22nd AIAA Aerodynamic Measurement Technology and Ground Testing Conference June 24-26, 2002/St. Louis, Missouri For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics, 1801 Alexander Bell Drive, Suite 500, Reston, VA 22091 AIAA 2002-2706 A Flow Quality Analysis for Future Hypersonic Vehicle Testing Carlos Tirres Arnold Engineering Development Center Arnold Air Force Base, TN 37389 Marty Bradley The Boeing Company Long Beach, California Calvin Morrison and Ray Edelman Boeing-Rocketdyne Canoga Park, California 22nd AIAA Aerodynamic Measurement Technology and Ground Testing Conference 24-26 June 2002, St. Louis, Missouri AIAA 2002-2706 This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.
Transcript
22nd AIAA Aerodynamic MeasurementTechnology and Ground Testing
ConferenceJune 24-26, 2002/St. Louis, Missouri
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics, 1801 Alexander Bell Drive, Suite 500, Reston, VA 22091
AIAA 2002-2706A Flow Quality Analysis for FutureHypersonic Vehicle TestingCarlos TirresArnold Engineering Development CenterArnold Air Force Base, TN 37389
Marty BradleyThe Boeing CompanyLong Beach, California
Calvin Morrison and Ray EdelmanBoeing-RocketdyneCanoga Park, California
22nd AIAA Aerodynamic Measurement Technology and Ground Testing Conference24-26 June 2002, St. Louis, Missouri
AIAA 2002-2706
This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.
A Flow Quality Analysis for Future Hypersonic Vehicle Testing*
Carlos Tirres† Marty Bradley‡
Arnold Engineering Development Center The Boeing CompanyArnold AFB, TN 37389 Long Beach, CA
andCalvin Morrison§ and Ray Edelman**
Boeing-RocketdyneCanoga Park, CA
AIAA 2002-2706
Abstract
A current R&D program at AEDC plans toexplore advanced technologies to develop a hyper-sonic test facility, followed by a multiyear prototypefacility development effort. The AEDC StrategicPlanning Office (AEDC/XPX) has an effort under-way to define the future test conditions that willdefine the performance requirements for an aero-propulsion hypersonic test facility. The BoeingPhantom Works is under contract to provide infor-mation on projected hypersonic vehicles anddevelop test requirements for them. The first phasedefined advanced hypersonic tactical missile,hypersonic cruise, and space access vehicles thatmay be developed within 15 to 20 years with air-breathing engines. Size, shape, weight, aerody-namic and propulsion performance, and trajecto-ries were defined for four baseline vehicle con-cepts. The first phase of the study provided ananalysis of the ground test facility requirements forrun time, model size, Mach number, enthalpy, andpressure for performance, operability, and durabil-ity testing. The second phase has provided ananalysis of the ground test facility flow qualityrequirements by conducting sample calculationsand discussing contamination effects, chemicalkinetics, vibrational relaxation, and condensationissues. Our analyses in this study indicated that thepresence of combustion products in the freestreamflow would not seriously undermine the validity oftest results and conclusions for testing at Mach 8and above for typical test conditions. Test facilitiesthat use combustion heating vitiation to simulatehigh Mach conditions can produce useful resultsbut require careful posttest analysis. Our analysesin this study also indicated that the presence of
NOX species did not cause significant test or post-test analysis difficulties.
1.0 Introduction
This effort is intended to define flow simulationrequirements for a future aeropropulsion hyper-sonic test facility. These test simulation require-ments are essential drivers of the facility perfor-mance requirements, which in turn will drive manyaspects of facility research, development, anddesign. Although these requirements have beenthe subject of much conjecture and speculation,definitive quantitative technical analyses have notbeen conducted to conclusively define the facilityrequirements. However, the study is limited to aMach 8 to 15 flight regime for airbreathing systems.The purpose of this effort is to define the require-ments for analyses that will lead to credible, defen-sible facility performance requirements that can beused to guide the development of this class of facility.
The systems to be tested in this facility remainvisionary in nature and are not always well defined.Existing design concepts for 1) a tactical missile,dual-use, two-stage-to-orbit (TSTO)/cruise vehicle,and 2) space launch vehicles using scramjet andoblique detonation wave air-breathing engines(ODWE) are used in this analysis.
The first phase of this study defined advancedhypersonic tactical missile and space access vehi-cles which may be developed within 15 to 20 yearswith scramjet and oblique detonation wave air-breathing engines; it also defined a dual-useTSTO/cruise vehicle. Vehicle size, shape, weight,aerodynamic and propulsion performance, and tra-
* The research reported herein was performed by the Arnold Engineering Development Center (AEDC), Air Force MaterielCommand. Work and analysis for this research were performed by personnel of U. S. Air Force, Boeing Phantom Works, Boeing-Rocketdyne, and personnel of Jacobs Sverdrup AEDC Group, technical services contractor for AEDC. Further reproduction isauthorized to satisfy needs of the U. S. Government.
1American Institute of Aeronautics and Astronautics
This paper is declared a work of the U. S. government andnot subject to copyright protection in the United States.
jectories were defined. Maximum use has beenmade of existing baseline designs to define thedesign concepts. This phase of the study also pro-vided an analysis of the ground test facility thermo-dynamic (Mach number, pressure, and enthalpy),run duration, and size requirements for perfor-mance, durability, and operability testing of thesefour vehicle concepts.
The second phase is providing an analysis of theground test facility flow quality requirements (chem-ical composition, spatial uniformity, and flow stabil-ity) for the vehicles defined during the first phase.
2.0 Phase I
2.1 Baseline Hypersonic Concepts
Top level parameters summarizing the four con-cept vehicles are listed in Table 1. The conceptsrange in size from 14 to over 200 feet long. Air-breathing engines operate at speeds as high asMach 15 using various engine types and operatingmodes. The baseline concept vehicles include thefollowing:
Mach 8 Scramjet Cruise Missile — This is ahypersonic, air-to-ground missile. Primary issuesaccounted for in the design and performance
included: launch platform constraints, rocketbooster integration, airbreather transition Machnumber, acceleration and cruise performance,thermal management, and packaging require-ments; structural, payload and subsystem weights;and missile stability, guidance, and control. Thismissile is shown in Fig. 1.
SSTO Space Access Vehicle with RBCC —The second study concept was a horizontal take-off/horizontal landing (HTHL) vehicle powered byAerojet Strutjet rocket-based, combined cycle(RBCC) engines that was cooperatively designedand assessed by Boeing and Aerojet for a single-stage-to-orbit (SSTO) mission.
This concept incorporated systems and aninfrastructure capable of achieving “aircraft-like”operations. The vehicle is shown in Fig. 2. Moreinformation on this concept can be found in Refs. 1through 5.
SSTO Space Access Vehicle With ODWE — Thethird concept considered for this test requirementsstudy was a SSTO vehicle powered by an obliquedetonation wave ramjet (Wavejet) engine and DualFuel/Dual Expander (DF/DE) rocket engines thatwould deliver 10,000 lb of payload to polar, circu-lar, low earth orbit.
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The shape of the Wavejet SSTO vehicle, pic-tured in Fig. 3, reflects the requirement for a sub-stantial amount of the trajectory to be “flown” at rel-atively small flight path angles to take advantage ofthe efficiency of the airbreathing Wavejet engine inthe Mach 6 to 16 range. The Wavejet engine con-sists of twelve engine modules that circumferen-tially “wrap” the upper half of the vehicle just aft ofthe payload bay.
Dual-Fuel, Dual-Role Mach 10 Cruise/TSTOSpace Access Vehicle — The fourth design con-cept was the DF-9 Dual-Fuel, Dual-Role Mach 10Cruise/TSTO Space Access Vehicle. This vehicledesign resulted from a study conducted by TheBoeing Company and NASA Langley ResearchCenter (LaRC). The first-stage vehicle wouldlaunch a rocket-powered-second stage, which inturn would boost a small payload into low earthorbit. The incentive for the study was the significantcost savings obtainable if one vehicle could bedeveloped to accomplish both missions.
This vehicle is powered by a hydrocarbon-fueled Air-Core-Enhanced Turboramjet (AceTR)low-speed propulsion system and a hydrogen-fueled ramjet/scramjet high-speed propulsion sys-
tem. The DF-9 configuration is shown in Fig. 4.More information on this concept can be found inRefs. 6 through 14.
2.2 Test Facility Requirements Analysis
Test facility requirements were established forthe four baseline design concepts. Run time,model size, Mach number, enthalpy, and pressurerequirements for performance, operability, anddurability testing have been documented.
A sample test requirements table for perfor-mance testing for the first design concept is shownin Table 2. Detailed notes (not shown) were alsodeveloped for each cell in the table. Similar tableswere developed for operability and durability test-ing and testing of the other three vehicle concepts.
Figures 5 through 7 plot run time, model size,and Mach number for all of the performance, oper-ability, and durability tests considered for the fourbaseline vehicle concepts. The vertical bars inthese figures represent the numerical values fromthe appropriate columns in the twelve test require-ments tables. (Table 2 is a sample of one of thesetables.) For each vehicle concept, values areshown for performance testing, then operabilitytesting, and finally durability testing.
Table 2. Sample Performance Test Requirements for Mach 8 Cruise Missile
Fig. 3. Summary of ODWE “Wavejet” SSTO SpaceAccess Vehicle
AECD_02_004
Figure 4. Dual-Fuel, Dual-Role Vehicle
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Tables 3 through 5 summarize maximumrequirements for performance, operability, anddurability testing, respectively.
For full simulation, testing is required at Machnumbers up to 15 with freestream dynamic pres-sures up to 2000 psf and run times as long as 30
minutes. Scale, pressure, enthalpy, and time com-promises were considered. Maximum run times are0.03 sec, 20 sec, and 30 min for performance,operability, and durability testing respectively. Max-imum model length is 40 ft, and the minimum rec-ommended scale for testing is 20 percent, with atleast one combustor test conducted at 50-percent
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Fig. 5. Run Time Summary Plot for All Test Cases
Fig. 6. Model Size Summary Plot for All Test Cases
Fig. 7. Mach Summary Plot for All Test Cases
PerformanceLegend Operability Durability
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4American Institute of Aeronautics and Astronautics
scale. Enthalpy and Mach number maximums arederived from the SSTO ODWE vehicle at Mach 15conditions. The SSTO RBCC vehicle is just belowthese levels, with a maximum speed of Mach 14.
The first phase of this study defined advancedhypersonic tactical missile and space access vehi-cles which may be developed within 15 to 20 years
with scramjet and oblique detonation wave air-breathing engines, and also a dual-use TSTO/cruise vehicle. The study was limited to a Mach 8to 15 flight regime for airbreathing systems. Vehi-cle size, shape, weight, aerodynamic and propul-sion performance, and trajectories were defined.This phase of the study also provided an analysisof the ground test facility thermodynamic (Mach
Table 3. Maximum Requirements for Performance Testing
Table 4. Maximum Requirements for Operability Testing
Table 5. Maximum Requirements for Durability Testing
Dual-Fuel, Dual-Role Mach 10 Cruise/TSTO Vehicle Indicates Multiple Concepts Have Same Maximum Requirement
31.1 × 3.8 × 18.9
w/Premixed Fuel-Air 20 17 × 3 × 3
5American Institute of Aeronautics and Astronautics
number, pressure, and enthalpy), run duration, andsize requirements for performance, durability, andoperability testing of four vehicle concepts.
