NASA/TM--2002-211888

/, :._7 NASA \,o \

__(ICOMPX?)

Multiple-Cycle Simulation

a Pulse Detonation Engine

of

Ejector

S. Yungster

Institute for Computational Mechanics in Propulsion, Cleveland, Ohio

H.D. Perkins

Glenn Research Center, Cleveland, Ohio

ICOMP-2002-05

AIAA-2002-3630

October 2002
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NASA/TM--2002-211888

, 7f- ---'4,_;_

/ >_7 NASA \,o \

__(ICOMPX?)

Multiple-Cycle Simulation

a Pulse Detonation Engine

of

Ejector

S. Yungster

Institute for Computational Mechanics in Propulsion, Cleveland, Ohio

H.D. Perkins

Glenn Research Center, Cleveland, Ohio

ICOMP-2002-05

AIAA-2002-3630

Prepared for the

38th Joint Propulsion Conference and Exhibit

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Indianapolis, Indiana, July 7-10, 2002

National Aeronautics and

Space Administration

Glenn Research Center

October 2002
The Aerospace Propulsion and Power Program at

NASA Glenn Research Center sponsored this work.

NASA Center for Aerospace Information

7121 Standard Drive

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Available from

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Available electronically at _sa. _ov
MULTIPLE CYCLE SIMULATION OF A

PULSE DETONATION ENGINE EJECTOR

S. Yungster

Institute for Computational Mechanics in Propulsion

Brook Park, Ohio 44142

H.D. Perkins

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

This paper presents the results of a study involving single

and multiple-cycle numerical simulations of various

PDE-ejector configurations utilizing hydrogen-oxygen mixtures. The objective was to investigate the thrust, impulse

and mass flow rate characteristics of these devices. The

results indicate that ejector systems can utilize the energy

stored in the strong shock wave exiting the detonation tube to

augment the impulse obtained from the detonation tube

alone. Impulse augmentation ratios of up to 1.9 were

achieved. The axial location of the converging-diverging

ejectors relative to the end of the detonation tube were shown

to affect the performance of the system.

Introduction

Ejectors are thrust augmentation devices whose performance depends on the efficient energy transfer between the

primary and secondary flows. Most of the past research on

ejectors has focused on steady primary flows, however, some

experimental studies have shown that the introduction of

unsteadiness into the primary flow can enhance the energy

transfer process and consequently improve the thrust

augmentation 1. The reason for the improvement in performance is that, while a steady ejector relies on viscous shear

mixing for the energy transfer, the unsteady ejector achieves

part of the energy transfer through more efficient flow

entrainment mechanisms that are essentially inviscid. In

addition, unsteady ejectors generally require shorter lengths,

as compared to steady ejectors, for completing the energy

transfer, and therefore have the potential to achieve a more

efficient structural design.

It is important to point out that two types of ejector augmentation systems have been considered in the past. They

are based on different physical principles to obtain thrust

augmentation, and are effective at different flight speeds.

The low-speed ejector systems rely on the low pressures

generated by the secondary flow accelerating around the

inlet contour 2. The thrust augmentation is primarily a result

of suction forces on the leading edge of the ejector shroud.

Such ejector systems, which include those investigated by

Lockwood 1, generate thrust gains at static conditions. However, they rapidly lose their thrust augmentation ability with

forward flight. The reason being that, with forward flight,

the secondary flow is no longer accelerated around the inlet

leading edge. This causes a decrease in the leading edge suction force and a reduction in thrust augmentation. The

changes in the secondary flow with flight speed may actually

produce thrust losses at cruise conditions 2.

The second type of thrust augmentation ejector system is

the ejector-ramjet, which is most effective from high subsonic to low supersonic speeds 3. Two classes of ejector-ramjet systems have been proposed; the simultaneous mixing

and combustion cycle (SMC), and the closely related independent ramjet stream cycle (IRS) recently proposed by

Trefny and Yungster 4'5.

