IAF-01-S.5.01

Role of Air-BreathingPulse Detonation

Engines in High Speed Propulsion

L. Povinelliand J.-H. Lee

NASA Glenn Research Center

Cleveland,OH

and

M. Anderberg

Massachusetts Institute of Technology

Cambridge, MA

52nd International Astronautical Congress

1-5 Oct 2001 /Toulouse, France

For permission to copy or republish, contact the International Astronautical Federation

3-5 Rue Mario-Nikis, 75015 Paris, France

https://ntrs.nasa.gov/search.jsp?R=20020009025 2018-05-15T10:45:07+00:00Z

NASA/TM--2001-211163 IAF-01-S.5.01

Role of Air-Breathing Pulse Detonation

Engines in High Speed Propulsion

Louis A. Povinelli and Jin-Ho Lee

Glenn Research Center, Cleveland, Ohio

Michael O. Anderberg

Massachusetts Institute of Technology, Cambridge, Massachusetts

September 2001
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NASA/TM--2001-211163 IAF-01-S.5.01

Role of Air-Breathing Pulse Detonation

Engines in High Speed Propulsion

Louis A. Povinelli and Jin-Ho Lee

Glenn Research Center, Cleveland, Ohio

Michael O. Anderberg

Massachusetts Institute of Technology, Cambridge, Massachusetts

Prepared for the

52nd International Astronautical Congress

cosponsored by the International Astronautical Federation, French Academy of Sciences,

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Toulouse, France, October 1-5, 2001

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September 2001
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Available electronically at http://gltrs.grc.nasa.gov/GLTRS
IAF-01-S.5.01

ROLE OF AIR-BREATHING PULSE DETONATION ENGINES IN

HIGH SPEED PROPULSION

Louis A. Povineili* and Jin-Ho Lee +

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio

Michael O. Anderberg*

Massachusetts Institute of Technology'

Cambridge, Massachusetts

ABSTRACT

In this paper, the effect of flight Mach number on the

relative performance of pulse detonation engines and

gas turbine engines is investigated. The effect of ram

and mechanical compression on combustion inlet

temperature and the subsequent sensible heat release is

determined. Comparison of specific thrust, fuel

consumption and impulse for the two engines show the

relative benefits over the Mach number range.

INTRODUCTION

In a previous publication, l the impact of dissociation

and the resultant sensible heat addition effect on pulse

detonation engine (PDE) performance was investigated.

Comparisons were carried out with a gas turbine engine

using a thermodynamic cycle analysis. The results

showed, for the first time in the literature, 1 a decrease in

the sensible heat available for the PDE, which generally

caused lower performance. In this paper, a more

detailed analysis is made of the effect of changing inlet

conditions on the sensible heat release in a PDE and a

gas turbine, as well as investigating the performance

changes associated with high Mach number flight.

ANALYSIS AND PROCEDURE

equilibrium properties of the products and the sensible

heat release, heR. A non-dimensional heat release, _,

was defined and then determined from the relationship:

= f hp_/Cp To

where f is the fuel air mass ratio.

The resulting _}was used to perform cycle calculations

for the PDE and Brayton cycles. The thermodynamic

cycle analysis was based on a modified version of the

work of Pratt and Heiser. 4 The modification provides

for variable specific heat, specific heat ratio and

reference temperature. The cycle analysis yields the

maximum theoretical performance for an ideal PDE and

Brayton cycle. The analogy with the Otto and Diesel

cycle analyses is noted. Further, it is noted that the PDE

performance is based on a closed form solution for the

leading edge normal shock wave Mach number and

entropy rise of the detonation wave, and does not rely

on a surrogate analysis such as the Humphrey cycle.

The relationship of _ to the Chapman-Jouget

parameters is given by:

A stoichiometric propane-air mixture was specified at

an initial temperature and pressure. The detonation and

deflagration properties were obtained using the method

by Pratt, 2 which is based on the CEA program of

McBride and Gordon) The results yielded the

where _is the temperature ratio Tv'To.

The approach used herein is a closed thermodynamic

cycle analysis rather than an integration over a cycle of

an unsteady solution of the instantaneous pressures

* Chief Scientist, Turbomachinery and Propulsion Systems Division

t Research Engineer, Combustion Branch, Turbomachiner)' and Propulsion Systems Division

*SunmaerIntern

NASA/TM--2001-211163 1
acting on the PDE tube. Further, it is assumed that

isentropic compression and expansion occurs with the

detonation device. These processes will give the upper

limit of the performance potential for a PDE.