3.0 Phase II Flow Quality Requirements
The purpose of the Phase II effort was to iden-tify flow quality requirements in a test facility. Theeffort included three concepts. The vehicle con-cepts included a Mach 8 scramjet cruise missile, aSSTO space access vehicle with RBCC, and adual-fuel, dual-role Mach 10 cruise/TSTO spaceaccess vehicle.
The second phase is now providing an analysisof the ground test facility flow quality requirementsfor chemical composition, spatial uniformity, andflow stability for vehicles defined during the firstphase. The effort has included developing baselineflowpath models for further analysis, conducting aliterature search, performing sensitivity studies,and making recommendations for chemical com-position flow quality requirements.
The effort also has included the establishmentof baseline flowpath performance models for aMach 8 missile and for 20- and 100-percent scalescramjet-powered vehicles (representing bothlarge space access and cruise vehicles), a discus-sion and analysis of flow combustion effects, andrelated recommendations for a ground test facility.
The work was divided into two tasks. The firsttask was to develop baseline engine performancemodels, and the second was to conduct a literaturereview plus sensitivity analyses using the flowpathmodels developed in the first task, and to makerecommendations for facility flow chemical compo-sition.
3.1 Baseline Flowpaths
Baseline flowpath designs are necessaryfor use in flow quality sensitivity studies. Inaddition, in future work, more complex CFDanalysis may be conducted, and a precisegeometric definition of the baseline flowpathsmust be available.
Three baseline flowpaths were developedfor the sensitivity analysis: 1) full-scale, Mach8 hydrocarbon-fueled scramjet missile; 2) full-scale, hydrogen-fueled scramjet RBCC forboth SSTO and TSTO for “truth” calculations;
and 3) 20-percent scale scramjet-powered vehicleto represent flight vehicle flowpaths being tested ina hypersonic facility.
Freestream and flowpath flow parameters wereselected to represent important operating points forhypersonic scramjet operation. A typical cruisecondition was selected for the Mach 8 missile. Forthe Baseline Vehicle flowpath, typical accelerationpoints at Mach 8, 11, and 14 were selected. Whenappropriate, values for the 20-percent vehicle flow-path were obtained by simple geometric scaling(multiply by 0.04 for mass flows and areas).
Detailed flowpath geometries were generated.SCURRY and HyCAD, standard analysis tools inuse at Boeing in the SEASHELL analysis system(Fig. 8), were run to generate operability and perfor-mance information for these flowpaths. Some minoradjustments were necessary to account for opera-bility and scale effects. For example, the flame-holder sizes were adjusted to obtain appropriateoperability at the chosen conditions, and cowl lead-ing-edge radii were held to a practical size limit.
These flowpath models were then used in thesensitivity analyses that are discussed in Section 3.2.
Inlet Design Overview — FloGeo is a geometryengine was used for this study that constructs two-dimensional (2-D) flowpath configurations. Each ofthe three flowpaths was designed using the Flo-Geo geometry engine (Fig. 9) based on a set ofconstraints developed to ensure they were repre-sentative of configurations and missions of interest.These included: vehicle and engine scale, design(shock-on-lip) Mach number, forebody/inlet turningangles, overall contraction ratio, and nozzle exit-to-inlet capture area ratio.
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Fig. 8. SEASHELL System Schematic
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SCURRY is a 2-D Parabolized Navier Stokes(PNS) CFD code specialized for hypersonic fore-body/inlet solutions. This code was developed dur-ing the late 1980s for use on the NASP program. Itis the standard inlet analysis tool at Boeing forhypersonic conceptual and preliminary design.
Cruise applications, such as this missile,require high inlet performance at the designcondition. This is quantified by total pressurerecovery at low Mach number, and by ηke at highMach numbers – usually above Mach 6. Somewhatlower performance can usually be tolerated underacceleration/deceleration conditions because ofthe excess thrust available.
Accelerator applications require good inlet per-formance throughout the flight regime. Ideally, thecowl shock is canceled at the bodyside shoulder asin the missile inlet. However, unless extreme vari-able geometry techniques are employed, this is anunreasonable goal due to variations in flow condi-tions along the trajectory.
The features of FloGeo were used to rapidlydesign inlets with the desirable characteristicsmentioned above. Parametric analyses using up to
nine computers in parallel were performed to sat-isfy the combustor entrance constraint for the mis-sile configuration. These were followed by severalfine-tuning iterations to improve performance.
Missile — A distributed shock compression sys-tem was achieved by providing three external com-pression ramps. Shocks from the first two rampsfocus on the cowl lip at the design point, while thethird shock impinges on the cowl. The impinge-ment point was located downstream of the comple-tion of turbulent transition. This was done to avoidboundary-layer separation. Weak shock reflectionsoccur after the inlet shoulder radius and are dissi-pated by the shoulder expansion field. The pres-sure field is nearly uniform at the combustorentrance, thus promoting uniform air mass flow.
Because its small size and low cruise dynamicpressure, the inlet is confronted with a low Rey-nolds number viscous environment. This results inlarge boundary layers and high viscous losses.Comparison of reflected shock strengths in thissolution with empirical data shows that separationshould not occur in this inlet at the design condi-tion. Further evaluation would be required for off-design conditions.
Fig. 9. FloGeo Flowpath Definition Parameters
Hypersonic Flowpath Geometry Engine
Flowpath Stations
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HSFI (High speed fuel injection) Station
* Input as increment from θchordal
7American Institute of Aeronautics and Astronautics
Overall, this design performs well (mass cap-ture, pressure recovery, distortion) and is a goodexample of this class of inlet.
Vehicle — This design was analyzed at Machnumbers of 8, 11, and 14. The shock-on-lip Machnumber was 10; however, some margin in shocklocation was included. Only the Mach 14 case is“oversped,” with the bow shock impinging justdownstream of the cowl leading edge. Because ofthis feature, and the generous shoulder radius, theinlet flow field is acceptable throughout this widerange of operation. Comparisons of the 20- and100-percent scale computed flowfields were made,and, as expected, the 100-percent scale configura-tion is operating in a higher Reynolds numberregime and thus has proportionately thinner bound-ary layers.
Performance — Hypersonic inlet compressionefficiency is most often quoted in terms of the ηkeparameter.16-17 A value of 1.0 would correspond toisentropic compression.
As ηke increases for a given contraction ratio,static pressure increases, static temperaturedecreases, and inlet exit Mach number increases.All of these are related to reduced oblique shockstrength or reduced entropy increase. Overall cycleefficiency and combustor performance and opera-bility are functions of static pressure ratio. Com-
bustor entrance conditions were area averaged.Generally, the 100-percent vehicle has higher per-formance and lower distortion than the 20-percentvehicle. The combination of larger scale and higherReynolds number reduces the effects of viscosity,which plays a major role in inlet performance.
3.2 Flow Vitiation Calculations
This section contains a series of calculationsperformed with the HYCAD code (Fig. 10) to helpassess the impact of flow vitiation on scramjet flow-path performance parameters. Two types of flowvitiation are examined: instream combustion andNO contamination.
HYCAD is a one-dimensional (1-D) multi-stream, space-marching finite-rate, kinetic com-bustion code supplemented by distortion theorywhich solves a system of “N” co-flowing streamsincluding effects of finite-rate chemical kinetics,multistream mixing, wall drag, and heat transfer.Up to 17 co-flowing streams can be used with akinetic mechanism of 2500+ reactions. Ignition andflame holding are modeled by an integral, well-stirred reactor and shock trains modeled using theSUPDIF or McLaffertty correlations. The calcula-tions are applicable to both subsonic and super-sonic conditions. In addition, through special treat-ment at the sonic point, HYCAD is capable of mod-eling dual-mode ramjet thermal throat conditions.
Fig. 10. Multistream HyCAD Model Features Hypersonic Combustor Analysis and Design Code
INLET
• Supersonic
Compression
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• Distortion
• Length
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Recovery
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• Drag
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8American Institute of Aeronautics and Astronautics
Mach 8 20-percent Scale Vehicle — Numeroussets of facility test medium flow “match” conditionshave been used in propulsion testing. These condi-tions match various combinations of pressure, tem-perature, enthalpy, velocity, or Mach number enter-ing the inlet or entering the combustor. The addi-tion of water and carbon dioxide to the flow in acombustion-heated facility results in an alteredmolecular weight, gamma, and heat capacity of thesimulated air stream relative to the flight condi-tions. The resultant facility test medium flow is viti-ated or contaminated. Because of the altered prop-erties of the combustion-heated flow, the streamvelocity, Mach number, and temperature cannot besimultaneously matched. This is illustrated in con-sideration of the flow Mach number:
The flow velocity and static temperatureparameters are of obvious importance in acombustor simulation. These parameters allowmatching of the residence time and chemicalreaction rates. In a velocity-matched flow, if theMach number is also to be simultaneouslymatched (critical for shock train and inletperformance simulation), then the gamma-to-molecular weight ratio (γ/WG) must remainidentical to that of air. Instream combustionheating, particularly with hydrogen, results in analtered gamma to molecular weight ratio andprecludes simultaneous matching of the flowvelocity and Mach number. The mismatch ingamma also affects the static temperaturesoccurring in the near-stagnated recirculation zonesof flame holders and wall steps. In these high-temperature regions, the variation of gamma withtemperature, in addition to the inherent mismatchin gamma, results in local flow conditions that aredifferent from that of the true flight case.
Instream combustion heating with ahydrocarbon-fueled heater reduces the flowmismatches caused by hydrogen combustionheating. The heavier molecular weight of CO2,generated by hydrocarbon combustion, partiallyoffsets the lighter molecular weight of water,reducing the discrepancy in molecular weightrelative to combustion heating using hydrogen. Theheat capacity of CO2 is similar to that of air on amass basis; however, the water formed throughthe combustion process results in an overallincrease in heat capacity and a discrepancy in
gamma. This is illustrated in Table 6 for typicalMach 8 combustor entrance conditions.
In stream combustion heating with propane orisobutane produces molecular weights that arevery close to that of air. However, the contaminantsproduced by the combustion process result in areduced gamma and a gamma-to-molecular weightratio that is lower than the pure air case. The viti-ated flow studies presented here are based oncombustion of isobutane with oxygen, whichresults in a molecular weight that is slightly greaterthan that of air and a Mach number that is slightlyhigher than the velocity-matched pure air case dueto the reduced gamma-molecular weight ratio.
The following combustion-vitiated flow condi-tions are based on the temperatures of ambientinlet gas supplied to the heater using isobutane/aircombustion and the necessary makeup oxygenaddition to bring the heater exit gas oxygen con-centration to a mole fraction of 0.2096. Themakeup oxygen is also added at ambient tempera-ture.
Two sets of typical test medium “match” condi-tions frequently used in combustor testing areincluded in Table 7. The first set matches the com-bustor entrance enthalpy (H), pressure (P), andMach number (M) and will herein be referred to as"HPM conditions." These conditions are intendedto simulate engine performance from an overallenergy addition standpoint and to simulate theMach number-based phenomena, which are partic-ularly important throughout the ramjet operationregime where precombustion shocks and shocktrains have important effects on the engine opera-tion. For these Mach 8 flight conditions theenthalpy-matched conditions result in a combustorentrance static temperature that is higher than thatof the flight case. The higher temperature has anobvious impact on augmenting the ignition processand flame propagation rates in the combustor. Theenthalpy-matched conditions also result in a lower
Mach u
γ RTWG-------------
-----------------=
Table 6. Flight Compared to Vitiated FlowConditions
Medium γ WG, lb/mol γ/WG
Air 1.332 28.965 0.046
H2 Vitiated 1.298 25.778 0.050
Propane Vitiated 1.288 29.223 0.044
Iso-Butane Vitiated 1.288 29.328 0.044
9American Institute of Aeronautics and Astronautics
velocity at the Mach number-matched condition.The lower velocity extends the flow residence timein the combustor, further augmenting the combus-tion process. Thus evaluation of the effect of testmedium HPM match conditions on combustor per-formance requires consideration of the combinedeffect of increased temperature and reduced veloc-ity. Temperature has an exponential effect on igni-tion which is well documented in ignition delaydata,18 and velocity has a linear effect. The HyCADfinite-rate kinetics combustor performance calcula-tions in the following figures evaluate the combinedeffects of these parameters as well as the effect ofthe test medium chemical composition on engineperformance. These comparisons are made rela-tive to the engine performance calculated for flightconditions.
