AFRL-PR-WP-TP-2006-271

SUPERSONIC COMBUSTION EXPERIMENTS WITH A CAVITY-BASED FUEL INJECTOR (POSTPRINT) Tarun Mathur, Mark Gruber, Kevin Jackson, Jeff Donbar, Wayne Donaldson, Thomas Jackson, and Fred Billig NOVEMBER 2001

Approved for public release; distribution is unlimited.

STINFO COPY

The U.S. Government is joint author of the work and has the right to use, modify, reproduce, release, perform, display, or disclose the work. PROPULSION DIRECTORATE AIR FORCE MATERIEL COMMAND AIR FORCE RESEARCH LABORATORY WRIGHT-PATTERSON AIR FORCE BASE, OH 45433-7251

i

REPORT DOCUMENTATION PAGE Form Approved OMB No. 0704-0188

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS.

1. REPORT DATE (DD-MM-YY) 2. REPORT TYPE 3. DATES COVERED (From - To)

November 2001 Journal Article Postprint 07/01/1998 – 06/30/1999 5a. CONTRACT NUMBER

In-house 5b. GRANT NUMBER

4. TITLE AND SUBTITLE

SUPERSONIC COMBUSTION EXPERIMENTS WITH A CAVITY-BASED FUEL INJECTOR (POSTPRINT)

5c. PROGRAM ELEMENT NUMBER 62203F

5d. PROJECT NUMBER

3012 5e. TASK NUMBER

AI

6. AUTHOR(S)

Tarun Mathur (Taitech, Inc.) Mark Gruber, Kevin Jackson, Jeff Donbar, Wayne Donaldson, and Thomas Jackson (AFRL/PRAS) Fred Billig (Pyrodyne, Inc.) 5f. WORK UNIT NUMBER

00 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION

REPORT NUMBER Taitech, Inc. Beavercreek, OH 45430

Propulsion Sciences Branch (AFRL/PRAS) Aerospace Propulsion Division Propulsion Directorate Air Force Research Laboratory Air Force Materiel Command Wright-Patterson AFB, OH 45433-7251

Pyrodyne, Inc. New Market, MD 21774

AFRL-PR-WP-TP-2006-271

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY ACRONYM(S)

AFRL-PR-WP Propulsion Directorate Air Force Research Laboratory Air Force Materiel Command Wright-Patterson AFB, OH 45433-7251

11. SPONSORING/MONITORING AGENCY REPORT NUMBER(S) AFRL-PR-WP-TP-2006-271

12. DISTRIBUTION/AVAILABILITY STATEMENT Approved for public release; distribution is unlimited.

13. SUPPLEMENTARY NOTES Journal article published in the Journal of Propulsion and Power, Vol. 17, No. 6, November–December 2001. The U.S. Government is joint author of the work and has the right to use, modify, reproduce, release, perform, display, or disclose the work. PAO Case Number: ASC/PA 99-0851; Date cleared: 19 Apr 1999. 14. ABSTRACT Recent results from combustion experiments in a direct-connect supersonic combustor are presented. Successful ignition and sustained combustion of gaseous ethylene have been achieved using an injector/flameholder concept with low-angle, flush-wall fuel injection upstream of a wall cavity. Two interchangeable facility nozzles (Mach 1.8 and 2.2) were used to obtain combustor inlet flow properties that simulate flight conditions between Mach 4 and 6 at a dynamic pressure of 47.9 kPa. Mainstream combustion was achieved at equivalence ratios between 0.25 and 0.75 using only a spark plug and no other external ignition aids. Delta-force levels between 667 and 1779 N were measured, with corresponding combustor pressure ratios between 3.1 and 4.0. Video records of the flame zone show an intensely active combustion zone with rapid flame spreading. One-dimensional performance analysis of the test data indicates a combustion efficiency around 80%with an average combustor skin friction coefficient of 0.0028.

15. SUBJECT TERMS Supersonic combustion, cavity-based flameholder

16. SECURITY CLASSIFICATION OF: 19a. NAME OF RESPONSIBLE PERSON (Monitor) a. REPORT Unclassified

b. ABSTRACT Unclassified

c. THIS PAGE Unclassified

17. LIMITATION OF ABSTRACT:

SAR

18. NUMBER OF PAGES

14 Mark R. Gruber 19b. TELEPHONE NUMBER (Include Area Code)

N/A

Standard Form 298 (Rev. 8-98) Prescribed by ANSI Std. Z39-18

JOURNAL OF PROPULSION AND POWER

Vol. 17, No. 6, November–December 2001

Supersonic Combustion Experimentswith a Cavity-Based Fuel Injector

Tarun Mathur¤

Taitech, Inc., Beavercreek, Ohio 45430Mark Gruber,† Kevin Jackson,‡ Jeff Donbar,§ Wayne Donaldson,§ and Thomas Jackson¤¤

U.S. Air Force Research Laboratory, Wright–Patterson Air Force Base, Ohio 45433and

Fred Billig††

Pyrodyne, Inc., New Market, Maryland 21774

Recent results from combustion experiments in a direct-connect supersonic combustor are presented. Successfulignition and sustained combustion of gaseous ethylene have been achieved using an injector/� ameholder conceptwith low-angle, � ush-wall fuel injection upstream of a wall cavity. Two interchangeable facility nozzles (Mach 1.8and 2.2) were used to obtain combustor inlet � ow properties that simulate � ight conditions between Mach 4 and 6at a dynamic pressure of 47.9 kPa. Mainstream combustion was achieved at equivalence ratios between 0.25 and0.75 using only a spark plug and no other external ignition aids. Delta-force levels between 667 and 1779 N weremeasured, with corresponding combustor pressure ratios between 3.1 and 4.0. Video records of the � ame zoneshow an intensely active combustion zone with rapid � ame spreading. One-dimensional performance analysis ofthe test data indicates a combustion ef� ciency around 80% with an average combustor skin friction coef� cient of0.0028.