In the SMC cycle, exhaust from a primary fuel-rich

rocket provides some fraction of the fuel required for combustion with the entrained secondary airflow. The rocket and

air streams mix and burn simultaneously. This process generally results in thermal choking where mixing is complete,

followed by expansion through a nozzle. However, the

requirement for complete mixing of the rocket and air

streams may result in very long mixing/combustor ducts.

In the IRS cycle complete mixing of the rocket and ramjet streams is not required. In this cycle, the airstream is

fueled independently using fuel injectors located upstream in

the inlet. The rocket serves as a pilot for the fueled airstream.

The IRS cycle has several potential advantages over the

SMC cycle which are described in ref. 5.

Numerous methods for introducing unsteadiness into the

primary flow have been proposed, including the Spin-Jet,

Oscillating-Primary-Jet and Pulse-Jet ejectors 1. In recent

years, the Pulse-Detonation-Engine (PDE) has been recognized as a promising propulsion system that offers advantages in thermodynamic cycle efficiency and hardware

NASA/TM--2002-211888 1
simplicity6'7.SincePDEsarehighlyunsteadydevicesthere

isconsiderable interest ininvestigating theirperformance in

anejectorconfiguration.

Theobjectiveofthispaperistopresent aninitialstudy

oftheperformance ofseveral PDEejectorconfigurations at

staticonditions, andforsingleandmultiplecycles.

Numerical Method

The analysis was carried out using an in-house developed time-accurate CFD code 8. The code solves the axisymmetric Navier-Stokes equations for a nonequilibrium mixture

of thermally perfect gases, using an implicit, total variation

diminishing (TVD) algorithm.

Since the main flow entrainment mechanisms in

unsteady ejectors are essentially inviscid, we neglect the viscous terms, and therefore, use the Euler equations with finite

rate chemistry. In our formulation, the global continuity

equation is replaced by ns species conservation equations,

where ns denotes the number of species.

The numerical method used for solving the governing

equations is described in detail in Ref. 8, and briefly summarized here. The equation set is solved using a fully implicit,

first-order-accurate in time, variable-step backward differentiation formula (BDF) method. The numerical fluxes are

evaluated using a second-order spatially accurate TVD

scheme. The resulting equations are then linearized in a conservative manner and solved iteratively, by using a

lower-upper relaxation procedure consisting of successive

Gauss-Seidel (LU-SGS) sweeps.

The chemical reaction mechanism for hydrogen-oxygen

combustion was based on Jachimowski's model (Ref. 9,10),

and consists of 19 elementary reactions among 9 species.

Reactions involving N 2 were neglected.

In order to maintain adequate numerical resolution of the

detonation wave front without the need to use thousands of

grid points, a multi-level, dynamically adaptive grid is utilized. Figure 1 shows a section of the grid at three different

times as the detonation wave moves from left to right. The

grid constantly adapts to keep the detonation front within the

finest grid level. An arbitrary number of levels can be specified. Nine grid levels were used in the present study, and 100

points were included in the finest grid level.

Results

Finite rate chemistry calculations were used to compute

the flow in various PDE ejector configurations. In this paper,

only development of detonations with direct initiation were

considered. A high pressure, high temperature driver gas,

consisting of H2-O 2 equilibrium combustion products, was

used in a small region next to the closed end (head-end) of

the detonation tube, as described in Ref. 10. A driver pressure ratio of 50 was used in the present study.

To verify that this computational approach yields Chapman-Jouguet detonations, the detonation velocity was plotted as a function of time for a stoichiometric H2-O 2 mixture,

at Po 0.4 atm. and To 298 K, and compared the results

with predictions from the CEA equilibrium code of Gordon

and McBride 11. It is observed in figure 2a that after an initial

overshoot during the short transient phase, the detonation

speed reaches a nearly constant value which is in very good

agreement with that predicted by the CEA code. Figure 2b

plots nondimensional pressure and temperature profiles

immediately behind the detonation front, showing that the

von Neumann spike conditions are computed accurately.

