NASA/TP--2001-211412

Development of a Gas-Fed Pulse

Detonation Research Engine

R.J. Litchford

Marshafl Space Flight Center, Marshafl Space Flight Center, Alabama

December 2001
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NASA/TP--2001-211412

Development of a Gas-Fed Pulse

Detonation Research Engine

R.J. Litchford

Marshall Space Flight Center, Marshall Space Flight Center, Alabama

National Aeronautics and

Space Administration

Marshall Space Flight Center ° MSFC, Alabama 35812

December 2001
Acknowledgments

This work was sponsored by the Third Generation Space Transportation Project Office (John Hutt, Project Manager),

Advanced Space Transportation Program, Marshall Space Flight Center, National Aeronautics and Space Administration.

The NASA Principal Investigator was Ron J. Litchford, Project Scientist, Space Transportation Research Project Office,

TD 15/ASTP. The author also wishes to acknowledge the excellent engineering and test support provided by

Bryan R. Thompson of TMET, Winchester, TN.

Available from:

NASA Center for AeroSpace Information

7121 Standard Drive

Hanover, MD 21076 1320

(301) 621 0390

National Technical Information Selwice

5285 Port Royal Road

Springfield, VA 22161

(703) 487 4650

ii
TABLE OF CONTENTS

1. INTRODUCTION ....................................................................................................................... 1

2. PROJECT OBJECTIVES ............................................................................................................ 3

3. DESIGN STRATEGY ................................................................................................................. 4

3.1 Integrated Injector Head Assembly .................................................................................... 4

3.2 Primary Engine Tube .......................................................................................................... 5

3.3 Ignition and Valve Control Circuits .................................................................................... 5

3.4 Microprocessor-Based ControlSystem .............................................................................. 6

4. SHAKEDOWN TESTING .......................................................................................................... 7

4.1 Initiator Test Firings ........................................................................................................... 7

4.2 PDRE-G-1 Test Firings ..................................................................................................... 8

5. CONCLUSIONS ......................................................................................................................... 12

REFERENCES .................................................................................................................................... 13

iii
LIST OF FIGURES

. Schematic of integrated injector head assembly. The initiator tube containing

the primary coaxial injector is inserted through the center of the injector head face.

Four auxiliary coaxial injectors axe distributed at 90 ° intervals around the central

initiator tube ............................................................................................................................. 15

. Photograph of integrated injector head assembly in preparation for preliminary

shakedown testing (side view) ................................................................................................. 16

. Photograph of integrated injector head assembly in preparation for preliminary

shakedown testing (injector face view) ................................................................................... 17

. Schematic of the igniter control circuit using fiber optic coupling from a TTL

optical transmitter to a TTL optical receiver ........................................................................... 18

. Schematic of the propellant valve control circuit using fiber optic coupling from

a TTL optical transmitter to a TTL optical receiver ................................................................ 19

. Microprocessor control signals acquired at the output of the TTL optical receivers

during a 40-Hz initiator shakedown test .................................................................................. 2O

7. Theoretical C-J detonation characteristics as a function of H 2 volume fraction ..................... 21

8. Theoretical flame front trajectories on the position-time wave diagram ................................. 22

. Cross section of the mild steel initiator model with pressure sensor ports

(dimensions in centimeters) ..................................................................................................... 23

10. Measured pressure waveforms in the initiator tube with smooth bore. Indicated flame

front velocity was 540 m/sec ................................................................................................... 24

11. Illustration of Schelkin spiral and its effect on turbulent mixing and flame acceleration ....... 25

12. Measured pressure waveforms in the initiator tube with Schelkin spiral No. 1.

Indicated flame front velocity was 1,400 m/sec ...................................................................... 26

13. Measured pressure waveforms in the initiator tube with Schelkin spiral No. 2.

Indicated flame front velocity was 2,400 m/sec ...................................................................... 27

V
LIST OF FIGURES (continued)

14. Schematic of PDRE-G-1 research engine on the test stand. Sensor port locations

and indices axe indicated .......................................................................................................... 28

15. Measured pressure waveforms in the primary tube with smooth bore. Indicated flame

front velocity was 505 m/sec ................................................................................................... 29

16. Measured pressure waveforms in the primary tube with Schelkin spiral installed along

its entire length. Indicated flame front velocity was 2,200 m/sec ........................................... 3O

17. Measured pressure waveform distribution in the primary tube with Schelkin spiral

installed along its entire length ................................................................................................ 31

18. Idealized PDRE cycle processes: (1) Detonation initiation; (2) propagation of detonation

wave through chamber; (3) propagation of rarefaction wave from tube exit to tube head;

(4) tube blowdown; (5) chamber exhausted to atmospheric pressure; and (6) injection

of fresh propellant mixture ...................................................................................................... 32

19. Idealized position-time wave diagram for PDRE blowdown process ..................................... 33

20. Idealized PDRE tube head pressure for a simple straight tube configuration ......................... 34

21. Long exposure photograph of the exhaust plume from the PDRE-G-1 engine

illustrating the firing characteristics for a filled tube condition. Obtained during

a night firing with hydrogen/oxygen propellants .................................................................... 35

22. Long exposure photograph of the exhaust plume from the PDRE-G-1 engine illustrating

the firing characteristics for an underfilled tube condition. Obtained during a night

firing with hydrogen/oxygen propellants ................................................................................ 36

