AIAA-90-1391

Design and Q~erationalFeatures of Low­Disturbance Wind Thnnels at NASALangley for Mach Numbers From3.5 to 18

I. E. Beckwith, F.-J. Chen, S. P. Wilkinson,M. R. Malik, and D. G. Thttle, NASA LangleyResearch Center, Hampton, VA 23665-5225

I,.,!

AIAA 16th Aerodynamic Testing ConferenceJune 18-20, 1990, Seattle, Washington

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/11111111 /1/11/111/1/11111111111/111111/1/1111111/TRP00237

DESIGN AND OPERATIONAL FEATURES OF LOW-DISTURBANCE WINDTUNNELS AT NASA LANGLEY FOR MACH NUMBERS FROM 3.5 TO 18

I. E. Beckwith·, F. -J. Chen.··,S. P. Wilkinsont, M. R. Malik*, and D. G. Thttle'C

NASA Langley Research Center, Hampton, Virginia 23665-5225

~x,~Y,ta. length, height, and width of quiet test

of aerodynamic codes. The helium facility will be usedprimarily for validation of boundary-layer and freeshear-layer stability and transition prediction codes.Shakedown runs of these facilities are scheduled in1990 for the helium tunnel andin 1991 for the Mach8 tunnel.

Abstract

The experimental and theoretical program atNASA Langley to develop high-speed low-disturbancewind tunnels for transition research is reviewed. Animportant objective of this research is to providereliable· predictions of transition from laminar toturbulent flow on supersonic and hypersonic vehicles.The design details will be presented for three new low­disturbance wind tunnels: the Mach 8 Variable DensityTunnel (M8VDl), the Mach 18 Quiet Helium Tunnel(M18QHT}, and the High-Speed Low DisturbanceTunnel (HSLDl). Some of their operational featureswill be discussed, particularily for the new largeHSLDT facility, which will have relatively large massflows and run times from about 6 to 25 minutes. TheHSLDT will accommodate nozzles for Mach numbersfrom 2.5 to 6. However, the current project is limitedto nozzles for Mach numbers of 3.5 and 6. ThisfacUity is designed to provide direct simulation oflow-disturbance flight conditions in the atmosphere.Shakedown runs are scheduled for 1992. Uniformtest flows required for validation of codes for vehicleperformance and transition prediction will be provided.The test sections will be up to 3 feet wide so thatexperimental techniques for the control and reductionof viscous drag and heat transfer can be studied indetail.

Extensive modifications are described of twoexisting facilities at Langley that will providelow-disturbance test conditions at Mach number 8(M8VDl) in air and at Mach numbers from 12 to18 in helium (M18QHT). The Mach 8 facility willprovide uniform flow conditions suitable for validation

·Senior Research Engineer, flAFS, The George WashingtooUnivenity, Hampton, Virginia, Associate Fellow AIAA··Senior Engineer, High Technology Corporatioo, Hampton,

Virginia, Member, AIAA.'tLeader, Quiet Thnnel Group, Experimental Methods Branch, FluidMechanics Division, Member, AIAA,*P~ident, High Technology Corporation, Hampton, Virginia,SenIor Member AIAA"Technical Project Engineer snd Project Manager, Facility Systems

Branch, Facilities Engineering Divisioo

1

ACenFfh.k

LMmNp

MqRReRk

R,rSTAtTuWXc,

xY,Z

Nomenclature

disturbance amplitude, or areaconstantcoefficient, eq (5)dimensionless frequency,~

U efrequency

throat heightpeak-ta-valley height of a surface

roughness defect or a scratchlength of nozzle from throat to exitMach numbermass flowIn of amplification ratio, -tpressurepressure dropdynamic pressureunit Reynolds number, pu/JJlocal unit Reynolds number (pu/p)olocal Reynolds number in undisturbed

laminar boundary layer at y=k, eq (1)momentum thickness Reynolds numberradiusaxial station, inchestimetemperaturestreamwise velocitymass flow per unit areaupstream tip of uniform flow test

rhombusaxial distance from nozzle throatvertical and horizontal distance from

nozzle centerline

Introduction

root-mean-square

Superscripts

The powerful adverse effect of wind-tunnelnoise on boundary-layer transition was conclusivelydemonstrated about 2 decades ago by the remarkablecorrelations of Pate and Schueler1• They showedthat Reynolds numbers for the end of transition onplanar models was accurately correlated by noisedependent parameters in 9 different wind tunnels atMach numbers from 3 to 8. Pate2 then successfullyapplied the same correlating parameters to transitionon sharp-tip cones in 11 different tunnels over theMach number range from 3 to 14. While stream noisedata were not yet available in all these tunnels, theparameters used in the correlations were related to theacoustic noise mdiated into the free stream by turbulentboundary layers on the nozzle walls as based on theearlier analysis and measurements in the JPL 2o-InchTunnel by Laufei3.4. These correlating pammeterswere the mean skin friction and displacement thicknessof the tunnel wall turbulent boundary layers and thetest section circumference. PateS later extended thesecorrelations to include data on planar models from 13

tunnels and data on cones from 16 tunnels for Machnumbers from 3 to 20. 'd

Stainback6•7 first showed that a quantitativerelationship exists between supersonic wind tunnelrms pressure levels and transition. The local Reynoldsnumbers at the onset of transition on sharp coneswas correlated directly with the noise measured bya pressure transducer flush with the cone surfaceunderneath the undisturbed huninar boundary layer.The correlations for data from 5 different windtunnels with free-stream Mach numbers from 6 to20 depended also on the surface-to-total tempemturemtio and whether the test gas was air or helium. Thehelium data at high local Mach numbers of Fischerand Wagner7•8 were later shown by Pate9 to agree withhis previous noise pammeter correlation for cones.

Following Morkovin's comprehensive review10

of the high-speed transition problem and hisrecommendation for the development of a low­noise hypersonic tunnel (based in part on therevelations of Pate and Schueler), the Boundary­Layer Transition Study Group (BLTSG), chairedby Reshotko, formulated a progmrn for transitionresearchll•12 that included the development of quietwind tunnels. Progress reports on this part of theBLTSG Progmrn by a group of researchers at NASALangley are available13- 18•

A more recent development for the BLTSGProgmm at Langley is a Mach 3.5 Pilot QuietTunne119,20 that exhibits very low-noise levels inupstream regions of the test rhombus (the "quiet testcore;" see figure 1). The original data showed thatthe low-noise levels could be maintained up to astream unit Reynolds number, Roo, of about 7.6 xlOS/in. These low-noise levels are only observed whenthe corresponding upstream "acoustic origin" regionsof the nozzle wall boundary layers are maintainedlaminar as illustrated in Figure 1. In the upper partof this figure, the downstream boundaries of the quiettest core are Mach lines that emanate from transitionlocations in the contour-wall boundary layers. In thelower part of the figure, the downstream boundaryMach lines originate on the flat sidewalls where Mw< 2.5 so that even for turbulent boundary layersthere, the noise mdiation levels are < 0.1 percent20•

The upstream boundaries of the quiet test core arethe Mach lines that form the upstream wedge ofthe uniform flow test rhombus. The laminar wallboundary layers occur on the contour walls when theturbulent boundary layer in the subsonic approach tothe nozzle is removed by suction slots upstream ofthe throat, although contributing factors are the verylow acoustic and vorticity disturbances in the settling

region, Fig. 1boundary-layer coordinate normal to

surfaceboundary-layer thicknesswavelengthdynamic viscositykinematic viscosity, pIpmass density

adiabatic wallRigimesh components (Figure 16(a»average leading edge thickness

centerlineat edge of boundary layerisentropic stagnation conditionsat diffuser exiton sidewall or projected from sidewall,

or conditions in vacuum spheresettling chambertransitionupstreamat wall or surfacefree stream

Y

6

"J.lvP

SCTuw00

Subscripts

awa, b, cbceo,oS

2

rd

chamber and the highly polished walls of the nozzle21 •Recent fmther improvements in the wall finish andwall cleaning procedures have been shown22 to extendthe quiet operation in the Mach 3.5 Pilot Quiet Tunnelup to Roo ~ 2 x 106/inch.