The second set of test medium “match” condi-tions is based on a velocity (U), pressure (P), tem-perature (T) (UPT) matched flow. These conditionsmatch the local static environment of the flow andare intended to simulate the flow from a chemicalkinetics standpoint in terms of matching the localstatic temperature, pressure, residence time, andoxygen concentration on a mole fraction basis. Theeffect of free radicals and combustion intermedi-ates generated by the combustion heater on aug-menting the ignition and combustion processshould be highlighted by these flow conditions. Forthe above Mach 8 conditions, the UPT match con-ditions result in an increased molecular weight andthe highest engine mass flow, which will directlyimpact the engine thrust.
The HyCAD computed pressure distribution forthe 20-percent scale vehicle operating under Mach8 flight conditions and stoichiometric hydrogen
fueling is shown in Fig. 11. The mixing length wasset to 20 gap heights, where gap is the duct heightat the combustor entrance (isolator exit). The 20-gap mixing length was selected as a mixing lengthtypical of modern engine designs and has beendemonstrated in previous combustor testing suchas the Marquardt19 hydrogen-fueled scramjetengine. This mixing length is approximately 1/3 ofthe length typically quoted on the basis of the 60-gap Langley20 sidewall injector design correlationand is the result of careful configuration of theinjector.
The HyCAD calculation is initialized at theentrance to the isolator (23.3 ft) with the flow condi-tions determined by the forebody/inlet analysisshown in Table 7. The duct walls in the isolator areslightly diverging, with a 0.36-deg angle on eachwall. The duct divergence creates the droppingpressures observed in Fig. 11 before the initiationof the shock train.
Fuel is injected at 24.7 ft from the nose in the20-percent vehicle. A 0.050-in. step downstream ofthe injectors is used to stabilize the flame. As indi-cated by the pressure distribution, all three casesignite upon fuel injection. For the flow conditions ofthe 20-percent vehicle configuration, hydrogen iscapable of autoignition without the aid of a flameholder. Under these conditions the presence of aflame holder makes little difference on the ignitionprocess, and the effects of trace contaminatesfrom combustion heating do not influence ignitionto any measurable extent. However, alternate vehi-cle configurations required to fly at lower dynamicpressures or lower contraction ratios can be influ-enced by the contaminants at Mach 8 conditions.
10American Institute of Aeronautics and Astronautics
This unfortunately requires a case-by-case evalua-tion for vehicle configurations that depart signifi-cantly from the “typical” configuration and operat-ing conditions selected for the study.
The computed HyCAD pressure distributionsshow that the clean air flight conditions achieve thehighest peak combustor pressure. The higher pres-sure is caused by higher combustion temperaturesfor the clean air case, which are a result of thelower specific heat for the clean air. Both the HPMand UPT combustion heating achieve similar peakcombustor pressures. In detail, the HPM casereaches a slightly higher pressure because of thecombined effects of lower combustor entrancevelocity and higher static temperature.
The exit stream thrust and associated air ISP(stream thrust divided by air mass flow) for thesethree cases are shown in Table 8. The HPM viti-ated flow case has the lowest mass flow and thelowest thrust. Normalizing these three cases on anair ISP basis reduces the differences among thesecases to approximately 2 percent, with the clean aircase offering the best performance. The improvedperformance of the clean air case is created by theincreased combustor peak pressure, which is theresult of the lower heat capacity of dry, clean air.
For the 20-percent vehicle scale Mach 8 condi-tions used in this study, the effects of composi-tional mismatch created by combustion heatingresult in only modest differences in engine perfor-mance. The hydrogen-fueled scramjet configura-tion investigated is capable of auto ignition of theinjected fuel in a relatively short distance. Thus, forthese conditions, the effect of trace contaminantson the ignition and flame propagation processeshas little influence on combustor performance. Thedifferences in heat capacity attributable to wateraddition from combustion heating limit the combus-tor peak temperature and pressure. The reducedcombustor peak pressures under vitiated test
medium conditions also reduce the wall heat trans-fer rate, which is an important consideration in thescaling of ground test data to flight conditions. Thereduced combustor pressure under vitiated testmedium conditions also influences the speciesrecombination rates in the nozzle. For the condi-tions of this study, the differences in peak combus-tor pressure between vitiated flow and flight condi-tions resulted in an 0.8 percent increase in the noz-zle exit combustion efficiency for the flight case. Inaddition, the occurrence of precombustion shocksand upstream interactions into the isolator aredirect results of the peak combustor pressure.Evaluation of engine operability limits demon-strated under Mach 8 vitiated test medium condi-tions requires consideration of the effect of vitiationon the peak combustor pressure and the demon-strated operability limits. These differences, how-ever, are small and can generally be correctedthrough adequate analysis.
Mach 11 20-percent Scale Vehicle
The vitiated flow conditions in Table 9 are basedon ambient inlet gas temperatures to the heaterusing isobutane/air combustion and the necessarymakeup oxygen addition to bring the heater exit gasoxygen concentrations to 0.2096 mole fraction. Theflow conditions are set to match the isolator inletflow conditions determined from the inlet analysis.The HPM conditions are set to match the isolatorentrance enthalpy, pressure, and Mach number.The UPT conditions are set to match the isolatorentrance velocity, temperature, and pressure.Combustion heating with isobutane for these casesresults in the addition of a water mass fraction of0.174 for the HPM case and a water mass fraction
11American Institute of Aeronautics and Astronautics
of 0.182 for the UPT case. While this represents asubstantial addition of water to the flow, the possi-bility of water condensation in the facility nozzle isremote since the exit static temperature of the noz-zle, over 2600°R, is well above the critical tempera-ture of water (Tc = 1165°R).
The enthalpy-matched conditions have a lowervelocity and higher static temperature, both ofwhich can potentially augment ignition and com-bustion. Because of the effect of combustion prod-ucts on the molecular weight and gamma, thevelocity-matched conditions result in a higherentrance Mach number.
The computed pressure distributions for the 20-percent vehicle Mach 11 flight conditions areshown in Fig. 12. The features of the pressure dis-tribution are similar to those obtained under Mach8 conditions. From the isolator entrance to the fuelinjection station, the drop in pressure distribution isattributable to the area divergence of the isolator.Fuel is injected at 24.7 ft. All three cases showimmediate combustion directly after fuel injection.At the elevated temperatures of these Mach 11flight conditions, there is little difficulty in ignitingthe injected hydrogen in the combustor. Underthese conditions, the trace combustion productintermediates have little impact on the ignition orcombustion process.
The higher specific heat of the vitiated flowreduces the peak combustor temperature, which,in turn, reduces the peak combustor pressure. TheUPT and HPM vitiated flow conditions producealmost identical pressures and appear as onecurve in Fig. 12.
In addition to the effect of increased specificheat, the vitiated flow also has a reduced oxygenmass fraction because of the oxygen displacementby water in these mole fraction-matched heaterconditions. The reduced oxygen mass fraction lim-its the peak combustor pressures of the vitiatedflow. To quantify this effect, the UPT match condi-tions of Fig. 12 were rerun with the oxygen massfraction increased from the mole fraction-matchedconditions (resulting in an oxygen mass fraction of0.225) to an oxygen mass fraction of 0.2315, iden-tical to that of the flight case. These results areshown in Fig. 13. With the heater flow oxygen massfraction increased to 0.2315, the combustor peakpressure is slightly increased; however, the peakcombustor pressure is still far below the pressuresachieved by the flight condition. This includes theeffect of the increased oxygen concentration on thekinetics.
The concentration, and hence the magnitude ofthe effects, of the products of instream combustioncan be reduced if some of the energy required forhigh Mach number simulation is added by heatingthe air (and oxygen) prior to the combustion in thefacility. A preheater temperature of 4000°R isbeyond demonstrated capabilities, but some recentwork21-22 suggests that temperatures approachingthis level may be achieved in the not too distantfuture. It is appropriate to consider the effect of theanticipated preheating capabilities.
Figure 13 also includes HyCAD calculationresults for the case of pebble bed preheating priorto combustion heating. For this case, the combus-tion heater airflow and makeup oxygen were pre-heated to 4000°R. This preheated flow was then
30 35 40Feet
Pre
ssur
e
20 25
Flight
Vitiated (HPM)
Vitiated (HPM)
Fig. 12. 20-percent Scale Vehicle Mach 11Pressure
20 25 30 35 40Feet
Pre
ssur
e
Flight
Vitiated (UPT)
(UPT O2 Mass
4000˚R Preheat
Fig. 13. Mach 11 Pressure Compared to O2 MassFraction Match and 4000-deg PreheatConditions
12American Institute of Aeronautics and Astronautics
burned with isobutane to bring the heater exit totaltemperature to 6359°R. When expanded throughthe facility nozzle under equilibrium conditions, theresulting flow was set to match the velocity, temper-ature, and pressure of the flight conditions. Throughpreheating of the combustion heater entrance flow,the water mass fraction decreased from 18 to 11percent. The resulting reduction in specific heatprovides a closer match to the flight conditions.However, even with preheating, the exit specificheat is still higher than that of air and still results inthe suppressed peak combustor pressures asshown in Fig. 13.
The results of Figure 13 at least partially isolatethe effect of combustion heating, showing that thereduction in peak combustor pressure is caused bythe increased specific heat of the wet vitiated flow.This effect can be reduced by preheating of theheater entrance flow; however, the water added bycombustion in the heater to bring the final totaltemperature up to the design point increases thevitiated test medium flow specific heat and leads tosuppressed peak combustor pressures as shownin Fig. 13.
The exit stream thrust and associated air ISP(stream thrust divided by air mass flow) for theflight and vitiated flow cases are shown in Table10. The HPM vitiated flow case has the lowestmass flow and the lowest thrust. When normalizedon an air ISP basis, the differences between thesethree cases are on the order of 3.5 percent, withthe clean air offering the best performance.Increasing the oxygen mass fraction in the UPTvitiated flow case to match those of the flight caseincreases the engine thrust and ISP performance.These levels, however, are still below thoseachieved by the flight condition. Preheating thecombustion-heater entrance air and makeup oxy-gen to 4000°R reduces the required combustion inthe heater and the resulting water mass fraction inthe test medium. These conditions bring the com-bustor peak pressures and air ISP closer to thoseof the flight case; however, the air ISP is still 1 per-cent below those of the flight case.