RESULTS

Static condition, V0_=0

A stoichiometric mixture of propane and air was

selected at an initial reference temperature of 400 °R

and a pressure at 3.8 psi. These conditions correspond

to those at an altitude of 6.2 nil. The heat release was

calculated from the equilibrium code.-" The calculations

for the sensible heat release yielded heR values of

1.6 x 10 4 (PDE) and 1.8 x 10 4 BTU/Ibm (Brayton) at a

gt (=TdTo) value of 1. The initial sensible heat addition,

figure 1, is about 12 percent lower for the detonation

process than for the deflagration. The temperature ratio,

_/, was then varied over the range from 1 to 5. The

results are shown in figure 1.

The thermal efficiencies for the two cycles were

calculated and are shown in figure 2 and is defined as

eff-th = 1

heat rejected

heatadded

It is noted that the PDE efficiency exceeds that of the

Brayton over the range of temperature ratios. However,

the sensible heat release for the Brayton exceeds that of

the PDE as shown in figure 1. These two varying

parameters determine the relative performance of the

two cycles.

The specific thrugt values calculated from the sensible

heat release and thermal efficiency values for the

Brayton and PDE cycles are shown in figure 3 as a

function of _. The significant feature of figure 3 is that

the PDE specific thrust crosses over the Brayton cycle

curve at a temperature ratio of 2.25. This result is due to

the lower sensible heat available in the PDE, and to the

decreasing value of thermal efficiency for the PDE

relative to the Brayton. Since, the specific thrust is

given by:

- + 2r/,h#c,,T 0 - Vo

hi g ,-

it may be seen that for V0 = 0, when the products of the

thermal efficiency and heat release become equal, the

specific thrusts have a common value. As the thermal

efficiency difference diminishes between the two

cycles, and with a larger Brayton cycle heat release

value, the Brayton cycle performance will eventually

exceed that of the PDE. As shown in figure 3, the

performance advantage for the PDE occurs at a

temperature ratio value below 2.25.

2.0 1

___ Bmyton

_1.5 i '

i

1.0 ,

1 2 3 4 5

T3FFo

Figure 1. Sensible heat release (× 104 BTU/Ibm) for the

Brayton and PDE cycles.

1.0

0.9 ]

0.8

0.7 ]

0.6

0.5

0.4 i

0.3 /

0.2

• /

0.1 ,_ 0.0

_ i _-i .... !

•U ,,

: -- Brayton

r i ,

1 2 3 4 5

T_o

Figure 2. Thermal efficiencies for the PDE and

Brayton cycles, y=1.25.

200

I

i -.._ _ ...... 7 - - _ _

..... PDE

-- Brayton

I I ,

3 4 5

Ts/T0

150 , •/,//

50 --_ --

0

1 2

Figure 3. Specific thrust (lbr-s/lbm), y= 1.25.

NASA/TM--2001-211163 2
Effect of Forward Velocity

The specific thrust calculations were performed for flight

Mach numbers from 1 to 4, where 980 ft/s is equal to the

speed of sound at the altitude of 6.2 miles and a freestream gamma value of 1.4. For these calculations it was

assumed that the combustor entrance temperature value

was equal for both cycles, and the gamma for the

combustion was 1.25. The results are shown in figure 4.

In all cases, it is seen that the PDE offers only a small

potential benefit over the Brayton cycle over the

investigated speed range. The PDE benefit exists at

values less than 2.25, which is the same value

observed for static conditions. This crossover value

remains fixed since all of the performance parameters

derive from the same thermal efficiency curves.

The relationship between Mach number and

temperature ratio is given by the expression

M0= _/2(_-1)/7_-1,

and the Mach number variation for yo..=1.4 is shown in

figure 5.

180

160 "-

1404

120

- 100

80

60

40

20

0

/ Mach 1

i

5 Mach_

..... Math 3 1

. i- - - Mach 4

/

i

-- Brayton

- - PDE

1 2 3 4

TdTo

Figure 4. Mach number effect on specific thrust.

40 il / .....

35 ! ...........

