)

II"'" U{.D; /1/ II. r.

ADVANCES IN HYPERSONIC EXTRAPOLATION CAPABILITY

WIND TUNNEL TO FLIGHT

J. A. Penland and D. J. Romeo

In the past a major problem has been the extrapolation of conven-

tional wind tunnel data to full-scale flight conditions. This is due

primarily to the limited Reynolds N~~ber capability of hypersonic wind

tunnels. The nature of the problem presented by this limitation is

shown in the first slide.

Shown are the relatively low Reynolds numbers

available on realistic configurations during the mid-

sixties.

The boundary layers over small subscale test

models were mostly laminar and transitional with a

wide untested region up to flight Reynolds numbers.

Over full-scale aircraft (having lengths of 100 m.

or' more), a turbulent boundary layer will cover virtually

the whole aircraft.

This difference in wind tunnel and flight Reynolds number has

always existed at lower speeds, but the problem has been overcome by

--(]J

attaching small roughness elements on the model to artificially produce

a turbulent boundary layer. At hypersonic speeds, however, usable

bounlury layer trips have not been developed.

Because of the inability to develop a turbulent boundary layer

over test models at hypersonic speeds, the viscous effects on skin

PEN00068

-rJ

friction could not be determined, nor could research be performed to

ascertain if there were any variations of these other important perform­

ance parameters with Reynolds number --- variations in other speed ranges,

in the absence of separation, are not severe. As a result, reliable extra­

polation of full-scale aerodynamics from Yind tunnel results over this

wide range of Reynolds number could not be made.

In the meantime, several developments have occurred to increase

the range of available Reynolds numbers. These include modifications to

existing conventional tunnels and improvement in shock tunnel capabili­

ties. In particular, the Cornell Aeronautical Laboratory Hypersonic

Shock Tunnel has been extremely useful in studying high Reynolds number

effects.

_____.1-,\ On the next slide are shown the test conditions of this facility at ~

a M,..,8.

The Reynolds number range varies from less than 0.1 million to ful1-

scale flight Reynolds numbers of 160 million ---

With corresponding range of boundary layers from fully laminar to

nearly full turbulent without forced transition.

This wide range of Reynolds numbers at a nearly constant Mach

nlmIDer is due to the many controlable variables of the shock tunnel,

the drive gas mixture, its temperature and pressure; the initial test

gas pre3sure, and the resulting stagnation pressure and temperature as

well as the size of the throat of the contoured nozzle. Fast response

2

) instrumentation has been developed to measure all necessaryl 3-tl ;md7es

for force balance tests during test times of 4 to 15 milliseconds. Mach

8 was selected for the present tests because the shock tunnel reached its

highest Reynolds number at this speed, using peak values of stagnation

pressure and low stagnation temperatures just high enough to avoid liqui-

faction.

The present test range varied from 0.5 to 160 million on the test

configuration shown on the bottom of the slide. M~7.5 to 8.1. ~

Configuration representative of~current hyper-

sonic transport concept having a length of about 100

meters.

Solid lines indicate the configuration tested.

Vertical tail and propulsion system were eli-

minated to reduce zero lift-drag, and make results

more sensitive to Reynolds number.

The model was 2/3 meters long or 1/150 scale

of full-size.

The lift-to-drag ratio, as a function of Reynolds number is shown

on the next slide.

Note, shown is LID at a constant 0:: = 3°,

not (Lin) max.

This is 0:: for (L/n) at the highest test m

Reynolds number and the model was held constant for

the lower Reynolds number tests.

3

Predictions for all laminar boundary layer

using TI and all turbulent boundary layer using

Spalding-Chi ~OUftdary leye~are shown.

No boundary layer studies were made on this

configuration, but all indications are that the

boundary layer is predominately laminar at the lower

Reynolds numbers, and predominately turbulent at

Reynolds numbers above 15 to 20 million.

The question then is what performance parameters have major effects

on wind tunnel data extrapolations.

First let's consider the normal force which is closely related to

the lift-force for low-drag configurations at low angles of attack.

