Louis A. Povinelli

Glenn Research Center, Cleveland, Ohio

Shaye Yungster

Institute for Computational Mechanics in Propulsion, Cleveland, Ohio

Real Gas Effects on the Performance

of Hydrocarbon-Fueled Pulse

Detonation Engines

NASA/TM—2003-212211

March 2003

ICOMP–2003–02

AIAA–2003–0712

NI STITUTE FOR COMPUTATIONAL MECHAN CI S

RP NI OPULS OI N

Glenn Research Center

NASA

Ohio Aerospace

Institute

https://ntrs.nasa.gov/search.jsp?R=20030032940 2018-05-15T15:57:57+00:00Z
The NASA STI Program Office . . . in Profile

Since its founding, NASA has been dedicated to

the advancement of aeronautics and space

science. The NASA Scientific and Technical

Information (STI) Program Office plays a key part

in helping NASA maintain this important role.

The NASA STI Program Office is operated by

Langley Research Center, the Lead Center for

NASA’s scientific and technical information. The

NASA STI Program Office provides access to the

NASA STI Database, the largest collection of

aeronautical and space science STI in the world.

The Program Office is also NASA’s institutional

mechanism for disseminating the results of its

research and development activities. These results

are published by NASA in the NASA STI Report

Series, which includes the following report types:

• TECHNICAL PUBLICATION. Reports of

completed research or a major significant

phase of research that present the results of

NASA programs and include extensive data

or theoretical analysis. Includes compilations

of significant scientific and technical data and

information deemed to be of continuing

reference value. NASA’s counterpart of peerreviewed formal professional papers but

has less stringent limitations on manuscript

length and extent of graphic presentations.

• TECHNICAL MEMORANDUM. Scientific

and technical findings that are preliminary or

of specialized interest, e.g., quick release

reports, working papers, and bibliographies

that contain minimal annotation. Does not

contain extensive analysis.

• CONTRACTOR REPORT. Scientific and

technical findings by NASA-sponsored

contractors and grantees.

• CONFERENCE PUBLICATION. Collected

papers from scientific and technical

conferences, symposia, seminars, or other

meetings sponsored or cosponsored by

NASA.

• SPECIAL PUBLICATION. Scientific,

technical, or historical information from

NASA programs, projects, and missions,

often concerned with subjects having

substantial public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific

and technical material pertinent to NASA’s

mission.

Specialized services that complement the STI

Program Office’s diverse offerings include

creating custom thesauri, building customized

databases, organizing and publishing research

results . . . even providing videos.

For more information about the NASA STI

Program Office, see the following:

• Access the NASA STI Program Home Page

at http://www.sti.nasa.gov

• E-mail your question via the Internet to

help@sti.nasa.gov

• Fax your question to the NASA Access

Help Desk at 301–621–0134

• Telephone the NASA Access Help Desk at

301–621–0390

• Write to:

 NASA Access Help Desk

 NASA Center for AeroSpace Information

 7121 Standard Drive

 Hanover, MD 21076
Louis A. Povinelli

Glenn Research Center, Cleveland, Ohio

Shaye Yungster

Institute for Computational Mechanics in Propulsion, Cleveland, Ohio

Real Gas Effects on the Performance

of Hydrocarbon-Fueled Pulse

Detonation Engines

NASA/TM—2003-212211

March 2003

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

41st Aerospace Sciences Meeting and Exhibit

sponsored by the American Institute of Aeronautics and Astronautics

Reno, Nevada, January 6–9, 2003

NI STITUTE FOR COMPUTATIONAL MECHAN CI S

RP NI OPULS OI N

Glenn Research Center

NASA

Ohio Aerospace

Institute

ICOMP–2003–02

AIAA–2003–0712
Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Available electronically at http://gltrs.grc.nasa.gov
NASA/TMó2003-212211 1 

REAL GAS EFFECTS ON THE PERFORMANCE OF HYDROCARBONFUELED PULSE DETONATION ENGINES

Louis A. Povinelli*

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Shaye YungsterÜ

Institute for Computational Mechanics in Propulsion 

Cleveland, Ohio 44135 

 

ABSTRACT 

This paper presents results for a single-pulse 

detonation tube wherein the effects of high 

temperature dissociation and the subsequent 

recombination influence the sensible heat 

release available for providing propulsive 

thrust. The study involved the use of ethylene 

and air at equivalence ratios of 0.7 and 1.0. 