A schematic of the PDE-Ejector configuration considered in this study is shown in fig. 3. The boundary conditions

at the ejector inflow plane depend on the direction of the

flow. If the flow was into the ejector, total pressure and temperature, Ptot and Ttot were specified (subsonic inflow boundary). If the flow was out of the ejector, the static pressure was

specified and the remaining variables were extrapolated from

the interior (subsonic outflow boundary). At the ejector exit

plane, the subsonic outflow boundary condition was specified. All calculations considered a stoichiometric H2-O 2

mixture atpo 1.0 atm. and To 298 K, andptot/Po 1.05.

The ambient pressure was set to 1.0 atm. All the calculations

were carried out for a detonation tube diameter, dr, of 2.6 in.

and for ejector dimensions L 1 13 in., and L2 26 in.

The first calculation considered a detonation tube having a length, Lt, of 72 in. The ejector cone angles were 01

10° and 02 = 3.5 °. The location of the ejector throat, Lth, was

set at Lth 2.6 in. downstream of the detonation tube end.

The ejector throat diameter was set at Dth 8.22 in., corresponding to an ejector area ratio, RAt h 10, where RAt h is

defined as

RAth

_ ejector area at throat

detonation tube area

Figure 4 shows the contribution to the thrust force and

impulse from the detonation tube, ejector shroud and the

small base area at the end of the detonation tube, for a single

PDE cycle. The total force and impulse on the PDE-ejector

system is also shown. The contribution of the base area was

always negligible in all cases considered. All forces were

computed by integrating the instantaneous pressure over the

surface area.

The force on the detonation tube (fig. 4a) shows an initial

large value that arises from the ignition mechanism used in

NASA/TM--2002-211888 2
thenumerical simulations. Noattempt tocorrectforthisartificialforcewasmadeinthepresent sudy,sincethefocusis

toinvestigatetherelativethrustaugmentation obtainedinthe

variousconfigurations studied.Thisinitialshortduration

spikeisfollowedbyalongerplateauregionof2.5msdurationandasubsequent decaytozero.It isprimarilyduring

thislevelpressuredurationthatPDEthrustisgenerated. The

ejectorforceplotshowsasharpnegativespikenear0.7ms

thatiscausedbythestrongshockimpingingontheconvergingsectionof theejector.Figure5bshowsthatthetotal

impulseissmallerthanthatobtainedbythePDEtubealone.

Thatis,theejectorisexertingadragforcethatreducesthe

performance ofthesystem.

Thepoorperformanceinthiscaseis theresultof the

strongshockexitingthedetonationtubereflectingfromthe

convergingsectionoftheejectorshroud. Therefore, anobviousimprovement couldbeachievedbyslidingtheejector

upstream. Also,onecouldtakemoreadvantage ofthisstrong

shockbyincreasingthedivergingangle02. The second case

therefore considers the same ejector, but moved upstream

relative to the tube, such that the ejector throat is 1.4 in.

upstream of the end of the detonation tube (that is Lth -1.4)

In addition, the divergence angle was increased to 02 = 10°.

Figure 5 shows the force and impulse results for the second case. These changes had no effect on the detonation tube

force and impulse, however, there was a substantial effect on

the forces acting on the ejector shroud. At 0.7 ms, there is

now a sharp positive spike in the ejector force. The positive

force continues until around 2.2 ms followed by alternating,

smaller negative and positive forces. Figure 5b shows that in

this case there is a significant increase in the total impulse of

the system, (29.96 N-s compared to 15.70 for the PDE tube

alone), corresponding to an impulse augmentation, g_, of

I!/ 1.9.

Additional cases were computed to examine the effect of

ejector area ratio, RAth, on the performance of the PDE-ejector. The results are shown in fig. 6. The delay in the arrival

of the shock front at the ejector shroud for increasing area

ratio can be seen in the first peak in total force between 0.7

and 1.0 ms approximately. This results in an steeper rise in

total impulse initially for the RAt h 10 case. After nearly 8

ms, however, there are small differences in the total impulse

generated. There is no general trend in performance over this

range of area ratios.

