Research on PDE Air Induction System Analysis

8/4/2019 Research on PDE Air Induction System Analysis

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PULSE DETONATION ENGINE AIR INDUCTION SYSTEM ANALYSISRobert J. Pegg*NASA Langley R esearch CenterHampton, VAB. D. Couch,t and L. G. Hunter^Lockheed Martin Tactical Aircraft SystemsFort Worth, TX

AbstractA preliminary mixed-compression inlet designconcept for potential pulse-detonation engine (PDE)powered supersonic aircraft was defined and analyzed.The objectives of this research were to conceptuallydesign and integrate an inlet/PDE propulsion system

into a supersonic aircraft, perform time-dependent CFDanalysis of the inlet flowfield, and to estimate theinstalled PD E cycle performance.The study w as baselined to a N A S A M a c h 5Waverider study vehicle in w h i c h th e baselineover/under turboramjet engines were replaced with asingle flowpath PDE propulsion system. As muchcommonality as possible w as maintained with th ebaseline configuration, including the engine locationan d forebody lines. Modifications were made to theinlet system's external ramp angles and a rotating cowll ip was incorporated to improve off-design inlet

operability and perform ance. Engin es were sized tomatch the baseline vehicle study's ascent trajectorythrust requirement at Mach 1.2. The majority of thisstudy was focused on a flight Mach number of 3.0.Two generalized concerns were addressed about theoperability of non-steady engines operating in theenvironment of a steady flow inlet. These concernswere (1) as the valving in the PDE's close in guillotinefashion, will i t cause a hammershock, which willunstart the inlet and (2) will a given arrangement ofPDE's allow the establishment of stable shock systemterminating with a normal shock in the inlet throat. Anon-steady CFD analysis was initiated to answer thesetw o concerns.

Asst. Manager, Systems Analysis Office, HVO.' Engineering Specialist, Sr.$ Engineering Specialist, Sr.Copyright by the American Institute of Aeronautics and Astronautics, Inc.No copyright is asserted in the United States under Title 17, U. S. Code. TheU. S. Government has a royalty-free license to exercise all rights under thecopyright claimed herein for Government Purposes. All other rights arereserved by the copyright owner.

Th e time-dependent Navier Stokes CFD analysesof a two-dimensional approximation of the inlet wasconducted for the Mach 3.0 condition. The LM TAS(Lockheed Martin Tactical Aircraft Systems)-developedFALCON C FD code with a two-equa t ion 'k-Pturbulence model was used. Th e downstream PDE wassimulated by an array of four sonic nozzles (valves) inwhich the flow areas were rapidly varied in variousopening/closing combinations, similar to the operationof multi-duct PDE rotary valves. Results of the CFDstudy indicated that the inlet design concept operatedsuccessfully at the Mach 3.0 condition, satisfying masscapture, total pressure recovery, an d operabili tyrequirements. Approx imately 5.0% inlet bleed wasrequired to stabilize the terminal shock train in theisolator region. Time-dep endent ana lysis indicated thatpressure an d expansion waves from the simulated valveperturbations did not effect th e inlet's operability orperformance.Installed PDE cycle performance analyses wereconducted using a LMTAS-developed PDE cycleanalysis code at Mach numbers of 1.2 and 3.0.

Realistic component performance values were used inth e analyses, an d ethylene (C 2H 4) was used as thebaseline fuel. Calculated effective Isp values werehigher for the W averider design powered w ith PDE'sthan w ith the baseline propulsion system.INTRODUCTION

Th e pulse detonation engine (PDE) is being evaluatedand developed as a potentially high-payoff newaeronautical propulsion system. The PDE represents apotential propulsion technology leap beyond the gasturbine engine. Based on the results of several studiesto date (Refs. 1-4), the airbreathing PDE offerspotential performance an d life cycle cost payoffs fo rboth subsonic and supersonic vehicle applications.Pulse detonation rockets and hybrid airbreathing/rocketPDE systems are also recognized to offer large potentialbenef i ts , especially for missile applications. Themaxim um operating co ndition for the pure airbreathing

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PDE is projected to be in the Mach 3 range, dependingon fuel autoignit ion cons tra ints . Hy br id pulsedetonation systems may have higher Mach numberapplication. Potentia l applications of interest arepropulsion systems fo r tactical aircraft (manned orunmanned) and missiles and subsonic/supersonicpropulsion sources for future hypersonic aircraft.