The real gas effects on the sensible heat 

release were found to be significantly large so 

as to have an impact on the thrust, impulse 

and fuel consumption of a PDE. 

INTRODUCTION 

In our previous publication (ref. 1) it was 

shown that the high detonation temperatures 

associated with hydrocarbon-air and 

hydrogen-air mixtures caused a substantial 

amount of dissociation, and was accompanied 

by high concentrations of intermediate 

species. This process was accompanied by 

approximately an 11% decrease in the amount 

of energy available for thrust relative to a 

ramjet. The consequence of this loss was to 

reduce the thermal efficiency of the PDE 

below the value of the Brayton cycle 

efficiency at low values of inlet (ram) air 

temperature ratios. 

*

Chief Scientist, Turbomachinery and Propulsion Systems Division, 

Fellow AIAA. 

Ü

Senior Research Associate, Ohio Aerospace Institute, Senior 

Member AIAA. 

 The effect of stoichiometric propane-air 

dissociation loss on the performance 

parameters of a PDE was calculated as a 

function of temperature ratio as well as Mach 

number (0 to 5) in reference 2. The conditions 

investigated were for an ideal PDE (isentropic 

inlet and nozzle) at an altitude of 33,000 ft 

with inlet pressure of 3.8 psia and temperature 

of 400 °R . In these calculations, both the PDE 

and the Brayton (gas turbine) cycle had the 

same ram compression, but the Brayton had 

additional compression from a mechanical 

compressor. Therefore, at a given flight Mach 

number, the PDE has a lower temperature at 

the detonation chamber entrance than does the 

turbojet. 

The thermodynamic cycle calculations in 

reference 2 show that (a) the PDE had higher 

specific thrust than a ramjet up to Mach 2.3 

and (b) the PDE had higher performance than 

a turbojet over a wider Mach range, provided 

that the mechanical compression ratio in the 

turbojet was less than 4. The advantage of the 

PDE disappears as the mechanical 

compression ratio of the turbojet exceeds 4. 

Some advantages of a combined PDE/Ramjet 

cycle where the PDE provides static thrust and 

acceleration with ramjet takeover at Mach 2.5 

to 3 and scramjet operation at Mach 5 have 

been discussed in reference 2. 
NASA/TMó2003-212211 2 

Since the products of the detonation process 

persist in the chamber for a short, but finite 

length of time prior to flowing out through the 

exhaust, we studied the effect of 

recombination on the sensible heat release 

(ref. 3). Unsteady, axisymmetric finite-rate 

chemistry computations were carried out for 

an open ended tube filled with hydrogen-air 

mixtures at standard conditions. Hydrogen-air 

mixtures were used due to the complexities of 

finite rate chemistry modeling for 

hydrocarbon-air mixtures and the lengthy 

computational times required. It was found 

that the recombination that occurred in the 

tube reduced the loss of sensible heat. At an 

equivalence ratio of 1.0, the sensible heat 

release loss with recombination was 

10.8 percent relative to the heating value of 

the mixture, whereas, previously, the loss was 

16.7 percent when chemical equilibrium was 

assumed. At an equivalence ratio of 0.6, 

dissociation effects reduced the sensible heat 

release by 4.7 percent when equilibrium was 

assumed, and recombination reduced the loss 

to 0.6 percent. The study (ref. 3) also showed 

that the fuel specific impulse with an 

equivalence ratio of 0.6 was higher than that at 

an equivalence ratio of 1.0, although the thrust 

at 0.6 was lower. 

