Qualitative Assessment of the Acoustic Disturbance Environment in the NASA LaRC 20-Inch MACH 6 Wind Tunnel

12 Pages·2001·0.88 MB·English

by  Horvath, Thomas J.

STAI 8.1
QUALITATIVE ASSESSMENT OF THE
ACOUSTIC DISTURBANCE ENVIRONMENT
IN THE NASA LaRC 20-Inch MACH 6 WIND
TUNNEL
Thomas J. Horvath,
Scott A. Berry, and H. Harris Hamilton
NASA Langley Research Center
Hampton, Virginia
Supersonic Tunnel Association, International
95 th
April 29 - May 2,2001/Hampton, VA.

QUALITATIVE ASSESSMENT OF THE ACOUSTIC DISTURBANCE
ENVIRONMENT IN THE NASA LaRC 20-1nch MACH 6 WIND TUNNEL
Thomas J. Horvath'. Scott A. Berry*, H. Harris Hamilton ''_
NASA Langley Research Center
Abstract
An experimental investigation was conducted on a5-degreehalf-angle cone with aflare in a conventional Mach
6 wind tunnel to examine the effect offacili O, noise on boumtar 3"laver transition. The effect of tunnel noise was
inferred by comparing transition onset locations determined from the present test to that previously obtained in a
Mach 6 quiet tunnel. Together, the two sets of experiments are believed to repi'esent the first direct comparison of
transition onset bem,een a conventional and a quiet hypersonic wind tunnel using a contmon test model. In the
present conventional hypersonic tunnel experinwnt, adiabatic wall temperatures were measured atut heat tratL_'er
distributions were inferred on the cone flare model at zero degree angle of attack over a range of length Reynolds
numbers (2xlO e'to lOxlO _) which resulted in laminar eau.tturbulent flow. Wall-to-total temperature ratio fi_r the
transient heating measurements and the adiabatic wall temperature measurements were 0.69 amt 0.86, respectively.
The cone flare nosetip radius was varied from 0.0001 to O.125-inch to examine the effects of bluntness on transition
onset. At comparable freestream conditions the transition onset Reynolds number obtained on the cone flare model
in the conventional "noisy" tunnel was approximately 25 %lower than that measured in the low disturbance tunnel.
Nomenclature w wall
d base diameter (in)
h heat transfer coeff. (Ibm/ft2-sec), q/(H,_- H,0 Introduction
where H,,,,= H,._, Thermal effects resulting from boundary layer
H enthalpy (BTU/lbm) transition during hypersonic flight can represent an
L_f reference length based on sharp tip model (in) important constraint for optimum reentry vehicle
M Mach number design. Design strategies for aerospace vehicle
P pressure, psia Thermal Protection Systems (TPS) that include
q heat transfer rate (BTU/ft:-sec) empirical methods derived from ground based
q dynamic pressure (psi) measurements are generally considered conservative.
R radius (in.) This is based upon the fact that high disturbance
t time (sec) levels in the freestream flow of conventional
Re unit Reynolds number (l/ft) hypersonic ground based test facilities can cause
S_f reference area (in2) transition to occur at lower values of Reynolds
T temperature (°F) numbers than may actually take place in flight t.
x axial distance from cone apex (in.) Quiet wind tunnel technologies developed at, arfl
angle of attack (deg) incorporated into, wind tunnels at the NASA LaRC
0 cone half angle (deg) in the 1980's and 1990"s have clearly established the
p density (lbm/in 3) influences of noise on transition and have "advanced
"_ ratio of specific our understanding of transition via stability
Subscripts experiments conducted in these t.a...Cllltles'-._7 By
aw adiabatic wall eliminating the overpowering effects of tunnel noise,
b base these studies conducted in the low disturbance
f flare supersonic facilities have shown that boundary layer
free-stream conditions transition develops from a linear growth of natural
n model nose instabilities that can be modeled by linear theory.