PDE development is still in a ve ry ear lydevelopment stage with many potential developmentrisks to be faced before operable engines will be areality. As with any high risk, high potential payofftechnology, critical enabling technologies should beidentified an d addressed early to reduce th e risk ofencountering show-stopper issues and to accelerate thedevelopment of an optimum concept. One of thecritical, enabling technologies is the integration ofsupersonic air induction systems with the intermittentflow PD E cycle. It has been estimated that in order forth e PD E cycle to be competitive with conventionalturbojet/turboramjet systems, they will be required toopera te in the 75 to 100 Hz range wi th nearstoichiometric fuel/air mixtures. This represents a cycletime of approximately 10 msec, requiring a propellantrefill time in the 5 msec, range. Developing com patibleai r induct ion systems that wi l l satisfy th e aboverequirements, as well as provide adequate sealing fromthe high pressure, high temperature exhaust products,represents a major technology challenge.

The objective of this study project was to init iatethe development of PDE/vehicle integration expertise,especially as related to air induction systems, throughthe conceptual design, analysis, and integration of asupersonic mixed-compression inlet system that iscompatible with the PDE cycle. The approach was asfollows:1. Select or define a notational study vehicle conceptt ha t wi l l permit replacement o f conventionalpropulsion systems with PDE systems.2. Define the vehicle flight requirements and determinepropulsive thrust requirements.3. Define a PDE modular design concept that iscompatible with the selected notional vehicle andwhich can be reasonably sized to satisfy thrustrequirements.

4. Define and size an inlet system design concept.5. Integrate the inlet system and PDE modules into thenotional vehicle.6 . Perform steady-state an d t ime-dependent fullNavier-Stokes CF D analysis of the inlet at a selectedsupersonic flight condition for evaluation of inletperformance, inlet operational com patibility with theintermittent flow PDE, an d assessment of bleedrequirements for terminal shock stabilization.7. Compute PDE installed performance.

PD E DESIGN AND VEHICLE INTEGRATIONNotional Vehicle Description

Th e first task of this program was to establish anotional vehicle definition to serve as a baseline for thisstudy. A N A S A Waverider study vehicle designconcept (Ref. 5) was selected. This 500,000-lb. class(TOGW) vehicle conceptual design, illustrated inFig. 1, designed for Mach 5 cruise, w as powered byfour Pratt & W hitney over/under turboramjet engines,which were integrated with the vehicle's lower surface,and fueled with an endothermic hydrocarbon fuel. Th edecision was made to replace the turboramjet engineswith PDE's, thus reducing the dual flow paths with asingle f lowpath and limit the PDE application toMoo = 3.0.Engine Sizing

PD E sizing requirements were based on ascenttrajectory aerodynamic drag and thrust histories as pre-sented in Ref . 5. (No additional vehicle performanceanalysis was conducted during this study.) The vehicledrag and thrust histories from the above reference areshown in Fig. 2. The engines were sized for a Mach1.2/25,000-ft. flight condition. As indicated in theabove figure, at M = 1.2, the total installed turbojetengine net thrust for the vehicle was approximately90,000 Ibs. with a Thrust/Drag ratio of approximately1.2. Also, th e ascent t rajectory installed thrust atM = 3.0 (end of turbojet/ramjet transition region) w asapproximately 120,000 Ibs. with a Thrust/Drag ratio ofapproximately 1.7. For this study, it was assumed them a x i m u m f l ight Mach n umber f or PD E operationwould be 3.0. Stoichiometric fuel/air ratio (equivalence

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ratio (0) = 1.0), near-instantaneous Chapman-Jouguet(C-J)-quality detonation w a v e generation, andcomponent efficiencies, i.e., inlet ram recovery, T|j.,combustion efficiency, T I C , an d nozzle gross thrustcoefficient (CFG) values were estimated at each flightcondition analyzed in the engine sizing study.PDE Module Design Concepts