This parameter could be sensitive to the nature of the boundary layer

through the viscous effects on lee-side flows and component interference,

and at the lower Reynolds numbers through the boundary layer displace-

ment effects.

The next slide shows the normal force

coefficient with angle of attack for various

Reynolds numbers over the test range.

Some viscous effects may exist at the lower

Reynolds numbers, but they are very small.

6 Above Reynolds numbers of 10 x 10 there are

no noticeable viscous effects at all.

4

This result is further borne out in the next slide where normal . ® force coefficient is shown at a constant angle of attack of 3° with

Reynolds number,

It is essentially independent of Reynolds

numbers.

This result implies that if viscous effects change the local flow

characteristics, then their overall effect is negligible for this con-

figuraticn, and the normal force par ameter may be neglected in the

extrapolation process.

Next, let's consider the axial force coefficient on the next slide. ~

This parameter is, of course, strongly affected by viscous effects on

skin friction and boundary layer displacement effects.

The large variation of CA with Reynolds

n~ber is shown for the various angle-of-attack

s.leeps that were traversed.

MOst of the change in CA is due to skin

friction changes as boundary layer displacement

effects would be confined to the very low Reynolds

numbers,

The variation of CA with R~olds number at a

constant ~ on the next slide is a more graphic

example of the effects of skin friction.

The skin friction then, as it has always been, is a strong factor

in extrapolations of wind tunnel results to flight Reynolds numbers.

The drag-due-to-lift is the last parameter that will be considered. OM ~7'-

This f actor would be most affected by viscous effects on both lee-side

flows and component interference. The variation of 0 Cr/dCL 2 with

Reynolds number .is-shown in- the nexb slida. (;)

Slight increase with Reynolds number same as

predicted trend which accounts for the small variation

in shock tunnel Mach number with Reynolds numbers.

For all practical purposes, Reynolds number has no effect on the

dr"ag-due-to-lift which indicates again, as with the normal force, that

the overall viscous effects are negligible for this configuratiGn.

Considerations of the drag-due-to-lift parameter in hypersonic wind

tunnel data extrapolation are therefore not required.

Of the parameters considered then, only the viscous effects on

skin friction need be accounted for in correcting subscale hypersonic

wind tunnel data to flight values.

6

J I;' It .

---®

I .

On the next slide, let us examine how accurately we can predict the

data at the highest Reynolds number by making these viscous corrections

to low Reynolds number data.

The correction simply am~unts to replacing the skin friction in

the data from the faired curve at lower Reynolds numbers with the skin

friction at the high Reynolds number.

Percent error in prediction of CL, CD and

LID at the highest test Reynolds number is shown

at the bottom. The CL as with the CN is

essentially independent of Reynolds number.

For curves shown, CF was assumed either all

laminar using TI or all turbulent using Spading-

Chi.

The predictions made under laminar boundary

layer conditions are poor, but above about 15

million, where turbulent boundary layers cover most

of the aircraft1s surface, the predictions are

satisfactory.

These results, of course, apply to the constant angle of attack

condition, and are not representative of the optimum, or (Lin) max

case.

7

The dete~nation of (L/D)max and the optimum values of CL and

CD requires that the axial force data be corrected to flight condi­

tions, that new lift and drag coefficients be calculated and plotted

in polar form for interpretation.

The accuracy with which the optimum case can be predicted is shown

in the next slide.

Shown are the errors in predicted flight Reynolds -----~------~--~--

number values of (L/D)max and optimum CL and CD'

dete~ned from the lower Reynolds number shock tunnel

data.

Using the 0.6 million tunnel data with a complete

laminar boundary layer the (LID) at R = 160 max e

million is well predicted, but the optimum values of

CL and CD are not.

These errors are partly due to the inability to accurately predict

skin friction on a complete configuration, particularly in the 1amin~

region and partly due to the critical nature of the determination of the

optimum CL and CD from faired data. By correcting data in the

Reynolds number range from 15 to 30 million where the boundary layer is

predominantly turbulent, adequate optimum performance values at the

high Reynolds numbers can be predicted.