s surface quantity Despite advances in quiet wind tunnel
T transition onset technology, relatively few of these low disturbance
t, I reservoir conditions supersonic or hypersonic facilities exist. Those in
2 stagnation conditions behind normal shock existence today are typically deficient in Reynolds
number relative to representative flight conditions,
and are generally not operated in a manner conducive
*AerospaceTechnologist, AerothermodynamicsBranch,NASA for aeroheating assessment/screening studies. Thus,
LangleyResearchCenter, Hampton,VA. conventional hypersonic wind tunnels continue to
Presentedatthe95'hMeetingoftheSupersonicTunnel serve as the primary source for experimental data from
AssociationInternational, Hampton,VA.,April29-May2.2{_)1.
The documentherein isnonreferencable. which to develop empirical methods TM for fight
transitionprediction.Furthermorweh,entransition the Mach 6 NTC Quiet Tunnel. Together, the two
bypassmechanism(as termcommonlyusedto sets of experiments are believed to represent the first
identifytransitionmodeswhichbypassthe linear direct comparison of transition onset between a
processsu)chasvehiclesurfacreoughneasrsepresent, conventional and quiet hypersonic wind tunnel using
it hasoftenbeenarguedthatfacilitynoisefrom a common test model at comparable freestream
conventiontaulnnelshaslittleeffect.Experimental conditions.
studieshavesuggestethdatwhilenoisemayhave
Experimental Methods
littleeffecftorroughneshseightslargeenoughtobe
considereefdfective(_tu2rbulencineitiatedimmediately
downstreaomftheroughneesslemenstite),therestill Details of the cone flare model (designation 93-
maybe an influenceof windtunnelnoiseon 10) originally built and instrumented for testing in
transitiononsedtataderivefdromroughnesthsatare the LaRC Mach 6 NTC Quiet Tunnel can be found in
lessthaneffective
]3_ Ref. 3. A schematic diagram of the model geometry
is shown in Fig. 1. Measuring 20-inches in length
Present day TPS technology suggests that
(sharp tip) the model base diameter is 4.6-inch. The
boundary layer transition on reentry vehicles will
first 10-inch section of the model consists of a 5
continue to be roughness dominated. Traditional
ceramic TPS tiles such as used on the Shuttle Orbiter degree half angle cone followed by a 10-inch sectirn
comprised of an outward flare. Tangency was specified
often suffer launch-induced damage and/or develop
at the cone flare junction. An arc radius of 93.07-
protruding gap fillers. Both forms of local surface
inches defined the flare curvature. As discussed in
roughness have been responsible for the occurrence of
Ref. 3, the purpose of the flare was to promote
early boundary layer transition in flight _6. Stitching
boundary layer instability via an adverse pressure
patterns found on thermal blankets produce another
gradient to insure transition would occur on the model
form of local roughness. Metallic TPS panels that
within the limited quiet flow Reynolds number
were proposed for use on the X-33 _tTts could have
capability of the Mach 6 NTC Quiet Tunnel.
been susceptible to thermally induced
expansion/bowing producing roughness in the form The 15-5 stainless steel model was constructed
of a wavy wall. Based upon experimental evidence t_ with a thin wall to reduce surface heat conduction
_ssuggesting the susceptibility of less than effective effects and to bring the model to thermal equilibrium
roughness elements to acoustic disturbances, as quickly as possible during a run. The model was
quantification of the conventional facility disturbance constructed with five interchangeable nosetips (R,=
environment is essential. 0.0001,0.03125, 0.06250, 0.09375, and 0.!25-inch)
fabricated from 13-8 stainless steel (see Fig. 2a). The
The purpose of this paper is to present
surface finish was originally highly polished to
preliminary experimental results from a study that is
minimize effects on the model boundary layer
part of a larger ongoing effort to assess the acoustic
disturbance environment of the NASA LaRC 20-1nch stability and, for the present test, appeared to be free
Mach 6 Air Tunnel. This initial experimental study of large surface scratches (despite several entries into
the quiet tunnel and the subsequent long-term
is considered qualitative and no attempt was made to
measure freestream fluctuations. The relative storage).