Significant advances in our understanding of theprofound effects of wind-tunnel noise levels and spectraon transition have been provided by correlations23 oftransition data on cones obtained at NASA Langleyin two quiet tunnels at Mach numbers of 3.5 and5.0. The noise conditions were varied with on or offcontrol of boundary-layer bleed upstream of the nozzlethroats, by varying the test unit Reynolds number, andby moving the model upstream or downstream withinthe quiet test cores. By appropriate manipulation ofthe first two test variables, the nozzle wall boundarylayers were either partially laminar or fully turbulentover the same range of unit Reynolds number. Thecorresponding stream-noise levels were then eitherextremely low (P/P, < 0.1 percent) with no energyat high frequencies, or the levels were much higher,approaching those in conventional tunnels, but withenergy at much higher frequencies than in conventionaltunnels. The quantitative effects of these variable noiseconditions on the transition Reynolds numbers in thequiet tunnels were demonstrated23 by correlations withprevious data in several conventional wind tunnels andwith flight data obtained by Dougherty and Fisher'A.The local Mach number range was restricted to valuesbetween 1.5 and 5.0. The final correlation resultsshowed that transition in these quiet tunnels wasdominated by the noise incident on the cone boundarylayer upstream of the neutral stability point The noiseincident on the boundary layer downstream of theneutral stability point has an exponentially decreasingeffect on transition as the distance from the neutralpoint increases. The correlations also indicated thatWillI t.l:!e boundary-layer bleed off, t.l:!e resulting highfrequency energy of the quiet tunnel radiated noisehad large adverse effects on transition.

Clearly, the boundary-layer transition processthat occurs in atmospheric flight can be easily "short­circuited" if large initial disturbance fields are presentwhich, in the limit, can completely by-pass any linearprocessing of incident disturbances. Unfortunately,for the high-speed case, all previously existing groundfacilities have large ambient stream disturbances whichprevent the correct simulation of flight transitionbehavior in these facilities. As noted above, thisambient disturbance field is caused primarily byacoustic radiation from the turbulent nozzle wallboundary layers. Detailed studies have indicatedthat accurate flight transition simulation requires themaintenance of laminar nozzle wall flows, and that

3

the reduction of noise radiation levels from turbulentwall boundary layers is not sufficient17•20,23,2S-27 forfull simulation of flight transition behavior.

In summary, the causes of facility wall turbulenceare primarily (a) continuation onto the nozzle wallsof the turbulent boundary layers present on the wallsof the upstream piping and stagnation chamber,21and (b) destabilization of the nozzle wall laminarboundary layers including Ol»1ler vortex formationand amplification in the concave curvature region ofcontoured nozz1es.28 The acoustic field produced bynozzle wall turbulence not only reduces the levelsof the model transition Reynolds numbers but, fromrecent research, the trends are also modified (e.g.,effects on transition of bluntness, cone angle of attack,model configuration, unit Reynolds number, etc26,27).

Quantitative Examples of Wind Tunnel Noise Effectson Transition

'JYpical transition onset data in the Mach 3.5Pilot Quiet Tunnel are compared in Figure 2 with flatplate data from conventional wind tunnels and withpredictions from linear stability theory using the f!imethod26 for flat plates and cones. Wind tunnel noisein conventional tunnels reduces transition Reynoldsnumbers on flat plates by nearly an order of magnitude~Transition onset Reynolds numbers on sharp tip conesin the Mach 3.5 Pilot Quiet Tunnel for different noiselevels are compared in Figure 3 with flight data24 andconventional wind tunnel data20. For this situation, thehigh noise levels reduce transition by a factor of about1/3. An interesting example27 of the large changes inboth levels and trends of transition Reynolds numberwith leading edge bluntness Reynolds number causedby different noise conditions in the Langley Mach3.5 Pilot Quiet Tunnel and in conventional tunnels isshown in Figure 4. Clearly, the wind tunnel noisehas significant effects on the complex flows due tobluntness effects.

Successful approaches to the establishment andmaintenance of laminar '\Vall flows in high-speedfacilities20--23 have been discussed Transition behaviorobserved in the Mach 3.5 pilot facility utilizingthese approaches exhibits excellent agreement withflight data24 as shown in Figure 3. It should benoted that, along with maintenance of laminar nozzlewall flows, valid simulation of high-speed boundary­layer transition behavior in ground facilities requirescontroVminimization of stagnation chamber vorticityand acoustic fluctuation fields, entropy spottiness,particulates,29 and model vibration. Stainback,3o eLaI., first showed that reduction of settling-chamber

disturbances (presumably control valve noise andentropy fluctuations) by installation of a large Mporous plate ("Rigimesh")* in the settling chamberof the Mach 8VDT (before its conversion to a quiettunnel) significantly increased the measured transitionReynolds numbers on a cone.

Advances in Quiet Nozzle Design

In order to apply nozzle wall transition data fromthe quiet tunnel studies to the design of improvednozzles, it is essential to understand the instabilitymechanisms involved and to develop theoreticalmodels that can be used for predictive purposes. Thedevelopment of satisfactory models depended on thesuccessful application of linear stability theory, usingthe eN method, to test data on cones in the Mach3.5 Quiet Thonel and in flighf1; and to the correctlocal flow conditions in nozzles28,32 as illustratedby Figure 5. Here, N-factors from about 9 to 11predicted transition locations in these different flows.The principal result of the nozzle investigation wasto show that transition in the wall boundary layersof nozzles was caused by the GOrtler instabilitymechanism in the concave curvature regions of thewall rather than Thllmien-Schlichting waves, providedthat the nozzle walls were maintained clean andpolished and boundary-layer removal slots upstreamof the nozzle throats were used. Comparisons32 oftheoretical predictions with experimental data showedthat application of linear stability theory with the eNmethod gave N-factors from about 6 to 11 for transitioncaused by GOrtler vortices in the wall boundary layersof nozzles for Mach numbers from 3 to 5. Malikhad shown previously that the use of N~ 9 to 11predicted transition for different low speed flows 33,34

and, more recently, for cold wall conditions on cones3S

in supersonic and hypersonic flight. Another importantresult from the previous nozzle design studies28, 32

was to show that transition on the nozzle walls couldbe delayed considerably by inserting a long straightwall section corresponding to a region of radial flowexpansion according to the nozzle design methods ofBeckwith, et al.36,37

The use of N ~ 9 for the GOrtler instabilitythen predicts quiet test cores in the new Mach 3.5and Mach 6 axisymmetric-long pilot nozzles (Figs.6 and 7) that are 3 to 4 times longer than observedin the rapid expansion pilot nozzles. These newnozzles utilize a region of radial flow which movesthe wall inflection point far downstream and thereby

·Trade name of Pall Process Filtration Corporation, East Hills. NewYOlk

4

delays the onset and amplification of the G6rtlervortices21,22,28,32 due to the larger radii of curvatureand thicker boundary layers along the concave wallregions. Therefore, the predicted locations of transitionwith No ~ 9 occur far downstream at the indicatedXT,w distances (Fig. 6) which decrease with increasingunit Reynolds numbers38• The resulting predictedaxial lengths, ~X, of the quiet regions are then muchlarger than in the rapid expansion pilot nozzle (Fig.1) as shown by comparison with hot-wire data26 inthat nozzle, reproduced here as Figure 8. In thisfigure, the normalized rms pressure fluctuations fromhot-wire surveys along the centerline of the nozzleare plotted against X, the distance from the nozzlethroat. The length of the quiet test core, ~X, (definedin Fig. 1) decreases with increasing unit Reynoldsnumber, but tends to approach a constant value that isslightly less than 5 inches. The largest amplificationof Thllmien-Schlichting (TS) waves in the new Mach3.5 axisymmetric nozzle occurs at the highest unitReynolds number where NTs =2.3 (Figure 6).