! /

i /

30 ,,'

20 /

/

<i 15 /-

kc _'

0 1 2 3 4

Figure 5. Mach number effect on minimum temperature ratio.

It may be seen that for a given Mach there exists a

value of tg due to the heating (ram and/or mechanical

compression) occurring through the inlet and

compressor. For a given Mach number, therefore, a

minimum value of Ig exists, below which physically

meaningful results are not possible. Using the values

from figure 5 allows us to establish physical boundaries

for the results shown previously in figure 4. Figure 6

shows the data from figure 4 with these limits imposed,

and illustrates the effect they have on narrowing the

operating range over which the specific thrust of the

PDE is greater than that of the Brayton cycle.

180, ! I

r

140 ._, _

I

40 /_'

20 ' i

0 I

1 2 3 4 5

w'ro

Figure 6. Regions where the PDE cycle performance

exceeds that of the Brayton cycle.

It is seen that the PDE offers a specific thrust advantage

for a flight Mach number of 1 only above the lower

boundary where V =1.2. In addition, the upper bound is

established by the crossover point with the Brayton

cycle, or Ig = 2.25. Hence, only a narrow operating

range exists over which the specific thrust for the PDE

exceeds that of the Brayton cycle. At Mach 2 the

minimum allowable V value is 1.8 and the upper

boundary remains at 2.25. Therefore, the PDE

performance benefit range is reduced to the small range

shown in figure 6. At Mach 3, the lower V limit

increases to 2.8, which is greater than the crossover

point between the PDE and Brayton cycles. Therefore,

there is no operating range over which the PDE offers a

performance advantage at this velocity. At a Mach

number of 4, the minimum allowable inlet temperature

ratio of 4.2 greatly exceeds the crossover point. As with

the Mach 3 condition, no performance benefits accrue

with the PDE relative to the gas turbine. As discussed

previously, l other advantages such as system weight,

cost and complexity must exist in order to select a PDE

device instead of a gas turbine engine.

The comparisons presented in figures 3, 4, and 6 are

based on the assumption that the PDE and Brayton

cycles have identical inlet conditions, and therefore

NASA/TM--2001-211163 3
equalT3values.In reality,bothcycleswouldhavethe

sameramcompression, buttheBraytoncyclewould

havesomeadditioncompressionduetothepresenceof

a mechanical compressor. ThePDEis designedto

avoidthecomplexitiesof themovingpartsassociated

vdthmechanical compressors byrelyingsolelyonram

compression. Thisdifferenceinconfigurationresultsin

differentT3valuesatagivenflightMachnumber.The

effectof thesedifferentinletconfigurationsis shown

Figure7.Here,themaximumspecificthrustis plotted

againstflight Machnumber.Thecompressionratio

Otc= P_/P2)for theBraytoncycleis addedasan

additionalparameterfor this figure.In addition,a

compressor exit temperatureof 1250°F is setasa

materiallimit. Thebold linesshowthe rangeof

operationpermissiblewiththistemperaturelimit and

thefaintlinesareforthecaseofunlimitedtemperature.

Increasingthecompressor exit temperature limit allows

for slightly higher maximum operating speeds for all

configurations. This effect is reduced with increases in

compressor pressure ratios. The operating ranges for

the different configurations with compressor exit

temperature limits of 1300 °F and 1450 °F are shown in

figures 8 and 9.

The data presented in figure 7 assumes a constant value

of rcc for all flight Mach numbers. In reality, the

compression ratio is reduced at higher flight velocities

to achieve optimum performance. A realistic variation

for the compression ratio for the Brayton cycle is

shown in figure 8.

Reducing the pressure ratio at higher Mach numbers

increases the maximum specific thrust available from

the Brayton cycle. The effect of using this pressure ratio

schedule and the corresponding ram compression

180

160

140

12o

e_

_. 1oo

e_

ao60

40

20

0

\../r c = _ " " " PDE

-- Brayton

_ l0

4

l,

/

J /

0 l 2 3 4

M0

Figure 7. Specific thrust for the PDE and Brayton cycles,

T3 ,,_, = 1250 °F, constant values of novalues are shown in figure 11. It is noted that the

variation in pressure ratio only has a significant effect

on performance at high Mach numbers where the

compressor exit temperature limit is exceeded.

180 q

I - -- - POE Bravton

140 _ "

_ 120 _ _-__

2. 60 /

40 . .