Several generalized PDE designs have beenpatented, two of which are shown in Fig. 3 (Refs. 6 and7) . These designs feature rotary valves for both frontand side filling, spark and laser ignition, stratified gasignition (fuel/oxygen mixture in the imm ediate ignitionpoint vicinity for quick detonation ignition and fuel-airmixture for the primary detonation wave to burn into)pr e - m ixe d and pos t -mixed reac tants , annulus ,cylindrical, an d multi cylindrical designs, lubricationand cooling systems, purge air, etc.This paper assumes a PDE detonation volume an dvalve areas large enough to support detonationfrequencies up to 110 Hz. LMT AS studies have shownthat side valve designs provide enough fill area tosupport frequencies in the 100 Hz range, where the fill

area throats are choked. The results of this study arebased on a side-loading engine concept. Individualengines are clustered together in mod ules. For thisstudy three modules, each containing eight engines,were successfully integrated into the propulsion systemenvelope of the Waverider. This is shown in Fig. 4.PDE/Vehicle Integration

The intermittent combustion process is the mostunique feature of the PDE that effects integration. Thisintermittent process is composed of essentially threephases, i.e., (1) a filling phase in which fuel an doxidizer are injected into the duct and are mixed (if notpre-mixed), (2) the detonation phase in which th epropellants are ignited and a resulting high pressuredetonation wave formed which traverses the length ofthe duct and exits through the exhaust system, and (3) ablowdown phase in which the products of combustionar e pumped out the duct 's exhaust system. Positiveaxial thrust is produced in phases 2 and 3. In order toreduce the impact of intermittent combustion on the airinduction system, it is necessary for the PDE module tobe made of a group, or cluster, of pulse detonation ductsthat operate out of phase such that the airflow rate inthe PDE module 's common inlet duct is relativelyconstant.

Inlet Duct DesignTh e turboramjet air induction system for the studyvehicle is a two-dimensional, mixed compression inlet

design with th e vehicle forebody provid ingprecompression and a boundary-layer diverter providedat the beginning of the first inlet ramp. Inlet strokes(internal splitters) are provided to isolate the fourturboramjet flowpaths in case of an inlet unstart, andvariable geometry is provided through variable rampson the body side.

A major challenge of this study was to design arelatively short isolator an d diffuser to satisfy th eselected "most challenging" PD E modular integration,i.e., the two wide by four high engine module. Thisrequires a two-dimensional duct height change from the12.5-inch-high isolator to the approximately 75-inch-high engine face in a relatively shor t distance(approximately 180 inches). In order to allow as muchlength as possible for the diffuser, a decision was madeto retain the b aseline vehicle's forebody boundary-layerdiverter and to use an aggressively short isolator length,i.e., a length-to-duct height ratio of only 4.0 and todepend on throat bleed to stabilize the terminal shocksystem (shock trap) in the isolator. This leaves13 0 inches for the approximately six-to-one area ratiodiffuser. To obtain this aggressive diffusion rate with areasonable area distribution and acceptable local flowangles requires a diffuser center body and a plenumsection just forward of the engine face. Although thecenter body adds weight to the inlet system, its volumeca n be used advantageously for 02 storage and/or PDEvalve drive or ignition system hardware. The areadistribution is shown in Fig. 5. Note that a rapid localarea increase (plenum) is provided aft of the centerbody, after the inlet airflow ha s been diffused to anappropriately low Mach number ( 0.2). The plenum,in addition to further reducing the airflow Mach numberat th e engine face, prevents pressure pulses andexpansion waves caused by individual PDE va lveactivity from feeding upstream and disturbing theterminal shock system. Also note that the areadistribution from th e isolator exit to the maximumheight of the center body is near-identical to the areadistribution of a conventional equivalent 5.5 (halfangle) conical diffuser and that no severe local flowangles are encountered prior to the plenum section. Aschematic of the inlet is shown in Fig. 6.