The free-stream Reynolds Number, RAx, based onthe length ~X provides the best criteria for comparisonof different nozzles21,22,38 as shown in Figure 9. Here,RAx from test data for f61lr different nozzles and Re, T

(transition onset Reynolds numbers from flight data oncones) are plotted against the unit Reynolds number.The data for the axisymmetric and two-dimensionalrapid expansion (R. E.) nozzles show an increasingtrend with unit Reynolds number except for the largervalues of k (the maximum peak-to-valley defects inthe surface finish) measured in the throat regions.These large values of k > 40 p-inches, exceed theroughness Reynolds number criteria of Rk ~12 whichis based on profilometer data, microphotographs, andnoise data22 in the Mach 3.5 pilot nozzle. Recent RAJ{data obtained in the new Mach 3.5 long-axisymmetricnozz!e38 show similar trends with Re and k but areabout three times larger than for the other nozzles.These data are in excellent agreement with predictionsfor No =9, and with further improvement in thesurface finish (k < 60 J.l-inches) agreement at the largervalues of Re can be anticipated. Preliminary datashown in Figure 9 from the new Mach 6 pilot nozzle(Fig. 7) indicate that laminar wall boundary layerswere achieved up to Roo ~ 3.4 x lOS/in. with RAx ~6.5 X 106• Again, further improvements in the surfacefinish (design requirement: k<30 winches) is expectedto provide the design RAx ~ 13 X 106•

Mach! Variable Density Tunnel

The Langley Mach 8 Variable Density 1\mnel(M8VDT) is a conventional air blowdown tunnelwhich is being modified to function as a hypersonicquiet tunnel. Maximum tunnel pressure is 3000 psiaand the flow can be heated up to 10500 F to prevent aircondensation. This tunnel is important for hypersoniclaminar flow stability experimental studies since thedominant instability disturbance will be 2nd mode TSwaves. The major modifications consist of a 1 pm airfilter, new components for the settling chamber, a newaxisymmetric slotted nozzle, and a new two-strokemodel injection mechanism with its new test sectionenclosure as indicated on Figure lO(a). The settlingchamber will contain three Rigimesh elements andfour turbulence management screens (Fig. 1O(b». Thesettling chamber wall boundary layer will be removedvia a bleed slot upstream of the nozzle throat and anew exhaust duct system routed through the existingretaining block as shown in the figure. Conventionaltesting will be conducted by substituting a non-slottednozzle throat section rather than by closing the bleedsystem valve as is customary in the pilot quiet tunnels.This precaution is taken to prevent undue degradationof the delicate leading edge of the nozzle bleed scoopand of the highly polished throat The nozzle for thishigh temperature application presents unique designchallenges. The requirement for a non-corrosive,super-finish surface in the nozzle throat region dictatesuse of a highly non-oxidizing alloy or surface coating.An evaluation program of various alloys has resulted inInconel601 as the most likely material for the nozzle.

The required values of k in the throat region ofthis nozzle are shown in Figure 11. Values of k forPo=IOOO psia were computed from laminar boundarylayer solutions by the method described previously22assuming that Rk = 10. For the other two values ofPo, k was computed by the following approximatemethod: By definition:

Ric = (pu) /e (1)P y=1c

Then for k«6, we may assume T::::: Tw andu/Ue:::::C k/6. Also for a laminar boundary layer,

6/X ::::: f (Tw ,Me) IJP"'X (2)Taw Pe

if To » Sutherlands constant (I98°R), and thereference values are from a laminar boundary-layersolution for the nozzle. Hence, if the variation of kwith X for a particular nQzzle and a particular valueof Twrr.w is known from the boundary-layer solutionfor given values of Po and To. k is then computed forother values of Po and To from equation (3) (providedthat Me is not significantly affected by different valuesof 6). From Figure 11, the minimum values of k thenrange from about 15 p-inches at Po = 1000 psia up toabout 55 p-inches for Po = 200 psia. The best surfacefinish ever achieved on the Mach 3.5 two-dimensionalnozzle blocks corresponded to k ::::: 20 p-inches.Hence, the surface finish requirements for the Mach8 Quiet Nozzle at the higher pressures may not bepossible and even if they are, the maintenance of suchhigh quality finishes at these high temperatures willbe extremely difficult.

The new Mach 8 long axisymmetric nozzle for theM8VDT was designed by the methods described in theprevious section of this report. This nozzle is 11.7-ftlong and has an exit diameter of 18-inches (Fig. 12(a». As noted above, the nozzle is equipped with aboundary-layer removal slot upstream of the throatDue to the unusually long, gradual expansion of thisnozzle the Tollmien-Schlicting (TS) amplification nowbecomes appreciable at the locations where predictionswith linear stability theory indicate that the G6rt1ervortices would cause transition. (The design conditionsused for the boundary-layer stability analysis are givenin the figure.) Hence, for this nozzle, the possibilityof interaction between the two instability mechanismsmay reduce the length of the quiet flow test core.However, even if this interaction causes the effectiveN for transition of the G(~rtIer vortices to be reducedto 7, most of a 4 foot long cone would be exposed tolow noise conditions as indicated in the figure.

Figure 12(b) shows Mach number contours inthe transonic slotted region of the nozzle calculatedwith the 3-D Navier-Stokes code, CFL3IY9. Theflow through the converging nozzle and bleed channelappears to be free of any large disturbances. The flowin the diverging nozzle is also being analyzed withthis code to validate the design procedure38 for theMach 8 nozzle at these conditions.

Mach 18 Quiet Helium Tunnel

Inserting these relations into equation (1) and solvingfor k gives,

.\_k _ (To Po,re! ) •/ere! - Po To,re!

(3)

5

The Mach 18 Quiet Helium Tunnel (MI8QHT)is a helium blowdown tunnel fitted with a conicalnozzle in an open jet configuration . The tunnel's mainpurpose is to allow study of the effects of pressure

gradient on the stability of laminar boundary layersin hypersonic flows. Modifications of an existingMach 20 facility (Fig. 13(a» are underway to producequiet, hypersonic flow. The modifications consist ofreplacing the original nozzle throat section with anew highly polished slotted nozzle, adding screensand Rigimesh plates to the settling chamber, andinstalling a boundary-layer bleed system as indicatedin Figure 13(b). A briefstudy of the tunnel opemtionalcharacteristics with the Mach 20 nozzle was conductedprior to installation of the new settling chamber andbleed system. An immediate problem was the lack ofa model injection mechanism. 'funnel-start tests ona series of cone-flare blockage models showed thatonly very mild flares on small base diameter coneswould allow the tunnel to start successfully. Whetherthe starting problem is caused by nozzle blockage ordiffuser characteristics has not yet been determined.If blockage is the problem, a solution may be to usea hollow cylinder as a test model instead of a cone.Installation of the new settling chamber and bleedsystem is proceeding.

Figure 14 shows an enlarged axial cross sectionof the conical slotted nozzle for the Ml8 QHT. Thenozzle is 79.75 inches long from the throat minimumto the exit and has an exit diameter of 14 inches. Thecone half-angle is 5 degrees. The highly polishednozzle throat section is a separate piece and extendsdownstream from the bleed lip by 15.9 inches. Thethroat minimum diameter is 0.678 inch. The straight5 degree taper begins at X =0.647 inch. The throatsection was machined from type 304 stainless steelwith a +0.001 inch coordinate tolerance and a 2microinch rms surface finish. The surface wavinesstolerance was 0.0002 inch/inch. The slotted portion ofthe nozzle shown in Figure 14 consists of an annularinset between the nozzle bleed lip and the settlingchamber wall. The 0.005 inch bleed gap correspondsto 30 percent of the total mass flow diverted throughthe annular slot The conical design of the nozzlewas chosen for two reasons. First, it minimizedmodifications to the tunnel as described in Figure13. Secondly, the stmight taper prevents formation ofG6rt1er vortices on the nozzle walls thereby increasingthe chances of successful quiet opemtion. However,the actual streamlines may have some concavecurvature due to the mpid boundary-layer growth inthis type of nozzle. Also, the TS growth rates maybe appreciable after long mns of laminar flow in spiteof the favomble pressure gradients. Data on nozzleperformance are not yet available.