23. Time-resolved PDRE-G-1 engine thrust profiles for filled and underfilled

tube conditions ......................................................................................................................... 37

vi
LIST OF ACRONYMS AND SYMBOLS

AFB

C-J

CDI

CFD

DDT

H2

MSFC

02

PC

PDE

PDRE

PDRE-G-1

TTL

Air Force Base

Chapman-Jouget (detonation condition)

capacitive discharge ignition

computational fluid dynamics

deflagration-to-detonation transition

molecular hydrogen

specific impulse

Marshall Space Flight Center

molecular oxygen

personal computer

pulse detonation engine

pulse detonation rocket engine

pulse detonation research engine (gas-fed series, model No. 1)

transistor-transistor logic

vii
NOMENCLATURE

c 3

Ec

Eigr_

Lc

LDD_

P

Pl

P2

P3

t

tDD_

bl

UCJ

ble

X

speed of sound in burned gas

critical detonation initiation energy

ignition energy

chamber length

deflagration-to-detonation transition length

pressure

initial pressure in unburned mixture

pressure of burned gases immediately behind detonation wave

pressure of burned gases after detonation wave passage

time

deflagration-to-detonation transition time

velocity

Chapman-Jouget detonation velocity

exhaust velocity

distance

characteristic time

characteristic time for detonation wave propagation

characteristic time for rarefaction wave propagation

characteristic time for tube blowdown

ix
TECHNICAL PUBLICATION

DEVELOPMENT OF A GAS-FED PULSE DETONATION RESEARCH ENGINE

1. INTRODUCTION

Pulsed detonation architectures have the potential to revolutionize aerospace propulsion and power

systems. This assertion is derived from the unique performance gains associated with unsteady reacting

flow effects and from the cost reductions associated with a highly simplified design configuration. Consequently, there is considerable interest in obtaining an accurate assessment of representative performance

and cost benefits.

It should be noted that there are two fundamental architectural classifications for pulse detonation

engines (PDEs). That is, they can either obtain air from the surrounding atmosphere or acquire an oxidizer

from onboaxd storage tanks. In the former case, PDEs represent an attractive low-cost alternative to turbojets based on their potential for surpassing intrinsic limits on scalability, operating range, efficiency, and

manufacturing costs. In the latter case, the device is referred to as a pulse detonation rocket engine (PDRE).

Generally speaking, PDEs operate on an intermittent cycle by filling a chamber with reactants and

inducing ignition in an energetic fashion such that a combustion detonation wave is rapidly established.

The high propagation speed of the detonation wave implies orders-of-magnitude increase in combustion

rates relative to the deflagration mode, and the chemical energy is completely released within a very narrow region behind the leading shock wave. As a result, the combustion process takes place at nearly constant volume, which yields a higher postcombustion temperature (i.e., higher performance or specific

impulse (Isv)) than a constant pressure process). The detonation wave has the added benefit of attaining a

higher peak pressure.

In fact, contemporary interest in PDREs arises primarily from advantages that accrue from the

significant combustion pressure rise developed by the detonation process. Because PDREs do not require

extreme precompression of the working fluid to fill the chamber, it should be possible to eliminate the need

for large and heavy turbopumps. This potential advantage must be carefully assessed, of course, in view of

the added requirement for fast-acting robust valves. Furthermore, the wide throttling range of a PDRE

should not be ignored as a favorable attribute since it can play an important role in many mission scenarios.

Although PDREs are expected to achieve greater I_7_ (due to the higher peak temperature) and

greater design simplicity (due to the relaxed turbopump requirement), it is not yet clear whether they will

prove superior to conventional rocket engines from a systems point of view. Basic scientific feasibility has

been clearly established in several experimental and theoretical studies, as noted in recently published

review articles, 1 3 but the unsteady operational attributes of PDEs complicate their analysis when including nonideal effects, and it is difficult to make the needed system level assessments.
Surely,significantstridesarebeingmadewithrespectto thedevelopmentof designstrategiesand

performanceoptimization,4_ butit shouldberecognizedthattheyarestill at anearlystageof development.Furthermore,thereaxecriticaloperationalissuesthatmustbeaddressedbeforethePDEconceptwill

everberealizedin practice,andreliableempiricaldatais badlyneeded.

A numberof privatecompaniesaxeengagedin thedesignanddevelopmentof variousprototype

enginesto addresstechnicaluncertaintiesandreducerisk, buttheacquireddataaxenotwidelydisseminateddueto proprietaryconcerns.Therefore,a criticalneedexistsfor highly instrumentedlaboratory

researchenginesthatwill permittheestablishment of anempiricaldatabase,whichcanbewidelydisseminatedthroughoutindustry,government,andacademia.TheNavalPostgraduateSchoolandtheAir Force

ResearchLaboratoryat Wright-PattersonAir ForceBase(AFB),for instance,haveeachbuiltresearch

enginesandinitiatedfundamentalresearchprogramsin ordertohaveanindependentsourceof empirical

data.1°,11ThisTechnicalPublicationdescribesthe developmentof a low-costgas-fedpulsedetonation

researchengine,whichisintendedto serveinasimilarcapacityattheNASAMarshallSpaceFlightCenter

(MSFC).

2
2. PROJECT OBJECTIVES

The MSFC pulse detonation research engine, designated as PDRE-G-1 (gas-fed series, model No.

1), was developed to satisfy a number of diverse needs. The principal objective overriding all other factors

was the desire to obtain a simple and flexible research device with which to generate empirical data independent of proprietary concerns. It should be emphasized that the resulting device is a research engine and

is not intended as an evolutionary step toward a production engine.

The hope is that the PDRE-G-1 device will serve as a flexible research tool for investigating fundamental physical processes and basic operational issues. The resulting knowledge can then be widely

disseminated so that it may have the broadest possible impact on design tool development, design optimization, and system analysis efforts. The device could also be used as a test-bed for developing critical

enabling technologies that may one day find use on a practical production engine.