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INLET TIME-DEPENDENT CFD ANALYSISFull Navier-Stokes analysis was used to evaluate a2-D approximation of the inlet isolator/diffuser designat the Mach 3.0/50,000-ft. flight condition. The CFD

code used was a LMTAS-developed 2-D, first-order,time-accurate central difference finite-volume code(C D FALCON) (Ref. 8) using th e LMTAS-developed'k-F two-equation turbulence model (Ref. 9). Theturbulence model is structured with wall functions ,which allows y+ values to be greater than unity at thewall. The grid was generated using GRIDGEN 10.Both steady state and time-accurate unsteady analyseswere conducted to evaluate the effects of the PDEmulti-valve flow area var ia t ions on the operationdynamics of the inlet airflow characteristics. The time-dependent solutions were run at physical time steps of0.24 microseconds pe r iteration. If strong pressureperturbations, such as hammershocks develop as thePDE valves close, a fine t ime resolution is required totrace th e development and propagation of the wave .One of the principle investigations of this study was todetermine if a hammershock of sufficient strength isformed, which will propagate upstream to the terminalnormal shock in the constant area isolator, and unstartthe inlet.Time-dependent CFD analysis was performed on theinternal inlet system from the cowl lip plane to theengine face. The starting condition included pre-compression from th e forebody an d external ramps. Itwas assumed that the boundary-layer diverter of thebaseline vehicle removed all of the forebody boundarylayer. The boundary layer developed on the externalinlet ramps was assumed to be bled off at the cowl lipplace (approximately 1.5% of inlet captured mass flowestimated). Th e engine face consists of four simulatedrotary valves, which open and close to intermittentlyingest air for the pulse detonation cycle. Typical cyclet imes are approximately 10 to 16 msec. The fill timerepresents roughly 4 to 5 msec, (nominally 1/3 of thePDE cycle time). These valves were represented asnozzles with choked throats whose area controls theamounts of flow going through the nozzle. When theflow is sonic at the throat (choked flow), the mass flowcan be reduced when the throat area is reduced. Whenflow transfers between PDE valves (when one isopening, the other is closing) are considered in terms oftens of microseconds, the static an d total pressure doesnot have time to vary significantly. Therefore, thepressure is assumed to remain essentially constantduring th e flow transfer between valves. Thus, byvarying the area of the valves, flow through the valveswill be con trolled in the CFD analysis. Since the valves

will be run out of phase, a natural sequence of eventswil l supply the prescr ibed opening and c los ingschedules.Figure 7 shows the detailed grid used in the

analysis, consisting of 236 longitudinal points an d105 lateral points. A center body is included in thehighly divergent duct, as well as four valve throats atthe end of the duct which represent a bank of engines.The center body does not extend all the way back to theengine face, where a large gap (plenum) is provided totransfer flow when valves are shut down an d openedup. Flow transfer studies included flow betweenadjacent valves as well as some cases where the flowtransfer length was up to two engine diameters away.Thus, a sizable plenum gap was used to ensure adequatetransfer time for the wo rst case situation.The valve throat areas must be initially sized inorder to set a back pressure sufficient to set the terminalshock in the constant area isolator duct. The requiredtotal (cumulative) valve throat height was found to beapproximately between 11 and 12 inches, depending onth e viscous ef fects . T he throat height in Fig. 7 is11.3 inch es and the isolator duct heights is 12.5 inches,as determined in the inlet sizing study for the Mach3.0/50,000-ft. flight condition. This approximation setsth e terminal shock approximately half w ay down th e

isolator duct. A shock trap is set by bleeding top andbottom areas just upstream of the terminal shock. Theisolator bleed was set for approximately 3.5% of theflow entering th e inlet (for a total of approximately5.0% bleed which includes inlet ramp bleed at the cow llip plane). This bleed rate was sufficient to hold th eterminal shock and not allow it to propagate upstreamfor any intermittent condition that was applied to thevalves in their various sequencing patterns.