8

) L

CAL IN CONCLUSION studies in the Cornell Shock Tunnel at Mach 8 of a """"""-----

representative hypersonic aircraft have shown that skin friction is

the most significant parameter affecting the extrapolation of subsca1e

Reynolds number data to flight conditions.

If wind tunnel data can be obtained at Reynolds numbers of 20

million or higher where the turbulent boundary layer predomimn tes,

simple turbulent skin friction corrections to the data will allow

good prediction of flight Reynolds number performance data.

The negligible variations of the lift and drag due lift factors

with Reynolds number indicates that overall viscous effects on both

lee-side flows and component interference are secondary for this

blended wing-body configuration. Studies will be conducted on other

types of configurations to determine magnitude of these effects.

Additional work is required to determine the viscous effects on

control effectiveness with Reynolds number. An accurate knowledge of

these par~~eters is required to predict reliable trimmed performance

at flight Reynolds numbers, and because of possible flow separation

over controls, viscous effects could be significant.

THANK YOU

9

6

(LJD)max 4

MODEL TESTS

LAM INAR-TRANS ITIONAL B.L. (TR I PS I NEFFECT IVE)

o o 0 0

FULL SCALE

TURBULENT B.L.

--~ ~ o

21- ----I~-'---'--...................... .....I......I.:--.L...---'----'----'---L---'--'--a....a...------'200 X 106 2 4 6 10 20 40 60 100

REYNOLDS NUMBER

PARAMETERS AFFECTING FLIGHT EXTRAPOLATIONS OF (L/D)max

FRICTION DRAG CD

PRESSURE DRAG C F Dp

LIFT CURVE SLOPE CL DRAG DUE TO LIFT C a

DL CONTROL EFFECTIVENESS hoC m hoC

L hoCD

1500

1000 Po atm 500

2000

T oK 0'

1000

oL_~~--~F=-']--lio)OO~1I[O(o08 x 106 1 1.0 10 0.01 O. REYNOLDS NUMBER

I /I

b: ----

94.2 m. FLI GHT

L/ D LAMINAR ....... __ ----...... SPALDING CHI C

4 ~OQ <fJp OA C>... t:::l F"""\ .(11--66

__ lJ.. __ -~s:f-f}--~-'l..-__ ---- TURBULENT

. REFERENCE lEMPERATURE C

F

1 3 6 10 REYNOLDS NUMBER

30 300 x 106

.12

.10

.08 eN

.06

.04

.02 / /

/

0

o

~ I

~/'--THEORY

~ ~ jf REYNOLDS NUMBER

~/ 0 153.2 x 106

o 32.31

~ o 10.41 fl. 4.02 ~ 1.53 0 .597

2 4 6 8 a. deg

, L.i .

3-' Sf .S-

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a = 30

.04 - --CB--@{)~O..1().~ Oc5ttm~-@-O~ -~--­\...THEORY

.6 1 3 6 10 30 60 100 300 x 106

REVNoLD5- NUMBrR

.016

.012 o O,.../"--LAMINAR

o // REF. TEMP. CF o "