6

High-Speed Low-Disturbance Tunnel

An elevation view of the new High-Speed Low­Disturbance Tunnel (HSLDT) is shown in Figure 15.Several of the major components of the tunnel. areidentified in the figure. Further details of the settlingchamber, the two-dimensional nozzles for Mach 3.5and 6, the model injection system, and the variablearea diffuser will be presented in this section. The finalengineering design of this tunnel is now completed.Existing high pressure air and vacuum systems atNASA Langley will be used. The tunnel is designedto accommodate nozzles of various lengths with Machnumbers mnging from 2.5 to 6. The specificationsand design of the 1 pm"high pressure air filter forthe new facility are based on experiences and dataobtained in the Mach 3.5 (Ref. 29) and Mach 6 PilotQuiet Tunnels. The filter will be installed immediatelyupstream of the settling chamber.

Settling ChamberDetails of the 8 foot diameter settling chamber

and its internal components are shown in Figure 16(a).The inside diameter of the internal flow liner is 76inches. The liner channels the inlet flow through theconical Rigimesh inlet diffuser, the two Rigimeshplates, the honeycomb with its adjacent screen, and the7 main turbulence screens to the conv-erging nozzles.The three Rigimesh components will reduce the rmspressure disturbances, caused primarily by controlvalves to less than 0.01 percent of the stagnationpressure as shown by data21 from the Mach 3.5 PilotTunnel which provided the design requirements forthe new tunnel. The screens will provide uniformmean flow velocity and turbulence levels of less than1 percent to the nozzle contractions, as based againon pilot tunnel data21 • For operation at Mach 6, thesettling chamber liner and internal components willbe preheated to 500°F with electric heating elementsinstalled on the outside surface of the liner and bysmall mass flows of heated air through the tunnel.The man-hole entrance in the bottom of the settlingchamber pressure vessel and liner converging nozzleis provided to allow access for inspection and cleaningof the nozzle throats and bleed scoop leading edges.

The pressure drop through the Rigimeshcomponents is a critical design parameter becauseof the resulting high loads carried by these largecomponents. The principal functions of the Rigimeshcomponents in the HSLDT are to diffuse the inletjet and to attenuate the extremely high valve andpipe noise present in most blow-down wind tunnels.Measurements of fluctuating pressures21 in the

I,d

Pilot Quiet Tunnel showed that the normalized rmspressures, Pw/Po, were reduced by more than anorder of magnitude by the use of several Rigimeshcomponents. When a Rigimesh entrance cone andtwo Rigimesh plates were used in the Pilot Tunneland when

operation to the low flow rates and small ii.P valuesused by the manufacturer is done with equation (5)using values of n and Cn for the tunnel conditions.If desired, a correction for temperature may also beincluded40,41.

li.Pa =15.1, li.P" =14.1, aPe =10.8 psid

is then solved for Cn. Here, the subscripts 1 and2 refer to the upstream and downstream pressureson any single component. An iteration process isrequired since

ii.Psc = ii.Pa + ii.P" + ii.Pe = 40 psid

where the subscripts a, b, and c refer to the componentsas identified in Figure 16(a). Thus for the downstreamplate (item a in Fig. 16(a», ii.Pa = Pl - Po ~

o.4 ii.Psc was used as a first approximation. Thefinal values are

It!

NozzlesThe centerpIane contours and cross-sections of

the converging nozzles are shown in Figure 17. Thecontour curves and cross-sections of both nozzlesare identical from Sta-90 (the entrance) to Sta~50 asalso indicated in Figure 16(a). As in the case of theM8VDT, a 3-D Navier-Stokes code39 is being adaptedto analyze the flow in these converging nozzles.

A two-dimensional Mach 3.5 nozzle and a two­dimensional Mach 6 nozzle (Fig. 18) are included inthe final design of this project. The exit dimensionsof these two nozzles will be identical (36-inches wideand 18.64-inches high) to simplify the connection withthe closed test section and diffuser. The aerodynamicdesign methods used for these nozzles are the sameas for the long pilot nozzles28• 32. 38 as indicated inthe figures. The large width of the new nozzles isrequired to prevent acoustic noise radiation into thequiet test core from sidewall boundary layers whichmay be turbulent at the higher pressures.

The locations of the leading edges (L.E.) of theboundary-layer removal slots for the two nozzles areindicated in the figure. Details of the slot channelsand bleed mass flows are given elsewhere42• The newnozzles are scaled-up by a factor of about three timeslarger than the corresponding pilot nozzles. Therefore,one of the main advantages of the new larger tunnel ascompared with the pilot nozzles is that, for the samelength Reynolds numbers, the laminar boundary layerson test models and also on the nozzle walls will bethicker by the sa-me scale-up factor used on the nozzlesand models. Hence, these boundary layers can beprobed with much greater accuracy so that the physicsof boundary-layer stability and transition can then bestudied in detail. For these conditions the allowableroughness defects on the nozzle walls and modelswill also be larger by the same physical scale-upfactor. This eases the nozzle surface finish problemsconsiderably and also allows for much improvedduplication of model surface details such as suctionslots, holes, steps, and other fabrication details relatedto laminar flow control techniques on real aircraft.

The required values of k from boundary-layersolutions for these nozzles with Rk=lO are shown inFigure 19. For the Mach 3.5 nozzle (Fig. 19(a» at theapproximate maximum operational pressure of Po=150

(5)

(4)ii.Pscfqo > 1500

then the maximum practical reduction in Pw / Po wasachieved. Since qo = i Po Mjc' the maximumvalues of ii.Pse occur in the HSLDT for operation atMach 3.5. Inequality (4) is then satisfied for a selectedvalue of ii.Pse=40 psid with the effective maximumoperational conditions of Po=167 psis, To=660oR, andrh=340 Ib/sec. With this value of aPse taken as adesign requirement, the basic pressure drop equationfor porous materials40 (where only the inertial termis used here),

corresponding to Cn=5.38 x lOS.The value of n used for the present application

is n=1.865 rather than the theoretical value40 ofn=2 because data for the Pilot Quiet Tunnel with aRigimesh cone and two Rigimesh plates21 gives theformer value. These data are shown in Figure 16(b)where Pu is the measured pressure at the upstreamentrance to the pilot tunnel settling chamber.

The coefficient Cn in eq.(5) depends primarilyon the porousity of the Rigimesh components asdetermined by their thickness and mesh size which iscontrolled and tested (at low flow rates and ii.P values)during the manufacturing process as described in detailby Dillon, 1iimpi, and WilcoX41 • Extrapolation fromthe high flow rates and ii.P values required for tunnel

7

IIIIIIIIIII/1I1111111111llll1lllllll1ililTRP00232

psia, the minimum values in the throat are k~50 1'­inches which Catl certainly be provided since the bestfinish ever achieved on the pilot nozzles correspondedto k~20 p-inches (see Fig. 9). For operation ofthe new nozzle at P0=50 psia, the required minimumvalues in the throat are k~125 JJ-inches which, fromexperience with the pilot nozzles22 , corresponds toa finish of 2 to 3 rms JJ-inches. This finish can beobtained on a high quality machined surface withsome hand or machine polishing work. The valuesof k at the bleed scoop leading edge are arbitmrilyfaired to a small value which depends on the boundarylayer thickness there. This thickness, in tum, dependson the leading edge radius (now specified as 0.040inch) and the pressure or local unit Reynolds number.The scoop leading edge radius on the Mach 3.5 pilotnozzle was specified as 0.020 inch which was verifiedwith a shop microscope. Maintenance of this leadingedge radius and finish is, of course, another importantrequirement which in view of experiences with thepilot nozzles, can be satisfied.

The throat finish requirements for the Mach6 nozzle also correspond to k~50 p-inches at thelaminar design pressure of Po=150 psia (Fig. 18(b».However, for operation of this nozzle at Po=550 psia,the required values of k~20 JJ-inches will be muchmore difficult to obtain and maintain. The man-holeaccess in the settling chamber (Fig 16(a» will beessential for this purpose.