The mass flow rate for area ratios of 20 and 30 is shown

in fig. 7. This figure shows that the secondary flow is alternating between positive and negative mass flow rates for

both cases. On average (over the nearly 8 ms of operation)

there is a net positive secondary flow of 1.55 kg/s for fig. 7a,

and 1.17 kg/s for fig. 7b. The average primary mass flow rate

for both cases is 1.36 kg/s.

The next case considers the same detonation tube, but

instead of an ejector shroud, a diverging nozzle, having the

same length (L 2 = 26 in.) and divergence angle (02 = 10°) as

the previous ejector, is attached at the end of the detonation

tube. There is no secondary flow in this case, and only the

constant area tube is fueled. Since the "throat" area is identical to the detonation tube area, this case will be denoted as

having an area ratio RAt h 1.0.

Figure 8 shows the force and impulse for this configuration. The force on the detonation tube remains unchanged

from the previous cases. The force on the nozzle is also plotted. This force is initially zero until the shock wave reaches

the end of the detonation tube. At that time, a sharp rise is

observed. The force on the nozzle peaks at around 1.2 ms

and subsequently decreases and remains negative from

around 1.8 ms until the end of the calculation. Note that the

force on the nozzle is a smooth function of time as opposed

to the force on the ejector in the previous cases which show

periodic oscillations resulting from shock reflections in the

ej ector.

Figure 8b shows that in this case, there is also an increase

in the total impulse over that obtained for the PDE tube

alone. The impulse augmentation, however is g_ 1.4, significantly smaller than that achieved with the ejector systems. The lower performance for this case is a result of the

below ambient pressures acting on the nozzle once the shock

front leaves the nozzle. A comparison of the total impulse

obtained in the present case and in two of the previous ejectors is shown in fig. 9.

Note that other nozzle geometric configurations could

have been considered for comparison with the ejector (for

example, a nozzle having a different divergence angle but the

same exit area as the ejector). Such additional configurations

are currently being investigated.

Experimental studies of the effects of nozzles on the performance of PDEs have been carried out by Daniau et al. 12.

They reported impulse augmentation values ranging from

1.2 to 1.8 for different diverging nozzles.

The results presented so far have considered a single

PDE cycle. However, PDE-ejector systems normally require

several detonation cycles before they reach a "limit cycle"

operation. The final case presented in this paper attempts to

address this issue by carrying out a multi-cycle computation

of a PDE-ejector system.

In order to reduce the computational time, a shorter tube

(L t 39.4 in.) was considered, and the ejector convergence

angle was reduced to 01 = 5°. An ejector with an area ratio of

10 was considered.

NASA/TM--2002-211888 3
Resultsofthiscalculationarepresentedinfig.10,which

showsthetemporal evolutionofthePDEejectorflowfield

duringalmostthreecompletecycles.Thedetonationisinitiatedatt 0.0ms,andpropagates downstreamuntilit reaches

theendof thedetonationtubeat approximately 0.35ms.

Subsequently, a strongshockfollowedbythecombustion

productsexpandintotheejector. At4.52msafreshcombustiblemixtureis introducedintothetube(apureoxygen

bufferzoneisusedtoseparatethehotcombustionproducts

fromthefreshcombustible mixture).At7.63msthedetonationtubeiscompletelyfilledwiththenewcombustible mixture,andat8.21mstheseconddetonationcycleisstarted.

Subsequent times(t 8.42mstot 15.31ms)showthesame

sequenceof eventsdescribedfor the first cycle.At

t 16.37ms,thethirddetonationcyclehasbeenstarted, and

theremainingfigureshowthesubsequent propagationf

thedetonationwave.

Thethrustforcesoverthe5cyclesareshowninfig.11,

andtheimpulseandmassflowratesareshownin fig.12.