NomenclatureC f = skin-friction coef� cientH = nozzle exit heightL=D = cavity length-to-depthratioM = Mach numberP = pressureQ = dynamic pressureT = temperaturet = timeu = velocityx = streamwise coordinatey = transverse coordinate1F = change in force´c = combustion ef� ciencyÁ = fuel–air equivalence ratio

Subscripts

cb = combustorcv = cavityo = � ight conditiont = stagnation conditionve = vitiator exit station4 = combustor inlet station

Received 14 June 2000; revision received 13 June 2001; accepted forpublication 24 June 2001. This material is declared a work of the U.S.Government and is not subject to copyright protection in the United States.

¤Research Scientist; currently, Innovative Scienti� c Solutions, Inc., 2766Indian Ripple Road, Dayton, OH 45440-3638.Member AIAA.

†Aerospace Engineer, AFRL/PRA, Building18, 1950 Fifth Street. SeniorMember AIAA.

‡Aerospace Engineer; currently, Jackson Design LLC, Beverly Hills, MI48025.

§Aerospace Engineer, AFRL/PRA, Building 18, 1950 Fifth Street. Mem-ber AIAA.

¤¤Deputy for Science, AFRL/PRA, Building 18, 1950 Fifth Street. Mem-ber AIAA.

††Consultant, 11280 Panorama Drive. Fellow AIAA.

Introduction

F UEL–AIR mixing, � ameholding, pressure losses, and thermalloadingare among the major issues that need to be resolved for

the successful design and implementation of hydrocarbon-fueledsupersoniccombustionramjet (scramjet) engines.A successfulfuel-injection scheme must provide rapid mixing between the fuel andoxidizer streams, minimum total-pressure losses, and have no ad-verse effects on � ameholding capability or thermal/structural in-tegrity of the device. These requirements place somewhat con� ict-ing constraints on the design of a viable fuel-injectionscheme, andsolutionsto theseproblemsare beingactivelysought internationally.A need exists for the development of a system that effectively inte-grates fuel injection and � ameholding for supersonic combustion.Such a device would contributesigni� cantly to the present researchand industrial technology base.

Three recent publications have presented relatively comprehen-sive literature surveys on the subject of cavity � ows and their rele-vanceto � ameholdingin supersoniccombustionengines.1¡3 Experi-mental and analytical research to date has predominantlyexaminedthe role of cavities in external� ows (as wheel wells or bomb bays onsupersonicaircraft),4¡24 althoughtherehavebeen studiesexaminingtheir � ameholdingcharacteristicsin low-speed25¡27 and high-speed� ows.1¡3;28¡38 Low-speedcombustionstudieswith an axisymmetriccavity26 found optimum � ameholding performance using a cavitywith its length-to-depth ratio L=D sized for the minimum aerody-namic drag. Longer cavities producedvortex shedding that resultedin unstable � ames, and shorter cavities did not provide enough airentrainment to hold the � ame. Experimental and numerical resultswere shown to agree closely on this point.27 A study by Yu et al.29

in an unheated Mach 2 � ow, with fuel injection upstream of a vari-able L=D cavity, suggested that small-aspect-ratiocavities providebetter � ameholding capability than longer cavities with inclined aftramp angles.

Combustion experimentswith a cavity-basedfuel injector/� ame-holder and gaseous hydrocarbon fuel are described in this paper.The baseline fuel injector/� ameholder has low-angled fuel injec-tion upstream of a wall cavity. This con� guration representsa novelconceptfor injecting fuel and pilotinga � ame in a scramjet combus-tor in that all of the componentsare containedin the wall. In contrastto in-stream concepts that introduce additional friction drag, wave

13051

1306 MATHUR ET AL.

drag, and cooling requirements to the combustor, this con� gurationuses no in-stream devices, thereby minimizing these detrimentaleffects and simplifying the overall combustor and system designs.Concurrent studies involving � ush-wall injection upstream of sim-ilar cavities in nonreacting supersonic � ow have provided valuableinsights into the effects of cavity con� guration (L=D ratio, offsetratio, aft ramp angle), fuel injection pressure, and imposed back-pressure on drag, residence times, and fuel distribution within thecavity.37;38 The combustion experiments described here, as well assome numerical simulations36;39;40 of the cavity-based fuel injec-tor/� ameholder, have shown robust � ameholding and combustionperformance in a scramjet combustor simulating Mach 4–6 � ightconditions at a dynamic pressure of 47.9 kPa.

Experimental FacilityThe test cell receives continuousair� ow of 13.6 kg/s at 5.17 MPa

and a maximum temperatureof 922 K with 20.7 kPa continuousex-haust from the Research Air Facility. Both liquid and gaseous fuelsystems are available, including pumped JP-4, pressurized JP-7,ethylene, and hydrogen. Liquid- and gaseous-oxygen systems areavailable for make-up oxygen in the vitiated heater. A recirculatedcooling-water system provides 158 l/s at 483 kPa; raw dump waterat 2.4 MPa is also available.The entire � owpath is securedto a thruststand for directmeasurementsof the thrust produced.This measure-ment may be combinedwith wall static pressure measurements anda performance analysis code to estimate the combustion ef� ciency.Additionally, the energy losses through the various water-cooledcomponents,coupled with temperature measurements from a steamcalorimeter, allow calculation of combustion ef� ciency.41

For the tests described here, Mach 1.8 and 2.2 facility nozzleswere used with on-design � ight Mach numbers near 3.5 and 4.5,respectively (see Table 1). However, due to the � nite number of fa-cility nozzles available, off-design tests were also performed withthese nozzles by varying the stagnationtemperature to achieve com-bustor inlet static temperatures (or velocities) corresponding to arange of off-design � ight Mach numbers between 4 and 6. The in-creased range of � ight Mach numbers simulated with these testsis associated with a corresponding compromise in combustor inletMach number and velocity (or static temperature).42

