AIAA-2001-3677 Alleviationof Facility/EngineInteractionsin an Open-Jet Scramjet Test Facility C.W. Albertson Center NASALangleyResearch Hampton,VA S. Emami LockheedMartinEngineeringand ScienceCompany Hampton,VA 37th Joint PropulsionConference & Exhibit 8-11 July 2001 Salt Lake City, Utah For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 1801 Alexander Bell Drive, Suite 500, Reston, VA 20191-4344 ALLEVIATION OF FACILITY/ENGINE INTERACTIONS IN AN OPEN-JET SCRAMJET TEST FACILITY Cindy W. Albertson* NASA-Langley Research Center Hampton, Virginia 23681-2199 Saied Emami t Lockheed Martin Engineering and Science Company Hampton, Virginia 23681-2199 Abstract leading edge positioned 2 in. into the core flow, Results of a series of shakedown tests to however some interaction effects were still evident eliminate facility/engine interactions in an open-jet in the engine data. A new shroud and diffuser scramjet test facility are presented. The tests have been designed with the goal of allowing were conducted with the NASA DFX (Dual-Fuel fueled tests to be conducted with the engine eXperimental scramjet) engine in the NASA forebody leading edge positioned in the core Langley Combustion Heated Scramjet Test without facility interaction effects in the data. Facility (CHSTF) in support of the Hyper-X Evaluation tests of the new shroud and diffuser program. The majority of the tests were will be conducted once ongoing fueled engine conducted at a total enthalpy and pressure tests have been completed. corresponding to Mach 5 flight at a dynamic pressure of 734 psf. The DFX is the largest Introduction engine ever tested in the CHSTF. Blockage, in terms of the projected engine area relative to the Propulsion testing of the Mach 5 DFX (Dual- nozzle exit area, is 81% with the engine forebody Fuel eXperimental scramjet) engine is ongoing in leading edge aligned with the upper edge of the NASA Langley's Combustion Heated Scramjet facility nozzle such that it ingests the nozzle Test Facility (CHSTF) in support of Mach 5 boundary layer. The blockage increases to 95% flowpath development for NASA's Hyper-X with the engine forebody leading edge positioned program. Although the Mach 5 flight test has 2 in. down in the core flow. Previous engines been eliminated from the Hyper-X program, Mach successfully tested in the CHSTF have had 5 ground tests continue for the purpose of blockages of no more than 51%. Oil flow studies technology development. One of the primary along with facility and engine pressure purposes of the present test series is to provide a measurements were used to define flow behavior. database to directly compare engine performance These results guided modifications to existing and operability between two types of facilities aeroappliances and the design of new commonly used for scramjet propulsion research. aeroappliances. These changes allowed fueled Tests were initially conducted in the NASA tests to be conducted without facility interaction Langley Arc Heated Scramjet Test Facility effects in the data with the engine forebody (AHSTF), which has a test gas composition close leading edge positioned to ingest the facility to air (ref. 1). These tests were conducted at a nozzle boundary layer. Interaction effects were total enthalpy and pressure corresponding to also reduced for tests with the engine forebody Mach 5 flight at a dynamic pressure of 885 psf. In *Aerospace Research Engineer, Hypersonic Airbreathing Propulsion Branch tprincipal Engineer, Assigned to the Hypersonic Airbreathing Propulsion Branch Copyright © 2001 by the American Institute of Aeronautics and Astronautics, Inc., No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for governmental purposes. All other rights are reserved by the copyright owner. 