Whiletheoscillatorypatternissimilarfromcycletocycle,

somedifferences areclearlyobserved, suchasthepeakvaluesintheejectorshroudforce.Morecyclesmayberequired

toestablishsomekindoflimitcycle.Theimpulseplotshows

that,afterthefirstcycle,theejectoraugmentation issmaller.

Afterfivecyclestheimpulseaugmentation was_ 1.7.The

massflowrateplotshowsthesamealternatingbetween

inflowandoutflowforthesecondarystream.Theaverage

secondarymassflowratevariesfromcycleto cyclebutis

alwayspositive, asshownintable1.Notethathemassflow

ratesinthefirstcyclearesubstantially different fromtheothers.Thisisduetothefacthathefirstcycledoesnotinclude

thefillingprocess.Thefirstcyclestartedwiththetube

alreadyfilledwiththedetonablemixture.

Conclusions

Thereis a significant amountof energystoredin the

strongshockwaveexitingthedetonationtube.If noejector

ornozzleisaddedattheendofthetube,thisenergywillbe

simplydissipatedintothesurroundingair.By addingan

ejector(oranozzle)someofthisenergycanbeutilizedfor

theproductionofthrust.

Thepresentcomputations idicatethata PDE-ejector

configuration produces higher impulse than a PDE-tube-nozzle combination having the same length and divergence

angle. The higher performance of the PDE-ejector is partly

due to its capacity to entrain secondary air, which prevents

the sub-ambient pressures that develop in the PDE-tube-nozzle system once the shock wave exits the nozzle. Impulse

augmentations of 1.9 and 1.4 were obtained for the

PDE-ejector and PDE-tube-nozzle configurations respectively.

The axial location of the ejector shroud relative to the

end of the detonation tube is an important parameter. The

ejector shroud should be placed in a location such that the

shock wave exiting the detonation tube impinges on the

diverging section of the ejector shroud. The ejector area ratio

had a small effect on the performance of the PDE over the

range investigated in this study (10 < Rat h < 30).

The multi-cycle PDE-ejector calculation showed that

after 5 cycles, an impulse augmentation factor of 1.7 was

achieved. The average secondary mass flow rate remained

positive for each cycle.

References

1. Porter, J.L. and Squyers, R.A., "A Summary/Overview of Ejector Augmentor Theory and Performance," ATC

Report No. R-91100/9CR-47A, Sept. 1979.

2. Presz, W., Reynolds, G. and Hunter, C., "Thrust Augmentation with Mixer/Ejector Systems," AIAA paper

2002-0230, January 2002.

3. Heiser, W.H. and Pratt, D.T., Hypersonic Airbreathing Propulsion, pp. 446-472, American Institute of Aeronautics and Astronautics, Washington, D.C, 1994.

4. Trefny, C.J., "An Air-Breathing Launch Vehicle Concept for Single-Stage-to-Orbit," AIAA Paper 99-2730, 1999

5. Yungster, S. and Trefny, C.J., "Analysis of a New

Rocket-Based Combined-Cycle Engine Concept at Low

Speed," AIAA paper 99-2393, June 1999.

6. Heiser, W.H. and Pratt, D.T., "Thermodynamic Cycle

Analysis of Pulse Detonation Engines", Jrn. of Propulsion

and Power, Vol. 18, No. 1, January-February 2002.

7. Kailasanath, K., "Recent Developments in the

Research on Pulse Detonation Engines," AIAA paper

2002-0470, January 2002.

8. Yungster, S. and Radhakrishnan, K., ''A Fully Implicit

Time Accurate Method for Hypersonic Combustion: Application to Shock-Induced Combustion Instability," Shock

Waves, Vol. 5, 1996, pp. 293-303.

9. Jachimowski, C.J., "An Analytical Study of the

Hydrogen-Air Reaction Mechanism with Application to

Scramjet Combustion," NASA TP-2791, Feb. 1988.