In a recent paper (ref. 4) the impulse results of 

the analysis described above for hydrogen-air 

were compared to the experimental data 

obtained at the Air Force Research Labs 

(ref. 5). The AFRL data, obtained with an 

opened-ended detonation tube, were shown to 

be in excellent agreement with our unsteady 

CFD results over a range of equivalence 

ratios. In addition, a thermodynamic cycle 

calculation, with isentropic inlet and exhaust, 

yielded values that were 200 to 300 seconds of 

impulse, which may reflect the potential 

improvement associated with the addition of 

an exhaust nozzle (ref. 4). 

Based on the results obtained with hydrogenair, an estimate was made of the effect of the 

recombination effect that would occur in 

propane-air detonation (ref. 6). Since the 

dissociation losses reduced the sensible heat 

release by 16.7 percent relative to the heating 

value with equilibrium chemistry and 

recombination reduced the loss to 10.8, a 

decrement of 5 percent was chosen as the heat 

loss relative to the lower heating value of the 

propane-air mixture. The results (ref. 6) 

showed that the performance of the PDE 

improved relative to the Brayton cycle. The 

PDE specific thrust was comparable to the 

ramjet performance (with no recombination) 

out to a Mach number of approximately 3.0. 

Our work in reference 6 was only an 

approximation of the recombination based on 

hydrogen-air chemistry. In the current paper, 

the degree of recombination that occurs in 

hydrocarbon-air mixtures is computed using 

the unsteady, axisymmetric, finite rate method 

described above (ref. 3). 

PHYSICAL MODEL AND 

NUMERICAL METHOD 

Geometry and Test Gas

Single-pulse computations were performed for 

a constant area detonation tube having a 

length of 1.0 meter and a diameter of 

6.6 centimeters. The tube is closed at the head 

end and the other is open. Ethylene and air 

having an equivalence ratio of 1.0 and 0.7, 

with a temperature of 298 K were used as the 

reactants. Calculations were performed for 

zero forward flight velocity and at a pressure 

of 1 bar. At the start of the computations, the 

entire tube is filled with the ethylene-air 

mixture. The ambient pressure outside the 

detonation tube is set at 1 bar. 
NASA/TMó2003-212211 3 

To initiate the detonation, a small region of 

high temperature and high pressure nitrogen 

gas was used. For an equivalence ratio of 1.0, 

the ignition region was 0.5 centimeters, the 

temperature was 3000 K and the pressure was 

45 bar. These values are comparable to those 

used by Kailasanath (ref. 12). It is noted that 

these values are higher than we used 

previously with hydrogen-air mixtures (ref. 3), 

which reflects the higher ignition energy 

requirements for ethylene-air mixtures. For an 

equivalence ratio of 0.7, the ignition region 

was 1.0 centimeter, the temperature was 

3000 K and the pressure was 50 bar. In this 

direct initiation method, the shock wave 

generated at the interface transitions into a 

Chapman-Jouget detonation wave, as will be 

shown later. 

Computational Method

Single-pulse finite rate calculations were used 

to compute the species evolution and resulting 

thrust and impulse. The analysis was carried 

out using an in-house developed time-accurate 

CFD code (refs. 7 and 8). The code solves the 

axisymmetric Navier-Stokes equations for a 

non-equilibrium mixture of thermally perfect 

gases, using an implicit, total variation 

diminishing (TVD) scheme. It includes a 

generalized detailed chemistry capability, 

various options for turbulence models, and 

steady state or time-accurate algorithms. In the 

present study, as in reference 3, we excluded 

viscous effects, thereby using the Euler 

equations. 

The numerical method used for solving the 

governing set of equations is described in 

detail in reference 7 and is briefly described 

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 used for 

ethylene-air combustion was based on the 

short mechanism of Li, Varatharajan and 

Williams (ref. 9) and consists of 36 reactions 

among 20 species, with nitrogen being treated 

as an inert, i.e., non reacting species. This 

chemical modeling was designed to be applied 

with reasonable accuracy over the range of 

post-shock temperatures between 1000 K and 

2500 K, pressures between 0.5 and 100 bar 

and equivalence ratios between 0.5 and 3. The 

short mechanism was shown to retain the 

essence of the chemical pathways and to yield 

reasonable accuracy with the detailed 

mechanism of reference 9. 