1 American Institute of Aeronautics and Astronautics order to obtain as good a comparisonas fuel equivalence ratio reasonablypossiblet,hesameenginehardware andforcebalancethatweretestedintheAHSTF Subscripts weretheninstalledintheCHSTF.Howeverd,ue tofacilityinteractionproblemsencountereidnthe ba base of engine past with large blockageengines, it was c test cabin recognizedthat a seriesof shakedowntests e facility nozzle exit would be necessary in order to minimize e,d design condition at facility nozzle exit interactionbsetweenthefacilityandengine.The e,eng engine nozzle exit stepstakentominimizeinteractioneffectsinthe eng engine CHSTFarethesubjecotfthispaper. le engine forebody leading edge TheDFXisthelargestengineevertestedin n facility nozzle theCHSTF.Blockagei,ntermsoftheprojected t,h total condition in the facility heater enginearea(includinginteriorflowpatharea) t,2 total condition behind a normal shock relativetothenozzleexitarea,is81%withengine Sill4 silane mixture (20% Sill4, 80% H2, by cowlclosedandwiththeengineforebodyleading volumn) edgealignedwiththeupperedgeof thefacility oo static condition upstream of aircraft bow nozzlesuchthatit ingeststhenozzleboundary shock layer. Theblockageincreasesto 95%withthe engineforebodyleadingedgepositioned2 in. Description of Experiment down in the core flow. Previousengines successfullytested in this facility have had Test Facility blockageosfnomorethan51%. The CHSTF has been in operation since 1978 Theshakedowntestsdescribedinthisreport and has historically been used to test complete wereprimarilyconductedatatotalenthalpyand (inlet, combustor, and partial nozzle) subscale pressurecorrespondingto Mach5 flight at a scramjet component integration models. To date, dynamicpressureof 734psf. The dynamic nearly 2000 tests have been conducted with a pressurewasreducedfromthedynamicpressure variety of engines including the NASA-Langley 3- oftheAHSTFtestsbecauseofconcernsthatthe Strut, NASA-Langley Step Strut, NASA-Langley interactionloadsat the higherpressurecould Parametric, NASP Government Baseline, exceedtheforcebalancedesignloads.Oilflow Rocketdyne A2, Pratt & Whitney C, JHU/APL B1, studiesalongwithfacilityandenginepressure Rocketdyne A3 Hydrocarbon-Fueled Scramjet, measuremenwtsereusedtodefineflowbehavior. and most recently, the NASA-Langley DFX engine Theseresultsguidedmodificationsto existing under the Hyper-X program. aeroappliances and the design of new The CHSTF (shown schematically in figure 1) aeroappliances. can operate at stagnation enthalpies duplicating that of flight at Mach numbers ranging from 3.5 to Nomenclature 6.0. These enthalpy levels are obtained by burning hydrogen and air in the facility heater. h diffuser duct height Oxygen is added upstream of the heater such that m facility total mass flow rate the oxygen content of final test gas matches that p pressure of air. Currently, either a Mach 3.5 or 4.7 nozzle q dynamic pressure may be installed between the heater and test x axial distance cabin to expand the test gas to the desired y vertical distance conditions. Both nozzles are a two piece design Ar argon (throat and expansion sections), have square H enthalpy cross sections, and are contoured to exit H2 hydrogen dimensions of 13.26 by 13.26 in. In addition to H20 water vapor exit Mach number, the main difference between M Mach number the two nozzles is that the M3.5 nozzle is a heat N2 nitrogen sink design, while the M4.7 nozzle is water-cooled 02 oxygen in the throat section and along a portion of the T temperature expansion section. For the present tests series, 2 American Institute of Aeronautics and Astronautics theMach4.7nozzlewasused. Thisnozzleis Data Acquisition, Instrumentation, and Flow relativelynewandhasjustundergonea seriesof Visualization calibrationtests. Thenozzleflowexhaustsinto The data acquisition system (DAS) consists of thetestcabin,whichhasinteriordimensionosf42 a commercially available software package in.highby30in.wideby96in.long. Thejetfrom (AutoNet) running on a Pentium processor. The thenozzlepassesthroughandaroundtheengine DAS incorporates a NEFF 300 signal conditioner modelandthenintothefacilitydiffuser.Theflow and a NEFF 600 amplifier/multiplexer capable of istypicallyexhaustedintoa70-ft.vacuumsphere. supporting 128 channels of instrumentation. In (Theairejectosrhowninfig.1isnolongerused.) addition, up to 512 pressure measurements can Thefacilityistypicallyoperatedsuchthatflow be recorded using a PSI 8400 ESP system and conditionsattheengineinletentranceplaneare sixteen 32-port modules. matchedwiththatoftheflightvehicle(fig.2). To In addition to the standard instrumentation for accomplishthis,thefacilityheaterisoperatedat health monitoring and