Model Injection SystemFigure 20 shows the two-stroke model injection

system for the HSLDT. The nominal distances of thevertical and forward strokes are 34 and 67 inches,respectively. The total elapsed time to place the modelleading edge at the upstream tip of the uniform flowtest rhombus will be approximately 3.5 seconds. Heattransfer and transition data can then be obtained onmodels by standard techniques43 •

Probe surveys of model boundary layers andfree-stream flows will be done with a 3-axis surveymechanism mounted on top of the test sectionenclosure. Run times will be rather limited for thistype of survey work.

DiffuserThis discussion on diffusers is opened with a

brief review of pressure recovery in high-speed windtunnels. Data for pressure ratios, PolP~, from severaldifferent high-speed wind tunnels at Mach numbersfrom 2.2 to 9.6 are shown in Figure 21. Comparisonsof data for 5 closed test section tunnels44 with the openjet design curve (also from experimental data for Mach

8

numbers from 1.2 to 4.8 from many data points for twotunnels) reported by Howarth4S and a data point fromthe Langley Pilot Quiet Tunnel at Mach 3.5, show thatpressure recovery,P~, is much larger (smaller valuesof PolP~ ) for closed test section tunnels than openjet tunnels. From the faired curves in this figure, thepressure recovery for a closed test section is largerthan that for an open jet tunnel by factors of up to2.7 at Mach 3.5 and up to 3.4 at Mach 6. The ratherlarge mass flows and comparatively short run times ofthe HSLDT (see Table I) then dictated the closed testsection design as illustrated in Figures 15 and 20.

Typical run times for the HSLDT at Mach 3.5and 6 were computed from· the approximate relation,

()dp,

puA. = ~ dt

or

where V.. P.. and Pa are the volUOle, density, andpressure, respectively in the vacuum spheres. Subscript"." denotes conditions at the nozzle throats where M.is taken as unity. Inserting R = 1716/t2Isec2°F forthe gas constant, 1= 1.4 for the ratio of specific heats,and integration from time, t = 0, to the maximumpressure recovery PafP0 then gives the run time, 1,in seconds;

t = (.0353)lVT;[(P,). - (P,) J-T,A. . Po moot Po t=o

where Va = 6.73 X lOS ft3 is the total vacuumcapacity, T is in OR, and A. is the total throat areain ft2. Also, the sphere pressure is assumed to bethe recovery pressure, P, = P~ and Po, To, and T.are assumed constant. The high pressure and vacuumpumping rates are neglected. For blow-down to theatmosphere, the current storage capacity of the highpressure bottle field is taken as 170,000 Ibs. air fromthe initial maximum pressure of 4250 Ib/in2 down tothe minimum allowed pressure of 2000 Ib/in2•

Results are shown in Table I for; (A), smallblockage model or probe; (B), typical blockage due totest model and strut; and (C), open jet test sections.Values of P~IPo ~ p. I Po are taken directly fromFigure 21. The same values of To shown for blow­down to atmosphere are used for blow-down to thespheres. Clearly, the use of the closed test section andvariable area diffuser provide at least twice the rontimes of an open jet test section.

A plan view-cross section of the variable areadiffuser is shown in Figure 22. The minimum closureshown is for Mach 6 operation from data reported byHermann46• The design entrance ramp angles werebased on previous experience44-46.

Schedule for Quiet Tunnel Developments

The development and construction schedules forthe various tunnels described in this report are outlinedin Figure 23. Further details are given below:

Current estimates indicate that the Mach 6,Slow-expansion nozzle in the Nozzle lest Chamber(NTC) should be ready for transition research by July1990 provided that the surface finish can be improvedto provide k ~ 30 Jl-inches. The next facility tocome on line will be the Mach 18 Quiet HeliumTunnel in September 1990 for initial shakedown runs.Problems with starting the tunnel without a modelinjection system for flared cone models may, however,delay the start of transition research. The Mach 8Variable Density Tunnel is scheduled for completionin July 1991. The high temperature environmentcombined with the most stringent surface finishrequirements of all the quiet tunnels indicate thatthe full design laminar conditions (quiet test region.6.X~ 90 inches) at Po=l000 psia will be difficultto achieve. However, at Po <500 psia some laminarflow should be possible. If so, the sensitive neutralstability regions of the boundary layers on conesand flat plates will be exposed only to low noiselevels and stability research can proceed. The Mach6 nozzle for the HSLDT may not be included inthe current fabrication and construction phase of theproject due to possible budgetary restraints. However,all design and engineering work for that nozzle willbe completed under the current project The Mach3.5 nozzle of the HSLDT is scheduled for shakedownruns in November 1992 and no serious problems,either in design or operation, have been identified thatmight alter L'Ult date.

Conclusions

The techniques and requirements for thedesign and fabrication of low-disturbance super­sonic/hypersonic wind tunnels have been developedand tested. Design and operational details of three newlow-disturbance tunnels are presented. These tunnelsare the Mach 18 Quiet Helium Tunnel, the Mach8 Variable Density Tunnel, and the Mach 3.5/Mach6 High Speed Low-Disturbance Tunnel. Thesenew low-disturbance facilities will be essential for

9

evaluating and controlling viscous flows for a varietyof high-speed applications including the developmentand testing of laminar flow control techniques forapplication to the high-speed civil transport. Thenew capabilities for quiet operation at Mach numbersfrom 6 to 18 will be required for hypersonic stabilityand transition research. Boundary-layer stability andtransition prediction codes can be calibrated/validatedin these low-disturbance tunnels over the speed rangefor applications to supersonic and hypersonic aircraftof the 'l\venty-first Century.

Acknowledgement

Large, complex wind tunnels such as thosediscussed in this report require essential contributionsfrom many different disciplines. The authors extendtheir sincere appreciation to all members of the designteams at NASA Langley'futd Sverdrup Technology,Inc., Tullahoma, Tennessee, and in particular, to J.A. Osborn and G. M. Summerfield of the FacilitiesEngineering Division, NASA Langley Research Centerfor their leadership on the Mach 8 VDT and to P.E. Sensmeier of Sverdrup Thchnology, Inc., for hisleadership of the detail engineering design for theM8VDT and the HSLDT.

References

1. Pate, S. R.; and Schueler, C. J.: RadiatedAerodynamic Noise Effects on Boundary LayerTransition in Supersonic and Hypersonic WindTunnels. AIAA Journal, Vol. 7, No.3, March1969, pp. 450-457.

2. Pate, S. R.: Measurements and Correlations ofTransition Reynolds Numbers on Sharp SlenderCones at High Speeds. AIAA Journal, Vol. 9,No.6, June 1971, pp. 1082-1090.

3. Laufer, John: Aerodynamic Noise in SupersonicWind Tunnels. Journal of Aerospace Sci., Vol.28, No.9, Sept 1961, pp. 685-692.

4. Laufer, John: Some Statistical Properties of thePressure Field Radiated by a Turbulent BoundaryLayer. The Physics of Fluids, Vol. 7, No.8,August 1964, pp. 1191-1197.

5. Pate, Samuel R.: Dominance of Radiated Aero~

dynamic Noise on Boundary-Layer Transition inSupersonic-Hypersonic Wind Tunnels. AEDC­TR-77-107, March 1978.

6. Stainback, P. Calvin: Hypersonic Boundary LayerTransition in the Presence of Wind Tunnel Noise.AIAA Journal, Vol. 9, No. 12, December 1971,pp. 2475-2476.

7. Stainback, P. C.: Part I: Measurement ofTransition Reynolds Numbers and SurfacePressure. Fischer, M. C. and Wagner, R. D.: PartIT: Transition and Hot-WIre Measurements up toa local Mach Number of 16 in Helium. Effectsof Wind Tunnel Disturbances on HypersonicBoundary Layer Transition. AJAA Paper No.72-181, January 1972.