disturbance environment of this conventional tunnel Model Instrumentation
was expressed via differences in transition onset A total of fifty-one type K thermocouples were
locations measured on a cone flare model previously tack welded to the model thin wall backside along a
tested in the LaRC Mach 6 Nozzle Test Chamber
single ray. The thermocouples were spaced axially
(NTC) Quiet Tunnel. along this ray at 1-inch intervals between model
In the conventional hypersonic tunnel experiment stations x=2 and 9 inches and 0.25-inch intervals
two independent measurement techniques were used to from model stations x=9 to 19.75-inches. The local
identify transition onset. Adiabatic wall temperatures model wall thickness was nominally 0.030-inch
were measured and heat transfer distributions were along the location of the streamwise row of
inferred on the cone flare model at zero degree angle of thermocouples and 0.060-inch elsewhere. Twenty-
attack over a range of length Reynolds numbers nine 0.040-inch diameter static pressure orifices were
(2x106 to 10xl06) which resulted in laminar and located on the opposite side of the model.
turbulent flow. Wall-to-total temperature ratio for the
Facility Descriptions
transient heating measurements and the adiabatic wall
Although the Mach 6 NTC Quiet Tunnel has
temperature measurements were 0.69 and 0.86,
since been disassembled, at the time of the original
respectively. The cone flare nosetip radius was varied
from 0.0001 to 0.125-inch to examine the effects of quiet tunnel experiments, both facilities were located
bluntness on transition onset. Transition onset within the same lab and shared a common air supply
line and heater element. For a short period, both were
locations inferred from the present test were compared
managed under the Aerothermodynamic Facilities
to onset locations previously determined from tests in
Comple(xAFC).Thiscomplepxresentlcyonsistosf
of the quiet flow envelope and measurements of the
4 hypersoniwcindtunnelsthatrepresenatlarge disturbance environment are discussed in Ref. 20.
fraction of the nation's conventional
aerothermodynatmesictcapability. Collectively,
theyprovidea widerangeof Machnumberu,nit Test Conditions and Setup
Reynoldnsumbear,ndnormaslhockdensitryatioj_
Reservoir and corresponding free stream flow
Thisrangeof hypersonsicimulationparametei_s conditions tbr the present tests in the LaRC 20-Inch
duei,npart,totheuseoftwodifferenttestgase(sair, Mach 6 Air Tunnel are presented in Table 1. The
andtetraflouromethathneer)e,bmy akingthefacilities freestream properties were determined from the
uniquenationaal ssets.TheAFC facilitiesare measured reservoir pressure and temperature and the
relativelysmallandeconomicatol operateh,ence measured pitot pressure at the test section. The
ideallysuitedforfast-paceaderodynampiecrtormance standard procedure used to compute flow conditions
andaeroheatinagn,dtransitionstudiesaimedat for the AFC facilities uses the viscosity formulation
screeninags,sessinogp,timizinga,ndbench-marking given by Chapman-Cowling and is detailed in Ref.
(whencombinewdithcomputationflauliddynamics) 21. For the present test, the computed Reynolds
advanceaderospacveehicleconceptsand basic number was based upon Sutherland's formulation lbr
fundamenftlaolwphysicrsesearch. viscosity to maintain consistency with the method
20-1nch Mach 6 Air Tunnel: Heated, dried, employed to compute quiet tunnel conditions.