The inversion of large matrices is avoided by 

partitioning the system into reacting and nonreacting parts. Therefore, the matrices 

requiring inversion are of the same size as 

those in the point implicit methods. The 

significant advantage of the present method, 

because it is fully implicit, lies in the fact that 

it is stable for large CFL numbers, thereby 

enabling the use of relatively large time steps 

to minimize computational cost. 

In order to maintain high resolution of the 

wave front at all times, without the use of 

thousands of grid points, a multi-level, 

dynamically adaptive grid was implemented in 

which a very fine subgrid continuously slides 

along with the detonation wave front 

(ref. 7).The minimum and maximum grid 

spacings used in this study were 0.00195 and 

1.0 cm. 

RESULTS 

Initiation

Finite rate chemistry computations were 

performed for equivalence ratios of 0.7 and 

1.0. As shown previously in our computations 
NASA/TMó2003-212211 4 

(ref. 3), a high pressure zone is observed at the 

closed end of the detonation tube. At the 

interface of this region and the reactants, an 

expansion wave moves toward the head end 

and is reflected. At the same time a 

compression wave moves toward the open end 

of the tube which compresses and heats the 

ethylene-air mixture. Following a short 

induction period, combustion occurs at the 

interface boundary and shock waves propagate 

upstream and downstream. The compression 

waves overtake the shock and cause 

acceleration. The shock front and the 

combustion front couple and form a 

detonation wave. 

Detonation Velocity

The computed detonation velocities for 

equivalence ratios of 0.7 and 1.0 are shown as 

a function of time in figure 1. The velocities 

are compared to the predictions from the CEA 

equilibrium code of Gordon and McBride 

(ref. 10) for a Chapman-Jouget detonation. 

Following an initial overshoot during the short 

transient phase, the detonation speeds reach 

constant values at about 200 microseconds. 

The detonation speeds are slightly higher than 

those predicted by the CEA code. 

Heat Release

The corresponding values of the sensible heat 

release (for both equivalence ratios) are shown 

in figure 2 for the finite rate calculations and 

for the chemical equilibrium results from the 

CEA code. The differences between the finite 

rate and the equilibrium results when the 

detonation reaches the end of the tube amount 

to approximately 6.6% for an equivalence 

ratio of 0.7, and 10.8% for an equivalence 

ratio of 1.0. The higher heat release for the 

finite rate calculations is due to chemical 

recombination inside the tube. 

In our previous work on hydrogen-air 

detonation with an equivalence ratio of 1.0 

(ref. 3), we found a 5.8% difference between 

the finite rate calculations and the equilibrium 

results. 

If we now take into account the increase in 

sensible heat due to recombination in the 

ethylene-air reaction, and use the thermo cycle 

code as described in references 6 and 13, the 

performance of a PDE exceeds that of a ramjet 

up to a flight Mach number of 3.5. 

Pressure Evolution

The pressure evolution along the tube is 

shown in figure 3 for an equivalence ratio of 

1.0. Initially, a large overshoot occurs when 

the detonation wave is initiated. The overshoot 

quickly reduces in value and within 25 cm. the 

pressure reaches a value which is equivalent to 

the von Neumann spike as determined from 

the chemical equilibrium code, CEA. The 

detonation pressure remains constant over the 

remaining 75 cm. of the tube, finally decaying 

as the wave exits the tube. 

Flow Properties at Exhaust

Figure 4 shows the flow properties at the exit 

of the detonation tube for the ethylene-air 

mixture with an equivalence ratio of 1.0. The 

passage of the initial detonation pressure 

spike and temperature rise occurred at 

0.5 milliseconds, followed by a rapid decay 

within 1 millisecond. The plateau region 

where the pressure remains at a level value is 

seen to persist up to about 2.5 milliseconds, 

after which the pressure decreases further to 

near ambient levels. 