Hardware

Supply air enters a JP-4 fueled Marquardt Sudden Expansion vi-tiator capable of sustaining temperatures to 2500 K. The vitiator is� tted with a H2/air igniter system. A water-cooled instrumentation

Table 1 Simulation conditions

Mo Qo , kPa M4 T4, K P4 , kPa u4 , m/s

3.5 47.9 1.74 468 64.6 7533.5 95.8 1.74 468 132.4 7534.0 47.9 1.96 523 66.3 8924.0 95.8 1.96 519 126.8 8884.5 47.9 2.17 577 67.0 10364.5 95.8 2.17 568 136.8 10295.0 47.9 2.39 628 66.7 11855.0 95.8 2.39 616 134.9 11735.5 47.9 2.59 679 66.4 13355.5 95.8 2.59 666 134.3 13226.0 47.9 2.79 730 65.7 14876.0 95.8 2.79 715 133.4 1472

Fig. 1 Flowpath schematic.

section placed downstream of the vitiator permits stagnation tem-perature and/or pressuremeasurementswith traversingprobes,wallstatic pressure/gas sampling ports, and eight thermocoupleports inthe inlet and outlet � anges. A water-cooled transition � ange com-presses the vitiated air� ow from axisymmetric (254 mm i.d.) totwo-dimensional (57:2 £ 177:8 mm) as it enters the water-cooledfacility nozzle. Two removable water-cooled isolator sections arepositioned downstream of the nozzle. These sections are used tocontain the precombustionpressure rise. A variable-geometryheat-sink combustor follows the isolator sections. This component has a� exible upper wall permitting combustor exit-to-inletarea ratios ofup to four (con� gured with an area ratio of approximately 2.5 fortests described here). The combustor has removable inserts on allfour walls allowing optical access, installation of instrumentation,and a wide parametricdesignspace for fuel injectionand � amehold-ing concepts. Finally, a calorimeter instrumentation section hous-ing water sprays, rakes, and probes connects the combustor to asteamcalorimeter.An aerogrid(a metal platewith 1266converging–

diverging holes, providing approximately 85% area blockage) wasplaced in the calorimeter for some of the tests described here. Theaerogrid was designed to homogenize any nonuniformities in thecombustor exit � ow before the � ow encountered the thermocouplesat the calorimeter exit. The calorimeter exit connects to an elbowthrough which the � ow exits the test cell to an exhauster system.Figure 1 illustrates the internal � owpath of the rig from the entranceof the facility nozzle to the combustor exit.

The fuel injector/� ameholder used for fuel injection and com-bustion experiments in the combustor has four low-angle (15-deg)fuel injectors (fed from a common manifold) across the span of thecombustor, just upstream of a cavity-based� ameholder. The cavitygeometry can be varied by changing the L=D ratio, the distancedownstream from the fuel injectors, and the aft ramp angle. Forthe experiments described here, the cavity aft ramp angle and L=Dratio were approximately 22 deg and 4.8, respectively. Ports areavailable in the bottom wall for fueling, ignition sources (for exam-ple, spark plugs), pressure taps, thermocouples, and gas samplingmeasurements. The baseline cavity geometry used in the presentinvestigationis shown in Fig. 2. For experimentswith the Mach 1.8facility nozzle, the cavity was fueled from its bottom wall, through� ve spanwise ports located just upstream of the aft ramp; the cavitywas not independently fueled for tests with the Mach 2.2 facilitynozzle. Furthermore, to improve fuel penetration with the Mach2.2 facility nozzle, the area of the main injector holes was reducedby approximately 50% relative to the size used with the Mach 1.8facility nozzle.

Fig. 2 Fuel injector/� ameholder schematic.

2

MATHUR ET AL. 1307

Prior to combustiontests,bothnozzle– isolatorcombinationswerecalibrated to completely document the combustor inlet � ow� eld atvariousconditionssimulating� ight Mach numbers 4, 5, and 6 at dy-namic pressures of 47.9 and 95.8 kPa. The calibrationsconsisted oftraversingprobe surveys along the vertical centerlineof the aft iso-lator exit using four pitot pressureand two total temperatureprobes.Furthermore, at each condition, wall static-pressure pro� les wereacquired at several facility backpressures that were controlled by abutter�y valve. These conditions were also simulated numerically.Experimental and numerical data from these calibrations comparefavorably and have been published elsewhere.36;42¡44

Instrumentation

A CAMAC-based data-acquisitionand control system with 416channelsof analog input, 64 channelsof digital I/O, and 40 channelsof analog output was used. The CAMAC crates were connectedto two Sun 630MP workstations via a � ber optic SCSI interface(one each for control and data acquisition). A Pressure SystemsIncorporated (PSI) 8400 pressure scanning system consisting of400 channels with real-time display and data reduction was alsoused. The facility nozzle had 23 static pressure ports on its sidewall. Each isolator had 30 pressure taps on its top and bottom walls(total of 120 taps). The combustor top, bottom, and side walls hadmore than 1000 static pressure taps; nearly 200 of these taps wereinstrumented for these tests. A large array of thermocouples wasused for monitoring air, fuel, oxygen, cooling-water, and hardwaretemperatures. These measurements were also used to estimate theheat transfer from the various components. In addition, all � ows(air, fuels, water, and oxygen) were measured using either ori� ceplates or turbine � owmeters.

Remote monitoring of the tests was provided by six video cam-eras placed throughoutthe test cell. Four cameras placed around thethrust stand monitored the facility hardware. An infrared camerasystem provided real-time thermal health monitoring of the heat-sink combustor. Finally, a hand-heldSony 8-mm camera monitoredthe region near the fuel injector/� ameholder through a quartz win-dow. This camera captured visual records of the � ame front, � amespreading, and related combustion events.

