Daniel E. Paxson

Glenn Research Center, Cleveland, Ohio

Kevin T. Dougherty

QSS Group, Inc., Cleveland, Ohio

Ejector Enhanced Pulsejet Based Pressure Gain

Combustors: An Old Idea With a New Twist

NASA/TM—2005-213854

August 2005

AIAA–2005–4216
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Daniel E. Paxson

Glenn Research Center, Cleveland, Ohio

Kevin T. Dougherty

QSS Group, Inc., Cleveland, Ohio

Ejector Enhanced Pulsejet Based Pressure Gain

Combustors: An Old Idea With a New Twist

NASA/TM—2005-213854

August 2005

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

41st Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Tucson, Arizona, July 10–13, 2005

AIAA–2005–4216
Acknowledgments

This work was supported by the NASA Glenn Research Center’s Constant Volume Combustion Cycle Engine

Project (CVCCE). The authors wish to thank the laboratory technicians and mechanics, Robert Pelaez,

Edward Pluta, Mark Robinson, and Roger Scheman for their invaluable assistance in getting and keeping this

experiment operational.

Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

This report is a formal draft or working

paper, intended to solicit comments and

ideas from a technical peer group.

This report contains preliminary

findings, subject to revision as

analysis proceeds.

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

This work was sponsored by the Low Emissions Alternative

Power Project of the Vehicle Systems Program at the

NASA Glenn Research Center.
NASA/TM—2005-213854 1

Ejector Enhanced Pulsejet Based Pressure Gain 

Combustors: An Old Idea With a New Twist 

Daniel E. Paxson 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Kevin T. Dougherty 

QSS Group, Inc. 

Cleveland, Ohio 44135 

Abstract 

An experimental investigation of pressure-gain combustion for gas turbine application is described. The test 

article consists of an off-the-shelf valved pulsejet, and an optimized ejector, both housed within a shroud. The 

combination forms an effective ‘can’ combustor across which there is a modest total pressure rise rather than the 

usual loss found in conventional combustors. Although the concept of using a pulsejet to affect semi-constant 

volume (i.e., pressure-gain) combustion is not new, that of combining it with a well designed ejector to efficiently 

mix the bypass flow is. The result is a device which to date has demonstrated an overall pressure rise of 

approximately 3.5 percent at an overall temperature ratio commensurate with modern gas turbines. This pressure 

ratio is substantially higher than what has been previously reported in pulsejet-based combustion experiments. Flow 

non-uniformities in the downstream portion of the device are also shown to be substantially reduced compared to 

those within the pulsejet itself. The standard deviation of total pressure fluctuations, measured just downstream of 

the ejector was only 5.0 percent of the mean. This smoothing aspect of the device is critical to turbomachinery 

applications since turbine performance is, in general, negatively affected by flow non-uniformities and unsteadiness. 

The experimental rig will be described and details of the performance measurements will be presented. Analyses 

showing the thermodynamic benefits from this level of pressure-gain performance in a gas turbine will also be 

assessed for several engine types. Issues regarding practical development of such a device are discussed, as are 

potential emissions reductions resulting from the rich burning nature of the pulsejet and the rapid mixing 

(quenching) associated with unsteady ejectors. 

I. Introduction 

Gas turbine engines utilizing conventional ‘constant pressure’ combustion actually suffer a loss in total pressure 

across the combustor. This is due partly to unavoidable thermodynamic losses associated with heating a fluid at 

finite Mach number, and partly due to aerodynamic losses associated with the geometrically convoluted flow path 

(e.g., liner cooling, mixing requirements, etc.). Typical combustor losses range from 4 to 8 percent of the total 

pressure delivered from the upstream compressor (ref. 1). The loss in total pressure yields lower work potential for 

the gas entering the downstream turbine. The resulting performance loss to the engine in terms of reduced specific 

thrust and/or increased specific fuel consumption is not insignificant. The effect of this loss, or more to the point, the 

benefit of reducing it, is illustrated in figure 1. This figure shows calculated reduction in specific fuel consumption 

(SFC) as a function of combustor pressure ratio for a small, uncooled turbojet engine, and for a high bypass turbofan 

engine flying at an altitude of approximately 30 kft., and a Mach number of 0.8. The relevant engine parameters 

used in the calculations are listed in the appendix. The baseline engine in each case has a combustor pressure loss of 

5 percent. The calculations were done using a simple cycle deck based on the analysis methods outlined in 

reference 2. In each case the specific thrust was held constant by either lowering the turbine inlet temperature 

(turbojet), or increasing the fan bypass ratio (turbofan) as the combustor pressure ratio rose. Details of the 

straightforward calculations are omitted here, as the intention is to illustrate a trend rather than to absolutely 

quantify. 
NASA/TM—2005-213854 2

It can be seen that if the pressure loss can be reduced, 

or better still, if a modest pressure rise can be achieved 

across the combustor, substantial improvements in SFC 

can be realized. Of course, the term “substantial” is 

somewhat relative; however, it may be argued that SFC 

reductions of several percent are competitive with many 

concepts for improving gas turbine performance on the 

near-term horizon, when all are assessed realistically 

(e.g., flow control, advanced materials, etc.). For 

example, calculations done with the cycle deck described, 

show that increasing the combustor pressure ratio from 

0.95 to 1.04 has the same impact as increasing the 

compressor efficiency by 6 percent (0.9 to 0.96), or the 

turbine efficiency by 2.5 percent (0.900 to 0.925) in the 

turbofan engine. Both of these represent formidable 

challenges in the mature technology of turbomachinery. 

The recognized benefits of pressure-gain combustion 

have been the impetus for numerous investigations in the 

past (refs. 3 to 6) and near present (refs. 7 and 8). In every 

investigation, pressure-gain is achieved, either analytically or experimentally, through some approximation of a 

constant-volume combustion process. As a result, every concept is intermittent (unsteady), though cyclic, in some 

way. This fact yields a combustor exit flow that is unsteady to varying degrees. Since unsteadiness is generally 

regarded as a detriment to the performance of downstream turbomachinery, its presence tends to detract from the 

thermodynamic benefits of the pressure-gain. Therefore it is desirable in any intermittent combustion process to 

minimize unsteadiness in the exhaust path. 

Pressure-gain combustion concepts vary widely in mechanical complexity and theoretical (or realizable) pressure 

ratio for a given overall temperature ratio (i.e., the amount of fuel added). Generally, and arguably, overall pressure 

ratio scales with complexity. The higher the potential pressure-gain, the more complex the device tends. Practically 

speaking however, higher complexity scales with larger developmental risk, time, and difficulty. It is for this reason, 

perhaps, that the high payoff pressure-gain concepts such as wave rotor, and detonation-based systems remain 

largely on paper, while the lower payoff concepts, such as those using pulsejets have been demonstrated. Of the 

latter concepts, unfortunately, the experiments have tended to yield performance levels that are too low to justify 

further development, despite their relative simplicity. One reason seems to be that the effluent from a pulsejet is far 

too hot for any turbine to tolerate and must therefore be mixed with some bypass flow. This is typically a loss-filled 

process and substantially degrades the overall performance. 