Specifically, the MSFC PDRE-G-1 engine was designed to accomplish the following technical

objectives:

• Provide a nonproprietaxy empirical database with respect to PDE performance and operational

characteristics;

• Support development of theoretical/computational fluid dynamics (CFD) analysis tools as well

as engineering design tools;

• Develop/validate PDE design scaling laws;

• Address system integration issues;

• Improve understanding of detonation physics;

• Explore innovative design strategies and configurations related to valves, propellant injection,

nozzles, detonation initiation, electrical power extraction, etc.;

• Improve definition of system operational requirements related to propellant pumping, electrical

power requirements, control, etc.;

• Serve as a test-bed for subsystem component testing; and,

• Lay the groundwork for development of a liquid-fed pulse detonation research engine.
3. DESIGN STRATEGY

The guiding design strategy for the PDRE-G-1 engine was to achieve a simple and flexible research apparatus that would be inexpensive to build and operate. Therefore, the engine was designed as a

heat sink device and, as such, can only be operated in a burst mode with run durations of a few seconds.

From a research perspective, this is not a critical limitation since short duration tests are entirely adequate

for acquiring the data of primary interest.

In keeping with the low-cost design strategy, maximum use was made of standard off-the-shelf

industrial or automotive components wherever possible. For example, the injector head assembly and

primary tube were fabricated from standard pipe and flanges, and simple coaxial injectors were incorporated into the design. Standard industrial solenoid valves were incorporated to control the propellant flow

to the injectors.

In a similar simplistic fashion, a high-performance automotive electronic ignition system was adapted

for use, and spark plugs were mounted at various locations to permit exploration of alternative configurations for detonation initiation. Because PDEs require precise control over valve cycling and ignition timing, significant effort was expended to develop a microprocessor-based engine control system that incorporated fiber optic coupling of the processor digital control signals to the engine.

As a research device, it was deemed necessary to make adequate provisions for the engine to be

highly instrumented. Therefore, sensor and optical diagnostic access ports were incorporated throughout

the engine design. Additional preparations related to operational support included assembly of propellant

storage and supply lines, construction of a rigid test stand, and development of test procedures and special

precautionary measures due to the unique hazards associated with PDE operation.

3.1 Integrated Injector Head Assembly

The core of the engine is the integrated injector head assembly, which is fabricated entirely from

stainless steel components and materials. This assembly, shown schematically in figure 1, consists of the

initiator tube containing the primary coaxial injector, which is inserted directly through the center of the

injector head. The injector head face has four auxiliary coaxial injectors distributed at 90 ° intervals around

the initiator tube. The primary engine tube attaches to a flange, which is in turn mounted to the injector

head assembly.

The primary coaxial injector is attached to the rear of the initiator tube such that the propellants

thoroughly mix in the initiator and then exit into the primary tube. The initiator tube is 15-cm long and

currently has a 1.1-cm-diameter bore. The initiator tube wall was designed sufficiently thick to accommodate a larger diameter bore if necessary. A spark plug can also be inserted into the initiator tube at a point

2.5 cm downstream from the primary injector.

4
Theauxiliaryinjectorswereincorporatedintothedesignto provideadditionalflexibility andto

permitexplorationof alternativepropellantinjectionschemes.Forexample,theycouldbeusedto inject

theprimarypropellantor buffergasesasdesired.Theseinjectorshaveaverysimpledesignin whichthe

centerpostsaxesilversolderedto theinitiatorbodyandinsertedthroughholesin theinjector-headfaceto

formacoaxialconfiguration.Thecenterpostsaxefedthroughholesboredintotheinitiatorbodyandthe

annuliaxefedthroughholesboredinto theinjectorhead.It shouldbenotedthattheprimaryinjectoris

fabricatedin asimilarlysimplisticfashion.

Photographsof theintegratedinjectorheadassemblyaxeshownin figures2 and3 whileundergoing preparationsfor preliminaryshakedowntestingof theinitiatortube.Theindividualfeedlinesto the

injectorsaswell astheindividualauxiliarycoaxialinjectorscanbeclearlyseen.

3.2 Primary Engine Tube

The basis for the primary engine tube was a 90-cm-long section of 5-cm-diameter schedule-40

stainless steel pipe. This simple approach was expedient and inexpensive, and it provided sufficient design

margin for operation in a heat sink mode despite the severe thermal and structural load cycles. The pipe

was threaded at one end and could be directly mated with the integrated injector-head tube flange.

The only required machining was the drilling of holes to accept various bosses for spark plugs,

sensor ports, and optical diagnostic windows. Sensor ports were located at 10-cm intervals along the

primary tube axis and two opposing spark plug bosses were located near the end of the tube closest to the

injector head assembly. Diametrically opposed optical window ports were located at the same axial location as the sensor ports and the spark plug bosses. Subsequent shakedown testing of the device demonstrated the thermal and structural integrity of the primary tube design.

3.3 Ignition and Valve Control Circuits

Successful operation of any PDE demands precise control over valve and ignition timing. Therefore, considerable attention was devoted to the development of adequate control circuits. It was decided at

an early stage that a digital microprocessor would be used for basic engine control since this would provide

the highest degree of flexibility and would be readily adaptable for the exploration of advanced automatic

control techniques.

It was also decided to pursue a strategy whereby the digital control signals would be fiber optically

coupled to the engine. Fiber optic coupling has many attractive benefits including complete electrical

isolation of the microprocessor control system from the engine, high signal-to-noise ratio, low signal

attenuation over long transfer distances, high reliability, and greater simplicity in the engine wiring harness. Following this strategy, the first objective was to develop a simple low-cost optical transmitter and

receiver pair, which could be utilized in both the ignition and valve control circuits.
In theresultingdesign,thetransistor-transistorlogic (TTL) opticaltransmittertakesthecontrol

signalfromthemicroprocessorandconvertsit intoalightpulseusingasimpletransistorcircuitandafast

response(>10MHz) infraredlight emittingdiode(MotorolaModelMFOE71).Thelight pulsefromthe

transmitteris thencoupledto a 1,000-gmplasticfiber opticcable.At thereceivingend,asimple5-kHz

TTL opticalreceiver,basedonastandardphototraxlsistor(MotorolaModelMFOD72),detectsthelight

pulseemergingfrom the fiberopticandconvertsit into a crisplow-levelTTL signalusinga Schmidt

trigger.Thefinaldesigndiagramsfortheignitionandpropellantvalvecontrolcircuitsareshownin figures

4 and5,respectively.