ResultsCombustiontests usingthe cavity-basedfuel injector/� ameholder

were conducted using two separate facility nozzles (Mach 1.8 and2.2) to cover a rangeof combustorconditionssimulating� ightMachnumbers between 4 and 6 at a � ight dynamic pressure of 47.9 kPa.Gaseous ethylene was the combustor fuel for all tests. Wall staticpressure data were collected for all tests. Facility parameters suchas stagnation pressure and temperature, constituent � ow rates (JP-4, ethylene, air, and oxygen), wall temperatures, fuel pressures andtemperatures, and load-cell force measurements were recorded asfunctions of time. Video records of the combustion region near thecavity were obtained.Tables 2 and 3 show the operating conditions,peak pressure rise, and load cell data for tests with the Mach 1.8 and2.2 facility nozzles, respectively.

Table 2 Operating conditions and results (M4 = 1.8)

PeakTest name Tve,a K Pt , kPa Peak Á pressure ratio 1F , N

98205AC 1001 379.9 0.28 3.11 97498238AB 1000 379.2 0.45 3.30 115298238AC 1001 379.2 0.51 3.39 122898238AD 1001 379.9 0.54 3.44 132698238AE 1001 379.2 0.60 3.50 138398246AG 1000 378.5 0.62 3.57 147298260AF 1000 379.2 0.58 —— 137998260AG 1001 379.2 0.70 3.59 156698260AH 1001 380.6 0.75 3.64 1673

aTotal temperature at combustor inlet is lower than Tve due to heat losses throughwater-cooled sections.

Table 3 Operating conditions and results (M4 = 2.2)

PeakTest name Tve,a K Pt , kPa Á pressure ratio 1F , N

98323AE 1222 690.9 0.49 3.49 124598323AF 1222 691.5 0.55 4.02 132198323AH 1223 691.5 0.57 3.86 142398342ACb 1222 688.8 0.54 3.21 153098342AD 1056 692.9 0.50 3.64 169598342AE 1000 688.8 0.47 3.93 177998342AF 945 689.5 0.31 —— 135798342AH 1221 690.9 0.53 3.20 1517

aTotal temperature at combustor inlet is lower than Tve due to heat lossesthrough water-cooled sections.bAerogrid installed in calorimeter and ethylene preheated to 60±C for testseries 98342.

a) Temperature and pressure histories

b) Load cell force and fuel pressure histories

Fig. 3 Typical test sequence (test 98260AH).

Sample Test Sequence

Figure 3 provides a sample set of data obtained from test98260AH that illustrates the sequence of a typical test. Figure 3ashows the time history of three temperaturesand three pressures forthis test.The temperaturescorrespondto vitiatorexit stagnationtem-perature Tve, cavity ethylene temperature Tcv, and combustor ethy-lene temperature Tcb. The pressure data include facility stagnationpressure Pt , and ethylene supply pressures to the cavity (Pcv ) andcombustor (Pcb). Fuel temperatures and pressures were measuredinside the respective manifolds immediately upstream of injection.Figure 3b shows the measured load cell force and the fuel pressureson the same time scale. Vitiator ignition occurred around t D 5 s,corresponding to the sharp rise in vitiator exit temperature.A largeincrease in thrust, followed by a gradual decay to steady state, isassociatedwith vitiator ignition (due to the sudden increase in JP-4� ow into the vitiatormanifold). The steady-stateforcemeasurementhas been found to be repeatable to within a few pounds for theseconditions. Stable stagnation conditions occur after t » 50 s.

Once the desired combustor inlet conditions were established,fuel � ow was initiated to the cavity, and the cavity manifold pres-sure increased correspondingly. Shortly after a stable cavity fuelpressurewas reached (approximately t D 15 s), the sparkplugswere

3

1308 MATHUR ET AL.

energized, initiating combustion inside the cavity. Combustor mainfuel � ow was then started. Ignition of the main combustor fuel oc-curred almost immediately, as indicated by the sudden increase inthrust occurring at t D 68 s.

During mainstream combustion, the combustorpressure rise washigh enough to feed back into the cavity fuel manifold, resultingin the sudden increase in cavity fuel pressure. Because of the rela-tively high � ow rate and limited supply of ethylene, combustor fuelpressure and temperature both decayed over the course of the test(load cell force decay paralleled the fuel pressure decay), althoughcombustion of the main fuel remained very intense.Combustor fuel� ow was terminated at approximately t D 90–95 s, after which theload cell force level decreased back to the steady-state value thatexisted prior to mainstream combustion. The cavity manifold pres-sure also decreased as the mainstream combustionstopped. Finally,cavity fuel � ow was terminated around t D 105 s, and the vitiatorwas extinguishedat t D 110 s.

Similar test sequencing was used with the Mach 2.2 facility noz-zle, except,as notedearlier, the cavitywas not fueled independently.Instead, the main fuel was initiated at a very low � ow rate, such thatsuf� cient fuel was entrained into the cavity for ignition. Once thecavity was lit, the main fuel � ow rate was ramped up to the desiredlevel for mainstream combustion.