This paper explores the potential of a pulsejet-driven, pressure-gain combustor in an experiment which 

capitalizes on the demonstrated ability of properly designed, unsteady ejectors (i.e., augmentors) to efficiently 

entrain, mix, and smooth secondary flows (refs. 9 and 10). In previous experiments, ejectors driven by unsteady 

thrust sources have demonstrated exceptionally high thrust augmentation levels. Thrust augmentation is defined as 

the total thrust of the ejector and driver divided by the thrust of the driver alone. Levels near 2.0 have been reported 

in the past (ref. 11), and observed more recently when the driver was a pulsejet (refs. 9 and 12). It was purported that 

if both the pulsejet and ejector were placed within a shroud, the ejector pumping action which normally yields such 

high thrust augmentation could instead be used to raise the total pressure. This experiment was designed to 

investigate that proposition. The results which follow will demonstrate that it is indeed true. To date an overall 

stagnation pressure ratio of 1.037 has been demonstrated while operating with a total temperature ratio of 

approximately 2.2. This temperature ratio is typical for a many gas turbine combustors. While this is a modest 

pressure ratio, it appears to be substantially higher than what has been demonstrated in previous pulsejet-based 

experiments (1.005 to 1.02). 

The experimental setup is presented in detail, including relevant instrumentation. Results are being presented 

showing performance with various ejector configurations, and at different operating conditions. Additional results 

and discussion are presented from the ejector yielding the highest performance. The paper concludes with an 

argument suggesting that this type of pressure-gain combustor may also have the potential for low-emissions. 

0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

4.5

0.95 0.97 0.99 1.01 1.03 1.05

Combustor Total Pressure Ratio

SFC Reduction, %

Turbojet

Turbofan

γ=1.35

Figure 1.—Reduction in Specific Fuel Consumption as

a function of combustor pressure ratio for a small,

uncooled turbojet engine, and a high bypass turbofan 

engine at an altitude of approximately

30 kft., and Mach 0.8 cruise. 
NASA/TM—2005-213854 3

II. Description of the Experiment 

The experimental rig is shown schematically in figure 2, with relevant components labeled. Air enters the shroud 

on the left through a regulated 2.0 in. diameter supply line. A portion of the incoming air then passes through the 

pulsejet, where it is mixed with fuel and reacted according to the cyclic, self-aspirating, resonant process of 

operation. A large portion of the incoming air bypasses the pulsejet and enters the ejector. Here it is mixed with and 

pumped by the hot gas exiting the pulsejet. The mixed flow then passes from the ejector into the downstream portion 

of the shroud containing instrumentation for measuring temperature, time-averaged total pressure, and total pressure 

fluctuations. Details of the instrumentation and other major rig components will be given following the overall 

flowpath description. After passing over the instrumentation, the mixed flow exits the rig to the ambient 

surroundings via a small opening or gap between the terminating flange of the shroud and a specially designed 

flange blank, which also serves as a safety burst disk. The size of the exhaust gap can be varied using spacers and, in 

combination with the regulated overall airflow, serves to establish the pressure inside the shroud. Note that the split 

in airflow between pulsejet and ejector cannot be regulated in this experiment. It is governed by the operation of the 

pulsejet and the entrainment characteristics of the ejector. 

A. Pulsejet 

The pulsejet used in the experiment is based on a commercially available design made by the Dynajet 

Corporation. A schematic of the device is shown in figure 3, with relevant dimensions and components labeled. The 

valve body is made from aluminum. The combustion chamber and tailpipe section are made from Inconel. The valve 

‘petals’ are made from blued spring steel. The unit operates on gasoline, at a nominal frequency of 220 Hz., and 

generates approximately 4.25 lbf thrust statically in the open air. Starting requires a spark plug (driven by a 60 Hz. 

high voltage supply), and several short bursts of relatively high pressure air supplied through the starting air line 

shown in figure 3. Once resonant operation is 

established, both are terminated. Descriptions of 

resonant operation may be found in the literature 

(ref. 13), and as such, will not be presented here. The 

fuel supply system for the device is of the venturi-type. 

The fuel tank is external to the rig. A line leads from it, 

through a remotely operated, on-off valve, to the inlet 

junction shown in figure 3. The fuel tank head is 

pressurized to the shroud internal pressure via an air 

line as shown in figure 2. 

Total Pressure

Static Pressure

Total Temperature

Burst Disc

Total Temperature

gap

Total Pressure

Air flow

Pulsejet Ejector

Shroud

Perforated Liner

Struts

Static Pressure

Fuel

Fuel Tank Pressurization Line

Starting Air Line ∆P

Figure 2.—Pulsejet-driven, pressure-gain combustor rig schematic. The rendering is approximately to scale. 

Valve body

Spark plug

Combustion chamber

Tailpipe

pressure tap

Valve

Starting air 

line

Fuel line seam 2.5 in. 1.25 in.

21.25 in.

Figure 3.—Schematic of the pulsejet device used in the 

experiment. 
NASA/TM—2005-213854 4

B. Ejectors 

Several ejectors of various lengths and diameters 

were tested. All had circular cross-section, and were 

constructed from aluminum. Three diameters were 

used, D = 2.2, 3.0, and 4.0 in. Testing was also done 

with no ejector present which, given the 6.0 in. shroud 

diameter could be considered as a fourth ejector 

diameter. For each diameter tested, a modular ejector 

body was available as shown in figure 4(a). Two center 

sections of different length were constructed thereby 

allowing testing of four total ejector lengths (no center 

section, short center section, long center section, and 

coupled center sections). Testing was also performed 

on a tapered ejector as shown in figure 4(b). The 

diameter at the end of the inlet was 3.0 in. This ejector 

was also constructed from modular sections allowing 

variations in length to be examined. For all of the 

ejectors tested, an inlet nose radius of 0.75 in. was 

used, with the exception of the D = 4.0 in. ejector 

which had in inlet nose radius of 0.5 in. in order to fit 

within the shroud. Each ejector inlet was fitted with a 

flange that sealed it within the shroud, thereby insuring 

that all air and hot gas passed through the ejector (air 

could not go around the outside). 

It is noted that the dimensions of the ejectors used 

in this experiment were not selected on the basis of any 

analysis. They were originally fabricated for the thrust 

augmentation optimization tests referenced earlier 

(refs. 9 and 14), but were conveniently available for 

the present tests. On the unproven, but rational, 

supposition that high thrust augmentation 

configurations will yield relatively high pressure gain 

however, they were a logical set to use. 

C. Shroud 

The shroud was constructed from 6.0 in. diameter, schedule 40, stainless steel sections, with standard flanges. 