defining flow conditions, theflighttotalenthalpyw, hichisconstanat cross instrumentation is also provided to assist in thebowshock,andtheflowinthefacilitynozzleis assessing engine/facility interaction effects. This expandedto conditionsmatchingthoseat the instrumentation primarily consists of static vehicleengineinletentranceplane. Thefacility pressure measurements located along the facility normallyoperatesatheaterstagnationpressures nozzle, inside the test cabin, at the base of the between50 and 500psia andat stagnation DFX engine, and along the facility diffuser (fig. 5). temperaturesbetween1300and3000°R. The Pitot probes are located at the facility mixer flightdynamicpressurerangesfrom250to3500 entrance and exit. A single total temperature psf,dependingonMachnumber.Testgasmass probe is located at the mixer entrance. flowratesrangefrom10to60Ibm/s.Therangeof Oil flow studies were conducted during this operationis shownby the Machnumberand investigation to define flow behavior and to aid in altitudesimulationenvelope(fig. 3). The left locating flow fences and other modifications. The verticalboundaryoftheenvelopeisthe nozzle oil flow mixture consisted of a mixture of 50% (by exitMachnumberof3.5andthe rightvertical volume) of 80W-140 gear oil and 50% of a boundaryreflectsthemaximumheateroperating commercially available oil thickener manufactured temperatureof 3000°R. The upperinclined for automobile engines. Lampblack was added to boundaryrepresentsthe minimumoperating the mixture to provide visibility. pressureof 50psia,uptoan altitudewherea Model and Installation flightdynamicpressureof250psfisimposedasa The engine was installed in the CHSTF test limit. Thelowerinclinedboundaryreflectsthe cabin as illustrated in figure 6. The engine and maximummassflowrateto the heaterat the force balance were mounted at the top of the test Machnumbeorf3.5limitandthemaximumheater cabin through a series of attachments as shown. operatingpressureattheMachnumberof6limit. Initially the forebody leading edge of the engine Calculatedtest gas compositionsfor these was positioned 2.0 inches below the upper wall of conditionsareshownin figure4. The primary the nozzle at the exit plane (Yle = 2.0 in.). This contaminant in the test gas is water vapor, which position was chosen to match the engine position varies from 0.060 mole fraction at Mach 3.5 to for the majority of the runs from the AHSTF tests. 0.198 at Mach 6.0. The angle of the engine relative to the facility The facility currently has a gaseous engine fuel nozzle water line was adjusted such that the Mach system consisting of six independently controlled number just ahead of the cowl leading edge fuel circuits which can be connected to any matched that measured during the Mach 5 DFX combination of injection stations in the engine. tests in the AHSTF. One circuit supplies a pyrophoric mixture of silane The initial aeroappliances consisted of the and hydrogen (20 percent silane and 80 percent same nozzle extension skirt and catch cone hydrogen, by volume) for igniting and piloting the diffuser used successfully with previous tests of fuel. Gaseous hydrogen and ethylene have been lower blockage engines. The purpose of the used as the primary fuel in past engine tests. nozzle extension skirt is to position the Mach Additional details of the CHSTF are given in wave from the nozzle exhaust far enough refs. 2 - 4. downstream that it doesn't disturb the flow delivered to the engine inlet. The catch cone 3 American Institute of Aeronautics and Astronautics catchesthenozzleexhaustanddirectsitintothe the vacuum valve. Hydrogen is then allowed to diffuser. A washer(fig. 7) is locatedat the flow into the facility heater, increasing the entranceof the catch cone to preventany temperature and pressure. Once good separatedflowonthewallofthecatchconefrom combustion in the heater is established, the timer spillingoutintothetestcabin.Thewasheralso start circuit is activated, allowing oxygen to be restrictstheareaaroundtheengineatthecatch added upstream of the heater (See fig. 1). Facility coneentranceandenhancestheaspirationofthe steady-state flow conditions are obtained in about testcabin. Wateris injectedparalleltotheflow 5 seconds, which is considered the no-fuel