In theignitioncircuit,theTTLsignalfromtheopticalreceiverisfedintoatransistorswitch,which

shortsthetriggerterminalof a capacitivedischargeignition (CDI) controller(ModelMSD 6A). This

actiongeneratesmultiple400-Vdischargesacrosstheprimaryof a 100:1transformercoil,andtheresulting40-kVoutputinthesecondarycoil is transmittedto asparkplug.Multiple 110-mJgasdischargesare

thenformedatthesparkplugelectrodesyieldinganeffectiveandenergeticignitionsource.

In thesolenoidvalvecircuit,the TTL signal from the optical receiver is connected to a solid-state

relay, which directly controls propellant valve activation. The fuel and oxidizer valves axe controlled with

independent circuits, and additional independent circuits axe available for controlling auxiliary valves as

needed.

Prior to installation on the engine, these circuits were built and tested on the bench to confirm

anticipated performance. The test results demonstrated excellent response and signal-to-noise

characteristics.

3.4 Microprocessor-Based Control System

Engine operation was completely automated using a computer-based control system. This approach

gives precise control over engine timing variables and provides a suitable framework for implementing

advanced feedback control techniques as desired.

The basis of the control system is a National Instruments modular PXI platform with embedded PC

controller. Currently, the TTL engine control signals axe generated by an 8-chaxlnel 32-bit counter/timer

module with software control under a customized program written in the LabVIEW graphical programming language. This program permits precise adjustment of the duty cycle and phasing of each control

signal. For example, figure 6 shows the control signals acquired at the output of the TTL optical transmitters during a 40-Hz initiator shakedown test. These waveforms demonstrate a capability for producing

high-quality signals with very precise timing.

6
4. SHAKEDOWN TESTING

To date, a number of shakedown tests have been conducted on the initiator and the full PDRE-G-1

engine assembly. All of these firings have been conducted with hydrogen/oxygen propellant mixtures. The

principle objective of these firings was to validate system and subsystem functionality and to verify the

attainment of true detonation conditions. The basic approach was to acquire waveforms from dynamic

pressure transducers at selected locations. The combustion wave velocity could then be deduced from its

measured time of flight and compared against theoretical Chapman-Jouget (C-J) values.

For further reference, the theoretical C-J detonation characteristics are provided in figure 7. This

plot shows the detonation velocity and pressure ratio across the shock as a function of the hydrogen volume

fraction. For a stoichiometric mixture ratio, the detonation velocity is roughly 2.8 km/sec and the pressure

ratio is about 18.7, which is near the peak value for H2/O 2 mixtures.

4.1 Initiator Test Firings

In general, the flame front can evolve along three possible trajectories as indicated in the positiontime wave diagram of figure 8. If the ignition energy (EigT¢)exceeds the critical energy for direct detonation

initiation (Ec), then the detonation wave is immediately formed at the ignition point and propagates down

the tube at the C-J detonation velocity. If EigT_< E c and a mechanism such as turbulent mixing is available

to accelerate the flame, then the flame can gradually accelerate beyond the laminar flame speed and eventually undergo deflagration-to-detonation transition (DDT). If EigT¢ < E c and no mechanism exists for

accelerating the flame, then the flame will propagate down the tube at the laminar flame speed. Because

the energy deposition of the PDRE-G-1 ignition source (110 mJ) was much less than the critical initiation

energy for oxygen/hydrogen mixtures, it was necessary to rely on a DDT initiation strategy. Furthermore,

the time and length scales for DDT would have to be very short for practical utilization.

To facilitate evaluation of the initiator tube, an identical copy was machined from mild steel and

outfitted with three sensor ports as shown in figure 9. Each port was located 2.5 cm from its neighbor with

the last port located 1.9 cm from the initiator tube exit. Pressure waveforms were then acquired at a 20-Hz

firing rate while tuning the propellant injection rate to obtain the highest possible combustion wave speed.

The measured results for the smooth bore initiator tube axe shown in figure 10 with an indicated

combustion wave speed of 540 m/sec. This value was significantly lower than the theoretical C-J value

implying that DDT was not being attained. Clearly, the smooth bore tube was not providing sufficient

flame acceleration.

At this point, it was decided to take advantage of the phenomena discovered by K.I. Schelkin

during a series of careful investigations carried out between 1945 and 1949 in which he observed that
flamepropagationspeedwoulddramaticallyincreasein tubeswith roughwalls.12His observationgave

birthto thenowwell-knownSchelkinspiralforinducingrapidtransitionto detonationconditions.Later

workbyL.RKayushindemonstratedsimilareffectsusinggridsandorificesinthetube.13Thenowwidely

acceptedviewisthatobstaclesin theflow generateintenseturbulencemixingwhichinducesrapidaccelerationof theflame.TheSchelkinspiralandits effectonturbulentmixingandflameaccelerationaxe

illustratedinfigure11.