Wall Static Pressure Distributions

Figure 4 shows the wall static pressure distributions from test98260AH. This � gure shows the measured wall pressures from thefacility nozzle entrance through the isolators and combustor for atare condition (facility at test condition but no fuel � ow) and for themaximum fuel � ow condition for this test. An outline of the � ow-path is shown above the pressure traces to provide a spatial frame ofreference.The gaps in the pressure distributionsoccur at the physi-cal joints between the nozzle and fore isolator,between the fore andaft isolators, and between the aft isolator and the combustor. Minordisturbancesare apparent in the tare pressuredistribution;these cor-respond to interface shocks between the combustor and aft isolatorsections and the recompression shock generated at the aft ramp ofthe cavity � ameholder. Aside from these weak waves, the pressuredistributioncompareswell with resultsof an isentropicanalysis.42;43

Tare and reacting � ow pressure distributions were obtained for alltest conditionslisted in Table 2 with the exceptionof test 98260AF.In that test, the fuel supply was gradually reduced and no pressuredata are presented.The tare measurements for all cases were nearlyidentical.

Figure 5 contains the wall static pressure distributionsmeasuredfor several test cases. For clarity, each set of data includesonly mea-surements from the isolators and the combustor. Pressure data fromtest 98205ACshowthe presenceof main fuel combustionproducinga peak pressure ratio of approximately 3.1 near the cavity location.

Fig. 4 Normalized tare and reacting � ow pressure distributions (test98260AH, Á = 0.75).

Fig. 5 Isolator/combustor pressure distributions (M4 = 1.8).

For this case, the maximum fuel–air equivalenceratio is about 0.28,and the shock train generated as a result of the combustion-inducedpressure rise begins just upstream of the combustor entrance. Thiscase represents the limit of engine operation without appreciableisolator length requirement. As the fuel � ow is increased, the nor-malized pressures increase throughout the isolators and combustor.The shock system moves progressively forward until it is locatedjust downstreamof the facility nozzle exit in test 98260AH. At thiscondition, the maximum equivalence ratio is approximately 0.75,and the correspondingpeak pressure ratio is slightly over 3.6. Thisis representative of a normal shock pressure rise at the conditionsentering the combustor. Further increases in fuel � ow rate wouldcause the shock train to move into the facility nozzle section.

With the Mach 2.2 facility nozzle, tare and reacting � ow pressuredistributionswereobtainedat all testconditionslistedin Table3 withthe exceptionof test 98342AF. In each case, the tare measurementswere nearly identical, suggestingthat no appreciabledegradationorchangeoccurredin the hardware.Figure6a includesthe98323seriesof tests that were conducted without an aerogrid. Tests from series98342 were performed with the aerogrid installed in the calorime-ter, and are presented in Fig. 6b. The effect of the aerogrid on thecombustor exit pressure is clearly evident when the tare pressuredata between Figs. 6a and 6b are compared for x=H > 35. Over thisportion of the combustor,Fig. 6b shows a separatedregion resultingfrom the aerogridblockage.Upstreamof the x=H D 35 location, thetare pressure traces in Figs. 6a and 6b are nearly identical.The aero-grid has since been modi� ed to decrease the blockage and reducethe in� uence on the combustor exit � ow� eld.

Figure 6a shows an increase in peak pressure ratio with increas-ing equivalence ratio, and a corresponding upstream displacementof the start of pressure rise. With the highest fueling rate used in thepresent experiments, the start of pressure rise reaches the interfacebetween the two isolators. Experiments with higher fueling rateswere notpossibledue to excessiveheat loads that alreadylimited thetest runs to about 5 s after initiation of mainstream combustion.Theexpectedtrendof increasedpeakpressurewith lower facilitystagna-tion temperature (for approximately the same equivalence ratio) isobserved in Fig. 6b. Successful ignition and sustained combustionat the lower stagnation temperatures was achieved by preheatingethylene to 60±C prior to injection. With room-temperature ethy-lene (test series 98323), ignition and combustion did not occur atfacility stagnation temperatures below 1222 K.

Load Cell Force Measurements

Load cell force measurements for the Mach 1.8 facility nozzleexperiments appear in Fig. 7a as a function of equivalence ratio.The 1F values were obtained by subtracting the steady-state forcemeasured prior to combustionfrom the forcemeasured during com-bustion. A range of equivalenceratios existed for each test becausethe fuel pressure was not constant, as indicated by the legend in

4

MATHUR ET AL. 1309

a) No aerogrid, Tt = 1222 K

b) With aerogrid

Fig. 6 Isolator/combustor pressure distributions (M4 = 2.2).

a) M4 = 1.8

b) M4 = 2.2

Fig. 7 Variation of D F with equivalence ratio.

Fig. 7a. The data show a remarkably linear relationship over therange of equivalenceratios studied.At the higher equivalenceratios(Á > 0:6), the slope of 1F decreases (relative to the best-� t line),possibly due to decreasing combustion ef� ciency.

1F data with the Mach 2.2 facility nozzle are shown in Fig. 7b.For the three cases at 1222 K without an aerogrid, the relationshipis approximately linear, similar to that observed with the Mach 1.8facility nozzle. For the cases with an aerogrid, higher base pressureon the combustor is the most likely cause of the increase in the loadcell force measurement (see Fig. 6b). Pressure instrumentation hasbeen added to the base area of the combustor, and, as discussed ear-lier, the effects of the aerogridon thrust and pressure measurementsare currently being assessed. The increase in 1F with decreasingfacility temperatureis consistentwith the rise in peakpressuresseenin Fig. 6b. With the Mach 2.2 facility nozzle, the short run durations(unavoidable due to the high heat loads) and the highly unsteadynature of the combustion introduce a larger degree of uncertaintyinthe average 1F determined from experimental data.

Reaction Zone Images

Images of cavity combustion and mainstream combustion ob-tained from the video records of several tests with the Mach 1.8facility nozzle are presented in Fig. 8. Although these photographsare affected by the limitations inherent in video images (pixel sat-uration, time-averaging,and line-of-sight integration), they do pro-vide qualitativeinformationabout the overall� ow� eld features.The� ow direction is from left to right, and the window is approximately178 mm long. The leading edge of the cavity is visible in all images.From these photographs, it is evident that the cavity functions wellas a � ameholder/stabilizer. The � ame front observed in these pho-tographs appears anchored at the leading edge of the cavity. Someof the video records, such as Fig. 8d, suggest instability associatedwith the � ame front leading edge and occasional combustion for-ward of the cavity. Close examination of the video footage revealsthat the � ame excursions forward of the cavity are localized at thecombustor side walls and not in the core region.

