Undercontractfrom NASA, Russia’sCentralInstituteof AviationMotors(CIAM) designedandbuilt anaxisym-metric,dual-modescramjetengine.On February12,1998this engineflew on thenoseof a modifiedSA-5 missile.Itwas fueled with hydrogen for about 77 seconds,andachieved the longestduration,dual-mode,scramjet-pow-
ered flight-test up to date1,2.
1.1. Description of the experiment
Thedesignlayoutof theengineis shown in figure1. Itincludes:- an external/internal, axisymmetric, Mach-6 inlet,- a burner section with three fuel-injection stages,
Abstract
TheCIAM/NASA flight testwasnumericallyanalyzed.Theflowpath was divided into inlet and burner sections,andsolved sequentially. Initial simulation of the inlet at high-speedconditionsfailed to describethe behavior of the data,which indicatedthe existenceof separation.An analysisoflow-speedoperationshowed inlet unstart and subsequenthysteresiseffectswhich qualitatively approximatesthedata.Simulation of the burner predictsdual-modeoperation,asthe datasuggests,althoughpeakpressureswere somewhatunderpredicted.Effects of grid convergence, turbulentSchmidtnumberandchemistrymodelswereevaluated.It isconcludedthat currentCFD tools may be usedto anticipateeffects of design and construction in actual operations.
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- an expansion section with a partial nozzle.Exceptfor the external inlet, mostof the engineconsistsofanannularductbetweenthebodyof theengineandanexter-nal cowl. This cowl is held in place by two sets of struts.
Theexternalinlet beginsat thenose-tipwhich incorpo-ratesa pitot probe.The body itself consistsof threeconicalsegments with increasing half-angles.The internal inletstartsat the locationof theblunt cowl-lip. Body-sideexpan-sionsandcowl-side compressionsareusedto turn the flowparallel to the body center-line by the time it reachesadiverging isolator, upstream of the first row of injectors.
The burnerconsistsof threeinjector stages,denotedI,II, and III in figure 1,c). Each stagehas 42 injectors; theinjectors of stagesI and II are aligned,and interdigitatedwith thoseof stageIII. StagesI andII arelocatedneareachof thetwo body-sidecavities,while stageIII is locatedat thecowl step. Hydrogen fuel was intended to be injectedthroughall threestages.StagesII and III were to operateduring most of the flight regime; stageI was supposedtooperateabove Mach 5, when supersoniccombustion wasexpected.
As mentionedbefore,two setsof strutshold theenginetogether;each set consistsof four struts. The first set isplacedat the internal inlet; thesestrutshave a small cross-section,andpresumablyhave very little impacton the flowin the region. The secondset is locatedtowardsthe endofthe engine,near the exit-nozzle throat; thesestrutshave alarge cross-section.At this axial location,the cowl expandsto compensatefor theresultingareablockage.Thecowl endsby opening up into an exit nozzle.
During manufactureof theengine,the internalflowpathwasaltereddueto structuralreinforcements,weld beadsandsurfacedeformationresultingfrom thewelding.Thereis stillconsiderableuncertaintyregarding the final configuration.To add to this geometryuncertainty, post-testinspectionofthe engine showed combustor-liner deformations2; thesedeformationshave not beenquantified.In any case,the(pre-sumed)as-built pre-testflowpath was usedfor the presentcalculations.
A brief descriptionof the flight test follows. While theboosterwasstill burning,fuel additiontook placewithin thescramjet(approximately38 s into theflight) at a flight Machnumberof 3.5.Themaximum-velocity conditionoccurredatboosterburnout,at a Machof 6.4 (around56 s). After burn-out, the scramjet/missilecombinationfollowed a ballistictrajectory, with increasingaltitude(anddecreasingdynamicpressure)until a maximum altitude was reached(90 s).Afterwards,dynamicpressureincreaseduntil flight termina-tion (115s). Fueladditioncontinuedall theway to termina-
tion, except for a brief period around 90 s.