10. Yungster, S. and Radhakrishnan, K., "Computational Study of Near-Limit Propagation of Detonation in

Hydrogen-Air Mixtures," AIAA paper 2002-3712, July

2002.

11. McBride, B.J. and Gordon, S., "Computer Program

for Calculation of Complex Chemical Equilibrium Compositions and Applications. II. Users Manual and Program

NASA/TM--2002-211888 4
Description," NASA RP- 1311, 1996

12. Daniau, E., Zitoun, R., Couquet, C. and Desbordes,

D., "Effects of Nozzles of Different Length and Shape on the

Propulsion Performance of Pulsed Detonation Engines," in

High-Speed Deflagration and Detonation, Eds. G.D. Roy,

S.M. Frolov, D. Netzer and A. Borisov. Moscow, 2001:

ELEX-KM Publ. pp 251-262.

Table 1: Average mass flow rates (kg/s)

Cycle Secondary mass flow Primary mass flow

rate rate

1a 0.82 0.76

2 2.44 1.20

3 4.28 1.27

4 2.18 1.21

5 2.94 1.16

aDifferent starting condition

tl

t2

4500

4000

3500

g 3000

& 2500

._ 2000

1500

o

1000

5OO

G°(rd _nAM_jbe;ide

0

0 2 4 6 8

Time (#s)

(a)

10 12

40

35

30

25

_- 20

15

10

(b)

Yon Neumann spike (cea code)

5

, i

.02 3.04

, i

3.06

x (cm)

m i

3.08 3.10

Fig. 2. (a) Detonation speed as a function of time and

(b) nondimensional pressure and temperature behind detonation front. H2-O2;Po = 0.4 atm, T o = 298.0 K, 0=1.0

t3

Fig. 1. Computational grid at three different times.

NASAfFM--2002-211888 5
Detonation tube

Ej ector inflow plane Ejector shroud

Exit plane

gt

@ 2

L1 L2

Fig. 3. Schematic of PDE-Ejector configuration. (Only every third grid point is shown)

(a)

2

20.0

10.0

0.0

-10.0

Detonation tube 1

---- Base

Ejector shroud

0.0

(b) 40.0

30.0

?" 20.0

z

E 10.0

0.0

_ Detonation tube

Base

Ejector shroud " •

Total

 IIIZII

-10.0

2.0 4.0 6.0 8.0 10.0 0.0 2.0 4.0 6.0 8.0

Time (ms) Time (ms)

10.0

Fig. 4. Force (a) and impulse (b) as a function of time for a PDE-ejector. Lth = 2.6", @2 = 3.5°, RAt h = 10

NASA/TM--2002-211888 6
LL

(a)

20.0

10.0

0.0

-10.0

- - - Detonation tube7

"_ 11 Base ]

" _ _ Ejector shroud /

0.0

40.0

30.0

_" 20.0

z

E lO.O

o.o

-lO.O

i

[

.... 2...... 2 ....... L ..... [

2.0 4.0 6.0 8.0 10.0 0.0 2.0 4.0 6.0 8.0

Time (ms) (b) Time (ms)

10.0

Fig. 5. Force (a) and impulse (b) as a function of time for a PDE-ejector. Lth = -1.4", O2 = 10°, RAt h = 10

2

P_

(a)

20.0

10.0

0.0

-10.0

RAth

_. 1,,1i --.20

i\' ^%:' :7',..%. .,,:.,

i N,q2',J_,-i_.fl', _1 :'-!_

i " :: i&l -

!

0.0

40.0

30.0 ............................ _ [i[i_-_-_- _-[-[ [ .................. 22:-; _.-=-.__:--------_-i_i-_2_._£ _---

_" es ! ! !

z 20.0 ..................... -_-----------------------

/o : : •

E

10.0

P_

0.0 ---

-10.0

2.0 4.0 6.0 8.0 0.0 2.0 4.0 6.0

Time (ms) (b) Time (ms)

8.0

Fig. 6. Total force (a) and impulse (b) as a function of time for three PDE-ejectors. Lth = -1.4", O2 = 10°

rY

LL

(a)

160.0

120.0

80.0

40.0

0.0

-40.0

r _ ' -- Secondary stream

i _ ! ---- Primary stream

Exit plane

.......................... _...._._.................................. _...............