A noteworthy observation from the video records is the rate atwhich the � ame appears to spread across the height of the combus-tor duct.Average� ame-spreadingangles between 24 and 30 deg aremeasured. Equivalence ratio appears to have a minor in� uence onthe � ame-spreading angle, although the highest equivalence ratiocase displays the largest spreading angle. Predictions of jet pene-tration from the low-angle injectors suggested that approximatelyhalf the duct height would contain fuel.36 Thus, it appears that someother mechanism, additional to the transverse momentum of thefuel, causes the rapid spreading of these � ames. The video recordssuggest that the blockage caused by the presence of the combus-tion zone turns the airstream upward and toward the top wall of thecombustor. This turning of the airstream would naturally transportfuel further from the wall than the jet transverse momentum alone,and could explain the observed � ame spreading.Recent OH planarlaser-induced � uorescence (PLIF) measurements in the near-� eldof the injectors show that the � ame-spreading rate is reduced com-pared with the video images.45 Although the PLIF images indicatethat combustion is most likely to occur along the combustor sidewall, static pressure measurements on the cavity � oor do not showsigni� cant spanwise pressure gradients. Although it is speculatedthat shock/boundary-layer interactions and wall corner effects areamong the factors in� uencing the combustion process, our under-standing of the combustion process and its underlying mechanismsis still incomplete and warrants further investigation.

Figure 9 presents similar images obtained from the video recordsof test 98323AF with the Mach 2.2 facility nozzle. These includea typical image of cavity combustion (low fueling rate through themain injectors), followed by images of main combustion.The lead-ing edge of the cavity is again clearly observed in Figs. 9a, 9b,and 9d. The � ame front observed in these photographs is normallyanchored at the leading edge of the cavity. The image-to-imagevariation between Figs. 9b, 9c, and 9d is indicative of instabilityassociated with the � ame front.

5

1310 MATHUR ET AL.

a) Typical image, cavity fuel only

b) Cavity and main fuel, Á = 0.42–0.45

c) Cavity and main fuel, Á = 0.44–0.51

d) Cavity and main fuel, Á = 0.49–0.54

e) Cavity and main fuel, Á = 0.64–0.70

f) Cavity and main fuel, Á = 0.65–0.75

Fig. 8 Side view images of � ame zones (M4 = 1.8).

a) Low Á, only cavity lit

b) Mainstream combustion, Á = 0.55

c) Mainstream combustion, Á = 0.55

d) Mainstream combustion, Á = 0.55

Fig. 9 Side view images of � ame zones (M4 = 2.2, case 98323AF).

6

MATHUR ET AL. 1311

Fig. 10 RJPA analysis (M4 = 1.8).

Performance Analysis

Combustor performance is assessed using the RJPA code.46;47

This tool employs a one-dimensional integral simulation forpredicting and assessing engine performance. Test data input intothe program include combustor geometry, � ow rates, temperatures,pressures, and heat � uxes. Combustion ef� ciency and skin-frictioncoef� cient are parametricallyvaried to generate a performancemapor carpet plot of the particular test (Fig. 10). The ordinate and ab-scissa represent directly observable quantities. From the measuredthrust and exit static pressure, deduced values of combustion ef-� ciency and skin-friction coef� cient are 80% and 0.0028, respec-tively, for test 98260AH with the Mach 1.8 facility nozzle.

Various uncertainties are still being addressed, including moreaccurate estimates of the transient energy loss into the heat-sinkcombustor and measurements of combustor base pressure. A sensi-tivity analysis shows that a 50% change in the combustor heat lossresults in only a 5% change in combustionef� ciency and no changein the skin-friction coef� cient. Further re� nements to the analysisand instrumentationcapabilities are currently underway. Addition-ally, future data reduction will also include steam calorimetry as asecond, independent assessment of combustion ef� ciency. Closurewith these two techniqueswill providecon� dence in both the facilityinstrumentationand the estimates for combustor performance.

SummaryA cavity-based � ameholder with low-angle � ush wall fuel in-

jection upstream of the cavity was successfully demonstrated ina model scramjet combustor using gaseous ethylene. Two facilitynozzles were used to simulate combustor inlet � ow properties ap-propriate for � ight Mach numbers between 4 and 6 at a dynamicpressure of 47.9 kPa.

1) Combustor operation was sensitive to fuel temperature. Theoperational envelope expanded to lower facility stagnation temper-atures with heated fuel.

2) Ignition and mainstream combustion were achieved at all testconditions reported in this paper using only conventional sparkplugs; no other external ignition aids were required.

3) Ignition and combustion produced a precombustion shocktrain, resulting in dual-mode combustor operation. The shock trainbecame stronger and the starting location of the shock train movedprogressivelyupstream with increasing fuel–air equivalence ratio.

4) Video records showed a very intense combustion zone withrapid � ame spreading. On further inspection, it is believed that theobservedrapid � ame spreading is con� ned near the sidewalls of thecombustor.

5) The cavity-based � ameholder concept proved very effectiveover a wide range of operating conditions and combustor fuel–airequivalence ratios. Over the range of equivalence ratios between0.25 and 0.75, combustorpressureratios and delta-forcelevels weremeasuredbetween3.1–4.0 and 667–1779N, respectively.The delta-force measurements from the Mach 1.8 facility nozzle tests variednearly linear with equivalence ratio. Preliminary analyses of theexperimental results from one test case using the RJPA code yielded

an estimateof combustionef� ciencynear80% with a correspondingskin-friction coef� cient of 0.0028.