8. Fischer, M. C.; and Wagner, R. D.: Transition andHot-Wire Measurements in Hypersonic HeliumFlow. AIAA Journal, Vol. 10, No. 10, October1972, pp. 1326-1332.

9. Pate, S. R.: Comparison of Helium Tunnel Tran­sition Data With Noise-lhmsition Correlation.AIAA Journal, Vol. 12, No. II, November 1974,p. 1615.

10. Morkovin, Mark V.: Critical Evaluation ofTransition From Laminar to T1.ubulent ShearLayers with Emphasis on HypersonicallyTraveling Bodies. AFFDL-TR-68-149, March1969.

11. Reshotko, Eli: A Program for TransitionResearch. AIAA Journal, Vol 13, No.3,March 1975, pp. 261-265.

12. Reshotko, Eli: Boundary-Layer Stability andTransition. Annual Review of Fluid Mechanics,Vol. 8, 1976, pp. 311-349.

13. Beckwith, Ivan E.; and Bertram, Mitchell H.: ASurvey of NASA Langley Studies on High-SpeedTransition. NASA TMX-2566, July 1972.

14. Beckwith, Ivan E.; Harvey, William D.; Harris,Julius E.; and Holley, Barbara B.: Control ofSupersonic Wind-Tunnel Noise by Larninarizationof Nozzle-Wall Boundary Layers. NASA 'I'MX-2879, December 1973.

15. Beckwith, I. E.: Development of a High ReynoldsNumber Quiet Tunnel for Transition Research.AIAA Journal, Vol. 13, No.3, March 1975,pp. 300-306.

16. Beckwith, I. E.; Anders, J. B.; Stainback, P.C.; Harvey, W. D.; and Srokowski, A. J.:Progress in the Development of a Mach 5 QuietTunnel. AGARD-CP-224, October 1977, pp.28-1-28-14.

17. Anders, J. B.; Stainback, P. C.; and Beckwith,I. E.: New Technique for Reducing Test SectionNoise in Supersonic Wind Tunnels. AIAA Journal,Vol. 18, No.1, January 1980, pp. 5-6.

18. Creel, T. R.; and Beckwith, I. E.: Wind TunnelNoise Reduction at Mach 5 with a Rod-WallSound Shield, AIAA Journal, Vol. 21, No.5,May 1983, pp. 643-644.

19. Beckwith, Ivan E.; and Moore, William 0., III.:Mean Flow and Noise Measurements in a Mach

10

3.5 Pilot Quiet Tunnel. AIAA Paper 82-0569,March 1982. '.01

20. Beckwith, I. E.; Creel, T. R., Jr.; Chen, F.-J.; andKendall, J. M.: Free Stream Noise and TransitionMeasurements on a Cone in a Mach 3.5 PilotQuiet Tunnel. NASA TP-2180, 1983.

21. Beckwith, I. E.; Chen, F.-J.; and Creel, T. R.,Jr.: Design Requirements for the NASA LangleySupersonic Low-Disturbance Wind Tunnel. AIAAPaper No. 8~63-CP, 1986.

22. Beckwith, I. E.; Chen, F.-J.; and Malik, M.R.: Design and Fabrication Requirements forLow-Noise Supersonic/Hypersonic Wind Tunnels.AIAA Paper No. 88-0143, 1988.

23. Chen, Fang-Jenq; Beckwith, Ivan E.; and Creel,Theodore R., Jr.: Correlations of SupersonicBoundary Layer Transition on Cones IncludingEffects of Large Axial Variations in Wmd-TunnelNoise. NASA TP 2229, 1984.

24. Dougherty, N. Sam, Jr.; and Fisher, David F.:Boundary-Layer Transition Correlation on aSlender Cone in Wind Tunnels and Flight forIndications of Flow Quality. AEDC-TR-81-26,February 1982.

25. Anders, J. B.; Stainback, P. C.; Keefe, L. R; andBeckwith, I. E.: Fluctuating Disturbances in aMach 5 Wind Tunnel. AIAA Journal, Vol. 15,No.8, August 1977, pp. 1123-1129.

26~ Chen, F.-J.; Malik, M. R.; and Beckwith, I. E.:Boundary Layer Transition on a Cone and FlatPlate at Mach 3.5. AIAA Journal, Vol. 27, No.6, June 1989, pp. 687-693.

27. Beckwith, I. E.: Effect of Tunnel Noise onTransition at High Speeds and Calibration of the~ Theory. Paper Number 58, Fifth NationalAerospace Plane Symposium, October 18-21,1988.

28. Beckwith, I. E.; Malik, M. R.; Chen, F.-J.;and Bushnell, D. M.: Effects of Nozzle DesignParameters on the Extent of Quiet Test Flow atMach 3.5. Laminar-Turbulent Transition, IUTAMSymposium, Novosibirsk 1984, Editor: V. V.Kozlov, Springer-Verlag, Berlin, Heidelberg,1985, pp. 589-600.

29. Creel, T. R., Jr.; Beckwith, I. E.; and Chen, F.-J.:Nozzle Wall Roughness Effects on Free-StreamNoise and Transition in the Pilot Low-DisturbanceTunnel. NASA TM-86389, 1985.

30. Stainback, P. Calvin; Wagner, Richard D.;Owen, F. Kevin; and Horstman, Clifford C.:Experimental Studies of Hypersonic Boundary­Layer Transition and Effects of Wind-TunnelDisturbances. NASA TN D-7453, March 1974.

31. Malik, M. R.: Instability and Transition inSupersonic Boundary Layers. Laminar-TurbulentBoundary Layers. (Ed.: E. M. Uram andH. E. Weber) Energy Resources ThchnologyConference, New Orleans, Lousiana, February12-16, 1984.

32. Chen, F.-J.; Malik, M. R.; and Beckwith, I. E.:Instabilities and Transition in the Wall BoundaryLayers of Low-Disturbance Supersonic Nozzles.AIAA Paper 85-1573, 1985.

33. Malik, M. R.; Wilkinson, S. P.; and Orszag, S. A.:Instability and Transition in Rotating Disk Flow.ArM Journal, Vol. 19, No.9, September 1981,pp 1131-1138. Also, AIAA Paper No. 814225.

34. Malik, M. R.; and Poll, D. I. A.: Effect ofCurvature on Three-Dimensional Boundary-LayerStability. AIAA Journal, Vol. 23, No.9,September 1985, pp. 1362-1369.

35. Malik, M. R.: Prediction and Control of Transitionin Hypersonic Boundary Layers. AIAA PaperNo. 87-1414, June 1987.

36. Beckwith, I. E.; and Moore, J. A.: An Accurateand Rapid Method for the Design of SupersonicNozzles. NACA TN-3322, 1955.

37. Beckwith, I. E.; Ridyard, H. W.; and Cromer,N.: The Aerodynamic Design of High MachNumber Nozzles Utilizing Axisymmetric Flowwith Application to a Nozzle of Square TestSection. NASA TN-2711, 1952.

38. Chen, F.-J.; Malik, M. R.; and Beckwith, I. E.:Advanced Mach 3.5 Axisymmetric Quiet Nozzle.AIAA Paper No. 90-1592.

39. Thomas, J. L.; Walters, R. W.; Rudy, D. H.; andSwanson, R. C.: Upwind Relaxation Algorithmfor Euler/Navier-Stokes Equations, NASA CP­2397, pp. 89-107. 1985.

40. Smith, T. S.: Experimental Investigationsof Compressible Air Flow Through FlatHomogeneous Porous Plates. AIAA-85-1571.

41. Dillon, James L.; Trimpi, Robert L.; and Wilcox,Floyd J., Jr.: Unexpected/Expected ResultsFrom the Langley 2Q-inch Supersonic WindTunnel During Initial Checkout. AIAA Paper88-1999-CP.

42. Wilkinson, S. P.; Beckwith, 1. E.; and Chen, F.-J.:Status of Langley Quiet Tunnel Developments forTransition Research (Unclassified). Eighth Na­tional Aero-Space Plane Thchnology Symposium,Monterey, CA. March 26-30, 1990.