This diameter was chosen because it was the smallest within which the existing test ejectors would fit. The 

assembled unit is shown in figure 5(a). Ideally, it is desirable to make the shroud as small as possible. Doing so 

results in a more compact, lighter weight, combustor for a given application. Furthermore, a small diameter in the 

region surrounding the pulsejet driver would result in higher speed bypass flow, which would provide more 

convective cooling effectiveness for the hottest section. In the present tests, the surface temperature of the pulsejet 

component exceeded 1700 °R in some regions. However, acoustic considerations must be made when selecting the 

shroud diameter. The diameter, or cross sectional profile of the shroud has a clear impact on its acoustic properties 

and these, in turn, can have a profound effect on the pulsejet performance. In fact, it was found that not only the 

shroud diameter, but its overall length and the hardware used to terminate each end (a crude diffuser on the inlet and 

a flat plate at the exhaust) produced an acoustic, or weak wave structure that profoundly affected the pulsejet 

operation to the point that it would not run, or stopped running shortly after commencing operation. In an attempt to 

alleviate some of this interaction, a perforated, mild steel liner was installed along the entire length of the shroud, 

with a standoff distance from the shroud wall of approximately 0.25 in. No rigorous selection of hole size or solidity 

was made in constructing the liner. Conveniently available materials were used. The liner can be seen installed in 

one of the shroud sections in figure 5(c). Here, several photographs of rig components, as well as the assembled 

shroud, are shown. The liner proved to be somewhat helpful in that it allowed operation with some configurations 

that was not otherwise possible. A more thorough design effort would likely have resulted in a much more effective 

article. 

(a) 

Diameter=D

θ=4°

(b) 

Figure 4.—Schematics of straight (a), and tapered

(b) ejectors used in the experiment. 
NASA/TM—2005-213854 5

D. Instrumentation 

The primary measures of interest for the experiment were the total pressure and temperature ratios across the rig, 

and the total pressure fluctuations in the flow downstream of the ejector. The total temperature ratio indicates the 

amount of chemical energy being added, the total pressure ratio indicates the benefit obtained from the concept 

(see fig. 1) and the total pressure fluctuation level indicates the degree of unsteadiness in the flowfield aft of the 

ejector. The pressure ratio was actually calculated from a pressure difference, namely: 

upstream Pt

P PR 1 ∆ = + (1) 

Where upstream

t downstream ∆P = Pt − P . The upstream total pressure was measured using a single total pressure probe*

located several inches upstream from the pulsejet inlet, and centered in the shroud (refer to fig. 2). The probe was 

connected to a low frequency transducer via approximately 4 ft. of 0.063 in. O.D. stainless steel tubing in order to 

damp any acoustic oscillations. The transducer signal was also filtered using a 10 Hz. low-pass filter, after which it 

was amplified and time averaged over a 5.0 second period. In a highly unsteady environment, mass-averaged values 

are more appropriate than time-averaging. As will be seen however, the unsteadiness in this region is relatively 

small (compared to the pulsejet combustion chamber for example), and time averaging is warranted. 

 

*

 Total pressure probes were constructed by forming an ‘L’ shaped section of 0.063 in. O.D. (0.031 in. I.D.) stainless 

steel tubing, and inserting it into a 0.13 in. O.D. sleeve to add rigidity. 

(b) (c) 

Figure 5.—Photographs of the pulsejet-driven, pressure-gain combustor rig: (a) assembled unit,

(b) downstream probe arrangement, and (c) view of liner. 

(a)
NASA/TM—2005-213854 6

The downstream total pressure was measured using 

a single total pressure probe mounted 1.125 in. radially 

from the shroud wall. The axial location is shown in 

figure 2. The radial position can be seen on the left side 

of figure 5(b). In this figure the probe is labeled 

Pt(d.c.). Note that there is another total pressure probe 

shown in figure 5(b). It is 2.0 in. from the shroud wall. 

A third probe was also temporarily mounted directly 

along the centerline. Preliminary tests for flow 

uniformity were conducted by taking pressure readings 

from these probes with the pulsejet operational and with 

a tapered ejector installed. Pressure ratio, as calculated 

from equation (1) varied by less than 0.1 percent 

between the three probes. Thus, the use of only one 

probe was justified. A tubing section of 4 ft. length, 

0.063 in. O.D. was connected to one side of a low 

frequency differential pressure transducer. The other 

side of this transducer was connected to upstream total 

pressure probe. Like the upstream total transducer, the 

differential transducer signal was 10 Hz. low-pass 

filtered, amplified, and averaged over a 5.0 second period. 

Total pressure fluctuations (unsteady pressures) were measured using two probes with high frequency Kulite 

(model# XTEH-7L) transducers installed directly inside. They can be seen in a head-on view in figure 5(b) where 

they are labeled as Pt(a.c.). A side view of one of the probes is shown in figure 6. The transducer face is rated to 

1210 °R; however, the body is only rated to 985 °R. Thus, a small amount of backside cooling air was required for 

the probe, and was supplied as shown in the figure. Although the transducer was rated for high temperature, it was 

designed for operation at a consistent temperature. The nature of the testing in this experiment entailed relatively 

short duration hot-runs followed by lengthy cool down periods. As such, the transducer experienced large thermal 

transients and did not operate at any single temperature. This, in turn, led to relatively large shifts in the absolute 

pressure reading that it supplied. However, the nature of the transducer thermal response is to shift the linear 

relationship between pressure and output voltage up or down. The slope remains valid. The transducers were 

therefore a.c. coupled to the acquisition system, where they were used to obtain valid pressure fluctuations readings 

as deviations from the time-averaged values measured with the low frequency probes. 

Upstream total temperature was measured using a single shielded Type-K thermocouple in the same axial 

position as the upstream total pressure probe. Downstream total temperature was measured using a four-probe 

arrangement of thermocouples, positioned at different radial locations, and connected in parallel to provide an area 

averaged reading. The radial positions can be seen in figure 5(c). Axially, the thermocouple probes were positioned 

just downstream of the downstream total pressure probes, as shown in figure 2. No filtering or time-averaging was 

applied to the thermocouple signal. 

In addition to the overall performance measuring instrumentation just described, instrumentation was also 

available to measure total air flow rate, time-averaged pulsejet combustion chamber pressure, fuel flow rate, and the 

pulsejet outer surface temperature. The air flow rate measurement system was a standard, low-loss, venturi-type. 

The pulsejet chamber pressure, a metric useful for determining the performance of the operational pulsejet 

independent of the overall system performance, was measured using a low frequency transducer coupled to a 

4 ft. long, 0.063 in. O.D., stainless steel tubing section, which penetrated the combustion chamber, and was 

terminated flush with the inside wall (see fig. 3). As with the other transducers of this sort, the signal was low-pass 

filtered at 10 Hz., and time-averaged over a 0.2 sec. interval. Fuel flow rate was measured by taking readings of the 

fuel tank level before and after each run, multiplying the calibrated volume change by the known density of 

gasoline, and dividing the result by the test run time. This was often an invalid method of measuring the fuel flow 

rate because the run time was measured from the moment that the pulsejet commenced self-sustained operation until 

it was shut down. The measured fuel consumed included that of the run and that used for starting. As mentioned, the 

pulsejet is started using brief bursts of high pressure air direct at the inlet. These induce both air and fuel into the 

pulsejet combustion chamber. Multiple bursts are often required for successful starting which results in fuel being 

consumed before the test time begins. This, in turn leads to average fuel flow rate calculations that are too high. As 

will be discussed later however, they provide a rough measure of fuel flow which, along with the air flow 

Kulite Transducer

Shroud Wall

Backside Cooling Air

CL

Low Frequency Probe

Figure 6.—High frequency pressure transducer

installation. 
NASA/TM—2005-213854 7

measurements, can be used to determine overall stoichiometry. This can in turn be used to predict downstream 

temperature and compare with measured values. 