CA

0 ...... "...... Rl = 0.597 x 106

0 - .... Q..g; ~ ~ / ~EyN9JDS NUMBER

~~~~ o / /

o 153.2· x 106 /

~a~~~ /

o 32.31 o 10.41

.004 _.B TURBULENT l::,. 4.02

-- SPALDING CHI CF ~ 1.53

RL = 153.2 x 106 0 .597

0 -2 0 4 6 8 10

a deg

. . 014

\

\ REFERENCE TEMPERATURE CF \~~

'<'§ 008 " 0"",-- TURBULENT C .• LAMINAR', ~o--_;Qm.CQ A An_

A '.... ° §b a ~~~ __ ~ a = 3

0

.006, .... ' ..... c:. ... ~ SPALDING CHI C

F

.010

INVI1.ClP _____ _ -- --------------

.6 6 JO. 30 1 3 REYNOLDS NUMBER

1.2: '----'---'--I. ............... """"-----i-----lI........l... .............. ....a....r...&...I....---..I-----II........I.......L.-L....r...I...L..L.....----'----'300 X 106 .3 ' .6, 1 3 6 10 30 60 100 REYNOLDS NUMBER

.010

CA

.008

a = 3° .006

.004, 111111 I I I I I II I I I "III

40 ACCURACY OF FULL SCALE EXTRAPOLATION AT CONSTANT ANGLE OF ATTACK

20 %

ERROR ASSUMED TURBULENT

O~~----------~--~----~~'------

-20 ,~SryMIEPI LAMI ~AR I I I I I II I

. 3 . 6 1 3. 6 10 30 I I I I II

60 100 I I ; 300 X l06

REYNOLDS NUMBER-

, ' . ., ~ . , .. . . .

(L/ D) 20

t ~ ~xoL-__ --____ ~==~~~~~t=-______ _

%ERROR ;p-$ --"'"' __ ... ___ :

-20

40 /ASSUMED LAMINAR '

CL 20 . ~ * ~ t ASSUMED TURBULENT OPT ~~~--------~====~~--T-------~-­% ERROR 0

-20

40 1 C D

OPT 20

%ERROROr----------------t===~====~-----------

-20~1 J-I~II~I~II __ ~ __ ~ __ ~II~'~II~~ __ ~I~I~I~II~I~I~I ___ I~I .3 .6 1 3 6 10 30 60 100 300 x 106

REYNOLDS NUMBER

I ,J! Vi,;

--------------------------------

IVDlma.4

HYPERSONIC WIND TUNNEL CAPABILITY CIRCA 1965

FULL SCAli

TURBULENT B.~ -"al- ~~ o ~;,

2 ~I -""'*'--''-+~6~U,1~0----t20;--~40,-L60:':-'-~1+'00;--~200 x 106

REYNOLOS NUMBER

10

PARAMETERS AFFECTING FLIGHT EXTRAPOLATIONS OF IliOlma•

FRICTION DRAG Co

PRESSURE DRAG C F Op

LIFT CURVE SLOPE CL DRAG DUE TO LIFT C a

°L CONTROL EFFECTIVENESS lICm lICL lICO

LIFT-DRAG RATIO M'B

~~_- SPALDING CHI CF~

LAMINAR15°~""~c>"§..o---aoo-- •

-c1tS~:5l.--·~ TURBUUNT ~.t:J1

-' REFERENCE' TEMPERATURE C

F

0~7-7---7-~~--~~~--~ .J .6 1 J 6 10 30 " 60 100 300 x 106

.(11

CN .06

a • )0 .04

REYNOLDS NUMBER

®

NORMAL FORCE

M' 8