Species Concentrations

The changes in gas species were determined 

during the detonation process. In figure 5, the 

species concentrations for an equivalence ratio 

of 1.0 are seen to decay rapidly within 

1.5 milliseconds, followed by a lower rate of 

decay. The concentrations of O and OH 

decline by one order of magnitude in the first 

1.5 milliseconds. 
NASA/TMó2003-212211 5 

Figure 6 shows the behavior for other species 

in the ethylene-air reaction. These species 

have very low concentration values during 

most of the detonation process except during 

the ignition phase. It is these changes in 

intermediate species that indicate 

recombination occurring within the tube and 

that give rise to additional sensible heat 

release. 

Since the decay of species occurs so rapidly at 

0.5 milliseconds, the use of a longer 

detonation tube would not be expected to have 

any significant effect on the heat release. 

Force and Specific Impulse Calculations

The corresponding thrust and impulse for the 

computations described above are shown in 

figures 7 and 8 for an equivalence ratio of 1. 

Initially, the sharp rise in pressure creates a 

short duration thrust spike, figure 7, which is 

followed by a plateau region of about 

2 milliseconds duration and a subsequent 

decay to zero. It is primarily in this level 

pressure duration that PDE thrust is generated. 

The corresponding impulse value is shown in 

figure 8. The value of the impulse at its 

maximum point is 1962 seconds. This value 

compares well with that of Wintenberger et al. 

(ref. 11) of approximately 1840 seconds. 

 

CONCLUSIONS 

Real gas effects, i.e., dissociation and 

recombination in detonating mixtures of 

ethylene-air mixtures have been calculated 

using unsteady CFD. These effects were found 

to be critical in determining the sensible heat 

of reaction so that propulsion performance 

calculations can be made for PDEís. The 

amount of heat release in each propulsion 

cycle is of critical importance in the 

comparison of PDE capability relative to other 

propulsion devices. 

It was found that the difference between the 

finite rate calculations and the equilibrium 

results for a stoichiometric mixture of 

ethylene-air was twice as large as those for 

stoichiometric hydrogen-air at an initial 

pressure of 1 bar and 298 K. 

The sensible heat calculations shown in this 

paper indicate that the losses associated with a 

typical hydrocarbon are less than those 

assumed previously (ref. 6), which were based 

on hydrogen-air detonation. Use of the new 

values determined in this paper would 

increase the competitive range of a PDE, 

relative to a ramjet, to 3.5. 

Additional work is required to establish the 

role of a PDE in flight applications, either as a 

stand alone device or as part of a combined 

cycle. Some of the issues are associated with 

static thrust capability, nozzle and exhaust 

systems and thrust per unit frontal area. 

As stated previously (ref. 6), the PDE may 

have practical advantages over other cycles, 

i.e., simplicity, fewer moving parts, and lower 

cost. In order to realize these benefits, the 

effects of high temperature and high internal 

flow velocity on heat transfer and viscous 

losses as well as fatigue and leakage issues 

related to multi-cycle operation and valving 

require attention. All of these items need to be 

addressed and await practical demonstration. 

REFERENCES 

1. Povinelli, Louis A., ìImpact of 

Dissociation and Sensible Heat Release on 

Pulse Detonation and Gas Turbine Engine 

Performance,î NASA/TMó2001-211080, 

July 2001, also 15th ISABE Symposium 

Paper 2001ñ1212, Sept. 2002. 

2. Povinelli, Louis A., Lee, Jin-Ho and 

Anderberg, Michael O., ìRole of AirBreathing Pulse Detonation Engines in 

High Speed Flight,î NASA/TMó2001-

211163, Sept. 2001, also 52nd IAF 

Congress Paper IAFñ01ñS.5.01, Oct. 

2001. 
NASA/TMó2003-212211 6 

3. Povinelli, Louis A. and Yungster, Shaye, 

ìAirbreathing Pulse Detonation Engine 

Performance,î NASA/TMó2002-211575, 

May 2002, 26th Air-breathing Propulsion 

Subcommittee, JANNAF Conference, 

Destin, FL, April 8ñ12, 2002. 

4. Povinelli, Louis A. and Yungster, Shaye, 

ìThermodynamic Cycle and CFD 

Analyses for Hydrogen Fueled AirBreathing Pulse Detonation Engines,î 

NASA/TMó2002-211698, June 2002, 

also AIAA Paper 2002ó3629, 

38th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference, Indianapolis, 

Indiana, July 7ñ10, 2002. 