Followingthislineofreasoning,anotherseriesof initiatortestswereconductedusingtwodifferent

Schelkinspirals.SpiralNo. 1wasconstructedfrom1/:5-in.-diameter wirewitha1/4-in.pitch,andspiralNo.2

usedconstructedfrom1/40-in. diameterwirewitha1/4-in.pitch.Themeasuredpressurewaveformsforthese

testsareshownin figures12and13,andthecombustionwavespeedwasobservedto increaseasanticipated.Themeasuredvaluefor spiralNo. 2 was1.4km/sec,whichwasstill faxlowerthetheoreticalC-J

value;however,themeasuredvaluefor spiralNo.2 was2.4km/sec,whichisin closeagreementwiththe

theoreticalvaluefor anH2volumefractionnear0.5.

4.2 PDRE-G-1 Test Firings

Preliminaxy testing ofthe fully assembled PDRE-G-1 reseaxch engine was initiated upon completion of the initiator shakedown tests. The engine was installed on the test stand, as shown schematically in

figure 14, and outfitted with dynamic pressure transducers at the indicated port locations. Also note that the

engine was equipped with a thrust transmission structure and a dynamic load cell for the purpose of making time-resolved thrust measurements. Optical window ports were also available for flow visualization.

In these tests, two additional spark plugs were mounted in the primary tube. This was done to

provide multiple ignition sources as a means of reducing the chance for misfire. It should also be noted that

the small fast-acting solenoid valves used during the initiator test firings were replaced with larger but

much slower industrial solenoid valves in order to minimize pressure losses. The slower valve response

limited the engine operating frequency to about 5 Hz. The entire propellant flow was injected through the

primary injector during all tests. The auxiliary injectors were not utilized. Following our established test

strategy, pressure waveforms were first acquired with a smooth bore primary tube while tuning the propellant flow to obtain high combustion wave speed.

4.2.1 Detonation Characteristics

The measured pressure waveforms near the end of the primary tube (ports 7 and 8) axe shown in

figure 15 for the smooth bore case. With an indicated flame speed of only 505 m/sec, it was clear that

detonation conditions were not being attained. Since the initiator tube with Schelkin spiral was known to

detonate, there was some hope of achieving detonation wave diffraction into the primary tube. However,

the critical tube diameter for hydrogen/oxygen is on the order of 2 cm, and the current initiator tube diameter is only 1.1 cm. It is clear from our results that the ratio of initiator tube diameter to primary tube

diameter must be increased in order to obtain successful detonation propagation into the primary tube.
Thisquestionis of greatinterestfor futureinvestigation,butthemostexpedientapproachwasto

investigatetheutilizationof aSchelkinspiralforinducementofdeflagration-to-detonationtransition.Along

theselines,asectionof 1⁄4-in.stainlesssteeltubewasshapedintoaspiralhavinga5-cmoutsidediameter,

a3/4-in.pitch,andalengthequalto theprimarytubelength.Thisspiralwasplacedinsidetheprimarytube

andtheenginetestswererepeated.

Theresultingpressurewaveformsneartheendof theprimarytube(ports7 and8) axeshownin

figure 16.In this case,combustionwavespeedwasobservedto increaseto 2.2 km/sec.This valueis

indicativeof detonationconditionsandcorrespondsto anH2volumefractionnear0.5asmaybededuced

frominspectionof figure7. Thisimpliedthattheenginewasoperatingwithaleanmixtureratio.

Unfortunately,the abilityto setanaccuratemixing ratiois difficult atthisjuncturebecausethe

currentsetuphasno provisionfor accuratelymeasuringflow rates.It is only possibleto makeestimates

basedonanapproximatevaluefor thevalveflow coefficientandtheknownsupplypressures.Therefore,

futureworkhasbeenplannedto addressthedevelopmentof anaccurateflowratemeasurement capability.

It wasuncertainwherethetransitionfromdeflagrationto detonationactuallyoccurredsoanother

seriesof testswasconductedto obtaintheaxialdistributionin flamespeedalongtheprimarytube.The

resultingpressurewaveformsfor ports2-8 axeshownin figure17,andit wasobservedthatthedetonation

velocityremainedvirtuallyconstantat2.2km/secalongtheinterrogatedlength.Thisimpliedthatdetonationtransitionoccurredbeforeport2,andthattheSchelkincoilcouldbeshortenedto alength<20cmand

still achieveasimilarresult.Thishypothesishasyet to beattemptedin practice,butwill be addressed

furtherasdevelopmentworkcontinues.

4.2.2 Thrust Characteristics

The engine thrust characteristics were of immense interest, and the availability of a fast response

dynamic load cell permitted the acquisition of time-resolved profiles for installed engine thrust. The total

engine impulse could subsequently be deduced through integration of these measured thrust profiles. The

term "installed thrust" is used because the engine and thrust transmission structure behave as an effective

spring-damper system. As such, the resulting thrust profile, measured by the dynamic load cell, depends

upon the dynamic response of the installed engine structure, which is in turn governed by the structural

mass, effective spring constant, and effective damping characteristics. Therefore, the installed engine thrust

profile does not necessarily correspond with the fluid dynamic force profile acting on the engine.

The ideal fluid dynamic forcing function may be deduced from a consideration of the ideal PDRE

cycle process as illustrated in figure 18. For the simple straight tube engine geometry, the only fluid

dynamic force acting on the engine is the surface pressure at the head of the engine. Once the detonation

has been initiated and the flame front begins to propagate through the tube, a centered expansion occurs

between the detonation wave and the engine head with the generation of right-running rarefaction waves.

The detonation wave exits the tube at time _CI = Lc/UCj , and left-running rarefaction waves then run back

down the tube at the speed of sound of the burned gases (c3). The time for the left-running rarefaction wave

to reach the engine head is "cii = Lc/C 3. At this point, the tube experiences an exponential blowdown of the

remaining combustion gases over a characteristic time _CIII. These detonation and rarefaction wave propagation trajectories axe illustrated in the position-time wave diagram of figure 19.
Basedontheidealizedblowdownprocessabove,theengineheadpressureprofilecanbeapproximatedasaconstantvaluefromthetimeof detonationinitiationtothetimefortheleftrunningrarefaction

to returnto thepointof ignition(_CI+_CII). The remaining portion of the thrust can then be approximated by

an exponential decay with a characteristic time of _CIII. This idealized engine head pressure profile is illustrated in figure 20.