Severalanomaliesanddeviationsfrom plannedflight-testconditionsoccurred.First, thetesttookplaceatanalti-tude lower than anticipated.In particular, the maximum-velocity conditionoccurredat 21.6km ratherthan24 km;theresultingdynamicpressureandmassinflow weredou-ble the designvalues.Furthermore,an apparentfailure inthe fuel control-systemresultedin excessive fuel flowrateand engine unstart for about 12 s. The control systemrespondedby drastically reducing the fuel flowrates ofstagesII andIII. Thisallowedtheengineto restartatabout50 s (Mach 5.0). However, the possiblepresenceof largeflow-separationnearthe inlet throatcausedthecontrolstokeepstageI shut.As a result,combustiontook placewithonly stages II and III active.
1.2. Previous Work and Present Approach
Previousto thepresentwork, ananalysisof theopera-tion of theinlet designat plannedtestconditionswasdoneby Hawkins3. It predictedinlet startat theplannedoperat-ing conditions;it alsofoundthepresenceof small separa-tion bubblesjust behindthecowl lip andin thebodyside,mainly becauseof shock impingement. Gaffney andSanetrik4 performeda CFD analysisof the full engineusing the VULCAN code. The design geometry andplannedfree-streamconditionsat the maximum-velocitypoint wereused;the total fuel flow-ratewassimilar to theexperiment.To reducecomputationalexpense,the entireflow was assumedaxisymmetric; the rings of injectorholeswere replacedby axisymmetricslots of equivalenttotal area. Their calculatedMach contours within theburner suggeststhat the averageMach was subsonic,ornearly so.
The presentwork documentsan ongoinginvestiga-tion of the CIAM flight testasit actuallyoccurred,basedon thebestavailableinformation;someearly resultswere
reportedelsewhere5,6. As mentionedbefore, the as-builtgeometry will be modeled; no attempt was made toaccountfor possiblein-flight deformation.The operatingconditions chosenfor analysiscorrespondto the maxi-mum-velocity point in the flight-test trajectory. Asexplainedbefore,only stagesII andIII actuallyworkedatthe chosenoperatingcondition. Therefore,for the pur-posesof this paper, “inlet” will denotethedomainfrom apoint upstreamof the nose,to a sectionin the duct justupstreamof stageII; theremainingductwill beconsideredthe “burner”. Note that, with thesedefinitions, the inletthroatwill betheminimumareajustaheadof stageI injec-tors. The inlet and burner were solved separateandsequentially.
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TheVULCAN codewaschosenfor thepresentinvesti-gation. VULCAN is a generalpurposeCFD codethat cansolve theReynolds-averagedNavier-Stokesequations.It hasa wide array of physical, turbulenceand chemistrymodelsavailable.A full descriptionmaybefoundin theliterature7.The specificsof its applicationto the presentwork will bedescribed throughout the paper.
2. Inlet
2.1. Solution Procedure
Since the inlet geometry(not accountingfor stageIinjectors) is axisymmetric,the axisymmetric form of thegoverningequationswassolvedwith VULCAN for theobvi-ouscomputationalbenefits.Theactualgrid wastwo-dimen-sional, with approximately192,000control volumes(CV)dividedamong15 blocksto facilitatetheuseof VULCAN’ sMPI capabilities;100CVs wereusedalongtheductheight.The block configurationcan be seenin figure 2,a), with adetail of the grid in the neighborhoodof the throat alsoshown (b).The maximum wall spacingwas less than 0.03
mm; the maximumy+ wasbelow 30, andoccurrednearthethroat.The grid wasmostly C(0)-continuous,except in theneighborhood of the cowl lip.
All inflow boundarieswere set at free-streamcondi-tions, while extrapolationwasusedfor outflow boundaries.Initially, thefree-streamconditionsweresetto correspondtothemaximum-velocity point in thetrajectory(56.5s into theflight) (seetable1 - the valuesfor the turbulent intensityτandviscosity-ratioµT/µL wereassumed);otherfree-stream
conditionswereusedfor reasonsthatwill becomeclearlaterin this paper. All walls were modeledas no-slip with pre-scribed temperatures.The wall temperatureswere takenfrom the measuredvalues. For convenience,these wereassumedto be stepwise-constantalong segments of thewalls. No attemptwasmadeto exactly matchthe tempera-turevalueat theprobelocations;instead,thesmoothestpos-sible distribution was imposed.