5. Schauer, Fred, Stutrud, Jeff and Bradley, 

Royce, ìDetonation Initiation Studies and 

Performance Results for Pulse Detonation 

Engine Applications,î AIAA Paper 2001ñ

1129. 

6. Povinelli, Louis A., ìPulse Detonation 

Engines for High Speed Flight,î NASA/ 

TMó2002-211908, Sept. 2002, also Paper 

17ñ5169, 11th International AIAA/AAAF 

Conference on Space Planes and 

Hypersonic Systems and Technologies, 

September 29ñOctober 4, 2002. Orleans, 

France. 

7. Yungster, S. and Radhakrishnan, K., ìA 

Fully Implicit Time Accurate Method for 

Hypersonic Combustion: Application to 

Shock Induced Combustion Instability,î 

Shock Waves, Vol. 5, 1996, pp. 293ñ303. 

8. Yungster, S. and Radhakrishnan, K., 

ìSimulation of Unsteady Hypersonic 

Combustion Around Projectiles in an 

Expansion Tube,î Shock Waves, Vol. 11, 

2001, pp. 167ñ177. 

9. Li, S.C., Varatharajan, B. and Williams, 

F.A., ìThe Chemistry of Ethylene Ignition 

for Application to Pulse Detonation 

Engines,î AIAA Paper 2000ñ3475 (A00ñ

36678), 36th AIAA/SAE/ASME/ASEE 

Joint Propulsion Conference, July 19, 

2000, Huntsville, Alabama. 

10. Gordon, S. and McBride, B., ìComputer 

Program for Calculation of Complex 

Chemical Equilibrium Compositions and 

Applications,î NASA Reference 

Publication 1311, Oct. 1994. 

11. Wintenberger, E., Austin, J.M., Cooper, 

M., Jackson, S. and Shepard, J.E., ìAn 

Analytical Model for the Impulse 

of a Single-Cycle Pulse Detonation 

Engine,î AIAA 2001ñ3811, A01ñ34455, 

7th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference and Exhibit, 

July 8ñ11, 2001, Salt Lake City, UT. 

12. Kailasanath, K., ìA Review of PDE 

Research-Performance Estimates,î AIAA 

2000ñ0474, 39th AIAA Aerospace 

Sciences Meeting and Exhibit, January 

8ñ11, 2001, Reno, NV. 

13. Heiser, W.H. and Pratt, D.T., 

ìThermodynamic Cycle Analysis of Pulse 

Detonation Engines,î Jrl. of Propulsion 

and Power, Vol. 18, No. 1, Januaryñ

February 2002. 
NASA/TMó2003-212211 7 

0 100 200 300 400 500

Time (μs)

0

1000

2000

3000

Detonation speed (m/s)

CFD finite rate (φ=0.7)

CEA equilibrium (φ=0.7)

CFD finite rate (φ=1.0)

CEA equilibrium (φ=1.0)

Figure 1. Detonation speed variation with time. Ethylene-air mixture, equivalence ratio 1.0, 

pressure 1 bar, temperature 298 K. 
NASA/TMó2003-212211 8 

0 200 400 600

Time (μs)

−10000

0

10000

20000

30000

40000

50000

60000

Heat Release (KJ/Kg fuel)

CFD finite rate (φ=0.7)

CEA equilibrium (φ=0.7)

CFD finite rate (φ=1.0)

CEA equilibrium (φ=1.0)

Figure 2. Heat release per unit mass of burned fuel as a function of time. Ethylene-air mixture, 

equivalence ratios 0.7 and 1.0, pressure 1 bar, temperature 298 K. 
NASA/TMó2003-212211 9 

0.0 50.0 100.0 150.0

Distance (cm)

0.0

10.0

20.0

30.0

40.0

50.0

p/p0

von Neumann spike (from CEA)

Figure 3. Pressure evolution traces. Equivalence ratio 1.0, pressure 1 bar, temperature 298 K. 
NASA/TMó2003-212211 10

0 1000 2000 3000 4000 5000

Time (μsec)