It should also be noted that the thrust profile depends on the tube fill conditions. If the tube is

completely filled with propellant, the detonation wave will propagate through the entire tube and fill it with

combustion gases. These gases axe expelled during the subsequent blowdown process. If the tube is only

partially filled, however, some quantity of air must be pushed out ahead of the combustion gases.

Clearly, maximum thrust will be obtained for the filled tube condition, but experiments carried out

by Professor V.V. Mitrofanov have demonstrated that maximum I_v can occur when the tube is underfilled. 14

That is, transference of combustion energy to the free additional air mass in the tube can be more effective

from a propellant consumption standpoint.

Following this line of thought, an attempt was made to operate the PDRE-G-1 engine for filled and

underfilled tube conditions. The filled tube condition had been demonstrated in the previous set of firings.

It was only necessary to fix the propellant supply pressure and increase the operating frequency to obtain

an underfilled tube condition.

The effect of the tube filling condition on the exhaust plume is demonstrated in figures 21 and 22,

which show long exposure photographs obtained during night firings for filled and underfilled tube conditions, respectively. The difference between the two cases is clearly observed from the plume luminosity.

For the filled case, the ejected combustion gases create a highly luminous expanding plume, which wraps

back around the nozzle. For the underfilled case, the luminous plume is confined to a dim narrow cone.

An attempt was then made to acquire time-resolved installed thrust measurements for these two

cases. The resulting ensemble-averaged thrust waveforms axe shown in figure 23. Both profiles show a

rapidly rising thrust pulse, which peaks at a time that is strongly correlated with the exit of the detonation

wave from the tube, and then exhibits a slow decay as the tube blows down. These results do not correspond to the idealized thrust profile of figure 20 due to the spring-damper response characteristics, but the

results axe consistent with previous time-resolved installed thrust measurements obtained by the author

using single-shot detonation tubes. 15

As previously noted, the timescale for passage of the detonation wave through the tube (_CI) appears

to be correlated with the rise time of the thrust pulse. This quantity is readily determined from the ratio of

the chamber length, L c, to the detonation velocity, uD, such that zI = Lc/u D -_ 400 _tsec. Due to the Taylor

wave, a centered expansion occurs between the detonation wave and closed end of the tube, where the

velocity must go to zero. The resulting pressure on the head of the engine determines the fluid dynamic

forcing function for the spring-damper system, but the shape of the installed thrust profile depends on

dynamic response of the system.

10
Thepeakthrustvaluefor theunderfilledtubecase(-_150lbf) is abouthalfthat obtainedfor the

filled tubecase(-_270lbf).By integratingtheseinstalledthrustprofiles,it wasalsopossibleto obtainthe

totalimpulseassociatedwith eachcondition.Theresultswere0.21lbfosecand0.41lbfosec,respectively.

Becauseoftheuncertaintyinthepropellantflowratesatthisstageof development,thereisconsiderableuncertaintyasto theexactquantitativevalueof themixtureratioandthefill amount.Therefore,no

attemptwasmadeto deducethespecificimpulsefromthisdata.Deductionof I_p will be deferred until an

accurate flow rate measurement becomes available.

11
5. CONCLUSIONS

A low-costgas-fedPDREhasbeensuccessfullydevelopedandplacedintooperationtoserveasan

independentsourceof empiricaldataandasatest-bedfor developingtherequiredenablingtechnologies

forthis classof engine.Preliminaryshakedowntestsusinghydrogen/oxygenpropellantsconfirmedthe

attainmentof detonationconditionswithmeasuredcombustionwavespeedsin therangeof 2.2-2.4km/sec

whenusingaSchelkinspiral.Time-resolvedthrustmeasurementswerealsoacquiredforfilledandunderfilled

tubeconditions.Peakthrustvalueswere270lbf and150lbf,respectively.Furtherdevelopmenteffortsare

plannedtoimproveengineperformanceandoperatingcapabilitiesandtoacquireimprovedcapabilitiesfor

moreaccuratemeasurement of engineoperatingparameters.

12
REFERENCES

1. Bussing, T.; and Pappas, G.: "Pulse Detonation Engine Theory and Concepts," Developments in

High-Speed-Vehicle Propulsion Systems, Murthy, S.N.B.; and Curran, E.T. (eds.), Vol. 165, Progress

in Astronautics and Aeronautics, AIAA, NY, pp. 421-472, 1997.

2. Lynch, E.D.; and Edelman, R.B.: "Analysis of the Pulse Detonation Wave Engine," Developments in

High-Speed-Vehicle Propulsion Systems, Murthy, S.N.B.; and Curran, E.T. (eds.), Vol. 165, Progress

in Astronautics and Aeronautics, AIAA, NY, pp. 473-516, 1997.

3. Kailasanath, K.: "Applications of Detonation to Propulsion--A Review," AIAA 99-1067, 1999.

4. Cambier, J.L.; and Tegner, J.K.: "Strategies for Pulsed Detonation Engine Performance Optimization," J. Propulsion and Power, Vol. 14, No. 4, pp. 192-196, 1998.

5. Cambier, J.L.: "Preliminary Modeling of Pulse Detonation Rocket Engines," AIAA 99-2659, 1999.

6. Li, C.; Kailasanath, K.; and Patnaik,G.: "A Numerical Study of Flow Field Evolution in a Pulsed

Detonation Engine," AIAA 2000-0314, 2000.