T [K] 203.5
p [Pa] 3968
M 6.4
τ .01
µT/µL 1.0
Table 1: Free-stream conditions at t = 56 s.
Thegaswasassumedto beasingle-specie,thermally-perfectair. To model the inviscid fluxes, Edwards’ low-dissipationflux-split schemewasused,togetherwith third-orderMUSCL extrapolationandVan Leer’s limiter. Wil-cox’ 1998k-ω model8 wasusedfor turbulencemodeling,coupled with Wilcox’s wall-matching functions at thesolid walls. The turbulentPrandtlnumberwassetat 0.90.Transitionfrom laminar to turbulent was imposedat thefirst changein slopein the external inlet. This locationiscloseto theonepredictedby theusualconical-flow transi-tion-criteria3 (Reθ/Me = 150).In VULCAN, transitioncanbe approximatedby using “laminar regions”, or regionsweresourcetermsin the turbulenceequationsare turnedoff. Time-integration was performed with the implicitdiagonalized approximate-factorization (DAF) scheme.For mostof thecalculations,thelocalCFL numberwassetat 2.0.Theentireflowfield wassolvedelliptically, in spiteof beingmostlysupersonic,in anattemptto captureall thepossibleseparationregionspresentin thedomain.As maybe recalledfrom theprevious discussion,datasuggestthepresenceof large separationat the beginning of the inter-nal inlet; smallerseparationbubblesmayalsobeexpectedat shock-impingementlocations.To reducecomputationaltimes,VULCAN’ s MPI capabilitywasused.Thecalcula-tions weredoneon an Origin 2000using12 R10000250MHz processors;the resultingparallelidealspeed-upwas11.40. Wall-time CPU was approximately0.165 ms perCV per iteration.Determinationof convergenceby resid-ual drop wasnot possiblebecauseof large oscillationsinthe residual.Most of theseoscillationsappearto occurinthe first two blocks, aroundthe pitot nose.Attempts toeliminateor reducetheseoscillationswere unsuccessful.Therefore,convergence was assumedwhen no changecouldbeobservedin theoverall domainandwall-pressuretrace. To accelerateconvergence,grid sequencingwasadoptedwith threesequences:coarse,medium,and fine.About 25,000 iterations were required (including about7,500for thefine sequence).Thisprocedurewasalsousedto give some measureof grid convergence,as will beshown in the results section.
2.1 Results
The Mach contoursfor the completeinlet solutionmay be seenin figure 3, a) andb). A conicalbow-shockforms at the noseof the spike, detachedfrom the body.Additional shock-waves are formed at the compressioncornersof theexternalinlet. All theseshockscoalesceandend at the cowl lip (where anotherbow shock is origi-nated). This flow configuration agrees with previous
solutions3, 4. A closer look at the internal inlet flowfield[figure 3, b)] shows the presenceof small recirculationbubbles,particularly at the cowl lip and at the structural
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reinforcement.However, thereis no evidenceof themassiveseparationsuggestedby the data.The lack of separationisapparentin the wall-pressuretrace (figure 4-a), where thenumerical distribution is compared with the data. Thenumericalresultsshow a dip in thepressuretracein coinci-dencewith thefirst body-sideexpansion(at about440mm).The data,on the other hand,show continuously-increasingpressureat that point which is consistentwith boundary-layerseparation.Overall, thecalculatedpressuresagreewellwith the data up to the point of the possibleseparation.Downstreamfrom there, the predictedpressurelevels areconsiderablylower than what the datasuggest.The calcu-lated mass-averageMach number at the exit is approxi-mately 2.67, which is much higher than the value of 2.0predictedby a one-dimensionalanalysisof the data2. Thecomparisonof the body wall pressuresfor the mediumandfinegrid sequencesis presentedin 4,b); thetwo solutionsaresufficiently close for the fine sequenceto be consideredacceptable.Clearly, a straightforward approachfails to cap-ture the behavior shown by the data.