43. Beckwith, I. E. and Miller, C. 0, III.: Aerother­modynamics and Transition in High-Speed WindTunnels at NASA Langley. Annual Review ofFluid Mechanics, Vol. 22, 1990. pp. 419439.

44. Lukasiewicz, J.: Diffusers for Supersonic WindTunnels. Journal of the Aero. Sci., Vol. 20, No.9, September 1953, pp. 611-626.

45. Modern Developments in Fluid Dynamics, Vol.II, Diffusers, pp. 501-507, Editor, L. Howarth,Oxford at the Clarendon Press, 1953.

46. Hermann, Rudolf: Supersonic Inlet Diffusersand Introduction to Internal Aerodynamics.Minneapolis-Honeywell Regulator Co., Aeronau­tical Division, Minneopolis, MN, 1956.

Tablt! 1Nominal RlIn 1'imt!s lor IlSIJ1Jl'

Run to atmosphere Run to spheres

A*, Diffuser I

To. tit '''Po, Ts, t,in 2 Type* PolPo Po, t,

MFig. 21 psia op Ib/sec. min. psia OF min.

A 3.1 47 80 135 21.1 40 80 8.23.5 95.4 B 4.0 60 172 16.5 t ! 6.3

C 8.3 124 355 8.0 3.0A 18 266 450 75 37.8 200 100 13.8

6.0 12.2 B 25 370 " 104 27.2 ~ l 9.9r lin - - - An-

·A: Variable area with small blockage model or probe (Cloged tegt section)B: Variable area with typical model and strut (Closed test section)C: Open jet test gectJon

11

--------_. -------_._----_.__ ..

Stability analysis condillons:Po" 300 psla, To= 360"F

R",,, 5.14 x 105nn

Transillon:GorllerN ,,9

TS N ,,3.6, Tw" Taw

8

6

N

-- Mach 3.5 quietwind tunnel 4

. rtlght

Tunnel Size ~t;> ~~s_~aToRC 6"i10" 3.5 -0

Onset {~ ~ ~ ~ SmallModerate

High{O AEDC 0 12" x 12" 3.0 H~hestEnd .1 ... A 40" x 40" ... Igh

2 o PWT 16' x 16' High

----- Mach lines

Throat diameter" 1.002 InchesExit dla",eter " 7.238 Inches

l--L-'-_'--.L_..J

o 2 3 4 5

F x105

2

10

12

Flow

N

THROAT TO EXITlENGTIl = 29.91 In.• EXIT OIA. = 6.86 In.. 1HROAT OIA. = 2.61 In.

6.990 Radial Flow XT, w - -- Mach lines

Flow --••• X07 __ :""-::.:~::_~.• ~.--L-_-I-_....:.J.::.~.L..d._---'-_-J:~~ 0 S W 6 m n my"

x, dl~I"nc(! from throat. In. Xc, T- '

/ Transillon predicted lor Gorller N "9.1. To = 5300R

Po Roofln. XT,w XT,c 6X= R6X TSN

p,la ><10"6 Inches Inches Xc. T -Xo.ln. ><W·6

100 1.03 2357 34.89 17.09 17.6150 1.55 21.11 31.40 14.60 22.6 1.3

Figure 6.- Predictions for quiet test core lengths. ~X. inthe new Mach 3.5 axisymmetric-long nozzle,

2

:-''--+--'''8~-~ 16 20 24Rei) x 10.2

Figure 4.- Effects of noise levels and leading edge bluntnesson flat plate transition.

N=lnA/Ao

TS waves on cones31 TG vortices on nozzle wall 3212 UCD/VCD

x 1O-6 /m010020t..320 40660

o I __LLJ---l

.4.6.81 2 4 6810 20()J/ilT

XVortex wavelength'rf Boundary·layer thickness

Figure 5.- N-factor calculations at measured locations oftransition on cones and the Mach 3.5 pilot nozzle wall.

--1.__-,-_.0-I 0

"'--------Figure 7.- Quiet test core prediction in the Mach 6axisymmetric-long nozzle.

Wind tunnel

LaRC low disturb.

AEDCJPL20 in.

Flat plnte predictionfor N ~ 10---~

Conepredictio~for N = 10 ~-_Jvo

Mach no.3.5 ,3.5 I3.03.7

_-- Tunnel A (40· x 40·)

_-----::.------:: JPL20"•-:::::.-- - "'-- Tunnel 0 (12" x 12")

L..::,----'--"'-"<---'--L--'-L--------12 4 6 8106 2

Refin.

2

M", " 3.5 Pilot quiet tunnel

XS , In. Noise ::.::::::..

a 5 lowest ~:::--::::.$ 8 lowIJ. 5 High ~ Flight data20v 8 Highest ~ Me" 1.4 to 4.6

Flight data 24 ~-Me",.6to,.8~~8!l

::-:-

ReRn.

ModelCone

Flat plateFlat plateFlat plate

2

ao

2

107. 8

Re,T 6

4

2

Figure 3.- Transition onset Reynolds numbers on cones inwind tunnels and in flight.

Figure 1.- Quiet test core in Mach 3.5 pilot quiet nozzle.

Figure 2.- Transition onset Reynolds numbers on conesand flat plates.

12.

11 .,::111'1"1'

Inneciion po'nt

X Inches 75.3 9~~llJir!I~,-NTS N 5 6.3

5 ~.'nch·l 67.1 89.6~~X10' 24.8 33.1R. 3280 3560

Radla' now

Figure 12.- Mach 8 quiet nozzle for M8VDT. .

(a). Nozzle contour and predicted quiet test core.

~_ ,-- 'L-.-_.-L......-__---l......_.-c':',------_:C:I,-,-='_--;t~-o 4 8 12 16 20 24 28 32

X,lnehesTr8nsl1l0n

1\,: 1000 PS'ATo" 1360'RTwlTaw= 0.73,R_no. = 3.1x10

',InchetSlot·· kc20Jl.lnchelLE. 6.5'

-"o..p·'·o 1000 From boundary. lay.r .olullon

n 500)• 200 Eq. (3)

.",,~:I~l.E.;L, . , . , . .-2.0 -1.0 0 1.0 2.0 3.0

X. dl.tanc",,{llm throat.lnche.

Figure 11.- Allowable surface roughness in throat regionof Mach 8 nozzle for M8VDT.

f,lnches

20~ exlldl•. :

.

.. Mach II... ~ __ .. * . -- 17.98 In...- - -••.•_ M~8

-~~r~~;- 2o·-k::~=ro·;,;:::=;~:::~:;:-~:.-;~:018. = 1.25 In.

Test section,and injection box

30" Horizontal stroke

L-----tl---

Nozzle

I!I

po.

psla

o 25

" 51o 75Ll 10)a 127c 154o 179o 20).6

M.ch NOlll. M.xk R.1.2 no. --!rl'!.. ":~,,oT- Axl. R.E. -20 32

I!l 3.5 2-D R.E. 100 IOrlgln.1 12003.5 2-D R.E. 40 blOCk.~ 3.5 2-D R.E. 205l New l2203.5 2-D R.E. 20 block.

~ III... 3.5 Axl. long

100 ~I> 3.5 AX/. long -60 38"'3.5 Axl. lOng}"'3.5 2·0 HSlOT ~..ory lor

t!l 6l 6 Axl. long GOt1lor N~ 9~

m 6 AX/. long -150

105 2 4 6 8106 2

Relin.~igure 9.- Quiet-test-core length Reynolds niimbeiS in fournozzles compared with transition data on cones in flight.

To ~ 84"F

RJln. Xr.C' ilX. Roo.ilX

x10-5 In. In. xlO-6---- --_.-2.5 11.0 12.0 2. 9~5.0 IH 8.9 U)U 12.1 7.1 5.21

10.1 11.1 6.1 6.1J12.5 10.8 5.8 1.2)15.1 9.9 H 1.4017.5 9.8 4.8 8.4019.9 9.7 4.7 9.36

~oagtii§i{> ~ 6 0

- - - - - - - 0-0 -­

!.II.dI~.......r.,8~~tI0~~12;!±-J'-"';1!14-U.J'""t16"""'~18,-o........;20O's'anc~ from throat. Xlin. I

-Figure 8.-Nonnalized nns pressures from hot-wire surveyson the centerline of the Mach 3.5 pilot quiet nozzle (Fig. I).