Mach 14 20-percent Scale Vehicle
The required enthalpy for simulation of Mach 14flight conditions cannot be achieved with isobutane/oxygen combustion without preheating theentrance flows. For isobutane combustion, theentrance air/oxygen flow must be preheated to at
least 3300°R to provide the required total tempera-ture. The vitiated flow conditions in Table 11 arebased on preheating the combustion-heatedentrance air and oxygen to a total temperature of4000°R. This study has not dealt with the issuesassociated with the high temperatures and pres-sures necessary to build such a facility. It hasquantified the aerothermal effects only. The com-bustion heater burns isobutane/air along with thenecessary makeup oxygen to bring the combustorexit gas oxygen concentration to a mole fraction of0.2096. The flow conditions are set to match theisolator inlet flow conditions determined from theinlet analysis of the flight configurations. The HPMconditions are set to match the isolator entranceenthalpy, pressure, and Mach number. The UPTconditions are set to match the isolator entrancevelocity, temperature, and pressure.
The enthalpy match conditions have a lowervelocity and higher temperature, both of which
13American Institute of Aeronautics and Astronautics
potentially augment ignition and combustion. TheUPT conditions are set to match the isolatorentrance velocity, temperature, and pressure.Because of the effect of combustion on the molec-ular weight and gamma, the velocity-matched con-ditions result in a higher entrance Mach number.
The computed pressure distribution for the 20-percent vehicle Mach 14 flight conditions is shownin Fig. 14. The features of the pressure distributionare similar to that of both the Mach 8 and Mach 11pressure distributions. From the isolator entranceat 23.3 ft, the area divergence of the isolator ductcreates a pressure drop. Hydrogen fuel is injectedat 24.7 ft and ignites immediately on injection dueto the high static temperatures present at the injec-tion station. The flame is stabilized by a 0.050-in.step downstream of the injectors; however, atthese Mach 14 conditions the presence of the stepmakes little difference in ignition.
The clean air flight conditions achieve the high-est combustor peak pressure. Both the HPM- andUPT-matched vitiated flow conditions present simi-lar combustor pressure distributions and similarpeak combustor pressures that are below that ofthe flight condition. The reduced peak combustorpressures are caused by the increased heatcapacity of the vitiated flow, which reduces thetemperature rise for a given level of heat release.The higher pressures of the flight case give rise toan increased integral of PdA down the nozzle, cre-ating higher thrust levels for the flight case.
Figure 15 presents the flow static temperaturedistribution through the engine. The temperaturedecrease due to fuel addition at the injection sta-tion is evident. After injection the fuel immediately
ignites, and the temperature rise from combustionis evident in the figure. The peak combustor tem-peratures for the vitiated flow cases are reducedbecause of the increased heat capacity of the viti-ated flow. Both flows continue to burn through thenozzle, and the recombination rates are such thatthe static temperatures at the nozzle exit are nearlyidentical for all cases.
As an alternative to combustion heating for theMach 14 flow conditions, these studies have alsoconsidered arc heating. Arc heating avoids themismatch in specific heat created by water additionfrom combustion heating. The mismatch in specificheat is the primary cause of the performance differ-ences noted in the Mach 8, 11, and 14 vitiatedflows. Arc heating, however, introduces its own setof complexities. Routine arc heater operation iscurrently limited to approximately 120 atm, which isadequate for Mach 9 to 10 simulation. Increasingthe pressure level significantly will be difficult; how-ever, achieving a Mach 11 pressure level might bea reasonable near-term goal. Hence, the Mach 9 to11 regime might be the most appropriate regimefor which to consider using an arc heater. Theaddition of NO and NO2 to otherwise clean air hasbeen observed by Slack and Grillo18 to enhanceignition of hydrogen/air mixtures. The effectsobserved by Slack and Grillo show the ignitiondelay reduction by NO addition to be maximized atmole fraction concentration levels of 0.005 to0.007. The effect of NO on ignition is discussed inmore detail in Section 3.3. See Ref. 23 for a dis-cussion related to Hyper-X.
In addition to the effects of NO on ignition, theapplication of an arc heater at Mach 14 conditions
14American Institute of Aeronautics and Astronautics
also requires consideration of effects of dissociatedspecies and electrically charged ions on the com-bustion and fluid dynamic processes within theengine. The Konnov kinetic mechanism does notcurrently address these phenomena. Additionalwork is needed to define the level of dissociationand charged concentrations in an arc-heated flowand the effect of these species on the combustionand fluid dynamic processes within the engine.
The current application requires an estimate ofNO and NO2 concentrations added by the arc forpreliminary estimation of its effect on engine perfor-mance. Under Mach 10 conditions the NOx levelsadded by an arc are typically in the 4 to 7 mole per-cent range.24 Figure 16 presents equilibrium NOand NO2 concentrations as a function of tempera-ture for clean air and heated by isobutane/oxygencombustion. The air heated by isobutane combus-tion shows peak NO concentrations of approxi-mately 2 percent and an abrupt fall off above tem-peratures of 3300K. The fall-off is attributable toreduced N2 concentrations as the isobutane heatedair approaches the composition of an isobutane/oxygen torch. For heated air, the NO concentra-tions reach a maximum concentration of 8 percentat temperatures of 4100K. Above this temperature,NO concentrations begin to fall, as expected, due todissociation/decomposition at high temperature.For the Mach 14 application, a total temperature ofapproximately 7200K is required, and for thesesample calculations an arc heater exit NO concen-tration of 8 percent was assumed. It was assumedthat the concentration was maintained through thefacility into the flowpath combustor region. This is arelatively high level of NO, and was picked to“bound” the effect. Actual concentrations could besignificantly lower.
The HyCAD-computed pressure distributionsfor flight, arc heated, and combustion-heated con-ditions are shown in Fig. 17. The arc-heated condi-tions, having a specific heat that is close to that ofair, result in combustor peak pressure and nozzlepressure distributions that are very close to thoseof the flight case. The small differences in the com-bustor peak pressure can be corrected through theaddition of makeup oxygen to compensate for theoxygen bound in the NO formed in the arc heater.For the combustion-heated case the combustorpeak pressure and nozzle pressure distribution aremuch lower because of the increased heat capac-ity and water concentrations of the combustion-heated flow conditions.
The exit stream thrust and associated air ISP forthe flight and vitiated flow cases are shown in Table12. The HPM vitiated flow case has the lowestmass flow and the lowest thrust. When normalizedon an ISP basis, the differences between these fourcases are on the order of 4 percent. The arc-heatedcase has a slightly higher air ISP than that of theclean air flight case because of NO decompositionin the combustor.
15American Institute of Aeronautics and Astronautics
Full-Scale Vehicle and Missile — Similar calcu-lations were completed for the full-scale vehicle atMach 8, 11, and 14 and the full-scale missile atMach 8.
Comparisons of Flowpath Performance — Acomparison of the flowpath performance (specificimpulse) for flight and ground testing of the hydro-gen-powered vehicle is shown in Fig. 18. In this fig-ure, the following cases are compared: 1) full-scalevehicle at flight conditions; 2) 20-percent scalevehicle at flight conditions (represents a subscaleflight test experiment); 3) hydrocarbon combustionmatching flight enthalpy, pressure, and Mach(HPM) at Mach 8 and 11; 4) hydrocarbon combus-tion matching flight velocity, pressure, and temper-ature (UPT) at Mach 8 and 11; 5) preheating to4000°R plus hydrocarbon combustion matchingflight enthalpy, pressure, and Mach (HPM) at Mach14; 6) preheating to 4000°R plus hydrocarbon com-bustion matching flight velocity, pressure, and tem-perature (UPT) at Mach 11 and 14; and 7) arc-heated flow at Mach 14.
These cases do not necessarily include all com-binations, but do give a feel for the comparativeperformance that can be obtained in various typesof facilities. Generally, all types of vitiation (fromhydrogen or hydrocarbon combustion or arc heat-ing) produce lower performance than flight, whichrequires posttest analysis to accurately predictflight performance. The closer a method is to theflight values, the less emphasis is placed on post-test performance extrapolations. We draw the fol-lowing conclusions: 1) based on this analysis, arc
heating, or facility methods that do not add combus-tion products to the “freestream” flow, will requiresmaller corrections than methods that add combus-tion products to the freestream flow; 2) the UPTflow-matching approach is superior to the HPMapproach for overall flowpath performance testing;and 3) subscale flight testing is an appropriatemethod for obtaining at least some flowpath data.
A comparison of the flowpath performance(specific impulse) for flight and ground testing ofthe hydrocarbon-fueled missile is shown in Fig. 19.Here, the effect of vitiation has a relatively smallimpact on the overall performance. With carefulposttest analysis, the ground testing can be madetraceable to flight results.
Overall, this study did not find that ignition was amajor driver on high Mach facility test requirements.Because of this, our sensitivities to vitiation effectswere relatively small. For vehicle trajectories atmuch lower dynamic pressures (and lower combus-tor pressures), these vitiation effects may be larger.
3.3 Discussion of Flow Vitiation
Problem Definition — A major issue in currenthypersonic ground test facilities is the composi-tional mismatch between the test medium and theair encountered under true flight conditions. Belowabout Mach 4 to 6 (Tt ~ 1600° – 3000°R), clean airheating is achieved by heating methods includingmetallic and ceramic storage/bed, tube and shell,and electric resistance heaters. The use of thesetypes of heaters presents no critical heated air sim-
80
82
84
86
88
90
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96
98
100
6 7 8 9 10 11 12 13 14 15Mach
ISP
, Th
rust
/Mdo
t Air,
per
cent
100% Scale Flight
20% Scale Flight
Vitiation (HPM)
Vitiation (UPT)
Preheat + Vitiation (HPM)
Preheat + Vitiation (UPT)
ARC Heated
Fig. 18. Comparison of Flowpath Performance forVarious Flight and Facility TestApproaches (Vehicle Flowpath)
80
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84
86
88
90
92
94
96
98
100
100% ScaleFlight
100% ScaleVitiation, HPM
100% ScaleVitiation, UPT
ISP
, Exi
t S
tre
am T
hru
st/A
irflo
w, p
erce
nt
Fig. 19. Comparison of Missile Flowpath Perfor-mance for Flight and Facility TestApproaches
16American Institute of Aeronautics and Astronautics
ulation issues. Accounting for flow property mis-matches in this test regime is straightforward, and,while the potential exists for minor vibrational relax-ation effects in the facility nozzle expansion pro-cess, there is no significant dissociation, and thegas is diatomic, allowing the effect of this mecha-nism to be explored by classical methods.25
Testing up to temperatures of about 5000°R,corresponding to the Mach 8 regime, in-situ com-bustion heating is commonly employed. Reference25 showed that by preheating the test medium upto 4000°R in a clean air stage, such as a storageheater, Mach 10 conditions could be achieved withsecond stage combustion using hydrogen or pro-pane. Hydrogen and hydrocarbons are used,including, but not limited to, methane/air (NASA-LaRC 8-ft HTT), propane/air (NAWC/China Lake T-Range), and iso-butane/air (AEDC APTU). Monom-ethyl hydrazine-NTO has also been considered to alimited extent (Aerojet ABP facility, currently out ofoperation). In each case, makeup oxygen is used tomaintain 21 percent oxygen by volume.
Testing at speeds above the Mach 8 regime cur-rently involves the use of arc or compression heat-ing. Arc heaters are capable of providing total tem-peratures up to 15,000°R but are typically consid-ered for testing up to Mach 12 to 14 (Tt ~ 7000° to12,000°R). However, required pressures can befrom 2000 to 10,000 atmospheres (depending ofthe trajectory simulated), posing serious difficul-ties for the operation of an arc facility. Compres-sion and expansion tubes have been employedto test beyond this range. Their use is appropri-ate for processes and concepts that can be stud-ied and evaluated in test times in the 1- to 10-msec range.