tare near the engine nozzle exit to reduce the data point. At 6 seconds into the run, the model temperatureandassociatedpressurerisefrom fuel sequence is initiated. Tests are terminated theengineexhausatndtestgas.Thislessensthe automatically by the run timer (typically 20 probabilityoffacilityinterferencewiththeforce seconds). balance measurementsand engine nozzle pressures. Results and Discussion As shown in figure 6, the DFX engine incorporateascowlthatrotatesaboutapointnear Baseline Configuration and Modifications, thecowlleadingedge,allowingthecontraction y_ = 2.0 in. ratioto bereducedtoallowtheenginetostart. Tests were initially conducted without engine Figure6showsthecowlopenatthe12°position fuel and with the baseline configuration at the Yie= as was usedduring the AHSTFtests. As 2.0 in. position in the facility (figs. 6 and 7). The discussedintheResultsandDiscussionsection, first set of tests were conducted at a reduced it wasfoundthatthis anglecouldbe reduced stagnation pressure and temperature of 100 psia significantwlyhilestillenablingtheenginetostart. and 1600°R, respectively, to minimize loads on Te_t Conditions and Procedure the force balance while determining the maximum The first set of tests was conducted at a load level from facility start-up to shut-down. reduced stagnation pressure and temperature of Tests continued at this condition to determine the 100 psia and 1600°R, respectively, to minimize minimum cowl-open angle necessary to start the loads on the force balance while determining the engine. Results showed that a cowl-open angle maximum load level from facility start-up to shut- of approximately 3° was sufficient for starting the down. Tests continued at this condition to inlet. The timing of the cowl closing was then determine the minimum cowl-open angle optimized to minimize loads on the force balance necessary to start the engine. while obtaining a started inlet. Results showed Follow-on tests were primarily conducted at a that the cowl could be closed about 2 seconds Mach 5 flight enthalpy to match the bulk of the after the facility vacuum valve opened (fig. 8) Mach 5 DFX tests in the AHSTF. The total while the heater pressure was still increasing. temperature corresponding to this enthalpy level Follow-on tests to minimize facility/engine was 2095 °R for the vitiated test gas. The interactions were conducted at the nominal nominal total pressure in the facility heater was condition of Pt,h= 175 psia and Tt,h= 2095°R. The reduced to 175 psia compared to 210 psia for the pressures measured along the facility nozzle wall AHSTF tests to minimize loads on the force for the baseline configuration (fig. 9) were high balance due to facility interactions. The and exceeded the predicted value near the nozzle corresponding flight dynamic pressure was 734 exit by a factor of 2.69 (table 3). Predictions were psf compared to 885 psf for the AHSTF tests. based on three-dimensional full Navier Stokes The nominal test conditions and test gas mole calculations assuming frozen flow down the fractions are summarized in tables 1 and 2, nozzle length (ref. 5). The measured cabin respectively. The test gas mole fractions were pressure was a factor of 1.97 higher than the calculated assuming complete combustion of the measured nozzle exit pressure (table 3) and 5.30 hydrogen and frozen flow along the facility nozzle. higher than the predicted static pressure at the A typical test sequence is illustrated in fig. 8. nozzle exit. Apparently, the cabin pressure was First the tunnel air flow is established (10 to 60 high enough to feed forward along the facility Ib/sec.). The heater ignitor is then activated and nozzle wall and separate the boundary layer. once good ignitor operation is verified, the timer The sidewalls of the nozzle extension skirt reset circuit is energized, initiating the opening of were then modified (fig. 10) to eliminate the shock American Institute of Aeronautics and Astronautics lossesandassociateddragcausedbyitsclose twice the nozzle exit pressure, as indicated by the lateralproximity(1.5in.)totheengineforebody static pressure distribution along the facility nozzle fences.Asaresult,theoveralllevelofthenozzle wall shown in fig. 17 and in table 3. Pressure