~~~.76~1--~·~~~6~10;--~~l~0~~~--'~ REYNtlDS NUMBER

CAL SHOCK TUNNEL TEST CONDITIONS; M - 8

001 2000

TO' IlK

1000

~~OI----~l~I----~I.~O~--~----~~~

.016

.012

C" .008

.004

~Z

.12

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942 m. Fli GHT

®

NORMAL FORCE M ·8

cp ~,(,. TliEORY

;S ~

., REYNOLDS NUMBER

/Y, 0 m.2 x 106

'r' 0 32.31 fl' 0 10.41 , lJ. 4.02

1./ t>. l.Sl .. 0 597

0,;1

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AXIAL FORCE M' 8

4 adeg

10

REYNOLDS r.JMBl:R

o lSl2"x I~ 0)231 o 10.41 lJ. 4.02 II. 1.S3 o .5'l7

·014

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AXIAL FORCE 'M· a

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C '.008 A

\~_c;-REf[RENC£ TEMPERATURE Cf

LAMINAR '\ o~ a • )0

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.004

"~o1r6b~ {lTURBULENT ..... ~ --~ ..., SPALDING CHI Cf~

.OOZ ...Itrll.iC.IJI ___________________ _

~~~.6~~~~)~~6~10~~~)~O~~~~100--~~)ooXI~ REYNOlDS NUMBLR

FAIRED EXPERIMENTAL DATA,Moe

CA

.008

a· ) • • 006 o

.OIOl

.004, I , II III ! 1 I II',,'

40

,20 'II

ERROR o

ACCURACY OF FULL SCALE EXTRAPOLATION AT CONSTANT ANGLE OF ATTACK

CD"'::-... "LID _-------- C

l ASSUMED TURB~LENT

I I II '" I

~ 100

2.0

L6

u

.8

, C.,'rt,A.l:' .' (." . ."

DRAG DUE TO LIFT M· •

o

.4

~~~~--~~~~~~~~~~~~xlf

®

ACCURACY OF FULL SCALE EXTRAPOLATIONS OF OPTIMUM VALUES • OF CL, Co. AN~ LID

ILl DI/AAX zor-.>=::::::;;::::=:::==---==?==t=--------' 'IIERROR 0, :::::=:::::: -20

4O~ (ASSUMED LAMINAR

ClOPT 20 r :== ~ (ASSUMED TURBULENT '1\ ERROR DI~-,!",,----'==::::::±:::f,.-"l------

-20

C 40~ ~ DOPT

20 Y . 'II ERROR 0!-------'1-====,...,-----

"20, , I , ,'1' , , I,ll , I ! , ! ,"

. 3 . 6 I 6 10 30 60 100 REYNOLDS NUMBER

January 25-27, 1971

increase in fuel vapor concentration; whereas the diffusion-thermo effect decreases the temperature by 1.'>%, to 35% and also effects adversely the rate of vaporization of the droplet.

TUESDAY/JANUARY 26

~l-"

l ) ,', { :..,'; D

29

·Work supported by Aerospace Research Lahoratories, Office of Aerospace Research, USAF, Wright-Patterson AFB, Ohio under the technical cognizance of Major John West, Hypersonic Research Lab" Contract No. F3361.568C1184.

1:30 PM/GRAND BALLROOM

Fluid Dynamics 5: Separated Flows

Chairman: ROBERT F. WEISS, Avco Everett Research Lab., Everett, Mass.

(71;127) Separation of a Supersonic Turbulent Bound­ary Layer by a Forward Facing Step.· WILHELM BEHRENS, California Institute of Technology, Pasa­dena, Calif.

The mean flow and fluctuations were measured in the separated flow region and on top of the step. The model was large enough to allow reverse flow profile measurements and to insure two-dimensionality of the separated flow. For an aspect ratio (span between fences/step height) of 9.4, reverse profiles show three-dimensionality. The mean flowfield is described in detail. In the constant pressure separated flow region the forward velocity profiles are similar. The reverse flow velocity reaches O.37u,. On top of the step the turbulent boundary layer attains nearly equilibrium shape within one step height. Most of the fluctuation energy is in random large scale low frequency fluctua­tions.

·Supported by JPL Director'. Discretionary Fund.

(71-128) Results of a Strong Interaction, Wake-Like Model of Supersonic Separated and Reattaching Turbulent Flows. L. G. HUNTER JR. and B. L. REEVES, Avco Systems Div., Wilmington, Mass.

Results of an interaction theory for supersonic separated and reattaching turbulent boundary layers are presented and compared with recent experiments for flow past a compression ramp. Effects of ramp angle, l\fach number, Reynolds number, and upstream pressure gradient are considered for situations where the critical point is located upstream of the trailing edge. When the critical point falls downstream of the trailing edge the whole region of separated flow is influenced by ramp length. In these "short ramp" flows the peak ramp pressure attains a maximum at a critical ramp angle and then decreases with increasing angle. It is shown that this effect is responsible for the spanwise pressure distributions measured by White­head and Keyes for flow over a delta wing with a trailing edge flap.