4 .

Verticalstroke

Figure 10.- Modifications 10 the Mach 8 Variable DensityTunnel 10 provide quiet test flow,(a). Elevation view of tunnel.

Rlglmesh coneand 2 plales 4 Screens

Figure 10.- Concluded.(b). Settling chamber and nozzle boundary-layer bleedsystem.

Bleedplenum

13.

-------- ...•----_.....

Figure 13.- Concluded.(b). Settling chamber and boundary-layer bleed system.

Po" too 1'0i .... To '" 1000 of • T. - Taw2

5

,. Modllled chAmbe,2. 'n'e. dlllu.e'3. Flowline'4. OI11lm..h pIA'e. (2)5. 9c,e,n. (2)

8. """ulnt b'eerl ring7. "Illhly.poll.hed nonle

Ihront s@clfon8. rr~~9Urf! port for ro9. ",hAlI.llo bleed manifold

Figure 12.- Concluded.(b). Mach number contours in throat region from Navier­Stokes solutions39•

1 4 3 2

~

1. Existing settling chamber. see figure13(b) for modlflcallons.

2. Mach 20 conical nozzle; 5' half-angle.3. ConIcal nozzle, highly polished throat

section.4. Inlet diffuser for exlsllng setlllngchamber. See figure 13(b) formodifications.

5. Test section enclosure.6. Diffuser.7. Model strut. .8. Model (cone tangent flare

shown)9. Model support stand (fixed).

10. Exhaust pipe to vacuumsphere.

Figure 13.- Modifications to the Mach 18 Helium Tunnelto provide hypersonic quiet test flows.(a). Elevation view of tunnei.

(AXisymmetric)

~I~.) 0.005"b Conical nozzle

~10 ~0::===2J:0==::L:=:4:;:0====6~0====80(ln.)X

Figure 14.- Conical nozzle and annular bleed ring.

14.

~ ,d

'.'

______..1i••t,~'~,·,·

.... 2 r 7t11

. -~~8-1- ~~3 4 6.... I'--rrr 8 f~O

~ !i JT::Qu TI I ~

~

I~I 111111 113 ..l1 8'v ~

([1 1() I /1 ~ II

n 1/ I ~ II ~

High speed low-disturbance tunnel1. High pressure filtered air.

Mass flow to 340 Ib/sec.2. Settling chamber.3. Rigimesh cone entrance diffuser.4. Two Rigimesh plates & supports.5. Honeycomb.6. 8 screens.Figure 15.- Elevation view and major tunnel componentsof High Speed Low-Disturbance Tunnel (HSLOn.

7. Bleed manifold.8. Interchangeable nozzles.9. Two-stroke model-injection

mechanism.10. Variable-area diffuser.11. Exhaust to vacuum spheres.12. Exhaust to atmosphere via

trench and muffler tower.

Figure 16.- HSLOT settling chamber.(a). Internal details.

STA -336

ISTA 0

I

-'" .".,,- -,,-- '-. . ". " ",--,,-- '\

a and b : Rlglmesh platesc: Rlglmesh cone

NoSIOdsoellons ·90

·70

~"I

M~VorlleolC!' -

« •• It ! • I , ,

-90 ·70 ·50 -30 ·10 STA 0

~

s·~.~·90 ·70 ·50 ·30 -10 STA 0

FIgure 17.- Converging nozzle centerplane contours andcross-sectlons for the HSLOT.

Pilot Quiet Tunnel

Symbol Nozzle Bleed Valve

o M =3.5 2·0 Open I<I>I> M =3.5 2-D Closed I<> M =3 Axlsym. Open I®.d M =3 Axlsym. Closed I

'-n = 1.87<I> Rigemesh cone,

2 Rigemesh plales,Honeycomb, 8 screens.

® Same as <D plus1 sheel of 1micronfiller paper.

102

.1 2 34567890

2 3456789',10 10 10

Specific mass flow, W,lb/sec ,,2

Figure 16.- Concluded.(b). Pressure drop across settling chamber components.

,(p/_ po2), 2

(lbnn 2)2

10397654

3

2

15.

•

Figure 18.- Contours and predicted quiet test cores oftwo-dimensional nozzles for the HSLOT.(a). Mach 3.5 nozzle. 32o

100

200

k,lI'lnches

500

400

Rk =10, To " 9100 R

300

V,ln.

Z,In.

70

I 20l" 65.4 In.

20 30 40 50 60X. dl.tance trom throat. Inches

10o

M••w " 2 2.5 '\00: Sidewall

·10

o 20

""" Sidewall

40 60 80X. distance Irom throat. Inches

Figure 18.- Concluded.(b). Mach 6 nozzle.

V" 9.318 In.10

Y.ln.o

10

Z.ln.I 20

l ,,93.6 In.

100

X, dIstance from throat, Inches

Figure 19.- Concluded.(b). Mach 6 nozzle.

Sta. 0.0 lor

~M6nOZZle

Sta.O.Olor!,,-i M3.5 nozzle

30.~1._

1 ~108' . ". I

~.[~=~1.--- ------;~o: -----_.IElevation View

X, Distance from throat, Inches

Closed test section: verlAbie areaFrom llik8slewlcz43

<> 1.t~ 1'1.1" tun".t.D1ogIM. N.O.l. (1951)

n 4.r I[ 4.7"lUnMl. Wl"~n., I tobb.N.O.l.(1952)

6. 2" .. 2.5"' tun".I, Hf!p~.Galeh (1947)

'11IIIII 11" .. "" fUm"'. S.l1rltm. N.A.C.A. (H150)

• t.3"1r 1.3"lunrtel. ""um_" & luttw"tII'."'.1.T.(1951)

OponlotX l8RC.2-Dqul@tIttnMI(19e3)

o W.Y.A.• Kochel (1152)44

4=:.....-:3:-'-~~,-"'-:~~'~6~g110Mach number

Figure 20.- Model injection system for HSLOT.

Figure 21.- Ratio of stagnation to diffuser exit pressures inwind tunnels.

400

200

'gg Norme' shoek -

60(thoory)

40 Hormel.hoek at --mlltX contn'ctlon

Po & lse-nlroplc flow"0' 20

~on leI de.lgn , I

'8 owerlh 1'9531~48

/7

4 WlIhmodel·· 7& support 7

~

2·

105o

500 f400

300

200

100Slot l.E.

o-5

k, ~l-inches

Figure 19.-Allowable surface roughness in nozzles for theHSLOT.(a). Mach 3.5 nozzle.

16.

----------_._~

6" Servo·hydrAuliccylinder wi 3" rod

,2'·6"

t

"I

_diM

Figure 22.- Plan view cross section of HSLDT diffuser.

f!STIMATP.D11JNNeL CURRENT CONsTRUCTION COMMeNTS

COMPLeTION DArn

FinAl de.lgn AndNo .eriotl. developmenl or

"SLOT" MAch 3.S November 1992 operAllonAI problem.englneeriog

Identified

FinAl design AndMAch 6 contracllon/no..le

HSLOT " Mach 6 1993 may not be Included Inengineering

ctllTenl projecI

No..le I. con.ldered

MgVOTI'Innl de.lgn and

My 1991 re.earch hem. Complellooengineering dale I. for In'lallAllon of

fi ..1 le.1 nozzleTunnel "Art problem for

MlgQ11T Under con.lrtJcllon Seplember 1990flAred cone. mAy reqtlire

AdditionAl Itlnnelmodification.

MAch 6 NTC PoII.h worlc underway My 1990No....le reqtlire. Improved

finhh

Figure 23.- Development and construction schedule forquiet tunnels at Langley.

I"

17.