The present focus is on the Mach 8 to 14flight speed range, where combustion and archeating are of interest for scramjet testing. Aspreviously cited, a major issue is the inability tomatch the composition and pressure of the airencountered in flight. Previous work in the Mach4 to 7 range provides some insight into this prob-lem. Water is the dominant species in combus-tion heating. The volume fraction is over 30 per-cent for hydrogen, while the hydrocarbons pro-duce a lesser amount, with smaller concentra-tions of CO2, CO, and NO. Arc heating also pro-duces NO, and very small amounts of water andCO2 are also present, presumably very close tothe composition of typical ambient air. However,
water and NO have been known for some time tobe two of the most important species in terms of theimpact of compositional mismatch effects onground testing. The effect on combustion kineticsand, potentially, nozzle recombination rates, aresignificant.18, 25-28 The studies cited have alsoshown that the high static temperatures encoun-tered in the heating of the test media will producehighly reactive radicals, including O, H, and OH,which, if present in the test air on entry to the testcombustor, can enhance ignition and flame propa-gation rates. However, the effects of the water andNO present in the test medium are more complexand can either enhance or impede the combustionrate in the test article. The problem is exacerbatedby the thermodynamic path typically followed by theairflow from the inlet to the combustor entrance intrue flight.
Figure 20, from Ref. 29, shows typical tempera-tures for a generic space access vehicle. Althoughconsiderable overlap exists for the regions identi-fied, it is useful in understanding the flow tempera-tures and physics. Flying above about Mach 10, theflow entering the combustor can contain atomicoxygen since the boundary layer will constitute asignificant fraction of the flowpath cross section andpeak boundary layer temperatures are sufficient todissociate the oxygen in the air. Therefore, match-ing freestream conditions for integrated engine
Tem
pera
ture
, ˚R
0 5 10 15 20 25Freestream Mach Number
Ideal GasRegion
VibrationalExcitation
Region
O DissociationRegion
NDissociation
Region
IonizationRegion
Temperature After Fi rst Wedge
Temperature After Isentropic TurnTemperature Entering Combustor
Maximum Te mpe
rature In Bou
ndary
Layer
Free
stre
am T
otal
Tem
pera
ture
1000900800700600500400
300
2000
3000
4000
50006000700080009000
10,000
20,000
30,000
40,00050,000
Fig. 20. Temperatures Required for Testing
17American Institute of Aeronautics and Astronautics
tests and, moreover, matching combustor entranceconditions for direct or semidirect combustor testsin the double-digit Mach number range must bedone with considerable care. While complete dupli-cation of conditions encountered in flight is not apractical goal, a reasonable understanding of thekinetics processes and the coupling with the fluiddynamics in each specific propulsion system con-cept is essential. This understanding is needed todefine the relationship between test and flightresults, including the question of which properties,and, hence, facility operating conditions to selectfor matching.
The present study emphasizes the kinetics pro-cesses with specific attention given to the notionalvehicles, including the Mach 8 missile at the cruisecondition and the full-scale and 20-percent scalespace access systems at Mach 8, 11, and 14.Although emphasis is on the kinetics, a compre-hensive engineering analysis was used to couplethe kinetics to the flowpath fluid dynamics (Section3.2).
Test Media Contamination Effects — The pres-ence of combustion products and NO in the testmedium causes several changes that affect thethermodynamic properties and reactivity of theflow, particularly in the combustion chamber of thetest article. The presence of combustion productschanges the mixture molecular weight andincreases its heat capacity, Cp. A comparative the-oretical study in Ref. 25 showed that the thrustfrom a freejet scramjet test conducted in hydrogencombustion-heated air simulating near Mach 10conditions is about 10 percent lower than would beexpected in flight. This calculation was carried outmatching static temperature, static pressure, andMach number in the test section, assuming equilib-rium flow including the facility nozzle expansionand test engine. In this singular case carried out inReference 25, where staged heating was usedassuming clean air at 4000°R via a pebble bed fol-lowed by combustion heating over a simulatedMach range up to Mach 10, it was observed thatthe lower thrust predicted for the ground test vs.flight is mainly due to the reduced mass capture,i.e., reduced density. The reduced effect of heatrelease in the test engine combustor also contrib-utes to reduced thrust because of the increase inspecific heat of the mixture. Propane combustionwas also studied in Ref. 25, but a correspondinganalysis for this freejet scramjet test was not car-ried out. However, based on the equilibrium
assumption, the reduction in the thrust measuredin ground testing would probably be substantiallyless for hydrocarbon combustion since the molecu-lar weight of equilibrium products of hydrocarboncombustion is close to that of air. The relativeimpact of property mismatches at other Mach num-bers and for different instream combustion heaterfuels remains to be evaluated. This includes theimportant question of the selection of test condi-tions to match with flight. This issue is stronglydependent on the kinetics of the combustion pro-cess in the test article.
Chemical Kinetics — As previously cited, thepresence of combustion products or NO can affectthe rate of combustion by either enhancing orimpeding it. This depends largely on the speciesthat are present at the entrance to the combustor,and the temperature, pressure, velocity (residencetime) and geometry, including scale. Even at tracelevels as small as 10−3 to 10−5 in mass fraction, theeffect of species such as O, H, and, particularly, OHcan be significant.
To evaluate the effect of the presence of radi-cals in the combustor entrance flow, a set of para-metric plug flow calculations for ignition delay andreaction times were carried out.26 Table 13 givesthe type and range of contamination levels used inthe plug flow calculations assuming stoichiometrichydrogen-air combustion. Hydrocarbon-air com-bustion results would be different. The parametricset includes mixtures of species relevant to hydro-gen and hydrocarbon combustion heating assum-ing complete combustion (H2O and CO2). Figure 21shows that at a static temperature of 1800°R(1000K) and a static pressure of 1 atm, the additionof 10−5 mass fraction of OH can reduce the ignitiondelay time by about a factor of 5 (compare cases 5and 6 vs. case 7 for pure air). In cases 5 and 6, H2Oand CO2 act as inert diluents and tend to increasethe ignition delay and reaction times and decreasethe combustion temperature in proportion to theireffect on the heat capacity of the mixture. In partic-ular, water has a specific heat nearly twice that ofair on a mass basis and accounts for the smallincrease in ignition delay for case 5 relative to case6 where OH is present in the initial mixture. Notethat the addition of CO2
has a negligible effect rela-tive to pure air (cases 3 and 4 vs. case 7). However,a small amount of water with no initial OH in theinlet air produces radicals by dissociation, and aslight reduction in ignition delay is observed (case1) whereas, doubling the water content for these
18American Institute of Aeronautics and Astronautics
conditions results in essentially cancelingthis effect.
The sensitivity to OH concentration isdemonstrated in case 8 with an addition ofOH at only 10−7 mass fraction. Thisresults in a more significant reduction indelay time (compare case 8 with cases 1and 2). It appears evident from theseresults that the rate of combustion isdependent on the composition of the flowentering the combustor in terms of thepresence of radicals or the ability to rap-idly generate them. Case 7 is the base-line, with clean air and hydrogen combustion. Thefirst set of columns is the incoming air. The secondset of columns is after hydrogen combustion. How-ever, these parametric results cannot be general-ized for application to more realistic vitiated testmedium environments.
As previously stated, nominal Mach 10 condi-tions were achieved by staged heating using a stor-age-type heater to raise the temperature to4000°R, followed by in-situ combustion. Results forthe test engine combustor temperature-time histo-ries at a Mach 9.5 flight condition are shown in Fig.22 for hydrogen combustion heating. They indicatethat the ignition delay time for clean air is shorterthan that obtained for vitiated air resulting fromcombustion heating. This is not in complete contra-diction with the previous parametric results on theeffect of radicals on reducing combustion time, butrather shows the important effect of the mismatch incombustor entrance conditions when only staticpressure, static temperature, and Mach number arematched in the freestream. This effect is illustrated
in Fig. 22 on the basis of equal static pressure andstatic temperature by adjusting the clean air state (p= 2.93 psia and T = 2057°R) to match the vitiatedair values of pressure and temperature (p = 2.63psia and T = 1865°R) at the combustor entrance.This reverses the relative trend and shows a slightlylonger ignition delay time for the clean air. How-ever, the overall reaction time for the clean air is still
Table 13. Cases Chosen for the Study of the Effect of Vitiation on Finite-Rate Stoichiometric Combustionof Hydrogen at 1000°K and 1 atm
See Table 13 forDescription ofDetails for theCases Shown
2
4
10−5 10−4 10−3 10−2
Time, sec
Fig. 21. Effects of Initial Free Radical Concentration on H2-AirCombustion Histories
5000
4500
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1500
Te
mp
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ture
˚R
10−5 10−4 10−3 10−2
Time, sec
b) Effect ofH2 - Vitiation
Basic CombustionHistories
Dry Air WithPressure AdjustedTo Vitiated Level
Dry Air WithTemperature andPressure AdjustedTo Vitiated Level
Vitiated Airp = 2.63 psia
Dry Airp = 2.93 psia
PO = 2000 psia
TO = 5803˚R
Xf = 0.023
Fig. 22. Effects of Vitiation of H2-Air CombustionHistories
19American Institute of Aeronautics and Astronautics
somewhat shorter, i.e., the rate of combustion tocompletion (reaction time) is faster in clean air forthis case. This result shows the importance of tun-nel operating conditions and the sensitivity of theresulting states of flow and combustion process tothe mismatch in composition between clean air andvitiated air. These results are also limited to hydro-gen combustion heating and to a single high flightMach number condition using hydrogen in theengine where autoignition occurred without coreflow ignition support. Furthermore, the steady one-dimensional flow and chemical kinetics analysis didnot require the use of a flame-holding device andnone was used in the engine flowpath model. Inreality, such devices would be required to achievesteady, controlled combustion in the engines burn-ing hydrogen or a hydrocarbon. These features ofcore flow ignition and flame holding are included inthe analysis of the notional vehicle flowpaths, previ-ously discussed.
The results previously discussed did not includeHO2, nor the generation or presence of nitrogenoxides, NOx, in the chemical kinetics analysis. Theimportance of HO2 depends on the operating condi-tions and can impede combustion by interferingwith chain branching reactions in the oxidation ofhydrogen as well as hydrocarbons. Nitrogenoxides, particularly NO, form in high temperaturecombustion and are important since they canenhance or impede combustion in the test engine,depending on their concentration and the operatingconditions. In addition, the arc-heating processtends to produce significant amounts of NO. There-fore, it is important to include these species andreactions in the kinetics models used in accurate,more comprehensive analyses of test media heat-ing methods and contamination effects.
The primary mechanism of importance in thekinetic propagation and quenching process is thecompetition between radical formation by branch-ing and the formation of species such as HO2,which act mainly to terminate the reaction process.The following two reactions are important in thisprocess for hydrogen and hydrocarbon combustion.