/

r fi

i

I RAt h = 20

i

......................... I.......:.:T ...................................................................................................

i i

i i

i i

_-_

0.0

rY

LL

2.0 4.0 6.0 8.0

Time (ms) (b)

160.0

120.0

80.0

40.0

0.0

-40.0

0.0

I',! !

'i -- Secondary stream

_r _'i Primary stream

i. Exit plane

........................... F{ .................................. '...............

i

i

I ! RAthi=30,

I i i

.......................... T'""_'"7 ................................................................ _.................................

I i i

, i i

, i i

, i i

J i i

2.0 4.0 6.0 8.0

Time (ms)

Fig. 7. Mass flow rate as a function of time for two PDE-ejectors. Lth = -1.4", 0 2 = 10°

NASA/TM--2002-211888 7
20.0

v

2

10.0

0.0

-10.0

' ' i '

A [_ Z_ Detonation tube

"_ //// / --" Base

_// /_ /- Nozzle

V/ /\ |. Total

_- L2_

(a)

?

z

E

0.0 2.0 4.0 6.0

Time (ms)

40.0 i

- - - Detonation tube

Base

30.0 _ Nozzle

Total

20.0

10.0

0.0

-10.0

0.0 2.0 4.0 6.0

(b) Time (ms)

Fig. 8. Force (a) and impulse (b) as a function of time for

a PDE with a diverging nozzle (Rat h = 1, L2 26", e2 = 10°).

40.0

?-

Q_

E

30.0

20.0

10.0

0.0

-10.0

RAth

0.0 2.0 4.0 6.0

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Fig. 9. Total impulse as a function of time for three PDE

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Fig. 10. Multi-cycle simulation showing nondimensional

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atm.; To=298 K; 0=1.0; Ptot=l.05 atm.

NASA/TM--2002-211888 8
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Form Approved REPORT DOCUMENTATION PAGE

OMB No. 0704-0188

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

October 2002 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Multiple-Cycle Simulation of a Pulse Detonation Engine Ejector

6. AUTHOR(S)

S. Yungster and H.D. Perkins

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

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

DlJ-708-48-13-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-13570

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--2002-211888

ICOMP-2002-05

AIAA-2002-3630

11. SUPPLEMENTARY NOTES

Prepared for the 38th Joint Propulsion Conference and Exhibit, cosponsored by the AIAA, ASME, SAE, and ASEE,

Indianapolis, Indiana, July 7-10, 2002. S. Yungster, Institute for Computational Mechanics in Propulsion, Brook Park,

Ohio 44142; and H.D. Perkins, NASA Glenn Research Center. Responsible person, Charles Trefney, Turbomachinery

and Propulsion Systems Division, NASA Glenn Research Center, organization code 5880, 216-433-2162.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 07 Distribution: Nonstandard

Available electronically at http://gltrs.grc.na.sa.gov

This publication is available from the NASA Center for AeroSpace Information, 301-621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

This paper presents the results of a study involving single and multiple-cycle numerical simulations of various PDEejector configurations utilizing hydrogen-oxygen mixtures. The objective was to investigate the thrust, impulse and mass

flow rate characteristics of these devices. The results indicate that ejector systems can utilize the energy stored in the

strong shock wave exiting the detonation tube to augment the impulse obtained from the detonation tube alone. Impulse

augmentation ratios of up to 1.9 were achieved. The axial location of the converging-diverging ejectors relative to the end

of the detonation tube were shownto affect the performance of the system.

14. SUBJECT TERMS

Pulsed detonation engine; Unsteady reacting flows; Computational fluid dynamics

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

16

16. PRICE CODE

18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT

OF THIS PAGE OF ABSTRACT

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