Voland et al2 suggestedthat the inlet separationmayhave beencausedby changesin the inlet geometry(withrespectto the design)or by a hysteresisin the inlet startingprocess(or a combinationof both); in the latter case,themassive separationcreatedduringtheinlet unstartmayhavesurvived after the restart.Previous numericalexperimenta-tion performedby the author5 showed that, after an artifi-cially-induced inlet unstart, a separation region wouldremain in place near the throat even after removal of thecause.Furthermore,analysisof the flight dataindicatesthattheinlet wasunstartedbeforeandupto thetimefuel flowratewasturnedon6. If all this is true, thena massive separationmay have beencreatedduring low-speedoperation(leadingto inlet unstart), and becauseof hysteresissome of itremainedin placeeven after reachingdesignflight condi-tions (and after the inlet restarted).
In order to have at least a qualitative insight into thephenomena,threepoints in the trajectorywererun sequen-tially (seetable2); t = 38 s correspondsto thepoint immedi-atelybeforefuel addition.Thefirst two conditionswererunfully-turbulent, and with constantwall temperatures(at anaverageof theexperimentaldata);thet = 56 s conditionwasrun exactly asdescribedbefore(but obviously with differentinitial conditions).Eachconditionwasrun for 15,000itera-tions and with the same convergence criteria as before.
Theresultsareshown in figures5 to 7. At t = 25s(fig-ure 5) the shockscoming from the external inlet (notshown here)arefar aheadof the inlet anddo not interferewith the cowl lip. As a result,the shockcomingfrom thelip is unobstructedandimpingeson the body side,result-ing in the separationof the boundarylayer (this processlikely startedearlier in the trajectory).The recirculationregion createsa shock that endsup aheadof the inlet.Therefore,thenumericalsimulationindicatesthattheinletwasunstartedundertheseflow conditions.The fair quali-tative agreementwith the datasuggeststhat this may alsohave beenthecaseduringtheflight. By t = 38 s (figure6)the inlet is still unstarted,asboth dataandCFD show. Itshouldbenotedthat,undertheseconditions,thenumericalflow appearedto be highly transient,with the body-wallrecirculationincreasingand decreasingin size; shown inthefigure is therecirculationat its smallest.Finally, at themaximum-velocity point of t = 56 s (figure7) theinlet hasrestarted,but a recirculationregion remainsat the throat;this seemsto be consistentwith the data. There was asmallunsteadinessassociatedwith therecirculationshape,but without any major changein size. There is a betterqualitative agreementwith the pressuredata, comparedwith the straightforward approachat the sameconditions(seefigure 4), but the pressurelevels are still low; themass-averaged exit Mach number is about 2.40.
Earlierattempts6 to modeltheconditionsbetween38sand 56 s failed to predict restartat 50 s (the inlet stillrestartedat 56 s); upstreaminteractionfrom dual-modecombustion may have played a part. In any case,it isacknowledgedthattheprocedureoutlinedabove is, atbest,a qualitative approximationto the full simulationof theflight trajectory. This would requirea time-accuratesimu-lation of the full engine,with all threeinjector stagesinoperation.The theorypresentedabove wasmeantto pro-vide a plausibleexplanationof thebehavior shown by thedata, which could not be reproducedby a more directapproach.
t [s] 25 38 56
T [K] 211 212 203.5
p [Pa] 21710 9740 3960
M 2.57 3.51 6.4
Table2: Free-streamconditionsat several points in theflight trajectory .
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As for the inlet conditions,the resultsfrom the inletcalculationswerenot deemedaccurateenoughto be usedasinflow conditionsfor theburnercalculation.Therefore,a uniform-inlet condition was used,suchthat it gave thesamemass-flowrate,mass-averagedtotal-temperatureandturbulenceconditionsas the inlet exit conditionsfor t =56s (with inlet separation),and with a Mach numberof2.0. These conditions are summarized in table 4.