AcknowledgmentsThe U.S. Air Force Hypersonic Technology Program, under the

guidance of R. Mercier, provided the funds that made this workpossible. The authors acknowledge the contributions of J. Bryant,M. Cox, W. Haendiges, C. Smith, and G. Streby for their techni-cal support. The support of the Research Air Facility at the U.S.Air Force Research Laboratory’s Propulsion Directorate is alsoappreciated.

References1Baurle, R. A., and Gruber, M. R., “Study of Recessed Cavity Flow� elds

for Supersonic Combustion Applications,”AIAA Paper 98-0938, Jan. 1998.2Ben-Yakar, A., and Hanson, R., “Cavity Flameholders for Ignition and

Flame Stabilization in Scramjets: Review and Experimental Study,” AIAAPaper 98-3122, July 1998.

3Davis, D. L., “Numerical Analysis of Two and Three Dimensional Re-cessed Flame Holders for Scramjet Applications,”Ph.D. Dissertation, Aero-nautics and Astronautics Dept., Air Force Inst. of Technology, Wright–Patterson AFB, OH, Sept. 1996.

4Baysal, O., and Stallings,R. L., “Computationaland Experimental Inves-tigation of Cavity Flow� elds,” AIAA Journal, Vol. 26, No. 1, 1988, pp. 6–7.

5Chokani, N., and Kim, I., “Suppression of Pressure Oscillations inan Open Cavity by Passive Pneumatic Control,” AIAA Paper 91-1729,June 1991.

6Clark, R. L., Kaufmann, L. G., and Maciulaitis, A., “Aeroacoustic Mea-surements for Mach 0.6 to 3.0 Flows Past Rectangular Cavities,” AIAAPaper 80-0036, Jan. 1980.

7Edwards, J. A., and Zhang, X., “Some Aspects of Supersonic Flow overa Cavity,” AIAA Paper 86-2025, Aug. 1986.

8Franke, M. E., and Carr, D. L., “Effect of Geometry on Open CavityFlow-Induced Pressure Oscillations,” AIAA Paper 75-492, March 1975.

9Hankey, W. L., and Shang, J. S., “Analyses of Pressure Oscillations inan Open Cavity,” AIAA Journal, Vol. 18, No. 8, 1980, pp. 892–898.

10Jeng, Y. N., and Payne, U. J., “Numerical Study of a Supersonic OpenCavity Flow and Pressure Oscillation Control,” Journal of Aircraft, Vol. 32,No. 2, 1995, pp. 363–369.

11Kim, I., and Chokani, N., “Navier–Stokes Study of Supersonic CavityFlow� eld with Passive Control,” Journal of Aircraft, Vol. 29, No. 2, 1992,pp. 217–223.

12Komerath, N. M., Ahuja, K. K., and Chambers, F. W., “Prediction andMeasurement of Flows over Cavities—A Survey,” AIAA Paper 87-0166,Jan. 1987.

13McGregor, O. W., and White, R. A., “Drag of Rectangular Cavities inSupersonic and Transonic Flow Including the Effects of Cavity Resonance,”AIAA Journal, Vol. 4, No. 11, 1970, pp. 1959–1964.

14Morgenstern, A., and Chokani, N., “Hypersonic Flow Past Open Cavi-ties,” AIAA Journal, Vol. 32, No. 12, 1994, pp. 2387–2393.

15Perng, S. W., and Dolling, D. S., “Passive Control of Pressure Oscilla-tions in Hypersonic Cavity Flow,” AIAA Paper 96-0444, Jan. 1996.

16Rizzetta, D. P., “Numerical SimulationofSupersonicFlowover a Three-Dimensional Cavity,” AIAA Journal, Vol. 26, No. 7, 1988, pp. 799–807.

17Rockwell, D., and Naudascher, E., “Review—Self-Sustaining Oscilla-tions of Flow Past Cavities,” Journal of Fluids Engineering, Vol. 100, No. 6,1978, pp. 152–165.

18Sarno, R. L., and Franke,M. E., “Suppressionof Flow-InducedPressureOscillations in Cavities,” JournalofAircraft, Vol. 31, No. 1, 1994,pp. 90–96.

19Vakili, A. D., Wolfe, R., Nagle, T., and Lambert, E., “Active Con-trol of Cavity Aeroacoustics in High Speed Flows,” AIAA Paper 95-0678,Jan. 1995.

20Zhang, X., and Edwards, J. A., “Computational Analysis of UnsteadySupersonicCavity Flows Driven by Thick Shear Layers,” AeronauticalJour-nal, Vol. 92, No. 919, 1988, pp. 365–374.

21Zhang, X., and Edwards, J. A., “An Investigation of Supersonic Oscil-latory Cavity Flows Driven by Thick Shear Layers,” Aeronautical Journal,Vol. 94, No. 940, 1990, pp. 355–364.

22Zhang, X., and Edwards, J. A., “Experimental Investigation of Super-sonic Flow over Two Cavities in Tandem,” AIAA Journal, Vol. 30, No. 5,1992, pp. 1182–1190.

23Zhang, X., and Edwards, J. A., “Computational Analysis of SupersonicJet and Instability Wave Interaction,” AIAA Paper 94-2194, June 1994.

24Zhang, X., “Compressible Cavity Flow Oscillation due to Shear LayerInstabilities and Pressure Feedback,” AIAA Journal, Vol. 33, No. 8, 1995,pp. 1404–1411.

25Huellmantel, L. W., Ziemer, R. W., and Cambel, A. B., “Stabilization of

7

1312 MATHUR ET AL.

Premixed Propane–Air Flames in Recessed Ducts,” ARS Journal, Jan. 1957,pp. 31–43.

