e r Hot wire Boltz, Kenyon, and Allen ( r e f . 5 6 )

----------- Heat t r a n s f e r Hot wire Dryden ( r e f . 57)

----------- Heat t r a n s f e r Hot wire Hall and Hislop ( r e f . 58)

=I .O Tunnel wall / LTV BLC l ~ e a t t r a n s f e r Hot wire Spangler and Wells

Wagner, Maddalon, and Weinstein ( r e f . 12)

Morkovln ( r e f . 60)

1 a 1 =18 t o 22 1.67

I I 0 I Z t 0 4 . 5 5 11.4 1 I

Shatsp f l a t JPL 20" SWT Heat t r a n s f e r Hot wire p l a t e

Cylinder f l a t p l a t e

0.68 t o 1.0

1 .O

-- - - - - - -

Laufer and Marte ( r e f . 11)

- Coles ( rep . 61)

Sharp f l a t p l a t e

Sharp f l a t p l a t e

Pa te and Schueler ( r e f . 13)

6.95 11.4 / 1.0 I Sharp f l a t / LUTT I Heat t r a n s f e r Hot wire Bergstrom and Ragi:unathan p l a t e ( r e f . 16)

LaRC 22.

CSKf

Unpublished I -- ~ - - ~ 1-

Heat t r a n s f e r

Pressure

Hot wire

Hot wir.?

10 - - 8 - - 6 - - 0 4 - -

2 -

1 - - .8 - - .6 - - - i

PC0 .4 - See table I for data -

percent source and test

conditions

Effect of diffdser

- Wall-hole resdnance

1

(d) Free-stream rm$ pressure f l uc tua t ions divided by mean s t a t i c pressure.

F i g w e 1.- T pica1 free-stream pressure f l uc tua t ions i n wind tunnels . 1.65 x 10: 1 R j n 5 55.2 x 106; 0.09 5 b/tt < 1.0. PLagged sp- bols represent da ta f ram acous t ic neusurements; unf lagged symbols i nd i ca t e da ta from hot-wire measurements; s o l i d symbols i nd i ca t e da ta from porous o r s lot ted-wal l tunnels ; double f l a g i nd i ca t e s ms densi ty .

ORIGINAL PAGE IS 25 OF POOR QU-

See table I for data source and test conditions

2

1

-8

.6

-4

.2

q 0 . I

percent .08

-06

-04

-02

-01

(b) Fr :e-stream rrps pressure f luc tuat ions divided by dynamic pressure.

- -

- - - - - - - - -

- - - - - - - - -

,

Figure 1 . - Concluded.

See table I1 for

-- Quation (1) data source anal

20 test conditions

- - - - Data hiring

10 8

6

4

2

T w Adiabatic (ref. 31)

1

-8

-6

-4 semiempirical (ref. 16)

.2

.1 .1 .2 -4 .6 .8 1 2 4 6 8 1 0 20 40 60

M,

Figure 2.- Comparison of rms wall and free-stream pressure f luc tuat ions . 0.09 5 t,,/tt $ 1.0. S o l i d symbols ind icate free stream; open sym- b o l s ind icate tunnel wal l ; f lagged symbols i n d i c a t e a c o u s t i c measurements; unflagged symbols i n d i c a t e hot wire.

ORIGINAL PAGE 1s OF POOR QUO

See table II]: for data saurce and test conditions

Figure 3.- Transi t ion Reynolds number on sharp cones for se l ,eral free-stream disturbance l e v e l s . a = 00; 0.1 =< t,/tt =< 1.0.

- j 4 Me 5 cone,

free flight (ref. 46)

Aeroballistic range cones,

2 . M, - 4.3 (ref. 1)

See table 111 for data source and test conditions

Figure 4 . - Influence of free-stream disturbance l e v e l on t r a n s i t i o n f o r cones , cy l inders , and f l a t p l a t e s . a = OO; 0 5 M, 2 4; 0 . 4 2 &Itt 2 1.0. Unflagged symbols ind icate cones; flagged symbols ind icate f l a t p l a t e s or

cy l inders ; s o l i d symbols ind icate @be.

t flight (ref.

Sharo cones

See table HI for data source and test conditions

D, / (unflagged symbols)

Sharp flat plates (flagged symbols)

fairing

Figure 5.- I n f l u e n c e o f free-stream d i s t u r b a n c e l e v e l on t r a n s i t i o n for cones and f l a t p l a t e s . a = 00; 4 1 Y,: 23; 0.3 2 t+/tt 1.0.

I

7 Author(sl 8 Pwtwm~nq Organuat~on Report No I

. 1 Report No. 2 Gornnrnmt Accasr~on No.

NASA Tt478635 4 T~t le and SuMltle INFLUENCE OF FREE-STREAM DISTURBANCES ON BOUNDARY-LAYER TRANSITION

3 Recop~nt's Catalog No

5 R e p m Date

April 1978 6 ~crtotrn~nq Orpan~zatum Code

William D. Harvey ,

9 Rnform~ng Orpm~za tm Name and Ad&-

NASA Langley Research Center Hampton, VA 23665

IS Supplmmtav Notes

L- 1 1905 10 b%wh U n ~ t No 505-06-4 1-0 1

1 1 Gntract or Grant Ma

I

12 Swnurmg Agency Name and AJdrnr National Aeronautics and Space Administration Washington, DC 20546

I Considerable experimental evidence exists which shows that free-stream disturbances (the ratio of root-mean-square pressure fluctuations to mean values) in con-

13 TI* o t . R r p n ma P o w d Cohered Technical Memorandum

14 S~ut'rv.rq Agency Code

I ventional wind tunnels increase with i~creasing Mach number at low supersonic to moderate hypersonic speeds. In addition to local conditions. the free-stream dis-

I turbance level influences transition behavior on simple test models. E3sed on this observation, existing noise-transition data obtained in the same test facility have been correlated for a large number of reference sharp cones and flat plates and are shown to collapse along a single curve. This result is a significant Improvement over previous attempts to correlate noise-transition data.

Transition Stream disturbances Noise Boundary layer

I9 k u r l t y a a u ~ f !of rh~s f e w r r 1 20 % , ~ r t r , (:,JSSI~ ( u l th(% wqe I

2 1 "40 at P-r 2 2 &,<*.

Unc lass i f i ed Unc lass i f i ed ! 30 , $4.50 - - I

search.jsp?r=19780012517 2018-05-06t13:01:13+00:00z · in conventional wind tunnels increase uith...

Documents