20 [ '/'

0

0 t 2 3 4

Mn

Figure 8. Specific thrust for PDE and Brayton cycles,

T_ma,= 1300 °F, constant values of re,..

180

160

140

_" 120

100

_- 80

60

4O

2O

\ - - - PDE

_ _ -- Brayton

/ ' /'

/

/,

f

0 1 2 3 4

Mo

Figure 9. Specific thrust for PDE and Brayton cycles,

T3 mat = 1450 °F, constant values of G.

35

30

25

'_

°20 \!

15

e_

10

5

0

0

N

\.

1 2 3 4

Mo

Figure I0. Compressor pressure ratio variation for the

Brayton cycle.

NASA/TM--2001-211163 4
180

160

140

120

10o

80

60

4O

20

0

Figure 11.

-rrc = - - - PDE

L. _'_ -- -Brayton

Y

i

,

I'

0 1 ,'_ 3 4

M0

Specific thrust, variable n_. T3m,,=1250 °F.

Clearly. the PDE cycle enjoys a performance benefit as

long as the Brayton cycle is operating at unrealistically

low compression ratios. It quickly loses that benefit as

the Brayton compressor pressure ratio exceeds a value

of 4. The PDE can only match the Brayton cycle if

higher PDE compression ratios can be obtained;

suggesting a combined cycle configuration which

would not be a pure PDE such as a PDE and dual-mode

ramjet.

The specific impulse = f/(F/mo) and the fuel

consumption are shown in figures 12 and 13 for the

corresponding variables of figure 11.

The performance calculations so far have been based on

ideal process efficiencies. Reference 4 developed the

relationships for non-ideal processes and that approach

is used herein. Figures 14 to 16 show the effect of

reduced expansion efficiencies on the relative

performance of the PDE and Brayton cycles. In all

these figures, the expansion efficiency is reduced to

0.95 while maintaining ideal compressor and burner

efficiencies. The ideal process efficiency curves are

shown in the background for reference. These results

indicate that the performance of both configurations is

reduced for all Mach numbers, but the PDE cycle

performance decreases more than the Brayton cycle.

This can be seen by comparing the crossover point

between the PDE curve and the Brayton curve for no=4.

For the ideal process efficiencies, the PDE exhibits a

slight performance advantage until a Mach number of

0.6. With r/e---0.95, the PDE curve remains below the

Brayton curve for all Mach numbers.

2500

2000

'-_ 1500

1000

n" =30

""_ - - PDE

-___ Brayton

500

II t

13 1 2 3 4

M_

Figure 12. Specific impulse, variable nc. T3max = 1250 °F.

PDE

-- Bra}ton

0 1 2 3 4

Mo

Figure 13. Specific fuel consumption, variable no,

T3max = 1250 °F.

180

160

140

120

"_ I00

80

60

40

20

0

tr c = 30 - - PDE

-- Brayton

0 1 2 3 4

M0

Figure 14. Specific thrust, variable n c, T3max= 1250 °F,

r/c=l.0, r/b=l.0, r/e--0.95.

NASA/TM--2001-211163 5
2500

2 ()o o

7 1500

2

[ 000

5OO

0

++ IO -''_-_ B ravlon

/

/

! •

1 2 3 4

Figure 15. Specific impulse, variable go, T3.m_ = 1250 °F.

G=I.O, t/b=l.O, r/c--0.95.

7 7

6 ] _LII_ i+ BPDErayton

5 \,

\

"_ 3

1

!

0

0 1 2 3 4

Mo

Figure 16. Specific fuel consumption, variable _¢.

T3,.a_ = 1250 °F, tJc=l.O, r/b=l.O, r/c=0.95.

Figure 17 shows the effect of reduced compressor

efficiencies on engine performance. It is apparent that a

reduction in compressor efficiency has the opposite

effect of the reduction in expansion efficiency by

slightly improving the relative performance of the PDE.

If the compressor and expansion efficiencies are both

lowered, the relative performance between the PDE and

Brayton cycles remains fairly constant. This is

illustrated in figure 18, where the compressor and

expansion efficiencies are both set at 0.95.

A reduction in burner efficiency, shown in figure 19,

has an even greater effect on reducing the performance

of the PDE relative to the Brayton cycle.