(71-129) Hypersonic Laminar Boundary·Layer Sep· aration on a Slender Cone at Angle of Attack. KEN­NETH F. STETSON and CAPT. EDWARD S. OJ DANA JR., Aerospace Research Labs., Wright-Patterson AFB, Ohio

Wind-tunnel experiments with a 5.6 degree half angle cone at l\1 (X) = 14.2 indicated that the three-dim en-

sional separation bubble concept was not the correct flowfield model for these data. Based upon data con­sisting of surface pressure measurements, pitat pressure surveys, and surface oil flow patterns a new model for hypersonic three-dimensional separation is proposed. Thi~ model contains symmetrical supersonic helical vortices with an attachment line on the most leeward ray. The vortices are in contact with the surface (at least up to a = 18°) and there is no subsonic reverse flow or singular points associated with thc vortex pattern.

(71-130) Separation Patterns of Boundary Layers over an Inclined Body of Revolution. KUO CHANG WANG, Martin-Marietta Corp., Baltimore, Md.

Based on recent symmetry-plane boundary-layer solutions as well as the limited experimental observations, new separation patterns over an inclined body of revolution are presented. Partial agreements and differences from other predictions in the literature are noted. Existing three-dimensional separation criteria are mostly inapplicable, and the resulting patterns found are compatible only with Maskell's general description of separation; i.e., a combination of the free vortex layer type and a bubble type. For a not-too-blunt body, a bubblc type prevails at low in­cidcnce, and a free vortex layer type dominates at high incidence. At extremely high incidence or for very blunt body, a bubble type prevails again.

·Work supported in part by AFOSR Contract F H620-70-C-0085.

(71-131) An Investigation of a Supersonic Surface Jet-Hypersonic Flow Interaction with Axial Sym­metry. R. P. SHREEVE, Boeing Scientific Research Labs., Seattle, Wash., and G. C. OATES and H. G. AHLSTROM, University of Washington, Seattle, Wash.

Physical arguments suggested that the jet angle and jet to surface prcssure ratio control major changes in the interaction structure. An experimental investiga­tion was made with a fully turbulent boundary layer in heated air at J[ = 6 using a .')0 cone With cir­cumferential slots. The injectant was cold air and a new fine wire thermocouple probe was used. From optical, surface and probe measurements it is shown that. thrce basic types of structure occur depending on the jet angle. Results for the dividing streamline displace­ment, separation and shock di~placements, and mass entrainment into the jets of each type are presented.

30 AIAA 9th Aerospace Sciences Meeting

TUESDAY/JANUARY 26 .... -~--~ ... ~ -"-- ........ " ... 1:30 PM/WEST ROOM

.......... .....-----... --.. _. , .. ",-""

.-- "., ... ,., -- Ground Testing and Simulation

I

f

! I Ii

I

I I i I

l

( ! \

Chairman: JACK D. WHITFIELD, ARO Inc., Arnold Air Force Station, Tenn.

(71-132) Advances in Hypersonic Extrapolation Cap­ability-Wind Tunnel to Flight. J. A. PENLAND, NASA Langley Research Center, Hampton, Va., and D. J. ROMEO, Cornell Aeronautical Lab. Inc., Buffalo,

goal/requirement for ground test facilities might be to increase their Heynolds number capability by a factor of ten at.lf "" 20.

In the past, the limited Heynold" number capability namic Behavior of Vehicles in Tunnels. M. SAJBEN, N. Y. D (70-135) Scaling Problems in the Study of Aerody-