H + O2 = O + OH, (1)and
H + O2 + M = HO2 + M, (2)
where M is a third body. Thus, reaction 2 is actuallya set of reactions, one for each third body whoseimportance depends on its concentration and chap-
eron efficiency. Since water can be more than 16times more efficient than nitrogen in catalyzingreaction 2, its effect on augmenting the kineticquenching of ignition in vitiated air may be very sig-nificant. However, the net effect of the competitionbetween reactions 1 and 2 depends on pressureand temperature, where the relative importance ofreaction 2 increases with increasing pressure. Thepressure is bounded by upper and lower limitsdefined by the explosion peninsula and decreaseswith increasing temperature above some relativelylow temperature on the order of 1500 to 2000°R,depending on residence time. These characteris-tics can be explained by the so-called explosionlimit boundary for the combustion of hydrogen. Theflammability boundary is represented in terms ofpressure vs. temperature and is similar in shape toa reversed “S”.30-32 The data on explosion limitsreported in Refs. 30 and 31 were obtained fromcombustion tests conducted at residence times upto several orders of magnitude longer than are typ-ically encountered in practical combustion systems.Therefore, for the “Explosion Limit” defined in thisway, time is essentially infinite, and the explosionlimit is considered independent of residence timeand to be a property of the combustible mixture.However, in practical propulsion system combus-tors, residence times are typically in the millisecondrange, and the lag, or delay, in ignition becomes animportant design consideration.
To quantify the design requirements to mitigatethe ignition delay problem it is useful to extend theconcept of the conventional explosion limit toinclude time as a parameter. Moreover, in testingwhere composition mismatches are present, it isalso necessary to include composition as a param-eter. Working charts can be constructed and usedto estimate the impact of vitiation on combustoroperability and performance, and, hence, design.This was done, to a limited extent, for a hydrogen-clean air system by Mitani et al.32 A one-dimensionalflow model incorporating a hydrogen-air kineticsmechanism was used to perform the calculations.The essential result showed that the explosionboundary curves, one for each selected residencetime, are similar in shape to the conventional explo-sion limit, but are shifted to higher pressures andtemperatures as residence time is decreased.
In Ref. 32, Mitani et al. presented test results ona hydrogen-fueled scramjet engine operating atnominal Mach 6 conditions, and, for comparativepurposes, used storage and combustion heating toobtain clean vs. vitiated air operability and perfor-
20American Institute of Aeronautics and Astronautics
mance data. A direct observation was that ignitionwas much easier to achieve using vitiated air and,hence, the formation of HO2 and its chemicalquenching effect was not a factor at their operatingconditions. The pressure was less than 1/3 atmwhere the pressure-sensitive three-body reactionforming HO2 is slow compared to the competitivebranching reaction. They also showed, by separateone-dimensional parametric analysis, that theamount of water present had very little effect at lowpressures. They concluded that the presence ofresidual free radicals from the combustion heatingwas the cause for the ignition augmentation. How-ever, once ignited, the clean storage heated flowburned more intensely, producing more thrust, wellbeyond the effect of reduced total captured massflow with vitiation. This may have been caused bythe mismatch in tunnel conditions between storageand combustion-heated flows where the freestreamtemperature was lower with combustion-heatedflow. In addition, the combustor inlet temperaturewith combustion heating would be further reducedrelative to the storage-heated flow, stretching outthe reaction time of the combustion-heated flow.This probably accounted for the “weak” combustionobserved with combustion-heating, where theflame was not anchored to the recirculation zone ofthe step flameholder, due, in part, to no burningthere, but rather was attached to the wall boundarylayer in the downstream portion of the combustor.Had the flame been attached to the intended flame-holder, the results may have been different, sincethe hot, reactive products of combustion entrainedfrom this recirculation zone region into the core flowcould dominate the incoming core flow contamina-tion levels.
The above example demonstrates the effect ofin-stream combustion heating in comparison to theclean air conditions of a pebble bed heater. Theresults of the tests and the demonstration of thefailure of the engine to flame hold when vitiatedwith in-stream combustion, provides an importantdemonstration of the effects of flow vitiation onengine performance. However, a complete data-base of the effects of flow vitiation, using purelyexperimental means, would require substantialadditional testing. For example, the fixed facilitynozzle area ratio used in these tests would pro-duce a slightly higher exit velocity for the in-streamcombustion test condition due to the lower gammainherent in a combustion-heated flow. These veloc-ity differences would be fed downstream throughthe inlet, resulting in a combustor entrance flow
condition that was different from the clean air case.Also, while pressure and temperature can be con-trolled by facility settings, the specific heat andmolecular weight and, hence, gamma and density,will be different. In addition, the Reynolds number-dependant phenomena of the inlet would beslightly different because of the differences in den-sity and velocity of the combustion-heated flow’sfurther altering the combustor entrance conditions.Capturing these effects in an experimental pro-gram would require fabrication of a second facilitynozzle with a slightly different area ratio and exitMach number in an attempt to account for theabove differences in inlet performance. Success ofthis approach, however, is unlikely because alter-ing the inlet entrance Mach number, compoundedby mismatches in gamma, will result in a differentwave structure in the inlet and an altered mass flowdistortion profile at the combustor entrance. Due tothe above effects, the role of analysis in addressingvitiation issues is critical in both the interpretationof ground test data and the scaling of ground testdata to flight conditions.
The above results of Mitani et. al. and those ofRef. 25 demonstrate the critical importance ofmatching the engine operating conditions, theselection of the particular properties to match, andthe presence of reactive species in the combustorinlet flow on the combustor performance leveldeduced from ground test data. These results alsoshow the importance of the engine design featuresand the associated fluid mechanics controlling flowdistortion, transport, and mixing that contribute toflameholding robustness and combustion effi-ciency.
More recently, and of more immediate interestto the current study, Kanda et al. published resultsof Mach 8 tests on the same engine used at Mach6.33 The test conditions were achieved throughstaged heating, where the first stage used storageheating followed by combustion heating. Thedynamic pressure at the entrance to the engineinternal inlet was about 1020 lb/ft2. The corre-sponding Mach number at the entrance plane wasM = 6.73, reduced from M = 8 to account for fore-body compression. Pt and Tt were 1500 psia and4680°R, respectively. Note that the low dynamicpressure corresponds to a trajectory point at Mach8 having a freestream dynamic pressure of about560 lb/ft2, above the upper altitude of a typical aero-space plane flight corridor. The resulting low enginepressure levels were compensated for by increas-
21American Institute of Aeronautics and Astronautics
ing the internal contraction ratio from a value of 5 toan atypical value of 8.33. This was accomplishedby using a very thick strut in the isolator section.However, this created an excessive base drag. Theessential results reported by the authors includedthe observation that good combustion efficiency(based on gas-sampling measurements), on theorder of 90 percent, was achieved. However, themeasured thrust was below the expected level, andthis result was attributed to the large strut basedrag. In addition, the data showed that the worstperformance occurred when the overall engineequivalence ratio was unity, leading to the conjec-ture that the flow was mixing, not kinetically con-trolled. Thus, it was concluded that the kineticquenching process involving HO2 was not signifi-cant under the conditions of this test. Unfortunately,no analysis was presented to corroborate this con-clusion. Further evaluation of the data in this papermay be warranted depending on a more extensivereview and determination of the availability of cer-tain flowpath details. Much useful additional infor-mation would be provided by pretest-type analysisand sensitivity study that included clean air vs. viti-ated air comparisons of expected engine operabilityand performance characteristics. Then, posttestanalysis should be done to compare with theexpected results. This would help to clarify theeffects of vitiation on combustion, nozzle perfor-mance, and the loss mechanisms in this particularengine and engine tests.
In addition to the effects of vitiation caused bythe usual combustion products involving CXHYOZ-type species, the presence of nitrogen oxides, ofparticular interest in arc-heated tunnels, is known toaffect combustion,18 and possibly the nozzlerecombination process. It should be noted that incombustion-heated tunnels encompassing Mach 5to 8 testing and in tunnels using staged storageplus combustion heating to achieve Mach 10+ test-ing, NOx levels relevant to this question may beproduced. In work by Slack and Grillo,25 an exten-sive shock tube study was carried out that includedthe effects of both NO and NO2 on the hydrogen-airignition process at stoichiometric conditions. Theoperating conditions for these tests were:
0.27 atm < P< 2 atm,
1440°R(800°K) < T< 2700°R(1500°K)
0 < [NO, NO2] < 4.5% mole fraction
The primary conclusion drawn from these testswas that ignition delay times were reduced with theaddition of either NO or NO2, with only minor differ-ences in the effect of adding NO or NO2, or mix-tures of these species. A maximum reduction inignition delay time of about an order of magnitudewas observed with a NO mole fraction of 0.5 to 0.7percent at a pressure of 2 atm and a temperature of1440°R (800K), (Fig. 23). Another important obser-vation was that the reductions in ignition delay timedecreased with decreasing pressure and increas-ing temperature. This is significant in terms of themechanism responsible for the effect of NO (andNO2) on the ignition process. The dominance of theformation of HO2 at the expense of the generationof highly reactive radicals via branching is the basisfor the reduction in ignition delay with the addition ofNO and NO2. The primary reactions believed to beresponsible for this sensitization of ignition are:
NO + HO2 = NO2 + OH (3)and,
NO2 + H = NO + OH (4)
P = 2.0 ATM
100 1 2 3
Mole Percent Nitric Oxide
102
103
104
T = 800˚K
T = 908˚K
T = 1000˚K
T = 1110˚K
Ind
uctio
n T
ime,
τ i
, µs
Fig. 23. Variation of Indction Time with Mole per-cent Nitric Oxide, at P = 2.02 × 105 N/m2(2.0 atm)
22American Institute of Aeronautics and Astronautics
Reaction 3 is relatively fast and produces anactive radical in the process of oxidizing NO to NO2.Reaction 4 produces additional NO and OH, aug-menting the sensitization process. Also, since OHis more effective than H in initiating and propagat-ing the hydrogen-air oxidation process, the neteffect of reaction 4 in removing H and producingOH is to accelerate ignition.
An important observation on the overall impactof added NO and NO2 on ignition is easily definedby noting that reaction 2 is a primary source of HO2and is a three-body reaction. Thus, as the pressureis lowered the production of HO2 decreases, andNO and NO2 lose their effectiveness in sensitizingthe ignition process via reactions 3 and 4. However,as pressure is increased, the production of HO2 willincrease up to some point, depending on tempera-ture and the concentrations of NO and NO2, whereother termination reactions and species will domi-nate and the ignition time will increase. This mech-anism was studied by Ashmore and Tyler.34 Thetermination reactions included in their mechanismwere:
NO + OH + M = HNO2 + M (5)and,
NO2 + OH + M =HNO3 + M (6)
Other reactions were also included, but reac-tions 5 and 6 are dominant in the Ashmore andTyler mechanism, helping to explain the dropoff inignition sensitization as the concentration of NO orNO2 increases. An important observation is that asthe static temperature exceeds a relatively narrowrange bounded by 1500 to 2000°R, the impedingeffect of HO2 and the effects of added NO and NO2on altering the ignition delay time rapidly drop off.This observation may be very significant in relationto hypersonic propulsion system operability andperformance where operating conditionsare in this range. Above about Mach 8,the static temperature entering the com-bustor will typically be on the order of2000°R and above. Therefore, minoreffects of vitiation on combustion kinetics,in general, and of NO and NO2 in particu-lar, might be expected above Mach 8.However, this depends on the details ofthe particular engine design where con-traction ratio is a critical design parametercontrolling combustor entrance tempera-ture and pressure, and the details of the
combustor/flameholding and nozzle designs deter-mine the importance of contamination on the oper-ability and performance of these components. Noz-zle recombination rates and the effects of contami-nation on this process must also be examined on acase-by-case basis over the entire flight range ofinterest. Thus, it is important to evaluate each situ-ation and to do so with the most comprehensiveanalyses that practical limitations allow. Additionalinformation on the effects of NO in scramjet testingis provided in Ref. 23.