26Hsu, K.-Y., Goss, L. P., and Roquemore, W. M., “Characteristics ofa Trapped-Vortex Combustor,” Journal of Propulsion and Power, Vol. 14,No. 1, 1998, pp. 57–65.

27Katta, V. R., and Roquemore, W. M., “Study on Trapped-VortexCombustor—Effect of Injection on Flow Dynamics,” Journal of Propulsionand Power, Vol. 14, No. 3, 1998, pp. 273–281.

28Niioka, T., Terada, K., Kobayashi, H., and Hasegawa, S., “Flame Sta-bilization Characteristics of Strut Divided into Two Parts in Supersonic Air-� ow,” Journal of Propulsion and Power, Vol. 11, No. 1, 1995, pp. 112–116.

29Yu, K., Wilson, K. J., Smith, R. A., and Schadow, K. C., “ExperimentalInvestigationon Dual-PurposeCavity in SupersonicReacting Flows,” AIAAPaper 98-0723, Jan. 1998.

30Vinogradov, V., Grachev, V., Petrov, M., and Sheechman, J., “Exper-imental Investigation of 2-D Dual Mode Scramjet with Hydrogen Fuel atMach 4–6,” AIAA Paper 90-5269, Oct. 1990.

31Vinogradov,V., Kobigsky, S., and Petrov, M., “Experimental Investiga-tion of Liquid Carbonhydrogen Fuel Combustion in Channel at SupersonicVelocities,” AIAA Paper 92-3426, July 1992.

32Vinogradov, V. A., Kobigsky, S. A., and Petrov, M. D., “ExperimentalInvestigation of Kerosene Fuel Combustion in Supersonic Flow,” Journal ofPropulsion and Power, Vol. 11, No. 1, 1995, pp. 130–134.

33Romankov, O. N., and Starostin, F. I., “Design and Investigation ofthe Stand and Flying Scramjet Models Conceptions and Results of Experi-ments,” AIAA Paper 93-2447, June 1993.

34McClinton, C., Roudakov,A., Semenov, V., and Kopehenov,V., “Com-parative Flow Path Analysis and Design Assessment of an AxisymmetricHydrogen Fueled Scramjet Flight-Test Engine at a Mach Number of 6.5,”AIAA Paper 96-4571, Nov. 1996.

35Mathur, T., Streby, G., Gruber, M., Jackson, K., Donbar, J., Donaldson,W., Jackson, T., Smith, C., and Billig, F., “Supersonic Combustion Experi-ments with a Cavity-Based Fuel Injector,” AIAA Paper 99-2102, June 1999.

36Baurle, R. A., Mathur,T.,Gruber, M. R., and Jackson,K. R., “Numericaland Experimental Investigation of a Scramjet Combustor for HypersonicMissile Applications,” AIAA Paper 98-3121, July 1998.

37Hsu, K.-Y., Carter, C., Crafton, J., Gruber, M., Donbar, J., Mathur, T.,Schommer, D., and Terry, W., “Fuel Distributionabout a Cavity Flameholderin Supersonic Flow,” AIAA Paper 2000-3583, July 2000.

38Gruber, M. R., Baurle, R.A., Mathur, T., and Hsu, K.-Y., “FundamentalStudiesofCavity-Based FlameholderConceptsforSupersonicCombustors,”Journal of Propulsion and Power, Vol. 17, No. 1, 2001, pp. 146–153.

39Eklund, D. R., and Gruber, M. R., “Study of a Supersonic CombustorEmploying an Aerodynamic Ramp Pilot Injector,” AIAA Paper 99-2249,June 1999.

40Eklund, D. R., Baurle, R. A., and Gruber, M. R., “Numerical Study ofa Scramjet Combustor Fueled by an Aerodynamic Ramp Injector in Dual-Mode Combustion,” AIAA Paper 2001-0379, Jan. 2001.

41Thompson, M. W., “Equilibrium Chemistry Steam Calorimetry DataAnalysis Procedure Using the Ramjet Performance Analysis Code,” Rept.RTDC-TPS-028, Johns Hopkins Univ./Applied Physics Lab., Baltimore,MD, 1997.

42Gruber, M., Donbar, J., Jackson, K., Mathur, T., Baurle, R., Eklund,D.,Cox-Stouffer, S., and Smith, C., “A Newly Developed Direct-Connect High-Enthalpy Supersonic Combustion Research Facility,” Journal of Propulsionand Power, Vol. 17, No. 6, 2001, pp. 1296–1304.

43Jackson, K., Gruber, M., Mathur,T., Streby,G., Smith, C., and Billig, F.,“Calibration of a Newly Developed Direct-Connect High-Enthalpy Super-sonic Combustion Research Facility,” AIAA Paper 98-1510, April 1998.

44Gruber, M., Jackson, K., Mathur,T., and Billig, F., “Experiments with aCavity-Based Fuel Injector for Scramjet Applications,” International Sym-posium on Air Breathing Engines, ISABE Paper IS-7154, Sept. 1999.

45Donbar, J. M., Gruber, M. R., Jackson, T. A., Carter, C. D., andMathur,T., “OH PLIF Imaging in a Hydrocarbon-FueledScramjet Combus-tor,” Proceedings of the Combustion Institute, Vol. 28, 2000, pp. 679–687.

46Pandol� ni, P. P., Billig, F. S., Corpening, G. P., Corda, S., andFriedman, M. A., “Analyzing Hypersonic Engines Using the Ramjet Perfor-mance Analysis Code,” APL Technical Review, Vol. 2, No. 1, 1990, pp. 68–

79.47Pandol� ni, P. P., and Friedman, M. A., “Instructions for Using Ramjet

Performance Analysis (RJPA)—IBM-PC Version 1.24,”Rept. AL-92-P175,Johns Hopkins Univ./Applied Physics Lab., Baltimore, MD, 1992.

8