The solution procedure was similar to the oneemployedfor the inlet simulation;only thosefeaturesthatweredifferentwill bedetailedin this section.Thegaswasassumedto be a mixture of thermally-perfectgases.Thechemistrymodel usedwas NASA Langley’s 7-specie/7-reaction(7x7) model9. At the operatingconditions,thismodelwasdifficult to autoignite4; therefore,botha 1-stepreaction model and VULCAN’ s ignition regions werealternatively usedwith successto initiate thereaction.Theturbulent Prandtl and Schmidt numberswere set at 1.0.Turbulencemodelingalso usedWilcox’s compressibilitycorrection8 to modelMachnumbereffectson mixing. Asbefore,a three-gridsequencewas usedfor convergenceacceleration.MaximumCFL usedwas5.0.Approximately60,000iterationswereused,of which about40,000corre-
Stage II Stage III
M 1.0 1.0
T0 [K] 716 771
T [K] 597 642
U [m/s] 1864 1934
ρ [kg/m3] 0.155 0.149
w [kg/s] 0.042 0.042
Table 3: Injectant conditions.
T0 [K] 1632
ρ [kg/s] 0.467
M 2.0
τ .06
µT/µL 350
Table 4: Inlet conditions for burner calculations.
3. Burner
3.1 Solution Procedure
Burner calculationswere doneon a three-dimensionalslice limited by the jet-centerplanesbetweenadjacentstageII and III injectors; this domain correspondsto about 4.3degreesof the annularcombustor (seefigure 1, b). To sim-plify the grid generation,andsincethe resultingwidth wasmuchsmallerthanthebodyradius,thedomainwasapproxi-matedasrectangularandwith thejet centerplanesparalleltoeach other and normal to the body and cowl walls.
Thegrid wasdiscretizedinto approximately2.8 millionCVs,distributedamong48 blocks(figure8). ThenumberofCVs rangedfrom 76 to 132 in the vertical direction, andfrom 28to 36 laterally;thehighernumberscorrespondto thevicinity of the injectors.The wall spacingvaried from 0.1mm at theinlet to 0.5mm towardstheexit. Theresultingy+wasmostly under100,exceptnearthe exit nozzleweretheflow acceleratedto supersonicconditions(aswill be shownlater),andwherethe y+ could be ashigh as200.Non-C(0)grid blocks were usedthroughoutthe geometry, especiallynearthe injectors,to reducecomputationaleffort. It shouldbe notedthat the areaincreaseimmediatelybeforethe exitnozzlein theexperimentalconfigurationis missingfrom thecomputationaldomain;asmentionedbefore,this expansionwasmeantto compensatefor therear-strutsblockage.Sincethesestrutsare not being modeled,the areawas held con-stant in the numerical simulation.
Unlessotherwisenoted,all calculationsweredonewiththeconditionsandprocedureto bedescribednext (hereafterknown asbaselineconditions).At the jet-centerplanes,sym-metry boundary conditions were imposed. No-slip, pre-scribedtemperatureconditionswereusedat the walls, withthe experimentally-measuredtemperaturesapproximatedinthesamewayasin theinlet. An extrapolationboundarycon-dition was usedat the exit. Hydrogenfuel was injectedatsonicconditionsthroughstagesII andIII, andat the anglesshown in figure 1,c). The resultingmassflowratesgave anoverall equivalenceratioof about0.60.At thelocationof theinjectors,fixedboundaryconditionswereimposedusingthevalues of table 3.
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spondedto thefine sequence;convergencewill bediscussedin thenext subsection.48R10000400MHz processorswereused,with a resultingidealspeed-upof 47.0.Wall-timeCPUwas approximately 0.264 ms per CV per iteration.