The Konnov chemical kinetic mechanism forpropane-air combustion,35 was selected for dataanalysis and for the comparative analysis of flightvs. ground test of the notional vehicle propulsionsystems of interest to this work. The Konnov mech-anism was selected on the basis of the experiencegained from comparisons with other available kinet-ics models and comparing the relative performanceof these models against selected experimentaldata. Propane is considered sufficient for the cur-rent application because its combustion character-istics in terms of ignition delay and reaction timesare similar to many higher hydrocarbons of practi-cal interest. The mechanism contains 130 activespecies participating in 1200 reversible elementaryreactions. It embodies a comprehensive hydrogen-air mechanism and includes a wide spectrum ofnitrogen-bearing species including the nitrogenoxides and an extensive array of reactions that cap-tures the so-called “quenching” chain involving HO2and H2O2. The full set of species included in thismechanism is given in Ref. 35. Nitrogen speciesare listed in Table 14. The experimental data pro-vided by Slack and Grillo18 were analyzed to vali-date the performance of the present model in pre-dicting these data on the effects of NO and NO2 onignition delay time as a function of concentrationlevels, pressure, and temperature.
Table 14. Konnov Mechanism
Nitrogen Containing Species
N HNO H2CN NH N2H2
N2 HONO C2N2 NCO HCNN
NO HCNO N2H3 NH3 NH2
NO2 HOCN NNH CNN HONO
NO3 HNCO N2H4 HCN N2O4
N2O3 NCO N2O HNO3
29 Species
23American Institute of Aeronautics and Astronautics
Vibrational Relaxation — In nozzles fed by highpressure, very low-velocity flow typical of wind tun-nels for high-speed testing, the residence timebetween the heater and the nozzle throat is rela-tively large, enabling both the chemical states andvibrational energy distributions to maintain equilib-rium in this portion of the nozzle. Beyond the throatthe situation is different. Since the pressure andtemperature are dropping rapidly, vibrational non-equilibrium effects (as well as chemical nonequilib-rium effects) can become significant and should beconsidered in the evaluation of thermal nonequilib-rium effects in clean air vs. contaminated air inrelation to test section and test article flow propertymismatches.
The study in Ref. 25 included a literature reviewand the development and application of a one-dimensional analysis to estimate the degree ofvibrational nonequilibrium in nozzle expansions.The basis of the model is the classical harmonicoscillator, which is strictly applicable to simplediatomic molecules, precluding important speciessuch as H2O and CO2. However, available empiri-cal information in the form of correlations was doc-umented and used to estimate the relaxation timesfor these species. A feature of the previously citedone-dimensional model was the inclusion of theeffect of vibrational relaxation on the dissociationrates for the “wet” CO mechanism involving a sys-tem of nine reversible reactions for the speciesincluding: H2, O2, O, H, OH, CO, CO2, and H2O.This is a reduced system since it does not containthe NOx species that are important at high temper-atures (T > 2500°R) in general, and in arc-heatedair, in particular. Although future calculations usingthis formulation would need to be upgraded, theconclusions from the work done at that time are rel-evant.
It was found that water is very effective in equil-ibrating the vibrational and translational energy lev-els. For example, Fig. 24 shows the relaxationtimes for O2 and N2, with and without the effect ofwater vapor. The expansion process starts at thethroat of a nozzle at a total temperature of 3200Kand a total pressure of 2000 psia. This state wasachieved with clean preheated air and makeup oxy-gen entering the combustion heater at 4000°R, fol-lowed by combustion heating using 2 percent massfraction of hydrogen entering at 520°R. The upperfour curves show the mutual effectiveness of oxy-gen and nitrogen in equilibrating these species. Thelower two curves show the dramatic effect of water
in effecting a two-order-of-magnitude reduction inrelaxation time. These results indicate that in typicalcombusted heated nozzle flows, vibrational non-equilibrium is not an issue. In addition, the self-relaxation times of water and CO2 were found to besufficiently fast to enable them to maintain vibra-tional equilibrium in practical nozzle expansions.However, in the dry air produced by arc heating, theeffects of vibrational relaxation may be significant.
In upgrading the analysis given in Ref. 25 totreat this problem, the work of Reddy et al.36 wouldbe appropriate. A similar two-temperature sub-model including the coupled vibration-dissociation(CVD) feature was used for vibrational relaxationapplied to air dissociating at a finite rate in a two-dimensional flow field surrounding a hypersonicreentry vehicle. There remains the question of bilat-eral coupling involving the added effect of dissocia-tion on the vibrational relaxation process. ThisCVDV mechanism adds significant complexity tothe problem, but little seems to have been done toquantify the impact on flow properties, presumma-bly because of the complexity of this bilateral cou-pling. However, the following seven species wereincluded in Reddy’s Ref. 46: N2, O2, N, O, NO,NO+, and e-. The production or consumption ratesof the species were determined based on a six-reaction mechanism, where, as previously stated,the dissociation rate expressions incorporated theeffect of vibrational nonequilibrium. Therefore, themodifications of the model given in Ref. 25 would
Fig. 24. Relaxation Times in an Isentropic Nozzlewith an H2O Mole Fraction of ~0.25
24American Institute of Aeronautics and Astronautics
be relatively straightforward. Such an analysiswould probably be sufficient to estimate the vibra-tional relaxation effects in the ground test modes ofinterest here. The laser technology area37 isanother potential source of information on vibra-tional relaxation that should be reviewed for itsapplicability in future model upgrades.
Water Vapor Condensation — Water vapor con-densation in wind tunnel nozzles can affect testsection and test article flows in two distinct ways:
1. Water has a relatively high latent heat ofvaporization, which, when released throughcondensation in the supersonic carrier gas, canhave a significant effect on pressure as well astemperature and gas phase composition. Thus,the potential exists for water vapor condensationto create property mismatches in the test sectionand test article.
2. The condensed water will be in the form ofa cloud of very small droplets, typically in thesubmicron range, depending on the water vaporconcentration, the saturation ratio, and theresidence time available for growth of the nucleionce they are formed. The presence of dropletscan interfere with instrumentation and may affectthe compression process in inlet and integratedengine testing.
To prevent condensation in clean air wind tun-nels, the air is typically dehumidified to eliminatethe natural humidity normally present in ambientair. In a combustion-heated flow the combustion-generated water vapor cannot be removed, and it isnecessary to define the acceptable limits of opera-tion to avoid compromising the validity of the testdue to water vapor condensation. This requiresanalysis of the condensation process to help indefining the facility operating boundaries.
In the expanding nozzle flow, a condition can bereached where the saturation point will be reachedor exceeded. A conservative estimate for this con-dition can be made in a straightforward processusing the well-known vapor-pressure curve forwater. For example, if equilibrium is assumed, thenfor each set of selected operating conditions, allflow properties, including the water concentration,are determined through the nozzle. Then, the arearatio corresponding to the water saturation condi-tion can be found on the basis of the vapor pres-sure curve for water in the test medium and thepath of the expansion in thermodynamic space.
However, because the expansion is rapid, the con-densation process can lag the changes in tempera-ture and pressure. The condensation rate is com-prised of a nucleation rate process in which criticalsize droplets are continuously formed, followedimmediately by a continuous growth process. Animportant result of this rate dependence is thatsupersaturation can occur, and, if the flow is devoidof significant particulate matter that might act asforeign nucleation sites, from 70 to 90 deg ofsupersaturation is typically observed in small noz-zles with rapid expansion rates. This can be signifi-cant in extending the operating range of a windtunnel using combustion heating. To exploit thispractical implication, a finite-rate condensationanalysis is needed. It remains to be determined ifthe required cleanliness for supersaturation isachievable in a practical, large-scale facility. Refer-ence 25 details an analysis of this process and pro-vides a solution technique and some results up toMach 10. For the current range of interest, themodel should be updated and applied to Machnumbers beyond M = 10. The results would also beuseful, for example, in trade studies designed toevaluate the relative merits of envelope expansionby dry preheating in a two-stage heater system vs.supersaturation in a single-stage combustion-onlyheated system.
4.0 Conclusions and Recommendations
4.1 Conclusions
1. Our analyses in this study indicates that thepresence of combustion products in the freestreamflow would not seriously undermine the validity oftest results and conclusions, but would necessitatecareful posttest analysis.
2. The increased heat capacity of combustion-heated flows reduces the combustor peaktemperature and pressure, directly affecting thecombustion process and resulting in a decreasedcombustor wall heat-transfer rate, which is animportant consideration in the scaling ofcombustion-heated ground test data to theexpected heat loads under flight conditions.
3. The increased heat capacity of combustion-heated flow conditions and attendant reduction inpeak combustor pressure also impacts thecombustor operability limits deduced from groundtest, which is important in the lower scramjettransition regime.
25American Institute of Aeronautics and Astronautics
4. Test facilities that do not use combustionheating (such as arc facilities) also require carefulposttest analysis to obtain good results. Theanalysis can be complicated because of thepresence of charged and NOX species.
5. Scramjet ignition/combustion was not foundto be sensitive to instream combustion effects or tothe presence of NOx in the Mach 8 to 14 regime atthe dynamic pressures analyzed.
4.2 Test Facility Requirement Recommendations
1. Our analyses in this study indicate that thepresence of combustion products in the freestreamflow would not seriously undermine the validity oftest results and conclusions. Therefore, norecommendation is made on an “allowable” level offreestream combustion products based on ignitionor combustor performance considerations.However, water vapor condensation limits shouldbe investigated.
2. Our analyses in this study indicate that thepresence of NOX species did not cause significanttest or posttest analysis difficulties. Therefore, norecommendation is made on “allowable” levels ofNOX concentration in the Mach 8 to 14 range.
4.3 Analysis Recommendations
1. The flight regimes should be extended tolower Mach numbers and higher altitudes to moreclearly define the relative effects of test mediacontamination as a function of Mach number andaltitude. For lower dynamic pressure trajectories,the ignition and combustion sensitivities tocontamination may be greater, and additionalanalysis is recommended to investigate theseissues in both ramjet and scramjet regimes.Additional HyCAD analysis runs should beperformed simulating the lower dynamic pressure(~500 psf) trajectories. This information will aid intrajectory optimization studies and will help definetest conditions and support data interpretation inrelation to true flight operability and performance.
2. Other hydrocarbon fuels should beevaluated for combustion heating. Although manyhigher hydrocarbons have similar ignition andreaction time characteristics, there are somepotentially important differences in the details of thereaction mechanisms. This includes different,potentially important, low concentration inter-
mediates and reaction paths. These differencescould play a role in effecting different responses intest article flow and combustion processes. Thus,while propane is used as a surrogate for higherhydrocarbons, it is, for example, substantiallydifferent in detail from the butanes. Also, methane,a light hydrocarbon, is used as a heating fuel, and itis different from all other hydrocarbons in relation tothe oxidation mechanism. Thus, there may bedifferences of one hydrocarbon over another thathave yet to be determined.
3. The understanding of arc heatingprocesses should be improved. Arc heaters can beused up to Mach 14 to 15, but there are pressurelimitations that may prevent application at thepressures desired. A systematic analysis of thelocal electrical discharge process is needed. Thisprocess couples with the chemical, vibrational, andpossible ion non-equilibrium kinetics and fluiddynamic processes governing the production andmixing of nitrogen oxides and other potentiallyimportant contaminating species (includingcharged species in the heater). Subsequent effectson the nozzle expansion process should be anintegral part of the study.