Reducing the nozzle efficiency further, to 0.90, for the

same parameters as in figure 14 shows a significant

drop in the performance. The results are seen in

figure 20, and underscore the importance of nozzle

performance to the PDE and Brayton cycles.

180

160

140

120

100

Y_

8O

,.7

60

4O

2O

0

e_

+_ - PDE ---- Bravlon

+ /

///

0 1 2 3 4

M_

Figure 17. Specific thrust, constant rG 0_=0.95,

rh,=1.0, r/_=1.0.

t80

160 _ .... PDE

140 _ -- Brayton

120 _

100

8O

60 "\

4020 t •//',``''

0 1 2 3 4

Mo

18. Specific thrust, constant no, r/c=0.95,

0h=l.0, tie-----'0.95.

Figure

40

20

0

.... PDE

-- Brayton

0 1 2 3 4

M0

Figure 19. Specific thrust, variable rtc+T3m_ = 1250 °F,

r/_=l.0, r/b----0.95,G=I.0.

NASA/TM--2001-211163 6
180

160

140

_" 12o

e-,

1oo

80

_, 60

4o

2o

.... PDE

q

_. -- Brayton

.... .__

/j"

,/'

/

//

' 1 i

1 2 3 4

Mo

Figure 20. Specific thrust, variable re,., T3m,_ = 1250 °F,

r/c=l.0, r/b=l.0, r/e----0.90.

Figures 21 to 23 show" the effect of reduced compressor,

burner and expansion efficiencies on the increase in

overall thermal efficiency of the two cycles. Figure 21

shows the effect of reduced compression efficiency. The

only difference between the two sets of curves is that the

background set has r/c=l.O0 whereas the foreground set

has rk.--0.95. The five percent reduction in compression

efficiency has almost no impact on the point where the

PDE and Brayton curves cross, indicating that it has very

little impact on the relative performance of the two

cycles.

Figure 22 examines the effect of the burner efficiency.

Here the background curves were calculated with

r/b=l.00 and the foreground curves with r/b--=0.95. The

reduced burner efficiency causes the crossover point of

the PDE and Brayton curves to move from a temperature

ratio around _u=3 to _=2.5. This means that a reduction

in burner efficiency has a significant impact on the

relative performance of the PDE and Brayton cycles.

Figure 23 shows the effect of expansion efficiency on

the overall thermal efficiency. The five percent

reduction in expansion efficiency causes the

intersection point of the PDE and Brayton curves to be

reduced by more than 1. This shows that the relative

performance of the two cycles is most sensitive to the

expansion efficiency.

1.0

0.9

0.8

0.7

0.5 -,

0.10"2_-- -- grayton

0.0 i

2 3 4

T_/To

Figure 21. Thermal efficiency, r/_=0.95, 0b=1.00, r/e=0.90.

Background: rlc=l.00, qh=l.00, r/_=0.90.

1.0

0.9

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

0.0

1

_+

i

] I

2 3

T _/T0

4 5

Figure 22. Thermal efficiency, r/_=1.00, 0b----0.95, r/e----'0.90.

Background: r/,=l.00, r/h=l.00, r/e---'0.90.

e1.0

0.9 t

0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

0.0

1

i

J

i

Bra)aon

2 3 4 5

Tfl" 0

Figure 23. Thermal efficiency, r/_=1.00, r/b=1.00, r/_--0.95.

Background: r/c=l.00, r/b=l.00, r/e=l.00.

NASA/TM--2001-211163 7
CONCLUSIONS

Performance parameters for a pulse detonation engine

were compared with that of a gas turbine engine. The

PDE exhibited a significant static (no forward velocity)

performance benefit when compared with the gas turbine

engine whose compression ratios were lower than 4.

However, that benefit quickly disappeared as the Bra_on

compressor pressure ratio exceeded a value of 4.

The sensible heat release for both the detonation and

the deflagration processes were determined over a

Mach number range from 0 to 5 using an equilibrium

calculation. As shown previously, _ there is a 12%

reduction in the sensible heat release for the detonation

process in the PDE compared to the deflagration

process in the Brayton cycle. This reduction is due to

the larger degree of dissociation in the PDE. The effect

of dissociation on sensible heat release is a primary

cause for reducing the specific thrust of a pulse

detonation engine relative to a gas turbine engine.