and filtered air is used as the test gas. Typical In the present test, the ratio of model base area
operating conditions for the tunnel are: stagnation to tunnel cross sectional area for the flared cone model
pressures ranging from 30 to 500 psia; stagnation was 0.04. In both wind tunnel experiments a base
temperatures from 760-deg to 1000-degR; and mounted cylindrical sting supported the cone flare
freestream unit Reynolds numbers from 0.5 to 8 model. Model angle-of-attack and sideslip were set to
million per foot. A two-dimensional, contoured zero in the tunnel using a combination of an
nozzle is used to provide nominal freestream Mach inclinometer and a laser alignment system. A
numbers from 5.8 to 6.1. The test section is 20.5 by photograph of the ting supported model is shown in
20 inches; the nozzle throat is 0.399 by 20.5-inch. A Fig. 2b. Details of the model installation in the
bottom-mounted model injection system can insert NASA LaRC Mach 6 NTC Quiet Tunnel can be
models from a sheltered position to the tunnel found in Ref. 3.
centerline in less than 0.5-sec. Run times of up to
Test Techniques
15 minutes were achieved for the adiabatic wall
temperature measurements. For the transient heat Adiabatic" wall temperature: In the
transfer tests, the model residence time in the flow original stability experiments conducted in the low
was limited to 20 seconds. A detailed description of disturbance tunnel, the individual thermocouple
this facility may be found in Ref. 19 temperature measurements were monitored with time
and used to determine when the model had obtained a
In an attempt to attenuate noise from upstream
state of thermal equilibrium. The resulting
piping/air control valves, the settling chamber was
temperature distribution was used to identify
recently retrofitted with a series of acousticall2_
transition onset. The test procedure for the quiet
damping porous screen elements. This technology-
tunnel measurements involved preheating the model
was based upon quiet tunnel experience at LaRC and
by exposing it to hypersonic conditions for 30 to 60
was designed to reduce pressure fluctuations in the
minutes to achieve thermal equilibrium. For the
settling chamber to approximately 0.005 % of the
adiabatic wall temperature measurements in the
stagnation pressure.
conventional tunnel test, the test procedure was
Mach 6 Nozzle Test Chamber Quiet designed to approximate as closely as possible this
Tunnel: Heated and dried air was used as the test gas. technique. That is, the model was injected into the
Typical low disturbance operating conditions for the hypersonic stream and allowed to thermally
tunnel were: stagnation pressures ranging from 80 to equilibrate. Higher mass flow rates with the present
130 psia, stagnation temperatures up to 810R, and a conventional tunnel tests limited total run times to
maximum freestream unit Reynolds numbers of 2.8 approximately 12 to 15 minutes depending on the
million per foot. A contoured axisymmetric slow desired Reynolds number. Temperature time histories
expansion nozzle was used to provide a nominal obtained during each run were monitored and indicated
freestream Mach number of approximately 5.9. The when thermal equilibrium was achieved. The model
nozzle exit diameter was 7.49 inches with the flow was not allowed to cool between runs in order to
exhausting into an open jet test section; the nozzle accelerate the time required to reach thermal
throat diameter was 1.0-inch. This facility had no equilibrium on subsequent runs.
model injection system. Run time for the adiabatic
Transient thin skin heat transfer: The
wall temperature measurements varied between 30 "and thin wall construction of the cone flare model made it
60 minutes. Details concerning the facility, the size
3
possiblteoapplythetransienthtinskincalorimetery
model was originally designed to measure adiabatic
measurementtechniqueto infer heattransfer
wall temperatures and wall thickness measurements
distributionsalongthe streamwisearray of critical for inferring accurate heating magnitudes for
thermocouples.Thelackof a modelinjection CFD benchmarking were not documented. However,
systemandplacemeonftthemodeilnsidethenozzle as the goal of the present test was the determination
preventetdhis typeof heatingtechniquteo be of transition onset, the heating distribution trends
exploitedin the lowdisturbanctuennel. In the were more than adequate for indicating the departure
currenctonventiontaulnnetlest,separattuennerluns from areference laminar boundary layer state. For the
wererequiretdo obtain transient heating data and transient thin skin heating measurements, the overall
adiabatic wall temperature data. For the transient experimental uncertainty is believed to be better than
technique, the hypersonic stream conditions were +-20%. Conduction effects near the model base may
established with the model in a sheltered position. have contributed to a larger uncertainty in this region.