T he init ial conditions at the internal in le t weregenerated by an inviscid analysis using FALCONwhere the ramps prior to the inlet were included. Thisis shown in Fig. 8. With this starting solution appliedto the grid in Fig. 7, a solution w as obtained for thesteady state mode. Then allowing valves 1 and 4 (seeFig. 7) to remain constant at 0.562 slugs/sec, an d0.99 slugs/sec., respectively, and changing valve 2 from0.861 slugs/sec, to 0.36 slugs/sec, and valve 3 for0.823 slugs/sec, to 1.6 slugs/sec, instantaneously, th eflow transfer was closely monitored by the finelyspaced t ime-dependent CFD analysis. Figure 9 alsoshows the flow drop in valve 2 to be almostinstantaneous, whereas th e flow in valve 3 has a moregradual buildup. The transfer was so rapid that therewas no possibility that a hammer shock could form.



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From Ref. 10, this requires approximately 12 msec.The distance from throat 2 to throat 3 was 65 inches,where Fig. 9 shows the bulk of the flow transfer takingplace in approximately .007 msec, and a flow balanceto be reached in 1.0 msec, at 17000 iterations. From13420 iterations to 14000 iterations, flow going out ofthe valves is slightly greater than flow coming in theinlet. Figure 10 shows velocity vectors colored byMach number just after the switch over, where the flowis supplied to valve 2 on the back side of the centerbody. Figure 11 shows the total stagnation pressure,referenced to the incoming inlet flow, which averagesapproximately 70%. (A total pressure recovery, T|r, of65% was assumed in the baseline installed performanceanalysis at this condition.) Static pressure at iteration17000 shows relative static pressures in the duct an dillustrates that constant pressure boundary conditionswould be difficult to implement at the diffuser exit.The flow transfer length from valves 2 to 3 was thelongest and represents the worst case. Flow transferlengths from valves 1 and 2 or 3 and 4 areapproximately 50 inches from throat to throat. SimilarCFD analyses were performed on valves 1 and 2 with 3an d 4 held constant, where similar time-dependentresults regarding flow transfer times were obtained.

The analytically demonstrated stable operation ofthe simulated inlet/PDE system for the selected "worst-case" integration concept is very encouraging. Itshould also be noted that 2-D CFD simulation of this ismore severe than for the actual integrated inlet/PDEgeometry since it does no t permit 3-D relaxation ofpressure pulses.

INSTALLED PERFORMANCEAnother impor tant task of this study was toestimate installed engine performance for the PDE. In

order to accomplish this task, an understanding of thePD E cycle, an acceptable cycle analysis code, andrealistic estimates of inlet, combustor, an d nozzleefficiencies are required.Cycle Analysis Procedure

Detonation in the PDE is a form of combustion thatdiffers significantly from deflagration, the type ofcombustion found in conventional gas turbine engines,pulse jets, an d rockets. Deflagration is characterized bysubsonic wave speeds, whereas the de tona t ioncombustion process occurs at high supersonic wavespeeds re la t ive to the u n b u r n e d reac tan ts(approximating C-J conditions). Detonation enginestake advantage of this supersonic wave speed to

produce a cycle of extremely short duration (about10 msec.) which allows frequencies on the order of100 Hz. The detonation acts as an aerodynamic pistonas it travels through the reactant gas mixture, raising theuseable pressure by a factor of 7 to 8. This constantvolume combustion process is thermodynamically moreefficient than the constant pressure def lagra t ioncombustion process and provides greater availableenergy for performing useful work.

To describe the entire PDE cycle requires no t onlymodeling detonation waves, but also the complexrarefaction waves. Utilizing both the detonation jum pconditions and the expansion wave relationship whichconsist of the characteristic Riemann invariant t imeintegration relationships, a time-dependent thermody-namic cycle ha s been defined. For the case withignition near the thrust wall, the following primarycycle events are illustrated in Fig. 12:1. Combustion chamb er (detonation duct) is filled withdetonable fuel/oxidizer mixture.2. Detonation is initiated near the thrust wall (closed

end of duct) and near C-J-quality detonation wavepropagates through the duct.

3. Detonation wave exits the duct. Th e duct is filledwith burned gases at pressure and temperature levelsconsiderably higher than ambient conditions.