pressuredistributiondecreasedtonearlythatof distributions along the facility diffuser are also theCFDprediction(fig.9). However,the test noticeably lower (fig. 18.), as were the pitot cabinpressurewasa factorof3.11higherthan distributions at the end of the 19 in. dia. duct (fig. the measurednozzleexit pressure(table3), 19). Pressures on the engine forebody and whichisstillhighenoughtocauseboundary-layer nozzle showed no indication of interactions with separation. Oil flow obtainedon the nozzle the facility. extensionskirtshowsometurningofthe flow, Based on these results, a fueled run was probablyas a resultofthehighcabinpressure attempted. A plot showing various pressures as a feedingforwardalongthecornersoftheskirt(fig. function of time for the current test configuration is 11).Oilflowpatternsobtainedonthecatchcone shown in figure 20. The data, taken at 20 Hz, topcoverextensionplateandcatchconewasher show that the inlet unstarted at 21.1 sec. into the (fig.12)indicatethatsomeoftheflowwasspilled run while the fuel flow rate to the engine intothetestcabin.Thediffuserdistributiongsiven combustor was slowly increasing. The engine infig.13showasignificanptressuredropwiththe base pressure increased about the same time. nozzleskirt modificationA. lso,pitotpressures However, the facility diffuser, cabin, and nozzle obtainedattheexitof the 19-in.diameterduct exit pressures were not affected until about 0.15 decreasedwiththenozzleskirtmodification(fig. seconds later. These results indicate that the 14), reflectinghigher Machnumbersat this DFX engine can be successfully tested up to the station.Inspiteofthedropin static pressures in point of inlet unstart without facility interaction the exhaust duct, the engine pressure effects at the Yle = 0.0 in. engine position in the distributions still showed signs of boundary-layer test cabin. separation near the engine nozzle exit. This indicated that the pressures in the catch cone and Facility Diffuser Modifications 19 in. duct were high enough to feed forward into Various sources were consulted in the the engine nozzle. redesign effort of the catch cone and diffuser. In an attempt to better direct the flow into the Unfortunately, the actual flow within these catch cone, flow fences were added to the catch components is complicated by three-dimensional cone top cover extension plate (fig 15). In spite of interactions of shock waves and boundary layers. this modification, the test cabin pressure was still The flow is further complicated with the addition of high at 3.16 times the measured nozzle exit a geometrically complex engine and its exhaust. pressure (table 3). Only slight changes were Because of the flow complexity and unknowns, noted in the nozzle pressure distribution (fig. 9), one must resort to empirically based guidelines. the diffuser pressure distribution (fig. 13) and the Literature Guidelines pitot pressure distribution (fig. 14) with the There are a number of reports describing addition of the flow fences. The engine pressure various guidelines in designing supersonic distributions improved slightly because of lower diffusers as summarized in refs. 6 and 7. Many of base pressure (table 3), but still showed signs of the reports listed in these summaries were written boundary-layer separation. Because of this, the in the 1960's and earlier. Of particular interest for fuel-off force measurement was unreliable. the present application are those that include Therefore, it was decided not to attempt to fuel the blockage effects in facilities in which the model is engine at the Y_e= 2.0 in. position with the current fixed (as opposed to injected) in the flow. Smaller diffuser and catch cone design. diffuser diameters tend to be more efficient, however, one must be careful not to overcontract Modified Configuration, y_ = 0.0 in. the flow. The first-order approximation of The engine was then raised out of the core such that the forebody leading edge was level Kantrowitz (ref. 8) is often used to provide a conservative estimate for the maximum geometric with the nozzle exit (Yle = 0.0 in.) (See fig. 16.) contraction ratio that permits supersonic flow. This reduced the blockage from 95% to 81% with Other guidelines include length/ diameter