3.2 Results
Figures9 and10 show the Mach andwatermass-frac-tion contours,respectively, for the vertical plane midwaybetweenthetwo injectors.A low-speedflow regionbeginsatthe stepon the cowl wall, andextendsjust pastthe cavity.The core flow is supersonicup to x ~ 950 mm, where itbegins to decceleratethrougha shocktrain to almostsonicconditions; this is confirmed by the mass-averagedone-dimensionalMach distribution (figure 12). The flow finallyreacceleratesto supersonicconditionsat thenozzleexit. Theextentof thereactionmaybejudgedfrom thewatercontours
andtheaxial distributionof efficienciesa (figure11).Mostofthe fuel from stageII appearsto have mixed by the time itreachesthe stageIII axial-location.Due to ignition delay,however, it startsto reactshortly before that location, andappearsconsumedby the time it reachesthe constant-areasection(x ~ 900 mm). StageIII fuel, on the other hand,seemsto start reactingin the constantsection,contributingperhapsto the chockingof the flow. All the fuel is mixed,and almost all (94%) reacted, by the time it reaches the exit.
Comparingthe calculatedwall pressure-traceswith thedata(figure13), CFD somewhatunderpredictsthepressuresup to x ~ 900. Furthermore,it shows a reaccelerationorsupersonicregion justdownstreamof thecavity, correspond-ing to theclosingof thecowl-wall low-speedregion(thiscanalso be seenin the one-dimensionalaverageof figure 12);thedatadoesnotshow reaccelerationat thatlocation.Down-streamfrom there,thenumericalpressuresrecover their val-ues before reacceleration,and remain fairly constant(incoincidencewith thesubsonicregion) until reachingtheexitnozzle;the experimentalpressuresalsoappearconstantbutat a somewhathighervalue.Therewassomeunsteadinessinthe fine-sequencesolution between 800 and 900 mm,approximately; a fully steady-statesolution was notachieved.Theorigin of this unsteadinesswasmostlikely therecirculationregion, and it resultedin changesin the pres-surelevelsof about10 Kpa in this region; thesevaluesalter-natedfor the last 25,000iterations,and the resultsshowncorrespond to the lowest pressures.
The effectsof grid sequencing(andthereforegrid con-vergence)areshown in figure14for thebodywall-pressures.
a. ηmix andηc aredefined10 asmixedfuel overtotalfuel, and reacted (water) fuel over total fuel,respectively (all evaluated at local axial planes).
Therearesmalldifferencesat theinlet andtowardtheexit;themaximumburnerpressureseemto bethesamefor bothgrids. Taking into accountthat thesedifferencescorre-spondto a factor of 8 in the numberof CVs, it can bearguedthat thesolutionis at leastcloseto beinggrid-con-verged.Sincethemediumsequenceappearsto giveaqual-itatively goodsolution it wasusedto performa seriesofparametricstudies.In what follows, all resultscorrespondto the medium sequence unless otherwise noted.
Lowering the turbulent Schmidtnumber(ScT) from1.0 to 0.5 (figure 15) andturning the compressibilitycor-rection off enhancesthe mixing and heat releaseandincreasesthe pressurein the near-field (x < 800 mm).However, it gives a much larger supersonicregion andlower pressureafterchoking.It would seemthattoo muchheatreleaseimmediatelyafter stageIII reducesthe pres-surelevels in the constant-areasection.Apparently, mostof the heat-releaseshouldoccur betweeninjector stagesanddownstreamof thecavity to have maximumeffect onthe pressurerise; it is not immediatelyobvious how toachieve this with a constant-Scmodel.Strongsensitivityof dual-modecombustion to turbulent transport coeffi-cients has been reported in the literature11.
Switchingfrom a 7x7 to a 9x18chemistrymodel9 (allother conditions left at their baselinevalues) does notappearto have a major impacton thesolution(figure16);theefficiency distributions(not shown here)aresomewhatlower in thenearfield, but otherwisevery closeto the7x7model.Sincemore ignition delaywasexpectedfrom the9x18 model, this issue may require further investigation.
4. Summary
The CIAM/NASA scramjetflight-test was subjectedto a CFD analysis.The datafrom the experimentshowsthattheinlet wasunstartedfrom a free-streamMachnum-ber of about2.5 to about5.0 (including the start of fuelinjection); even after restart,separatedflow remainedjustaheadof the throat up to the Mach 6.4 condition beingevaluated in the present paper.