The model was then injected into the flow and Repeatability for the normalized laminar heat transfer
thermocouple temperature time histories were measurements was found to be generally better than
acquired. The model was allowed to cool between ±4%.
runs in order to obtain isothermal conditions The ESP pressure measurement system was
necessary for the calculation of heat transfer. calibrated prior to each run. The measured surface
Surface static pressure: Surface pressure pressure was expressed in nondimensionat form,
measurements were obtained concurrent with the P)P_, and in terms of a pressure coefficient.
adiabatic wall temperature or heat transfer data and Measured reservoir values of Pt,1 and Tt,l are
were made with an electronically manned pressure believed to be accurate to within _+2 percent.
(ESP) transducer. The full-scale range of the Uncertainties in model angle-of-attack and sideslip are
absolute pressure transducer was 0.36 psia. The long believed to be ±0. Idegree.
run times associated with the adiabatic wall
temperature measurements provided more than
Results and Discussion
adequate time for settling. The relatively short test
times associated with the transient heating Conventional Tunnel
measurements did not provide enough settling time. Normalized surface pressure distribution for the
Flow Visualization: Flow visualization in sharp tip (R,=0.0001-inch) at ReL=4.8 x 106is shown
the form of schlieren was used to complement the in Fig. 3. The flare section produced a nearly
surface temperature and heating measurements. The constant adverse pressure gradient. While not the
LaRC 20-Inch Mach 6Air Tunnel is equipped with a focus of the present analysis, the nondimensional
pulsed white-light, Z-pattern, single-pass schlieren pressure gradient magnitude [d(P)PD/d(X/L)] from the
system with a field of view encompassing the entire present test was within the range produced by a
test core. The model length did not permit an entire family of cones tested to examine the effects of
view of the cone flare length. The light source was favorable and adverse pressure gradients on transition
pulsed for approximately a 3 ms duration. Images zone length 23. The value of the nondimensional
were recorded on a high-resolution digital camera and pressure gradient [d(PJP_)/d(X/L)] on the flare section
will be enhanced with commercial software for future was obtained by linear regression of the data and
analysis. calculated to be 3.3. As expected, the pressure
coefficient (not shown) distribution along the entire
Data Reduction and Uncertainty
model was in agreement with previous quiet tunnel
A 16-bit analog-to-digital facility acquisition pressure measurements reported in Ref. 6.
system was used to acquire data. The facility, model
Normalized heat transfer distributions for the
thermocouple, and pressure data was collected by this
sharp tip (R,--0.0001-inch) cone flare model are
system at a rate of 50 samples per second over a 20
presented in Fig. 4, over the length Reynolds number
second interval during each run. The raw data was
(ReL) range 2.0 x 106 to 10.3 x l0n. The onset of
transferred to a Hewlett-Packard 9000 computer for
boundary layer transition at any given Reynolds
data reduction and storage.
number has been interpreted as the departure from the
Thermocouple mV output was referenced to an laminar distribution measured for the ReL=2.0 x l0n
electronic ice point and was converted to engineering condition. The collapse of the heating distributions
units via a standard type K lookup table. Accuracy
with Reynolds number on the cone section and onto
of the surface temperature measurement is believed to the flare indicated the presence and extent of laminar
be better than --5 degrees. Heating rates under flow. The onset of transition was observed for
transient conditions were calculated from backside
Ret=3.8 x 106 on the flare section and progressively
surface temperature measurements as discussed in moved forward onto the straight cone section with
detail in Ref. 22. The inferred heating rates were increasing Reynolds number. The transition zone
normalized to a theoretical reference value. The
extendetodthebasoeftheflareatReL=3.x8106and
temperature ratio of 0.69 and 0.86 had little, if any,
4.8x 106.AtthehighelrengthReynoldnsumbers,
impact on the location of transition onset whether it
Ree=7.x5l06and10.3xl0s't,ransitioonnsemt oved
occurred on the straight cone or flare section.