7. Eidelman, S.; Sharov, D.; and Book, D.: "Aerothermodynamic Analysis of Pulse Detonation

Engines," AIAA 2000-3892, 2000.

8. Ebra_limi, H.B.; Mohanraj, R.; and Merkle, C.L.: "Multi-Level Analysis of Pulse Detonation

Engines," AIAA 2000-3859, 2000.

9. Mohanraj, R.; Merkle, C.L.; and Ebra_limi, H.B.: "Modeling of Pulse Detonation Engine Operation,"

AIAA 2001-0475, 2001.

10. Brophy, C.M.; Netzer, D.W.; Sinibaldi, J.O.; and Johnson, R.G.: "Operation of a JP10/Air Pulse

Detonation Engine," AIAA 2000-3591, 2000.

11. Schauer, F.; Stutrud, J.; and Bradley, R.: "Detonation Initiation Studies and Performance Results for

Pulsed Detonation Engine Applications," AIAA 2001-1129, 2001.

12. Schelkin, K.I.: Zhurnal Eksperimental'noi I Teoreticheskoi Fiziki, 25 (5) 1953.

13. Kayushin, L.E: "The Influence of External Friction and Heat Exchange on the Motion of an Ignition

Surface and a Shock-Discontinuity in a Chemically Reactive Medium," in Gas Dynamics and Physics of

Combustion, NASA-TT-F-79, pp. 53-65, 1962.

14. Mitrofanov, V.V.: Lavrentyev Institute for Hydrodynamics, Novosibirsk, Russia, Private Communication, March 2001.

15. Litchford, R.J.: "Integrated Pulse Detonation Propulsion and Magnetohydrodynamic Power," NASA/

TP 2001-210801, January 2001.

13
.m

ae-a

I..[-

t'-"

I..[-

Figure 1. Schematic of integrated injector head assembly. The initiator tube containing the primary

coaxial injector is inserted through the center of the injector head face. Four auxiliary coaxial

injectors axe distributed at 90 ° intervals around the central initiator tube.

15
Figure 2. Photograph of integrated injector head assembly in preparation for preliminary

shakedown testing (side view).

16
Figure3. Photographof integratedinjectorheadassemblyin preparationfor preliminary

shakedowntesting(injectorfaceview).

17
o_

•r-- .i

ODin

c:l) i

¢.9 ¢.,O - m

CO _.. O

.-_ o

CM

,r--

k...

_o

_ i

Lt_ _ ...........

+ ' i

¢.9

C)

i__

cOD

_...I

I----

I----

\ __.->

Figure 4. Schematic of the igniter control circuit using fiber optic coupling from a TTL optical

transmitter to a TTL optical receiver.

18
O

ii

> %

-tt'-"

_..1

I--- ......... i

I °

J

Figure 5. Schematic of the propellant valve control circuit using fiber optic coupling from a TTL

optical transmitter to a TTL optical receiver.

19
Figure 6. Microprocessor control signals acquired at the output of the TTL optical receivers

during a 40-Hz initiator shakedown test.

2O
Ld/ _d

IlilllililOilillilillii li

0

0

0

0

e--

0

ii

E

0

--rFigure 7. Theoretical C-J detonation characteristics as a function of H 2 volume fraction.

21
_o

k.-

.m

g_=

E

_d I---

Figure 8. Theoretical flame front trajectories on the position-time wave diagram.

22
t-"

O

CO

c_

0

0

t--"

CM

¢aO

t

-_- o

F

I

I

I

I

I

I

I

I

a,

O

"T'--

"T'--

O

t.O

Figure 9. Cross section of the mild steel initiator model with pressure sensor ports

(dimensions in centimeters).

23
O

,T--

O

O

c5

O

E

c5 "-"

E

.m

c5

O

CXl

O

,T--

O

oJnssoJdpo0nPoEI

0

0

C_I

0

I

Figure 10. Measured pressure waveforms in the initiator tube with smooth bore.

Indicated flame front velocity was 540 m/sec.

24
¢..O

I:D

¢--

O"9

t"-

¢..O

rr"

e

D

¢..O

@ eO..

¢..O

¢D

®

Figure 11. Illustration of Schelkin spiral and its effect on turbulent mixing and flame acceleration.

25
_- 0 0 C_ 0

0

0

cO

0

QO

0

m

0

cO

0

0

T--

0

0

o o o o o o o o

I I

oJnssoJdpoonPoEI

E

E

Figure 12. Measured pressure waveforms in the initiator tube with Schelkin spiral No. 1.

Indicated flame front velocity was 1,400 m/sec.

26
0 0"_ (:(D _ rE) LO _1" CO CM -_. 0

_- o o o o o o o o o o

oJnssoJdpo0nPoEI

0

m

0"_

0

CO

mw .

0

1

m

r..O

0

0

E

O'3

0

C_I

0

0

C_J

0 0

I I

Figure 13. Measured pressure waveforms in the initiator tube with Schelkin spiral No. 2.

Indicated flame front velocity was 2,400 m/sec.

27
m

m

e--

O

.m

0

0

ID._

m

cO0 O0 '-_

i_ _ ._._

_co

I--

Figure 14.

E

C.9

0

E

0

E

0

E

C.9

0

E

C.9

0

E

C.9

0

_E_

E

0

E

C.9

Od

m

_m

E_

_o_

L.A _..I

0 0

0

Schematic of PDRE-G-1 research engine on the test stand.

Sensor port locations and indices are indicated.

28
o

o

o

LO

CM

oo

CM

0

LO

g

E

i=

Figure 15. Measured pressure waveforms in the primary tube with smooth bore.

Indicated flame front velocity was 505 m/sec.