Calculations were performed to examine the effect of

Mach number on the performance of the PDE and gas

turbine engines. The PDE was found to have higher

specific thrust at subsonic Mach numbers relative to the

gas turbine engine whose compression ratios were

lower than 4. This trend is reversed as the gas turbine

compression ratios become higher than 4.

As the flight Mach number increases to sonic and

higher, the PDE performance decreases. At sonic

speeds, the PDE and Brayton with pressure ratio of 4

are equal. At a Mach number of 2, the PDE and

Brayton with a pressure ratio of 2 are equal and at

Mach 3, the PDE is lower than the Brayton with no

mechanical compression (i.e., only ram compression or

a ramjet). For reasonable Brayton compression ratios

t10 to 30) the PDE performance was always lower,

having lower specific thrust, lower impulse and higher

specific fuel consumption.

The performance calculations were also examined for

non-ideal process efficiencies, i.e., compression,

burning and expansion. For the ideal nozzle expansion,

the PDE had a slight performance advantage up to a

Mach number of 0.6 compared to a Brayton

compression ratio of 4. With a 0.95 nozzle efficiency,

the PDE performance remains below the Brayton for all

Mach numbers.

A reduction in compressor efficiency was found to

slightly improve the relative performance of the PDE,

and for a reduction in both the compressor and

expansion efficiencies, the relative performance

between the two cycles remained fairly constant.

A reduction in burner efficiency had a larger effect on

reducing the PDE relative to the Brayton cycle. Further

reducing the nozzle efficiency to 0.90 caused a large

reduction in PDE performance as shown previously. 4

Inspection of these trends using the thermal efficiency

curves also shows that the relative performance of the

two cycles is most sensitive to the expansion efficiency.

Finally, it is observed that the PDE may be best suited

for a combined cycle application. Coupling of a PDE,

which theoretically can provide static thrust, with a

ramjet starting at Mach 2.5 to 3.0, and with a scramjet

at Mach 5.0 to 6.0 may be such a possibility.

REFERENCES

1. Povinelli, L.A. _2001) Impact of dissociation and

sensible heat release on pulse detonation and

gas turbine engine performance, ISABE Paper

2001-1212, NASA/TM--2001-211080, July 2001.

2. Pratt, D.T. (2000) EQL calculation program,

personal correspondence.

3. McBride, B.J. and Gordon, S. (1996) Users manual

and program description, NASA RP-1311.

4. Heiser, W.H. and Pratt, D.T. (2001)

Thermodynamic cycle analysis of pulse detonation

engines, J. Propulsion & Power, in preparation.

NASA/TM--2001-211163 8

j1

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Role of Air-Breathing Pulse Detonation Engines in High Speed Propulsion

6. AUTHOR(S)

Louis A. Povinelli, Jin-Ho Lee, and Michael O. Anderberg

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

11.

Technical Memorandum

5. FUNDING NUMBERS

WU-706-31-23-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-13021

10. SPONSORING/MONITORING

AGENCY REPORTNUMBER

NASA TM--2001-211163

IAF-O 1-S.5.01

SUPPLEMENTARY NOI[=S

Prepared for the 52rid International Astronautical Congress cosponsored by the International Astronautical Federation, French Academy

of Sciences, Air & Space National Academy, Floral World Academy, and the Art & Literature Academy, Toulouse, France,

October I-5, 2001. Louis A. Povinelli and Jin-Ho Lee, NASA Glenn Research Center. and Michael O. Anderberg, Massachusetts

Institute of Technology, Cambridge. Massachusetts. Responsible person, Louis A. Povinelli, organization code 5000, 216--433-5818.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified- Unlimited

Subject Categories: 07, 20, and 28 Distribution: Nonstandard

Available electronically at htro://eltrs.grc.nasa.gov/GI,TRS

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

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

In this paper, the effect of flight Mach number on the relative performance of pulse detonation engines and gas turbine

engines is investigated. The effect of ram and mechanical compression on combustion inlet temperature and the

subsequent sensible heat release is determined. Comparison of specific thrust, fuel consumption and impulse for the

two engines show the relative benefits over the Mach number range.

14. SUBJECT TERMS

Propulsion: Air-breathing propulsion

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE

Unclassified Unclassified

NSN 7540-01-280-5500

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

15

16. PRICE CODE

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102