E. Test Operation 

Typical test runs were 15 seconds long once resonant operation was established. Resonant operation was 

detected audibly by the operator. As mentioned, portions of the pulsejet become red-hot during test runs and it was 

decided that a 15 second time limit would minimize hot, cyclic damage. Furthermore, the reed valves used in the 

pulsejet have a limited number of cycles before they fail. In an effort to avoid excessive valve replacement, which 

requires a teardown of the rig, short test times were sought. Test data was acquired during the last 5 seconds of each 

run. 

Subsequent to testing it was found that while 15 seconds is sufficient time to establish steady readings for 

average pressure, it is not sufficient for the thermocouples. This is seen clearly in figure 7 which shows transducer 

output of rig pressure difference, ∆P, and downstream temperature during a run of extended duration. Also shown in 

the figure are the normal run time window, and the data acquisition time during which time-averages are made and 

point measurements are obtained. The pressure difference, from which overall pressure ratio is obtained achieves a 

level profile relatively quickly†

. The thermocouple readings are still rising during the acquisition time and it appears 

that the true steady-state temperature is 100 to 150 °R higher than what was recorded. Analysis later confirmed that 

this rise-time was consistent with the dimensions of the thermocouple sensing element. 

The performance figure of merit for this system is overall pressure ratio as a function of overall temperature 

ratio. Therefore, errors in temperature measurement will impact results. As a crude correction for the error just 

described, 100 °F was added to each of the measured temperature data points to be presented. 

III. Results 

A. Ejector Optimization 

Each of the ejectors described was run in the shrouded rig over range of overall air mass flow rates (and 

therefore temperature ratios) and the corresponding pressure ratio was measured. The objective was to find the 

ejector which produced the largest pressure ratio for a given temperature ratio. In the context of this experiment, this 

will be referred to as an optimization process. For this portion of the testing, the largest of four available spacer sets 

for the exhaust flow gap setting was used. This produced the lowest shroud pressures for a given flow rate, and the 

least variation in shroud pressure as mass flow was varied. As will be discussed in subsequent sections, it also 

allowed for the greatest variation overall temperature ratio due to unexplained operability issues that arose at higher 

shroud pressures. The results of the tests are shown in figures 8 through 10. Here overall total pressure ratio is 

 

†

 Most runs actually showed a much smoother pressure trace than that of figure 7 during the acquisition period. This 

trace may indicate the early stages if valve failure. 

400

600

800

1000

1200

1400

0 10 20 30

Time, sec.

Temperature, deg. 

R

-0.10

0.10

0.30

0.50

0.70

0.90

P, psi

Temperature

Pressure Difference

Run Time

Acquisition Time

∆P, psi

Figure 7.—Traces of rig pressure difference, ∆P, and

downstream temperature during a run of extended

duration. 

1.00

1.01

1.02

1.03

1.04

1.90 2.10 2.30 2.50

Total Temperature Ratio

Total Pressure Ratio

No Ejector

D4-L3.1

D4-L7.4

D4-L12.4

D2.2-L3.4

Figure 8.—Total pressure ratio as a function of total 

temperature ratio for the 4.0 and 2.2 in. diameter 

straight ejectors. The number adjacent to L in the 

legend denotes the ejector length in inches. 
NASA/TM—2005-213854 8

plotted as a function of overall temperature ratio for each of the ejector configurations tested. Ejectors are denoted 

by a diameter and length specification (in inches) in the plot legend. For each plot the maximum temperature ratio 

measured was set by thermal limitations of instrumentation. The minimum temperature ratio measured was 

generally established by a failure of the pulsejet to operate at the required overall flow rate. No results are shown 

from the longest (L = 16.8 in.) straight ejectors because in these configurations the pulsejet would not start. For this 

same reason only results from the shortest length of the D = 2.2 in. ejectors are shown. The data points from each 

ejector tested have been, fit with a polynomial. The fit has no physical basis and is made in order to clarify trends. 

Examining figures 8 through 10, three conclusions are evident. First, there is a slight pressure gain obtained with 

no ejector present in the system. This was expected since the shroud itself, with a diameter of 6.0 in., is actually a 

large ejector, albeit one with poor performance. It is noted that a pressure drop was verified in every configuration 

when air was forced through the system without the pulsejet operating. Thus, the instrumentation was performing as 

expected. Second, the presence of an ejector in a pulsejet-driven combustor can improve performance. Third, the 

level of performance is sensitive to both the diameter and length of the ejector. 

The 3.0 in. diameter ejectors (tapered and straight) produced the highest pressure ratio for a given length. This is 

consistent with thrust augmentation measurements made with the same ejectors and pulsejet driver, and described in 

reference 9. In that experiment the highest levels of thrust augmentation were also found with a 3.0 in. diameter 

ejector. That high thrust augmentation and pressure-gain are related is a somewhat intuitive observation; however, it 

should be kept in mind that in the shrouded configuration of the present experiment, the ejector is not operating as it 

does in the open air (though the pulsejet is). For the reference 9 thrust augmentation experiment, entrainment ratios 

( ejector mpulsejet m /  ) in excess of 16 were measured. This is flow induced though the ejector by the driver and is 

directly related to thrust augmentation. In the present experiment, the entrainment ratio cannot be measured directly; 

however, it can be estimated from the known overall air flow rate and the calculated thrust of the pulsejet (described 

below). Typical estimated entrainment ratios were less than 5. Hence, the ejector is functioning somewhat ‘off 

design,’ but evidently effectively. 

In general it was found that for all the ejector diameters tested, the performance improved as the ejector length 

grew. However, for both the tapered and straight 3.0 in. diameter ejectors the differences in performance between 

the longest and intermediate length ejectors was much less than between ejectors of intermediate and shortest length. 

In fact, given the scatter in the data, the performance of the long and intermediate length, straight and tapered, 3.0 in. 

diameter ejectors can be considered comparable. For unexplained reasons however, starting the pulsejet was easiest 

when the intermediate length tapered ejector was installed, and thus it was chosen for further testing at higher 

pressures. Clearly it is a configuration that is representative of the highest performance attained, and will therefore 

be termed the optimal design. 

1.00

1.01

1.02

1.03

1.04

1.90 2.10 2.30 2.50

Total Temperature Ratio

Total Pressure Ratio

No Ejector

D3-L3.4

D3-L7.6

D3-L12.6

Figure 9.—Total pressure ratio as a function of total

temperature ratio for the 3.0 in. diameter straight

ejectors. The number adjacent to L in the legend

denotes the ejector length in inches. 