CFD analysisperformedat the Mach 6.4 conditionsuggeststhat the inlet would have startedandoperatedasdesigned,without major flow separation;this result is notconsistentwith the data.A qualitative studystartingfromthe Mach 2.5 conditionsshowed the presenceof massiveseparationat low-speedconditions.At reachingMach6.4conditions, the inlet restartedbut significant separationremained because of apparent hysteresis effects.
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The separatedflow in the inlet causedsignificanttotalpressureloss, lowering the combustor-entranceMach num-ber from an expectedvalueof about2.7 to an actualvaluenear2.0.With this lower Mach inlet condition,theresultingCFD solutionscomecloseto the measuredflight wall-pres-suredata.Bothdataandcomputationsshow theburneroper-atingin aclassicaldual-mode,with largeregionsof subsonicreactingflow dominatingthe combustor. The presentanaly-sisindicatesthatthefuel wascompletelymixed,andthatthecombustion efficiency was 94% at the combustor exit.
Underlyingassumptionsmeantto simplify calculations(i.e., steady-stateanalysis,uncoupledinlet andburner)mayhave beenresponsiblefor the numericalanalysisnot beingable to quantitatively match the data and may have to bereconsidered.A complete-enginetime-accuratecalculationof the flight test from Mach 2.5 to 6.4 may be neededtoaccuratelyreplicateinlet separationandcombustorentranceconditions.In addition,turbulenceandtransportmodelsforthe highly-distortedcombustor flow may needto be reas-sessed;therelaxationof theconstant-ScT assumptionshouldbe a priority. As it stands,the presentanalysisshowed thatCFD is ableto predictpotentialproblemsin a given design(andactualconstruction)so asto prevent themfrom occur-ring in an actual experiment or flight-test.
Acknowledgments
The author wishes to thank Randy Voland, MichaelSmartandAaron Auslenderfor their help in providing theexperimentaldata, as well as their suggestionsand com-ments.Rick Gaffney madeavailablehis unpublishedreportand gave useful advice. Robert Baurle assistedwith one-dimensionalprocessingof numericalresults.Jeff Whitecon-tinued to provide supportin the useof the VULCAN codefor thisandotherprojects.This researchwasinitially fundedby NASA grantNASW-4907throughtheNationalResearchCouncil, with Charles McClinton as technical advisor.
References
1 McClinton, C., Roudakov, A., Semenov, V., and Kope-henov, V., “Comparative Flow Path Analysis and DesignAssessment of an Axisymmetric Hydrogen Fueled Scram-jet Flight Test Engine at a Mach Number of 6.5”, AIAAPaper 96-4571, 1996.
5 Rodriguez, C.G. “CFD Analysis of the CIAM/NASAScramjet Engine (I): Inlet and Scramjet Analysis”, HXReport HX-977, NASA Langley Research Center, 2001.
6 Rodriguez, C.G. “CFD Analysis of the CIAM/NASAScramjet Engine (II): Inlet Unstart and Restart”, HXReport HX-977, NASA Langley Research Center, 2001.
7 White, J.A., and Morrison, J.H., “A Pseudo-TemporalMulti-Grid Relaxation Scheme for Solving the Parabo-lized Navier-Stokes Equations”, AIAA 99-3360, 1999.
9 Drummond, J.P., Rogers, R.C., and Hussaini, M.Y., “ADetailed Numerical Model of a Supersonic ReactingMixing Layer”, AIAA 86-1427, 1986.
10 Rogers, R.C., “A Study of the Mixing of HydrogenInjectedNormalto aSupersonicAirstream”,NASA TND-6114, 1971.
11 Eklund,D.R.,Baurle,R.A., andGruber, M.R., “Numer-ical Study of a Scramjet Combustor Fueled by an Aero-dynamic Ramp Injector in Dual-Mode Combustion”,AIAA Paper 2001-0379, 2001.
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a)
b)
c)
Figure 1: Inlet and combustor design geometry (units in mm; φ denotes diameter)a) inlet; b) combustor; c) injector stages (courtesy R.T. Voland).