the cowl in the closed position. No additional recommendations for the second minimum to changes were made to any of the areoappliances. provide good isolation and corresponding run time The cabin pressure was reduced to approximately 5 American Institute of Aeronautics and Astronautics (ref.6). Oneespeciallyusefulrecommendatioisn arrangement also allows for thermal expansion, toincorporataerearward-facinsgtepinthedesign eliminating the need for a new and larger bellows tohelpisolatetheflowinthevicinityofthetest section, and also provides a length adjustment. articlefromthespherepressure. This length adjustment eliminates the need for Thedesignof catchconesandshroudsto separate spoolpieces to accommodate the two directtheflowintothediffuserisdiscussedinrefs. different facility nozzles. A sliding seal is installed 7and9. Aerodynamitcailoringofshroudsurfaces at the face of the flange attached to the 25.25 in. isdiscussedinref.9. I.D. duct to minimize leaks. The Redesigned Diffuser The new diffuser components have been The objective of the new diffuser design is to designed and constructed. The new shroud will reduce the losses to the flow and associated be constructed and fitted in place following the effects in the engine data, without compromising completion of the DFX test series at the Yle = 0.0 run time and also without having to replace any in. location. Once all the new components are more components than necessary. The catch installed, some tests will be repeated with the cone and 19 in. I.D. (inside diameter) constant DFX engine at Y_e = 0.0 in. for comparison area diffuser were targeted for redesign, based on purposes. Tests will then be conducted with the the relatively high wall pressures measured in engine lowered into the core flow at Yie= 2.0 in. these areas (fig. 18). Other components that will be replaced include the original transition section, Concluding Remarks expansion bellows, and a small spool piece from the 25.25 in. I.D. dia. constant area section (fig. A series of shakedown tests to eliminate 21). The air ejector is no longer used and will be facility/engine interactions in an open-jet scramjet removed. test facility have been conducted. The tests were The strongest shock from the engine originates conducted with the NASA DFX (Dual-Fuel from the leading edge of the engine cowl. The eXperimental scramjet) engine in the NASA resulting two-dimensional shock pattern is Langley Combustion Heated Scramjet Test illustrated in figure 22 for the original catch cone Facility (CHSTF) in support of the Hyper-X configuration. A new shroud is being designed to program. Oil flow studies along with facility and replace the original catch cone (fig. 23). This engine pressure measurements were used to shroud will be aerodynamically tailored to reduce define flow behavior. These results guided the shock losses and will also have rectangular, modifications to existing aeroappliances and the as opposed to circular cross sections, consistent design of new aeroappliences. As a result of with the engine exterior geometry. This shroud these changes, fueled tests could be conducted will attach to a square-to-round contraction without facility interaction effects in the data with section that will fit into the transition section, as the engine forebody leading edge aligned with the illustrated in figure 24. Water injectors will be upper edge of the facility nozzle such that it incorporated into the new shroud to reduce the ingested the nozzle boundary layer. Interaction temperature and associated pressure rise from effects were reduced for tests with the engine the engine exhaust and test gas. forebody leading edge positioned 2 in. outside of The 19 in. I.D. constant area diffuser duct, the facility nozzle boundary layer, however some which forms the second minimum, will be replaced effects were still evident in the engine data. A with a 23.25 in I.D. constant area section (fig. 24). new shroud and diffuser have been designed with This cross-sectional area of the 23.25 in I.D. the goal of further reducing blockage effects and section is 50% larger than the original 19 in. I.D. allowing fueled tests to be conducted with the section. This reduces the "duct blockage" forebody leading edge of the engine positioned 2 (projected engine area/ second minimum area) in. into the core flow. Installation and evaluation from 50% to 34% with the engine positioned at Yle of the new shroud and diffuser will commence = 0.0 in. (engine cowl in closed position) and from once ongoing engine tests have been completed. 