4. The burned gases exit the duct in a blowdownprocess as rarefaction waves propagate forwardfrom the open end of the duct.5. The rarefaction eaves, after being reflected off theclosed end (thrust wall), exit the duct after most ofthe burned gases have been exhausted.6. After th e reflected rarefaction waves have been

exhausted, the duct is at a near uniform low pressurelevel and ready for the purge of the remainingburned gases an d subsequent refill of detonable fuel-oxidizer mixture; thus, beginning a new cycle.Previous research has provided suff ic ientunderstanding of detonation and rarefaction physics todescribe the PDE cycle and to develop an approximate

algebraic, 1-D cycle analysis code which comparesfavorably with CFD (Ref. 1). The cycle analysis codemodels planar detonations in tubes with a thrust wallan d ignition initiation specified anyw here in the tube. Aparametric engine performance model w as developedby LMTAS for PDE-powered vehicle analysis studies.



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The model provides the user with options to configure aPDE by select ing cycle design parameters in aconvenient namelist input format and then generatei ns t a l l ed eng ine performance at selected f l igh tconditions (Mach, altitude, and ambient temperature)an d power setting (detonation frequency). The primaryunique feature of the PDE cycle analysis model is thedetonator solver subroutine. The detonator solversubroutine analytically models the physics of C-Jdetonation waves and characteristic Riemann waves(see Ref. 1 for more detail). An exhaust nozzle system,w h i c h provides additional thrust, can be selected. Thecycle output data format has been developed fo ranalysis convenience, especially for parametric an dsystems analysis studies. The primary benefit of thiscycle analysis code is that it provides rapid, low-costpe r fo rmance predictions, w h i c h are needed foranalyt ical studies , such as this program, and forevaluat ing large matr ices of configurat ions , fl ightconditions, etc. fo r system-level studies.

To demonstrate the validity of the cycle code, itsresults have been compared with results from th eMOZART 1-D CFD code (Ref. 11) (with a complexchemical kinetics model). The time-dependent solu-tions were in good agreement as indicated in Fig. 13 .This agreement between two independently developedanalysis tools provide c o n f i d e n c e that the PDE cycleanalysis code adequately describes the pressure, tem-perature, an d mass flow versus t ime over th e entirecycle. From these basic relationships, time-averagedthrust w as computed with both th e cycle deck and CFDcode. Thrust comparison between the two codes agreedto within 1.5%. Additional information on the cycleanalysis code can be found in Refs. 1,2 and 3.

Basel ine cycle per fo rmance was computedassuming development of instantaneous C-J detonationwaves at the beginning of the detonation tube in astoichiometric fuel/air mixture. (In reality, oxygenenrichment may be required in the immediate vicinityof the ignition source to obtain a primary detonation.The pr imary detonation wave will then detonate th estoichiometric fuel/air mixture in the remainder of thetube. Another approach would be to use a highpower/high frequency ignitor system.)Cycle Performance Predictions

Until such time as actual pulse detonation enginesare on test stands, calculated performance numbers areonly estimates. However, in an effort to addressrealistic performance , th e fill valve coefficients havebeen estimated a t 80% and realistic component

efficiencies have been used. The airflow is alsoassumed to be injected through choked flow rotaryvalves into the combustion chamber. Frequencies inthe 70-100 Hz range are also assumed to be possible.(An engine des ign study es t ima tes tha t thesef r e q u e n c i e s are possible, but are at the upper end ofpossible frequencies fo r annular designs.) Based onthese assumptions, performance numbers have beengenerated for Mach/altitude f l i g h t conditions of1.2/25,000 ft. and 3.0/50,000 ft. (see Fig. 14). This firstorder analysis shows that thrust varies linearly withf requency with the detonat ion volume fixed. A sindicated from Fig. 14a, the required net thrust atMach 1.2/25,000 ft. is 30,000 Ibs. per module or90,000 Ib. total for 3 modules at a frequency of 90 Hz.Figures 14b and 14c also show the fuel flow and inletai rf low for these conditions. For the Mach3.0/50,000 ft . condition, the engine produced th erequired thrust at 74 Hz. The fuel-air ratio isstoichiometric. In addition, the net thrust nu m b e r srepresent thrust generated by expanding th e flow toatmospheric conditions and m u l t i p l y i n g that number bya CFG correction factor.