1.00

1.01

1.02

1.03

1.04

1.90 2.10 2.30 2.50

Total Temperature Ratio

Total Pressure Ratio

No Ejector

D3-L8.3

D3-L10.8

D3-L13.3

Figure 10.—Total pressure ratio as a function of total 

temperature ratio for the 3.0 in. diameter tapered 

ejectors. The number adjacent to L in the legend

denotes the ejector length in inches. 
NASA/TM—2005-213854 9

B. Pulsejet Performance 

When examining the overall performance of the system as presented in figures 8 through 10, it is not evident 

whether changes result from the different ejectors or altered performance of the pulsejet. This can be discerned 

however, since the pulsejet was equipped with a low frequency static pressure sensor mounted on the combustion 

chamber. It was found in previous studies (ref. 9) that the difference between the time-averaged combustion 

chamber pressure, pcc and the upstream or ambient pressure correlated well with the thrust according to the relation 

( ) F = Aexit pcc − Pupstream (2) 

where Aexit is the pulsejet tailpipe area. This relation assumes that the shroud Mach number is very low and that the 

pulsejet exit pressure is essentially the same as the upstream pressure. If the upstream pressure (and therefore 

density) in the shroud for all of the test points were constant, then equation (2) could be used directly to measure 

whether the pulsejet performance was changing, and thus isolate the mechanism responsible for the changes in 

overall pressure ratio seen in figures 8 through 10. As mentioned however, the overall temperature ratio is controlled 

by the overall air mass flow rate which, for a given fixed exhaust gap spacing, also alters the upstream pressure. 

Figure 11 illustrates this pressure variation with mass flow for each of the four exhaust gap spacers used. The 

optimization testing was performed with the largest exhaust spacing which, from figure 11, and the known near 

constant upstream temperature of 520 °R, indicates upstream density variations of nearly 40 percent. For the 

subsonic pulsejet flow the thrust can also be written as 

⎟

⎠

⎞ ⎜

⎝

⎛ ≈ ρ − ′ 2 2

c PJ

exit

upstream u u

g TR

A F (3) 

where ρupstream is the upstream density, gc is the Newton constant, TRPJ is the pulsejet (not the overall) total 

temperature ratio, and 2

u and 2u′ are the square averages of the mean and periodically fluctuating velocity 

components at the pulsejet exit plane. Thus, according to equation (3), the thrust of the pulsejet should change at 

different test points simply because the density is changing. In other words, the pulsejet thrust is increasing because 

more flow is passing through it. 

With this in mind, the thrust per unit area ( F Aexit ) is shown in figure 12 for all of the test points, both with and 

without ejectors present, as a function of upstream density. There is substantial scatter in the data, which is typical of 

pulsejet operation; however; it is apparent that a nearly linear trend exists. Each of the data sets (with and without 

ejectors present) has been fit with a linear regression, which is also shown. The dashed line in the figure is a linear 

regression fit of the combined data sets. The ejector optimization was performed with the largest exhaust gap 

spacing and it is evident that in this region of figure 12 (the lower left sector outlined in green) no substantial 

differences exist in the pulsejet performance with or without ejectors. Hence, the overall performance enhancements 

3

4

5

6

7

8

9

0.075 0.095 0.115 0.135 0.155

Upstream Density, lbm/ft3

 Thrust per Unit Area, psi

No Ejector

All Ejectors

gap=0.070 in.

gap=0.090-0.125 in.

Figure 12.—Thrust per unit area for all of the test 

points, with and without ejectors present, as a function 

of upstream density. 

0

5

10

15

20

25

0.15 0.35 0.55 0.75

Air Mass Flow Rate, lbm/s

Upstream Pressure, psig

gap=0.125 in.

gap=0.110 in.

gap=0.090 in.

gap=0.070 in.

Upstream Pressure, psig

Figure 11.—Upstream total pressure ratio as a function

of air mass flow rate for the four exhaust gap spacers

used. 
NASA/TM—2005-213854 10

seen in figures 8 through 10 are due to the ejectors, not the pulsejet itself. There is a more substantial difference 

between pulsejet performance with and without the ejector at the highest upstream densities when the smallest 

exhaust gap was used. The reason for this difference is unclear at the time; however, it is a maximum 10 percent 

discrepancy as assessed by the respective regression fits, which is acceptable. It is interesting to note as an aside that 

the overall temperature ratios shown in figures 8 through 10 decrease as the overall air flow rates, and hence 

upstream densities increase. But from figure 12, the pulsejet thrust increases with upstream density. Since overall 

pressure ratio generally decreases with decreasing temperature ratio in figures 8 through 10, it may be inferred that 

the lowest overall pressure ratios are attained when the highest thrust is being delivered from the pulsejets. 

Evidently, the additional thrust, or pumping energy is negated by the increased bypass flow which must be pumped. 

C. Higher Pressure Testing of Optimal Ejector 

Each of the three remaining exhaust gap spacer sets was installed in the rig and with each one the pulsejet was 

operated over a range of overall flow rates, both with and without the optimal ejector in place. The results are shown 

in figure 13. As in figures 8 through 10, overall pressure ratio is shown as a function of overall temperature ratio. 

The smaller the exhaust gap used, the higher the shroud pressure maintained for a given overall air flow rate. 

Smaller gaps also led to larger variations in shroud pressure for a given change in air mass flow rate, as seen in 

figure 11. Thus, as overall temperature ratio was changed by varying air mass flow, the upstream or shroud pressure 

was also changing for any exhaust gap spacer used. The trend of figure 13 demonstrates that the performance of the 

rig was not strongly affected by the upstream pressure and density variations. For a given temperature ratio 

achieved, the overall pressure ratio was the same regardless of the upstream pressure. It is noted however that there 

was a general rightward shift toward higher temperature ratios as the upstream pressure was raised. This may be 

explained as follows. 

The measured temperature rise across the rig for all of the test points is shown in figure 14 as a function of 

measured overall air fuel ratio. Also shown is the expected result as calculated for a C6H10 (a gasoline-like 

hydrocarbon)/air mixture using the STANJAN thermodynamic package (ref. 15). The measurement of fuel 

consumption was, as mentioned, crude and error prone, as evidenced by the broad scatter in the data. Nevertheless, it 

is clear from figure 14 that, on an overall basis, significantly less heat release is occurring than expected. The 

proposed (though not provable) reason for this is that the actual process taking place is a very rich reaction in the 

pulsejet followed by rapid mixing with the bypass flow. The mixed flow is sufficiently cool such that all further 

reaction, or heat release, virtually ceases. This process would give the appearance of the reduced fuel heating value 

seen in figure 14. It would also result in an exhaust flow filled with unburnt hydrocarbons. No measurements of such 

were taken during testing, but the odor that is characteristic of such flows was present in the test facility after each 

run, which tends to confirm the proposed process. Furthermore, it is well documented that pulsejets such as the one 

used in the present experiment do, indeed, run rich. Stoichiometric combustion yields higher exhaust temperatures 

than does rich combustion. Thus, if operation of the rig at higher upstream pressure led to leaner (more 

1.00

1.01

1.02

1.03

1.04

1.90 2.10 2.30 2.50

Total Temperature Ratio

Total Pressure Ratio

No Ejector

gap=0.125 in.

gap=0.110 in.

gap=0.090 in.

gap=0.070 in.