ontothestraighctonesectiont:hissuggestthsatthe
While primarily intended to provide data for
initialinstabilityandgrowthoccurreidn the zer_
comparison to the quiet tunnel results, the present
pressurgeradienrtegion.Theheattransfeartthe
conventional tests also served to assess the effects of
higheRr eynoldnsumberesxhibitetdhesameclassic
nose bluntness on transition onset. The effect of
transitionaolvershooatndsubsequerenltaxatiotno
turbulenletvelsastheconeflaresresultosfRef.23. nose bluntness at Ret=3.8 x l06 is shown in Fig. 7.
As expected, nosetip bluntness delayed boundary layer
Aftermanyyearsoftestingandhandlingth, e
transition. Transition onset for the sharp tip
qualityof thesharpnosetip(R,=0.0001-incwha)s
(R,=0.0001-in.) occurred at x=14-inches with the
initiallya concern.As a second(backups)harp
transition zone extending to the end of the model. An
nosetipwasavailablem,easuremeonntsit werealso
increase in the nosetip radius to 0.125-inches
obtainedT.ransitioonnseltocationvseryconsistent
stabilized and extended the laminar boundary layer to
withtheoriginasl harptip weremeasureadndhave the end of the model.
alleviatecdoncernasstoconditioonftheleadinegdge
radius. Comparison to Qui¢_ Tunnel
Themodewl all temperatuurenderadiabatic A direct comparison of conventional vs. quiet
tunnel transition onset locations at adiabatic wall
conditionwsas,naturallyh,ighetrhanthatmeasured
undetrhetransienhteatingtechniqueT.hisdifference conditions for comparable reservoir and freestream
in surfacetemperatuirseshownin Fig. 5, lbr conditions is shown, Fig. 8, for ReL=4.8 x 106. To
reservocironditionsselectetodmatchthoseachieved the author's knowledge this represents the first
inthequiettunne(lReL=4.x8106).Thebackside comparison of transition onset location between a
walltemperatudreistributionalongtheconeflare conventional and low disturbance hypersonic tunnel
obtained1.5secondasftermodelexposurewas utilizing a common model. The transition onset
essentialclyonstan(Tt,_--'9d5egF. :T,,/T.--0.69A).t Reynolds number (ReT-based on the length to
transition onset) as measured in the conventional
thesameReynoldnsumbeurndeardiabatciconditions
tunnel was 23% lower than that inferred from the low
the wall temperatuwreashigher(250deg.F;
disturbance tunnel. Prior to this direct comparison,
T_/T,=0.86w)ithsurfacteemperatuinrecrease(3s10
the influence of acoustic disturbances oh flare cone
deg.F: TJT,=0.95a)ssociatewdithboundarlyayer
transition.A stabilityanalysiosf thequiettunnel hypersonic transition was conservatively estimated by
coneflareresults running the quiet tunnel in a "noisy" mode (diverter
2_has indicated that a variable wall
valves normally open to promote a laminar nozzle
temperature distribution at adiabatic conditions
wall boundary layer are closed resulting in a turbulent
produced little effect on the instability amplification
nozzle boundary layer). In the "noisy" mode of the
rates on the model. The theoretical results of Ref. 24
low disturbance tunnel, transition onset occurred
also indicated that an increase in wall-to-total
earlier than that measured in the conventional tunnel.
temperature ratio from 0.51 to 0.86 would increase
Interpretation of the data presented in Ref. 3 yielded a
the frequency of the most amplified first and second
42% reduction in the cone flare transition onset
mode disturbances. The calculated second nw,de
Reynolds number when the Mach 6 low disturbance
frequencies were considered high (260kHZ range) ,'uad
tunnel was operated in the noisy mode as compared to
no conclusion was reached as to whether or not they
the 23% reduction via a conventional facility.