CONCLUSIONS The N A S A M a c h 5 Waver ider veh ic le , wi thwell-defined mixed-compression inlet lines for aturboramjet propulsion system, represents a goodbaseline vehicle for incorporation of PDE systems

for supersonic applications. Based on ascent trajectory thrust and drag historiesfrom Ref. 5, the required installed net thrust for theMach 1.2/25,000-ft. f l i g h t condition is= 90,000 Ibs . For the M a c h 3.0/50,000-ft.condition, 120,000 Ibs. of net thrust is required.The PDE system was sized to satisfy th e moredemanding Mach 1.2/25,000-ft. condition. PDE thrust is a direct function of engine volumeand operational frequency. Fo r propulsion systemsnecessary to power vehicles such as evaluated inthis study, operational frequencies on the order of

75 to 100 Hz are required. The in let isolator /dif fuser concep tua l designintegrated well with a three m odule, four-duct-highPDE concept. A single vehicle inlet duct suppliesair to all three PD E modules, with flow splittersbetween modules. The isolator's length-to-heightratio at the Mach 3.0 flight condition is 4.0. Ashock trap boundary-layer bleed system is used to

American Institute of Aeronautics an d Astronautics


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help stabilize the terminal shock train. The shortlength diffuser's area ratio (also height ratio) is six-to-one. A major feature of the diffuser is a centerbody which allows a conservative area distributionan d acceptable f low angles. The diffuser alsoincludes a plen um aft of the center body, justforward of the engine face, to dampen engine-induced pressure waves.

Time-dependent, ful ly viscous CFD analysisresults indicate that a terminal shock train ca nbe stabilized in the isolator at the Mach 3.0flight condition with a reasonable amountof isolator boundary-layer bleed (shock trap)and realistic PD E va lve-choked flow areas.Pressure perturbations and expansion waves causedby simulated PDE valve area changes do notdisturb the terminal shock system, thus analyticallyvalidating the inlet design/integration concept.Computed in te rna l inlet s tagnation pressurerecovery was approximately 70%, in agreementwith previous estimates.

The CFD analysis results show that for openplenum designs, the transfer of flow between side-valve PDE ducts due to phased valve timing occurswithin tens of microseconds. This does not allowtime for the formation of potentially destabilizinghammershocks. (Typical times for hammershocksformation is approximately 10 milliseconds.)ACKNOWLEDGMENTS

The authors are indebted to Neal Domel and JohnMaGuire of LMTAS an d Zane Pinckney of NASA-LaRC fo r their assistance in this study.REFERENCES

1. Hunter, L. G.; Couch, B. D.; Domel, N. D.; andWinf ree , D. D.: Pulse Detonation Concepts.General Dynamics Fort Worth Division ERR-FW-4339,1995.

2. Bratkovich, T. E.; and Bussing, T. R. A: A PulseDetonation Engine Performance Model. AIAA 95-3155, July 1995.

3. Bussing, T. R. A.; and Pappas, G.: An Introductionto Pulse Detonation Engines. A I A A 94-0263,January 1994.4. Hunter, L. G.; and Couch, B. D.: Air InjectionSystem Integration Study for Pulse Detonation

Engines. Lockheed Martin Tactical AircraftSystems FZM-8382, October 1995.5. Pegg, R. J.; Hunt , J. L.; et al.: Design of a

Hypersonic Waverider-Derived Airplane. AIAA93-0401,1993.

6. Hunter, L. G. and Winfree, D . D.: PulseDetonation Engine U.S. Patent 5,473,885; 1995.7. Bussing, T. R. A.: Rotary V a l v e Multiple

Combustor Pulse Detonation Engine. U.S. Patent5,345,758; 1994.