Figure 13.—Total pressure ratio as a function of total

temperature ratio for the 3.0 in. diameter,

10.8 in. long tapered ejector, with elevated upstream

pressure (reduced exhaust gap spacing). 

400

600

800

1000

1200

1400

40 90 140

Air/Fuel Ratio

Tdownstream-Tupstream, deg. R

measured

STANJAN-C6H10

Figure 14.—Measured temperature rise due to 

combustion as a function of air/fuel ratio. Calculations 

from STANJAN code for C6H10 fuel and air are also 

shown. 
NASA/TM—2005-213854 11

stoichiometric) operation of the pulsejet, then when this hotter flow mixed with the bypass flow the resulting 

downstream temperature would be higher as seen in figure 13. It has already been shown in figure 12 that the 

increase in pulsejet thrust at higher upstream pressures is linearly related to increased air mass flow through the 

device if the increased air mass flow is due to higher density, not higher velocity. This is proposed here. Recalling 

from figure 3 that the fuel delivery system is a venturi-type, it can be deduced from Bernoulli’s equation that the 

pressure differential which drives a venturi flow is linearly related to the density of the incoming air. However, the 

flow rate of fuel through the regulating jet of the fuel system is related to the square root of the venturi pressure 

differential. In other words, as the upstream density increases, both the air and fuel flow through the pulsejet 

increase; however, the air flow rate increases more than the fuel flow rate. Hence, the pulsejet operates leaner, emits 

a hotter flow, and causes the overall temperature ratio in the rig to increase as seen in figure 13. 

D. Unsteadiness Measurements 

Unsteady pressure fluctuations downstream of the 

ejector were obtained for all runs using the two high 

frequency probes described earlier. As described earlier, 

these transducers measured deviations in pressure from 

the mean value. Fluctuation levels were recorded as root 

mean square (rms) pressure readings, averaged over a 

5.0 second interval. The rms readings from the two probes 

were then averaged to provide a single measure of 

unsteadiness (the two readings differed from one another 

by less than 0.7 percent). These measurements are 

presented in figure 15. The rms pressure fluctuations have 

been normalized by the time-averaged downstream total 

pressure, and are plotted as a function of overall 

temperature ratio. Data for all upstream pressures 

(exhaust gap settings) are shown. Also shown in the 

figure is the baseline rms level obtained when there was 

no flow in the rig. This represents electrical or 

background noise. For configurations both with and 

without the so-called optimal, tapered, 3.0 in. diameter, 

10.8 in. long ejector, the rms pressure appears to stay at 

approximately 4.5 percent of the mean value regardless of 

the temperature ratio or upstream pressure. That is to say, the presence of the ejector doesn’t seem to mitigate or 

worsen the unsteadiness compared to having no ejector at all. For reasons that are not clear however, the straight 

3.0 in. diameter, 12.6 long ejector does show a reduction in unsteadiness to a level of approximately 3.2 percent of 

the mean. Given that the pressure gain performance of this ejector was comparable to the tapered ejector, it may be 

the better candidate for the label of optimal. Of the two however, it was still the more difficult with which to start 

the pulsejet. Fluctuation data from all of the other ejectors tested is also shown in figure 15. Some of this data shows 

levels somewhat below the 4.5 percent average of the optimal ejector; however, the pressure gain performance of 

these ejectors was too low for the reduced fluctuation results to be worthwhile. 

The worst of the fluctuations shown in figure 15 is, of course, well below the level seen in the pulsejet 

combustion chamber. While such measurements were not available in this experiment, they have been made 

previously (ref. 9). Typical normalized rms levels are 0.36, or nearly an order of magnitude above those seen in the 

shroud. Thus, the use of a shrouded ejector is an effective means for damping pulsations from unsteady combustion. 

Whether the pressure fluctuation levels seen in the shroud are compatible with efficient operation of a downstream 

turbine is an open question. However, those presented here suggests that an answer in the affirmative is at least 

plausible. 

It is worth noting that in a previous open-air pulsejet driven ejector experiment, the time-dependant velocity field 

near the exit of the ejector and at the exit of the pulsejet was mapped using digital particle imaging velocimetry 

(DPIV). The pulsejet was the same one used in this experiment. The ejector was not performance optimized, having 

a 2.5 in. diameter, and only a 7 in. length. It was found that the standard deviation of the centerline velocity near the 

ejector exit was 33 percent of the mean velocity. At the exit plane of the pulsejet, the standard deviation of the 

centerline velocity was 136 percent. Thus, it appears that both velocity and pressure fluctuations may be 

substantially mitigated by the use of a shrouded, pulsejet-driven ejector system. 

0.00

0.01

0.02

0.03

0.04

0.05

0.06

1.90 2.10 2.30 2.50

Total Temperature Ratio

rms pressure/ Pdownstream

No Ejector

D3-L10.8 Tapered

D3-L12.6 Straight

All Others electrical noise

Figure 15.—Measured normalized rms total pressure 

fluctuation downstream of the ejector as a function of

overall temperature ratio for all runs. 
NASA/TM—2005-213854 12

The power spectra from the high frequency total pressure transducers in the rig provide some insight into the 

content of the fluctuations. Figure 16 shows results from 1.0 second traces of the highest pressure configuration with 

and without the optimal ejector present. The data was sampled at 10 kHz. The quantities plotted represent the 

contribution to the signal, in terms of amplitude (not power) at a given frequency. The amplitudes have been 

normalized by the mean pressure. While the dominant frequency in both cases corresponds to the operational 

frequency of the pulsejet, there are significant contributions at other frequencies as well. This distribution is 

interesting as it differs markedly from that of the pulsejet itself. Figure 17 shows the normalized power spectra from 

a transducer mounted on the combustion chamber wall during previous open air testing (ref. 9). Here, the 

operational frequency completely dominates, and the relatively strong 900 Hz. signal seen in figure 16 is virtually 

non-existent. It is believed that the differences in frequency distributions between the two figures are the result of 

interactions with the shroud. The fact that the shroud fluctuations are not a “pure tone,” but are instead composed of 

multiple frequencies may be beneficial in terms of their impact on downstream turbomachinery performance in any 

practical application of this type of combustor. 

IV. Discussion 

From the results presented thus far, it is clear that a pulsejet-driven, ejector-based, pressure-gain combustor does, 

in fact, work. That is, it provides a pressure rise, which is a thermodynamic benefit to a gas turbine engine cycle. 

A. Issues 

Many issues remain before such a device can be implemented however, and it is worthwhile to note some of 

them. They include: 

1.Cooling.—The pulsejet in this paper measured a surface temperature of nearly 2000 °R during operation. 