could be excited by freestream disturbances in a given
environment. Measurements on sharp cones performed in
supersonic quiet tunnels typically yield values of Rex
Despite the wall-to-total temperature that are two to three times the values obtained in
differences in the present test, the two methods for
conventional tunnels (and nearly an order of
determining transition onset yielded equivalent
magnitude for flat plates) -'_. While this may be valid
locations as shown in Fig. 6, (representative of
at M=3.5 for sharp cones, the present cone flare tests
comparisons of transition onset for all Reynolds
at M=6 do not exhibit such a level of disparity
numbers). The onset point from the transient heating between a conventional and a low disturbance tunnel.
distributions clearly indicated adeparture from laminar
Recognition of the presence of different instability
heating (ReL=2.0 x 10") on the flare section at
modes and the influence of Mach number are equally
x=12.2-inches. The method for estimating transition
important parameters to consider.
onset via the adiabatic wall temperature consisted of
The reservoir temperature and pressure
determining the intersection of two straight lines
conditions for the quiet tunnel condition were 350 deg
passing through the laminar region and the sharp
temperature rise region as discussed in Ref. 3. While F and 130 psia, respectively. This relatively low
more subjective in nature, the onset locations agree temperature reservoir condition had been selected to
remarkably well. Differences in the wall to total reduce hot-wire overheat requirements for that test
series.Concerntshatthisquiettunneloperating facility noise has little effect. Recent studies have
conditionlietooclosetotheairliquefactiocnurve indicated that quantification of the disturbance
wereaddressinedtheconventiontaulnnelbytesting environment in a conventional hypersonic tunnel is
at different reservoir pressure/temperatureessential even when boundary layer transition bypass
combinatiosnosastoholdReLfixed.Nomeasurable mechanisms, such as surface roughness, are present.
effecton transitiononsetReynoldsnumberwere
The purpose of.the present study was to
observe(dnotshown).
present preliminary experimental results from a study
Transitiononseftorconesandflat platesat that is part of a larger ongoing effort to assess the
supersoncicondition(stakenfromRef.2)withthe acoustic disturbance environment of the NASA LaRC
presenctoneflaredataareplottedagainstunit 20-1nch Mach 6 Air Tunnel. An experimental
Reynoldsnumberin Fig. 9. Also includedis investigation was conducted on a 5-degree half angle
unpublishefladtplatetransitiodnatarecentloybtained cone with a flare in a conventional Mach 6 wind
in the conventionLaal RC20-InchMach6 Air
tunnel to examine the effect of facility noise on
Tunnel.TheMach6 coneflareandflatplatedata
boundary layer transition. The effect of tunnel noise
obtaineidnaconventionhaylpersontiucnneel xhibit was inferred by comparing transition onset locations
the,so-calleudn,itReynoldnsumbeerffectwhereby determined from the present test to that previously
radiatendoiseintensitfyromthetunnewl allboundary obtained in a Mach 6 quiet tunnel. Together, the two
layervarieswithReynoldnsumbearnddominatethse sets of experiments are believed to represent the first
smoothbodytransitionprocess.Asexpectetdh,e direct comparison of transition onset between a
presendtataalsoshowthattheconeflaretransition conventional and quiet hypersonic wind tunnel using
onseRt eynoldnsumbewrashigherthantheflatplate a common test model. At comparable freestreanl
dataforagivenunitReynoldnsumber.Thistrend, conditions the transition onset Reynolds number
commonlyobserveidn conventionafalcilities,is obtained on the cone flare in the conventional "noisy"
contrartyothatmeasureindsupersonqicuiettunnels tunnel was approximately 25% lower than that
andtothatpredictebdystabilittyheory.Asdiscussed measured in the low disturbance tunnel.