of hypersonic facilities has prevented reliable extra- California I nstitute of Technology, Pasadena, Calif. polation of data to flig;ht ne'ynol~" numbers. nec~nt, Problems of relating laboratory results to full scale resul.ts on a hypers~:)lllC crlllse atl'craft configUrahOn( vehicle-tunnel systems are discussed, accounting for obtamed at Mach 8 III a shock tunnel o,:,"er a Heynolds ,transient effects and dissimilar force-speed character­number range fro,m a completely lammar boundary .... istics of the respective propulsion systems. Data are layer ~o ?- predommately t.urbulent olle are present~d') I best utilized by the indirect process of evaluating The ,slglllficant factors whICh can affect extrapolation I certain empirical coefficients and use these in the \ o~ wmd-tun:lel da~a at. subscale neynold~ .numbers to 'appropriate theory to predict characteristics of full , fl!g~t values are Id~nhfied. The capability for pre- I scale systems. New ways are recommended to evaluate dl.ctmg turbulent flight lleyno~ds number data: from and present reduced data, allowing a substantial 1 wmd-tunnel ~ata ~lI!del: l.amma: and transItional; V r~duction of the experimental work needed to c?ver a boundary-layer conditions lil e sho» 11. t 0 ~171 i1-!. given range of parameters. The general questIOn of I (71-133) Hypersonic Low Density Cone Drag.· A. G. KEEL JR., University of California at Berkeley, Berkeley, Calif., and L. G. KRAIGE, R.D. PASSMORE, (71-136) Modeling of the Turbulence Structure of the and R. N. ZAPATA, University of Virginia, Charlottes- Atmospheric Surface Layer. T. R.SUNDARAM, G. ville, Va. R. LUDWIG, and G. T. SKINNER, Cornell Aeronau-

A th d f . d . f tical Lab. Inc., Buffalo, N. Y.

me· 0 or measurmg aero ynamlC orces on .,. axisymmetric bodies in the transition regime of low Hesults of th~orehc~l and expenmental studies on density, hypersonic flow hag been developed over the= the labo:-atory smlUlaliOl! of the structure of the. tur-past few years. The combination of a 3 degrees-of- bulen~e !I! the atmo~phenc surfac~ laye:- are d~scnbe4· freedom electromagnetic balance and a free jet afford~ The }'1I!lllitude reqlllrements are l,uvestll>ated III detail a high degree of precision, accuracy, and resolution and It IS shown that,. for proper SimulatIOn of the flow under strictly steady state conditions. Zero angle-of- over a mode}, a honzontally homogeneous, turbulent attack cone drag data will be presented to illustrate flow I!lust eX!.lt far upstream ?f the mo~el. Laboratory results of studies on the effects of nose bluntness, Mach exper.lments on the, generatIOn. of SUitable flows are number, axial flow gradients, and sting interference, descnbed, and de~m~ed compansons between the tur-The extension of the technique to nonzero incidence will bulel.lce charactenstlcs of the la:boratOl;~ and atm?s-he discussed phenc surface-layer flows are given, Ihe operatmg

. requirements of a facility required to achieve proper simulation are also discussed.

'Work supported under Grant No. AFOSR-69-179S.

(71-134) Flight Test Base Pressure Results at Hy­personic Mach Numbers and High Reynolds Num­bers in Turbulent Flow: Implications to Ground Test Simulation Requirements. J. M. CASSANTO, General Electric Co., King of Prussia, Pa.

Hecent slender cone flight data are presented which indicate 1) no significant radial base presslII'e gradients exist, 2) base presslll'e ratio (P b/ Pro) is relatively constant for order of magnitude changes (107 to 108) in Reynolds number, 3) data for small and large angle cones correlate when the ratio of base to cone pressure are plotted as a function of local Mach number preceding the base, and 4) full scale flight Heynolds numbers are an order of magnitude higher (108) than the simulation capability (107) of ground test facilities at J[ "" 20. Accordingly, this suggests that a

(71-137) Optical Crossed-Beam Measurements of Turbulence Intensity in a Supersonic Jet Shear Layer. M. Y. SU, Northrop Corp., Huntsville, Ala., and F. R. KRAUSE, NASA Manned Space Flight Center, Huntsville, Ala.

An analytical relationship which approximates the probabilitv density of flow velocity fluctuations through the cross-correlation of two sensor outputs of the optical cross-beam system is present. The optical system consists of two radiation source/sensor pairs, whose lines of sight pass through the local region of interest in the jet plume. Experimental results obtained by this technique for turbulent intensity measurements in the shear layer of a cold supersonic air jet (AI = 2.54) nre present; the probability density of the velocity fluctuations, the characteristic function and the tur­bulence intensity are given.

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