While the stresses are relatively low, this is an extremely hot, life limiting temperature that will only rise when 

the combustor faces the elevated discharge temperatures of an upstream compressor. It is possible that 

equipping the pulsejet with external fins and using a smaller shroud will allow the entrained bypass air to more 

effectively cool the walls. Of course, this will come at a performance penalty; as such cooling will carry with it 

a pressure drop for the bypass air. 

2.Turbine Cooling.—Any pressure rise device causes difficulties in this area because turbine cooling flows can 

no longer be supplied by the upstream compressor. Thus, the pulsejet itself must somehow supply the cooling 

air, or a small boost compressor must be installed. The latter option is extremely unlikely. 

3.Integration.—A combustor such as the one presented in this paper is volumetrically large, as well as long 

compared to its conventional counterpart. 

4. Valves.—The pulsejet used in this experiment operates with simple reed valves constructed from spring steel. 

These tend to fail often, and even more so at high pressure. Of course, the valve design is very crude, having 

undergone few, if any modifications since the 1950’s. It may be possible to extend life considerably using 

improved materials or design. It may also be possible to use valveless pulsejets, as was done by Kentfield in 

Ref. 5. 

0 500 1000 1500

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

Frequency, hz

Normalized Amplitude

D3−L10.8 Tapered

No Ejector

Figure 16.—Normalized power spectra obtained from

high frequency pressure transducers with and without

the optimal ejector present. 

0 500 1000 1500

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

Frequency, hz

Normalized Amplitude

Figure 17.—Normalized power spectra obtained from a 

high frequency wall-mounted pressure transducer on 

the combustion chamber of the pulsejet during open air

testing. 
NASA/TM—2005-213854 13

This somewhat daunting, and incomplete list should not be interpreted as a roadblock to the technology. The first 

three of the four items listed are common to nearly all conceivable pressure-gain systems (e.g., wave rotor, 

detonative combustion, etc.). Any efficiency enhancing concept comes with practical hurdles which must be 

identified, assessed in terms of risk, and overcome if the payoff is deemed worthy. The current concept is no 

exception. Furthermore, it is worthwhile noting that the combustion portion of a Brayton (gas turbine) cycle 

produces far more entropy than the compression of expansion portions. As such, the largest payoffs in terms of cycle 

efficiency gains are likely to come from concepts which reduce combustion induced entropy production such as the 

pulsed combustion concept introduced here and elsewhere(refs. 3 to 6). 

B. Low Emissions 

Pulsejets tend to operate rich, as mentioned. It is not clear whether this is an inherent necessity to their proper 

operation, or whether it results from poor control of the metered fuel. However, in the present combustor 

embodiment, it may be possible to use this operational characteristic to some advantage. The ejector which follows 

the pulsejet has the ability to rapidly mix the bypass flow with the pulsejet discharge. This may be inferred from the 

fact that good ejector performance relies on a high level of mixedness. This is how momentum is transferred from 

primary to secondary flow. As pointed out earlier, pulsejet driven ejectors have exceptional performance in terms of 

thrust augmentation, and do so with relatively short ejectors. Thus, there must be substantial mixing in a relatively 

short distance, implying that it takes place quite fast. This process may be seen to some degree in figure 18 which 

shows results from the reference 10 DPIV experiment using an open-air pulsejet and a transparent ejector. 

Figure 18(a) shows contours of vorticity downstream of the pulsejet and in the ejector just after the characteristic 

vortex ring has formed. The color spectrum is such that yellow indicates large positive (counterclockwise) vorticity 

and blue/black represents large negative vorticity (clockwise). Red indicates no vorticity. Note that the secondary 

flow is irrotational and cannot become rotational except by mixing with the flow from the pulsejet. The volume of 

the vortex ring corresponds closely to the volume of gas that has discharged from the pulsejet as measured by 

integrating the near-exit plane velocity profile. Thus, little mixing has taken place. The vortex travels quickly 

downstream and appears to disintegrate once inside the ejector. In that time however, much entrainment and mixing 

takes place. Figure 18(b) shows vorticity contours on the same scale, approximately 1.0 msec later. This point was 

chosen since it is the moment when the ejector becomes 

completely filled with rotational fluid. The volume of 

rotational fluid in figure 18(b) is approximately 2.3 times 

the volume discharged from the pulsejet by this time. 

Thus, in the span of 1.0 msec. the ejector has become 

filled with a mixture of pulsejet effluent and entrained air. 

With this sort of rapid mixing, it is possible that the 

configuration presented in this paper is a manifestation of 

a practical Rich-burn/Quench/Lean-burn (RQL) 

combustor (ref. 17). The RQL concept is intended to 

reduce NOx emissions by minimizing the time at nearstoichiometric conditions where the rate of NOx

production is the greatest. Instead it relies on a sequence 

whereby a low NOx producing, rich combustion process 

releases some heat, and produces unburnt hydrocarbons. 

The flow is quenched, or rapidly mixed with air, and 

completes the reaction of, and heat release from, the 

unburnt hydrocarbons via a lean process which also 

produces little NOx. Of key importance in this sequence is 

rapid mixing. In steady RQL combustors this is achieved 

through techniques such as cross-flow injection of the 

quenching air. Rapid mixing in the present system is 

affected by the vortex emitted from the pulsejet and 

possibly the highly turbulent, intermittent jet that follows 

it. The notion could not be tested in the current 

experiment, because the bypass air was unheated. As a 

result, the rapid mixing which did occur resulted in a lean 

mixture that was so cool as to render the reaction rate, 

which is exponentially related to temperature, essentially 

Vortex

Pulsejet Ejector

Inlet Region (not visible)

(a) 

Pulsejet Ejector

(b)

Figure 18.—DPIV measured contours of vorticity in a 

pulsejet-driven ejector flowfield at two different times 

in the cycle, (a) t = 0.0 msec., (b) t = 1.0 msec. Data is 

from reference 10.
NASA/TM—2005-213854 14

zero. With the current emphasis on low emissions aircraft, this feature of the pressure-gain combustion system, if 

realizable, may be of greater value than the increased efficiency afforded by the pressure-gain itself. 

V. Conclusion 

A laboratory rig demonstration of a pulsejet-driven, ejector-based, pressure-gain combustion system was 

presented. It was shown that at temperature ratios comparable to those of modern gas turbine combustors, pressure 

ratios near 1.035 were obtained. These pressure ratios, in comparison to those of conventional combustors, may 

provide a 2 to 3.5 percent reduction in specific fuel consumption when used in a gas turbine engine. It was also 

shown that the resulting unsteadiness levels, when measured in terms of rms pressure fluctuations, were only about 

4.5 percent of the mean total pressure. Such levels are near the acceptability limit in terms of reducing downstream 

turbine performance. Finally, it was shown that the concept may yield a low emissions combustor by affecting a 

practical, low loss, Rich-burn/Quench/Lean-burn sequence. 

Appendix: Baseline Engine Information for Figure 1 

The following tabulated parameters constitute the baseline turbojet and turbofan engine data used in the cycledeck described to generate the graph shown in figure 1. 