8. Reed, C. J.: Central Difference FALCON User'sManual . Lockheed Martin Tact ical AircraftSystems, 1994.9. Smith, B. R.: Prediction of Hypersonic ShockWave Turbulent Boundary-Layer Interactions with

the K-L Tw o Equation T urbulence Model. AIAA95-0232,1995.10 . M iller, D. R. and Hamstra, J. W.: A Com puta-

tional Assessment of Surge-Induced Inlet Over-Pressure Distr ibution fo r Preliminary Design.ASME 96-67-548.11 . Cambier, J. J. and Adelman, H . G.: PreliminaryNumerical Simulations of a Pulse DetonationWave Engine. AIAA 88-2960,1988.



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- Laser ignition

Air cooling(2 places)

Oil in

Purge air(4 places)

- Drive shaftfor rotary valve(variable RPM)Fuel/C>2 Fuel/air

Stratified mixing

(a) Side loading (U.S.Patent 5,473,885).

(b) Front loading (U.S. Patent 5,345,758).Fig. 3. Two types of PDE concepts.

Typical oiloutlet



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6.9 Forebody 14.5Movable10.1 Fixed ramp ramp atM a c h 3Equiv. engine flowarea 1289 sq. in.Shock trap Isolatorbleed ,*~ (LA = 4

Fig. 6. Inlet geometry.

Fig. 7. Grid for detailed inlet CFD valve closure study.

11



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I:IINIIIIIH LEWIS0.000000.200000.400000.600000.300001.000001.20000l.-KKHMi. J I K JDUi . o a o o o2. 200002.400002.6OOOO2.800003.000003.20000

3.ODDa.iici IJH,1.I8x1063.0-lxlD I197x8?

rtflCHHLPHHfluT I H EKHin

Fig. 8. External/internal Euler flow field solution at M = 3.

u.

13000 14000 15000Iteration count

16000 17000

Fig. 9. Valve flow transfer time history.

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ttJNIOUM L EUEL S0.00.100.20,30.

B l e e d

j.uuu nm:n0.00 DEC HLIHHI. lfiMia*fi ReI,3T10-I TIME

736X IDTi 6R1D

Valve 4Valve 3

Te rmi n a lnorm a l ./shock '"

Fig. 10. Velocity vectors jus t after valves are sequenced (closing valve 2, opening valve 3).

OONTOUft LEUfIS l.UDD IWCHO .QD OEG DLPHHI.IBxlO*6 ReI./IHIII..I rinc

lUj RR IO

Fig. 11. Total pressure for the case shown in f ig. 10 .

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Ignition point

(a) Initial co ndition.

f - -_ _ Detonation wave

(b) Ignition.

v = o

(c) Charge exits chamber.

Expansion wave -7V-/


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I Overall thrust, CFDOverall thrust, deckPressure thrust, CF DPressure thrust, deck60005000

4ooo3000

200010000

Static temperature, CFDStatic temperature, deck

.05 .10 .15 .20 .25Time, milliseconds .30 .35 .40

Fig. 13. Com parison of the LM TAS PDE cycle analysis with the 1-D CFD analysis of ref. 11.

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rtonon

< i f t n n n iiruvwi

orvwi j

-1

-

- . . . . . . -

Design point j.....$.......i

i i i i

I-"-'""

Desigi . . . . . . . -di . . - - i i i i

. . . . i

i point ..l>-""""

i i i i

'MO = 3.0/Alt = 50K

Mo=1.2/Alt = 25K

1 1 1 160 70 80 90Frequency, Hz

(a) Net thrust.100 110

35 -

25 -20115 -i n -

=

=^E . . - "t~ i i i i

. . . i

i i i i

, . . . . - v "''""

i i i i

...,-.-

.....-...----

i i i i

Mo = 3.0/Alt = 50K

Mo=1.2/Alt = 25K

l l l l60 70 80 90

Frequency, Hz(b) Airflow.

100 110

600

500g 400

=2 300

200

100

-

EI . . - i.-j r . - - - - - -~ 1 1 1 1

. . . - '

....-I

1 1 1 1

I--'"'

|.. ...-""

1 1 1 1

..-*

........-*

1 1 1 1

Mo = 3.0/Alt = 50K

M o= 1.2/Alt = 25 K

i i i i60 70 80 90Frequency, Hz 100 110

(c) Fuel flow.Fig. 14. Installed PDE performance parameters for one module.

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