4. Effects of contamination and propertymismatches on inlet flows should be evaluated. Forinlet testing, the compression process through theinlet requires adequate simulation of Mach number-thermodynamic property phenomena to track theoblique shock compression process. In addition,because of differences in the thermodynamicproperties of the simulated flow, the statictemperatures, flow speeds, and static pressuresthrough the inlet compression process are differentfrom that of the flight case. These differences resultin inlet exit flow properties that are different fromthose of the flight case, which are then feddownstream into the combustor. The mismatchesin temperature and velocity are further magnified bythe combustion process. These studies arerecommended to quantify the impact of thesedifferences created by vitiated flow on the inletcompression process, inlet performance, and theimpact of these differences on the combustionprocess and propulsion efficiency of integratedengines.
5. More comprehensive nozzle flow-fieldresults should be provided. More accurate thrustpredictions accounting for the wave structure andflow nonuniformities on nozzle recombination
26American Institute of Aeronautics and Astronautics
kinetics and nonequilibrium thermodynamic effectsare possible. This should be done for clean andcontaminated test media
6. A parametric study of water vaporcondensation should be conducted. This shouldinclude combustion heating with and without cleanair preheating. The limits of combustion heating arenot well established, and this study should includethe condensation boundaries imposed on the testenvelopes up to the Mach 14 upper flight speed.
7. A systematic study could be performed onthe selection of tunnel operating conditions todetermine if contamination effects can beminimized for specific test objectives. For example,inlet aerodynamics and combustor chemicalkinetics will each depend most heavily on differentflow properties indicating that it may be possible tooptimize the tunnel operating conditions for a "bestmatch" of flight conditions for a particular process.
8. It is imperative to continue to improve thecapability to model the flow physics of an advancedhypersonic test facility, including the facilitystagnation chamber, energy addition, facilitynozzle, and propulsion test flowpath. A quasi-one-dimensional facility model capability is adequate forinitial studies, but it needs to include models for anyenergy addition processes (electron beams, laser,etc.) and the impact on chemical species (includingcharged species) and reactions. This capability willallow an analysis of a complete “streamtube”through the facility and test hardware.
5.0 Acknowledgments
Several other Boeing personnel made signifi-cant contributions to the Phase I effort includingDaniel Farrell and Edward Eiswirth. Mel Bulman atAerojet supported the test requirements analysisfor one of the concepts. Others at Pratt & Whitney,Aerojet, Rocketdyne, and NASA Langley also con-tributed. For Phase II, Scott Halloran from BoeingRocketdyne made major contributions and manypeople at AEDC contributed significant commentsand additions to the original report.
6.0 References
1. Bulman, M. J., and Siebenhaar, A., “RocketBased Combined Cycle Propulsion for Space
Launch,” Paper No. IAF-95-S.5.02, 46th Interna-tional Astronautical Congress, Norway, 1995.
2. Siebenhaar, A., and Bonnar, D. K., “StrutjetEngine Paves Road To Low Cost Space Access,”Paper No. IAF-97-S.5.05, 48th International Astro-nautical Congress, Turin, Italy, 1997.
3. Siebenhaar, A., Bulman, M. J., and Bonnar,D. K., “The Role of the Strutjet Engine in New Glo-bal and Aerospace Markets,” Paper No. IAF-98-S.5.04, 49th International Astronautical Congress,Melbourne, Australia, 1998.
4. Bulman, M. J., and Neill, Todd, “Testing of theStrutjet RBCC Engine,” Invited paper, 35th AIAA/ASME/SAE/ASEE Joint Propulsion Conference,June 22, 1999.
5. Stemler, J. N., Bogar, T. J., Farrell, D. J., Bul-man, M. J., and Hunt, J. L., “Assessment of RBCC-Powered VTHL SSTO Vehicles,” AIAA-99-4947,AIAA 9th International Space Planes and Hyper-sonic Systems and Technologies Conference, Nor-folk, VA, November 1999.
6. Scuderi, L. F., Orton, G. F., and Hunt, J. L.,“Mach 10 Cruise/Space Access Vehicle Definition,”AIAA-98-1584, 8th International AIAA SpacePlanes and Hypersonic Systems and TechnologiesConference, April 27-30, 1998.
7. Orton, G. F., and Scuderi, L. F. “A HypersonicCruiser Concept for the 21st Century,” Paper985525, AIAA/SAE 1998 World Aviation Congress,September 28-30, 1998.
8. Scuderi, L. F., and Hunt, J. L., “Mach 10 Dual-Role Vehicle Concept,” Paper presented at theSpace Technology and Applications InternationalForum (STAIF-99), February 1-4, 1999.
9. Hunt, J. L., and Eiswirth, E. A., “NASA’s Dual-Fuel Airbreathing Hypersonic Vehicle Study,” AIAA-96-4591, AIAA 7th International Space Planes andHypersonic Systems and Technologies Confer-ence, Norfolk VA, November 1996.
10. Bogar, T. J., Eiswirth, E. A., Couch, L. M.,Hunt, J. L., and McClinton, C. R., “ConceptualDesign of a Mach 10 Global Reach Reconnais-sance Aircraft,” AIAA-96-2894, 32nd AIAA/ ASME/
27American Institute of Aeronautics and Astronautics
SAE/ASEE Joint Propulsion Conference, LakeBuena Vista, FL, July 1996.
11. Bonds, T. M., Espinosa, A. M., Hopping, B.M., and Scuderi, L. F., “Synthesis of ConfigurationComponents for Hypervelocity Airbreathing Vehi-cles,” Report No. WL-TR-93-3102, November1993.
12. Bogar, T. J., Alberico, J. F., Johnson, D. B.,Espinosa, A. M., and Lockwood, M. K., “Dual-FuelLifting Body Configuration Development,” AAIAA-96-4592, AIAA 7th International Space Planes andHypersonic Systems and Technologies Confer-ence, Norfolk VA, November 1996.
13. Weirich, T. L., Fogarty, W. C. Dry, K., Iqbal,A., and Moses, P. L., “Dual-Fuel Vehicle Airframeand Engine Structural Integration,” AIAA-96-4594,AIAA 7th International Space Planes and Hyper-sonic Systems and Technologies Conference, Nor-folk VA, November 1996.
14. Lau, K. Y., Scuderi, L. F., Petley, D. H., andYarrington, P. W., “Dual Fuel Thermal Manage-ment,” AIAA-96-4595, AIAA 7th International SpacePlanes and Hypersonic Systems and TechnologiesConference, Norfolk VA, November 1996.
15. Wilson, D., et al., “Analysis of a Pulsed Nor-mal Detonation Wave Engine Concept,” AIAA2001-1784, 10th International Space Planes andHypersonic Systems and Technologies Confer-ence, Kyoto, Japan, April 24-27, 2001.
16. Mager, A., “Simplified Analysis of Hyper-sonic Ramjet Performance,” Marquardt CorporationReport No. MR20, 054, July 1959.
17. Curran, E.T., and Bergsten, M. B., “Inlet Effi-ciency Parameters for Supersonic CombustionRamjet Engines,” TDR 64-61, Air Force Aeropro-pulsion Laboratory, Air Force System Command,Wright Patterson AFB, Ohio, June 1964.
18. Slack, M., and Grillo, A., “Investigation ofHydrogen-Air Ignition Sensitized by Nitric Oxideand Nitrogen Dioxide,” NASA CR-2896, October1977.
19. Burnette, T. D., Heins, A. E., Reed G. J.,”Dual Mode Scamjet Part II Combustor Design andPerformance Characteristics,” AFAPL-TR-67-132,Part II, The Marquardt Corporation, December 1967.
20. Northam, G. B., and Anderson, G. Y., ”Sur-vey of Supersonic Combustion Ramjet Research atLangely,” AIAA 86-0159, 24th AIAA Aerospace Sci-ences Meeting, Reno, NV, January 1986.
21. Powell, E. S., “Conceptual Design of a Mach7, True-Temperature Continuous-Flow Air Wind Tun-nel,” AIAA 2002-0444, 40th AIAA Aerospace Sci-ences Meeting and Exhibit, Reno, NV, January 2002.
22. Jue, Jan-Fong, and Virkar, Anil Vasudeo,“Method for Forming T-Phase Zirconia for High-Temperature Applications,” Patent No. US6,168,745 B1, January 2, 2001.
23. Fischer, K. E., and Rock, K. E., “CalculatedEffects of Nitric Oxide Flow Contamination ofScramjet Performance,” AIAA Paper 95-2524, 31stAIAA/ASME/SAE/ASEE Joint Propulsion Confer-ence and Exhibit, San Diego, CA, July 1995.
24. Limbaugh, C., and Pruitt, D.,”Calculations ofAir Chemistry in the Electron-Beam Heated Hyper-sonic Wind Tunnel. Initial Study,” AIAA 2000-2278,Denver, June 2000.
25. Edelman, R. B., Spadaccini, Louis J., andEconomos, C., “Analytical Effects of Vitiated AirContamination on Combustion and Hypersonic Air-breathing Ground Tests,” AEDC-TR-69-148, Octo-ber 1969.
26. Edelman, R. B., and Spadaccini, L. J., “The-oretical Effects of Vitiated Air Contamination onGround Testing Hypersonic Airbreathing Engines,”Journal of Spacecraft and Rockets, Vol. 6, No. 12,December 1969, pp. 1442-1447.
27. Chinitz, W., and Erdos, J. I, “Test FacilityChemistry Effects on Hydrocarbon Flames andDetonations,” AIAA-95-2467, 31st AIAA/ASME/SAE/ASEE Joint Propulsion Conference andExhibit, San Diego, CA, July 10-12, 1995.
28. Chinitz, W., and Erdos, J. I., “Test FacilityContaminant and Ozone Effects on HydrocarbonFlames and Nozzle Expansions,” AIAA 96-2917,32nd AIAA/ASME/SAE/ASEE Joint PropulsionConference, Lake Buena Vista, FL, July 1-3, 1996.
29. Billig, F.S., et al., “Proposed Supplement toPropulsion Management Support Plan,” JHU/APL-NASP-86-1, The John Hopkins University AppliedPhysics Laboratory, July 15, 1986.
28American Institute of Aeronautics and Astronautics
30. Lewis, B., and Von Elbe, G., Combustion,Flames and Explosions, Academic Press, NewYork, 1961, pp. 22-70.
31. Drell, I. L., and Belles, F. E., “Survey ofHydrogen Combustion Properties,” NACA RME57D24, July 26, 1957.
32. Mitani, T., et al., “Scramjet Engine Testing inMach 6 Vitiated Air,” AIAA-96-4555, 7th Interna-tional Space Planes and Hypersonic Systems andTechnologies Conference, Norfolk, VA, November18-22, 1996.
34. Ashmore, P. G., and Tyler, B. J., “TheNature and Cause of Ignition of Hydrogen and Oxy-gen Sensitized by Nitrogen Dioxide,” Ninth Sympo-sium (International) on Combustion, 1965, pp. 201-209.
35. Konnov, A. A., Detailed Reaction Mecha-nism for Small Hydrocarbons Combustion. Release0.4, http://homepages.vub.ac.be/~akonnov/, 1998.
36. Reddy, K., Fujiwara, T., and Murayama, T.,“Thermally and Chemically Nonequilibrium FlowAnalyzed by Park’s Two Temperature Model,” AIAA90-0142, 28th Aerospace Sciences Meeting, Reno,Nevada, January 8-11, 1990.
37. Anderson, J., Gasdynamic Lasers, An Intro-duction, Academic Press, New York, 1976.
29American Institute of Aeronautics and Astronautics
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