Parameter Turbofan Turbojet 

Overall Pressure Ratio (OPR) 30.00 8.00 

High Pressure Compressor Ratio 7.36 8.00 

Low Pressure Compressor Ratio 2.50 1.00 

Fan Pressure Ratio 1.63 ---- 

Fan bypass Ratio 6.0 ---- 

Compressor Efficiency 0.90 0.90 

Turbine Efficiency 0.90 0.90 

Fan Efficiency 0.90 0.90 

Flight Mach Number 0.80 0.80 

Ambient Temperature (°R) 410 410 

Combustor Exit Temperature (°R) 2968 2400 

Cooling Flow (fraction of core) 0.21 0.00 

Chargeable Cooling fraction 0.40 ---- 

Ratio of Specific Heats 1.35 1.35 

Real Gas Constant (ft-lbf/lbm/R) 53.57 53.57 

Burner Pressure Ratio 0.95 0.95 

Fuel Heating Value (BTU/lbm) 18400 18400 

Specific Thrust (lbf-s/lbm) 18.26 75.86 

Specific Fuel Consumption (lbm/hr/lbf) 0.585 1.109 

References 

1. Lefebvre, Arthur, H., Gas Turbine Combustion, Taylor and Francis, 1999, Chapt. 4. 

2. Oates, Gordon, C. “Aerothermodynamics of Gas Turbines and Rocket Propulsion,” American Institute of 

Aeronautics and Astronautics, 1997. 

3. Catchpole, B.G., and Runacres, A., “Constant Volume Gas Turbine Experiments with Gaseous Fuel in a 

Rotating Pocket Combustor,” Mechanical Engineering Note 353, Australian Defense Scientific Service 

Aeronautical Research Laboratories, October, 1974. 
NASA/TM—2005-213854 15

4. Mathur, Atul, “A Brief Review of the G.E. Wave Engine Program,” Proceedings of the 1985 ONR/NAVAIR 

Wave Rotor Research and Technology Workshop, May 1985. 

5. Kentfield, J.A.C., and O’Blenes, M.J., “Small Gas Turbine Using a Second-Generation Pulse Combustor,” 

AIAA Journal of Propulsion, vol. 6 no. 2, 1990, pp. 214–220. 

6. Gemmen, R.S., et. al., “Achieving Improved Cycle Efficiency Via Pressure Gain Combustors,” ASME 95-GT63, June 1995. 

7. Snyder, P.H., Alparslan, B., Nalim, M.R., “Gas Dynamic Analysis of The CVC, A Novel Detonation Cycle,” 

AIAA paper 2002–4069, July 2002. 

8. Paxson, D.E., Perkins, H.D., “Thermal Load Considerations for Detonative Combustion-Based Gas Turbine 

Engines,” AIAA–2004–3396, July 2004. 

9. Paxson, D.E., Wilson, J., and Dougherty, K.T., “Unsteady Ejector Performance: An Experimental Investigation 

Using a Pulsejet Driver,” AIAA paper 2002–3915, July 2002. 

10. John, W.T., Paxson, D.E., Wernet, M.P., “Conditionally Sampled Pulsejet Driven Ejector Flow Field Using 

DPIV,” AIAA paper 2002–3231, June 2002. 

11. Lockwood, R.M. “Interim Summary Report on Investigation of the Process of Energy Transfer from an 

Intermittent Jet to Secondary Fluid in an Ejector-Type Thrust Augmenter,” Hiller Aircraft Report No. ARD286, March 1961. 

12. Paxson, D.E., “2003 Pulse Detonation Engine Project: University/Government Spring Ejector Meeting,” 

unpublished presentation, 2003. 

13. Litke, P.J., Schauer, F.R., Paxson, D.E., Bradley, R.P., and Hoke J.L., “Assessment of the Performance of a 

Pulsejet and Comparison with a Pulsed-Detonation Engine,” AIAA paper 2005–0228, January 2005. 

14. Wilson, J, Paxson, D.E., “Unsteady Ejector Performance: An Experimental Investigation Using a Resonance 

Tube Driver,” AIAA paper 2002–3632, July 2002. 

15. M.W. Chase Jr., C.A. Davies, J.R. Davies, Jr., D.J. Fulrip, R.A. McDonald, and A.N. Syverud, JANAF 

Thermochemical Tables, 3rd edition, Journal of Physical and Chemical Reference Data, vol. 14, Supplement 1, 

1985. 

16. John, W.T., Paxson, D.E., and Wernet, M.P., “Conditionally Sampled Pulsejet Driven Ejector Flow Field Using 

DPIV,” AIAA paper 2002–3231, June 2002. 

17. Rizk, N.K.; Mongia, H.C., “Low NOx Rich-Lean Combustion Concept Application,” AIAA paper 91–1962, 

June 1991. 
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NASA TM—2005–213854

AIAA–2005–4216

E–15224

WBS–22–066–10–05

21

Ejector Enhanced Pulsejet Based Pressure Gain Combustors: An Old Idea With a

New Twist

Daniel E. Paxson and Kevin T. Dougherty

Pulse detonation engines; Combustion; Thermodynamic cycles

Unclassified -Unlimited

Subject Category: 07

Prepared for the 41st Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE, and ASEE,

Tucson, Arizona, July 10–13, 2005. Daniel E. Paxson, NASA Glenn Research Center; and Kevin T. Dougherty, QSS

Group, Inc., 21000 Brookpark Road, Cleveland, Ohio 44135. Responsible person, Daniel E. Paxson, organization code

RIC, 216–433–8334.

An experimental investigation of pressure-gain combustion for gas turbine application is described. The test article consists of an offthe-shelf valved pulsejet, and an optimized ejector, both housed within a shroud. The combination forms an effective ‘can’ combustor

across which there is a modest total pressure rise rather than the usual loss found in conventional combustors. Although the concept

of using a pulsejet to affect semi-constant volume (i.e., pressure-gain) combustion is not new, that of combining it with a well designed

ejector to efficiently mix the bypass flow is. The result is a device which to date has demonstrated an overall pressure rise of

approximately 3.5 percent at an overall temperature ratio commensurate with modern gas turbines. This pressure ratio is substantially

higher than what has been previously reported in pulsejet-based combustion experiments. Flow non-uniformities in the downstream

portion of the device are also shown to be substantially reduced compared to those within the pulsejet itself. The standard deviation of

total pressure fluctuations, measured just downstream of the ejector was only 5.0 percent of the mean. This smoothing aspect of the

device is critical to turbomachinery applications since turbine performance is, in general, negatively affected by flow non-uniformities

and unsteadiness. The experimental rig will be described and details of the performance measurements will be presented. Analyses

showing the thermodynamic benefits from this level of pressure-gain performance in a gas turbine will also be assessed for several

engine types. Issues regarding practical development of such a device are discussed, as are potential emissions reductions resulting

from the rich burning nature of the pulsejet and the rapid mixing (quenching) associated with unsteady ejectors.