0

10

20

30

40

p/p0; T/T0; ρ/ρ0

p/p0

T/T0

ρ/ρ0

Figure 4. Flowfield conditions at the exit of detonation tube (1.0 meter) as a function of time, 

equivalence ratio 1.0. 
NASA/TMó2003-212211 11

0 1000 2000 3000 4000 5000 6000

Time (μsec)

10−8

10−6

10−4

10−2

100

Mole Fraction

H

O

H2O

OH

O2

H2

HO2

CO

CO2

C2H4

Figure 5. Species mole fractions at the exit of the detonation tube (1.0 meter) as a function of 

time, equivalence ratio 1.0. 
NASA/TMó2003-212211 12

549 550 551

Time (μsec)

10−8

10−6

10−4

10−2

100

Mole Fraction

H2O2

HCO

CH2O

CH3

CH2CO

C2H2

C2H3

C2H5

CH2CHO

Figure 6. Species mole fractions at the exit of the detonation tube (1.0 meter) as a function of 

time, equivalence ratio 1.0. 
NASA/TMó2003-212211 13

0 1000 2000 3000 4000

Time (μsec)

0

1000

2000

3000

4000

Force (N)

Figure 7. Force at the exit of the detonation tube (1.0 meter) for finite rate chemistry calculation. 
NASA/TMó2003-212211 14

0 1000 2000 3000 4000

Time (μsec)

0

500

1000

1500

2000

2500

Specific Impulse (sec)

Figure 8. Specific Impulse at the exit of the detonation tube (1.0) meter for finite rate 

chemistry calculation. 

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

REPORT DOCUMENTATION PAGE

2. REPORT DATE

19. SECURITY CLASSIFICATION

 OF ABSTRACT

18. SECURITY CLASSIFICATION

 OF THIS PAGE

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102

Form Approved

OMB No. 0704-0188

12b. DISTRIBUTION CODE

8. PERFORMING ORGANIZATION

 REPORT NUMBER

5. FUNDING NUMBERS

3. REPORT TYPE AND DATES COVERED

4. TITLE AND SUBTITLE

6. AUTHOR(S)

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

11. SUPPLEMENTARY NOTES

12a. DISTRIBUTION/AVAILABILITY STATEMENT

13. ABSTRACT (Maximum 200 words)

14. SUBJECT TERMS

17. SECURITY CLASSIFICATION

 OF REPORT

16. PRICE CODE

15. NUMBER OF PAGES

20. LIMITATION OF ABSTRACT

Unclassified Unclassified

Technical Memorandum

Unclassified

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135–3191

1. AGENCY USE ONLY (Leave blank)

10. SPONSORING/MONITORING

 AGENCY REPORT NUMBER

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

National Aeronautics and Space Administration

Washington, DC 20546–0001

Available electronically at http://gltrs.grc.nasa.gov

March 2003

NASA TM—2003-212211

ICOMP–2003–02

AIAA–2003–0712

E–13819

WBS–22–719–10–01

20

Real Gas Effects on the Performance of Hydrocarbon-Fueled Pulse

Detonation Engines

Louis A. Povinelli and Shaye Yungster

Propulsion; Detonation engines; PDE

Unclassified -Unlimited

Subject Category: 07 Distribution: Nonstandard

Prepared for the 41st Aerospace Sciences Meeting and Exhibit sponsored by the American Institute of Aeronautics and

Astronautics, Reno, Nevada, January 6–9, 2003. Louis A. Povinelli, NASA Glenn Research Center; Shaye Yungster,

Institute for Computational Mechanics in Propulsion, Cleveland, Ohio 44135. Responsible person, Louis A. Povinelli,

organization code 5000, 216–433–5818.

This paper presents results for a single-pulse detonation tube wherein the effects of high temperature dissociation and the

subsequent recombination influence the sensible heat release available for providing propulsive thrust. The study

involved the use of ethylene and air at equivalence ratios of 0.7 and 1.0. The real gas effects on the sensible heat release

were found to be significantly large so as to have an impact on the thrust, impulse, and fuel consumption of a PDE.