inRef.2,thismaybeduetothefastebroundarlyayer
Acknowledgments
growthonaflatplaterelativetoaconeandthus,
Without the assistance of the following
strongerreceptivitoyfthefiatplatetotheincident
acoustifcield(intheconventiontaulnnel).Theflat individuals this work would not have been possible:
Steve Wilkinson, Charlie Greenhalgh, and Kevin
platetransitiononsetReynoldnsumberfsromthe
conventionLaal RCMach6 tunnelareconsistently Meidinger for model design information and model
refurbishment; John Ellis, Grace Gleason, Rowland
highetrhantheflatplateM=3and3.7datafromthe
AEDCandJPLfacilities.Sinceit hasbeenshown Hatten, and Steve Jones for wind tunnel support;
thatthedisturbancfieeldsoundintensitycorrelates Sheila Wright and Mike Difulvio for data acquisition
assistance; Steve Wilkinson and Brian Hollis for
with testsectionsize26onemightexpectbetter
analysis support; and Richard Wheless for
agreemeonftthepresenstmoothflatplateReTdatato
documentation assistance. The authors gratefully
theJPLresults. Despitethesimilartestsection
dimensiontostheJPLtunnel,otherfactorssuchas acknowledge their contributions and behind-the-
Machnumbeerffectosnradiatetdunnewl allnoise, scenes work.
boundarlayyerstabilitoyntheconeflare,orthenoise References
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settlingchambearppeatroshaveaninfluencoenReT.
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Fluid Mechanics, Vol. 22, 1990.
Concluding Remarks
2.Beckwith, i. E., Chen, F.J., and Malik, M. R.,
Despite advances in supersonic/hypersonic "Transition Research in the Mach 3.5 Low-
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Disturbance Wind Tunnel and Comparisons of Data
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3.Lachowicz, J. T., and Chokani, N.. "Hypersonic
relative to representative flight conditions, and are
generally not operated in a manner conducive for Boundary Layer Stability Experiments in a Quiet
Wind Tunnel with Bluntness Effects," NASA CR
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198272, Jan., 1996.
conventional hypersonic wind tunnels continue to
serve as the primary source for experimental data from 4.Blanchard, A. E., and Selby, G. V., "An Experimental
which to develop empirical methods for flight Investigation of Wall-Cooling Effects on
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mechanisms such as vehicle surface roughness are Tunnel," NASA CR 198277, Feb., 1996.
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Table h Flow Conditions
Facility Mx q_('psi) Pt.dpsi) 'T,a(°F) 92/9_ ReJft (xl0*)
20-In Mach6 Air 5.88 1.08 62 4 i7 5.24 1.20
5.94 2.15 129 446 5.27 2.30
5.94 2_ 5_6 2_9
5.98 4.10 254 443 5.26 4.48
6.00 5.87 369 469 5.27 6.17
Shaded conditions represent a match to Mach 6 NTC Quiet Tunnel reservoir conditions
Rn= 0.00010, 0.03125, 0.06250, 0.09375, 0.12500
.....:.:.:..........4.i i !L..+I
Rn w
R =93.071
" A
. 2 i
4.6 04 _L 0.6 0.8 1
Origin (tip)
(0,0) Pressure V Fig. 3. Static surface pressure distribution,
taps (29) M==6, Rec=4.8 x IOn,(z=0 deg, R,--O.0001-in.
Thermocouples (51)
Fig. I. Model 93-10 flared cone model and ReL
nosetip dimensions (inches) "-0-"2.0 x 106
'"tr- 398x 106
' 94.8x 106
.-A.-7.5 x 106
•-_-- 10.3 x 106
o2
z
0 5 10 15 20
x, inches
Fig. 4. Reynolds number effect on transition onset
location as inferred from transient heating distributions,
M==6, (z---Odeg, R.,=0.0001 -in.
400
Fig. 2a. Flared cone model interchangable nosetips
35O
3OO
,u. 250
z.
200
1-
150
Fig. 5. Comparison of adiabatic wall temperature
distribution with transient measurement,
Fig. 2b. Flared cone model installation in the NASA
M_=6, ReL=4.8 X106,(Z=0deg, R.=0.0001-in.
LaRC 20-Inch Mach 6Tunnel
8