29
CM

O

CM

-'rI

E

.m

W

I._ 0 I._ 0 I._ 0 I._ 0 I._

(tu|e) eJnsseJa

0

0

0

C"3

0

0

C"3

0

CM

oo

E

0

0

0

0

0

I

0

Figure 16. Measured pressure waveforms in the primary tube with Schelkin spiral installed

along its entire length. Indicated flame front velocity was 2,200 m/sec.

3O
CM

O

CM

-1-

I

L.-

E.m

L._

L.LI

0 LO 0 LO 0 LO 0 LO 0

(rule) aJnssaJd

o

o

o

o

o

o

co

o

o

co

C_l

=L

o

o.E_

o

o

o

o

o

o

o I

Figure 17. Measured pressure waveform distribution in the primary tube with Schelkin spiral

installed along its entire length.

31
¢,g)

>

m Err-

_iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiii

iiiiiiiiiiiiiiiiiiiiiiiTiiiiiiiiiiiiiiiii

iiiiiiiiiiiiiiiililiiiiilii[iiiiiiiiiiii

¢.-

¢,g)

Id- m

(,O

,o

t--

._o

_gt'_ .__

_iiiiiiiiiiiiiiiiiiiiiiiiiiiiiii

iiiiiiiiiiiiiiiiiiiiiiiiiiiiiii!i!_i_i_i_i_i_i_

iiiiiiiiiiiiiiiiiiiiAiiiiiiiiiiiiiiiiiiii

¢,g)

g== .m ..-.i

_cr"

iiiiiiiiiiiiiii_iii iiiililiiiiiiiiiiiii

iiiiiiiiiiii_i_iiiiiiiiiii

>

_A

::::::_::::

iiiiiiiiiiiiiiititiiiiiiiiiiiiI_iiiiiiiiiiiii

iiiiiiiiiiii_iii_iiiiiiiiiii

Figure 18. Idealized PDRE cycle processes: (1) Detonation initiation, (2) propagation of detonation wave

through chamber, (3) propagation of rarefaction wave from tube exit to tube head, (4) tube blowdown,

(5) chamber exhausted to atmospheric pressure, and (6) injection of fresh propellant mixture.

32
\

\

\

\

.i

aO!l

Figure 19. Idealized position-time wave diagram for PDRE blowdown process.

33
,--..I

II

z m_

t.a

-- i

Gk_

E

,--..I

II

t-a

aJnssaJ¢lPeaH aqn.I.

Figure 20. Idealized PDRE tube head pressure for a simple straight tube configuration.

34
Figure21. Longexposurephotographof theexhaustplumefromthePDRE-G-1

engineillustratingthefiring characteristicsfor afilledtubecondition.

Obtainedduringanightfiring withhydrogen/oxygenpropellants.

35
Figure22. Longexposurephotographof theexhaustplumefromthePDRE-G-1engine

illustratingthefiring characteristicsfor anunderfilledtubecondition.

Obtainedduringanightfiring withhydrogen/oxygenpropellants.

36
t'-" m

0 0 0 0 0 0 0

LO 0 LO 0 LO 0 LO

CO CO CM CM _

(Jql) lsnJq.I.

0

N

L.('3

CO

0

CO

Figure 23. Time-resolved PDRE-G-1 engine thrust profiles for filled

and underfilled tube conditions.

37
REPORT DOCUMENTATION PAGE FormApprovod

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1. AGENCY USE ONLY (Leave Blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 2001 Technical Publication

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Development of a Gas-Fed Pulse Detonation Research Engine

6. AUTHORS

R.J. Litchford

7. PERFORMING ORGANIZATION NAMES(S) AND ADDRESS(ES)

George C. Marshall Space Flight Center

Marshall Space Flight Center, AL 35812

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

8. PERFORMING ORGANIZATION

REPORT NUMBER

M-1036

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/TP--2001-211412

11. SUPPLEMENTARY NOTES

Prepared by the Advanced Space Transportation Program, Space Transportation Directorate

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Category 20

Standard Distribution

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

In response to the growing need for empirical data on pulse detonation engine performance and operation, NASA Marshall Space

Flight Center has developed and placed into operation a low-cost gas-fed pulse detonation research engine. The guiding design

strategy was to achieve a simple and flexible research apparatus, which was inexpensive to build and operate. As such, the engine

was designed to operate as a heat sink device, and testing was limited to burst-mode operation with run durations of a few

seconds. Wherever possible, maximum use was made of standard off-the-shelf industrial or automotive components. The 5-cmdiameter primary tube is about 90-cm long and has been outfitted with a multitude of sensor and optical ports. The primary tube is

fed by a coaxial injector through an initiator tube, which is inserted directly into the injector head face. Four auxiliary coaxial

injectors are also integrated into the injector head assembly. All propellant flow is controlled with industrial solenoid valves. An

automotive electronic ignition system was adapted for use, and spark plugs are motmted in both tubes so that a variety of ignition

schemes can be examined. A microprocessor-based fiber-optic engine control system was developed to provide precise control

over valve and ignition timing. Initial shakedown testing with hydrogen/oxygen mixtures verified the need for Schelkin spirals in

both the initiator and primary tubes to ensure rapid development of the detonation wave. Measured pressure wave time-of-flight

indicated detonation velocities of 2.4 km/sec and 2.2 km/sec in the initiator and primary tubes, respectively. These values implied

a fuel-lean mixture corresponding to an H 2 volume fraction near 0.5. The axial distribution for the detonation velocity was fotmd

to be essentially constant along the primary tube. Time-resolved thrust profiles were also acquired for both tmderfilled and

overfilled tube conditions. These profiles are consistent with previous time-resolved measurements on single-cycle tubes where

the thrust is found to peak as the detonation wave exits the tube, and decay as the tube blows down.

14. SUBJECT TERMS

pulse detonation engine, combustion, propulsion

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48

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