59% to 40% with the engine at Yle-- 2.0 in. The 23.25 in. I.D. duct slides into the existing 25.25 in I.D. duct, resulting in a rearward facing step with overlap to provide isolation as the sphere pressure increases during the run. This sliding 6 American Institute of Aeronautics and Astronautics Acknowledgements References 1. Guy, R. W.; Rogers, R. C.; Puster, R. L.; A special thanks to Bobby Huffman, Barry Rock, K. E.; Diskin, G. S.: The NASA Langley Lawhorne, and Clint Reese for their help with the Scramjet Test Complex. AIAA-96-3243, July model and instrumentation and also with the 1996. operation, diagnostics and repairs of the facility. 2. Andrews, Earl H., Jr.; Torrence, Marvin G.; Thanks to Don Harper for his help with all things Anderson, Griffin Y.; Northam, G. Burton; and that are DAS related, to Carl Davis for helping Mackley, Earnest A.: Langley Mach 4 with the instrumentation and diagnostics, to Earl Scramjet Test Facility. NASA TM-86277, Andrews for his helpful suggestions and 1985. schematics relating to the new diffuser design, to , Andrews, Earl H., Jr.: A Subsonic to Mach Troy Middleton for initiating the detailed design 5.5 Subscale Engine Test Facility. AIAA-87- and construction of the new diffuser, and to our 2052, June 1987. new facility safety head, Diego Capriotti, for his 4. Rock, Kenneth E.; Andrews, Earl H.; and facility help and also for his help with the DAS Eggers, James M.: Enhanced Capability of post-processing Fortran codes. Thanks also to the Combustion-Heated Scramjet Test Roger Jones for his help with many of the figures Facility. AIAA-91-2502, June 1991. (plus revisions) contained within this report. . Gaffney, Richard L., Jr.: Private communications with Richard L. Gaffney, Jr. Aerodynamics, Aerothermodynamics, and Acoustics Competency, NASA Langley Ms Pth Ht.h Tth q= Mn m Research Center, 2001. (psia) (B1u/ °R Ibm/s Ibm) 6. Handbook of Supersonic Aerodynamics, 5.0 175 573 2095 734 4,7 12,3 Section 17, Ducts, nozzles and Diffusers. NAVWEPS Report 1488, January, 1964. Table 1. Nominal test condition for the present 7. Andrews, Earl H., Jr.: Nozzle/Tunnel Starting series of shakedown tests. in Free-Jet Engine Test Facilities. CPIA Publication 585, June 1994. Kantrowitz, Authur and Donaldson, Coleman , duP.: Preliminary Investigation of Supersonic 5.0 .6477 ,2095 I ,1338 I .0090 Diffusers. NACA WR L-713, 1945. 9. Bulman, M.; Leonard, J.; Keenan, R.; and Table 2. Test gas mole fractions at the nominal Wade, M.T.: Advancing the State of the Art test condition given in table 1. in Hypersonic Testing; HYTEST/MTMI. AIAA Paper 93-2023. Configuration Yle Pe/P e.d Pc/Pe Pba/P e.enQ Facility/engine interactions? (in.) Forebody Engine nozzle Original (U*) 2.0 2169 1.97 1.16 Yes Yes Modified skirt (U) 2,0 1,07 3.11 2.27 Yes Yes Modified skirt 2.0 1.07 3.16 1,28 No Yes +fences (U) Modified skirt 0.0 1,07 228 1,20 No No +fences (U) Modified skirt 0.0 1.07 2,45 1.04 No No +fences (F*) Notes: *U = Unfueled run *F = Fueled run just prior to inlet unstart Table 3. Blockage results for various test configurations (Pt.h= 175 psia and Tt.h= 2095°R-) 7 American Institute of Aeronautics and Astronautics Figure1. SchematiocftheNASALangleyCombustioHneatedScramjeTtestFacility,Dimensionasrein feetunlessstatedotherwise. Flight 175 --Stagnat on 0._3' / 0._5' / t.0• 1.5 20 _,, I_t pressure,psia, / ', ---Stagna!ion i/ , / 150 - entSalpy/1000,',/ ' / 8t.,_ Y't ',/ 0 Tsehstap.o_inot_fol//',_ .://,." i ..4_ ! Fore 125 1000 100 Altitude, ._-Z, _°°° kft ///._, ._ ' tO 5000 75 ,i:iii:: Simulationingroundfacility i/_II i ../ / :i:: ' 5O I / /, .. ' si_ : Facility bounda_ layer 25 i,/.,/I..'." /,, _'_/ I" i /.; 1111/ _" .' / ,I . ,r 2 4 6 8 tO ea_ Engine modularly,. Flight Mach number Figure 3. Flight simulation envelope for the LaRC Combustion Heated Scramjet Test Facility. Figure 2. Matching _ght conditions. 8 American Institute of Aeronautics and Astronautics