Robert J. Santoro and Sibtosh Pal

Pennsylvania State University, University Park, Pennsylvania

Thrust Augmentation Measurements Using a

Pulse Detonation Engine Ejector

NASA/CR—2003-212191

March 2003

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Robert J. Santoro and Sibtosh Pal

Pennsylvania State University, University Park, Pennsylvania

Thrust Augmentation Measurements Using a

Pulse Detonation Engine Ejector

NASA/CR—2003-212191

March 2003

National Aeronautics and

Space Administration

Glenn Research Center

Prepared under Grant NAG3–2657
Available from

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This report contains preliminary

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The Propulsion and Power Program at

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Available electronically at http://gltrs.grc.nasa.gov
NASA/CRó2003-212191 iii 

TABLE OF CONTENTS 

1. Introduction ........................................................................................................................... 1 

2. Key Accomplishments in FY02 ............................................................................................ 1 

 2.1. Thrust Stand Description........................................................................................... 2 

 2.2. Thrust Measurements for 1.3 in. Diameter Detonation Tube ................................... 4 

 2.3. PDE-Ejector Setup .................................................................................................... 6 

3. Measure of Technical Performance....................................................................................... 9 

4. Proposed Schedule and Relative Progress ............................................................................ 10 

5. Collaborations ....................................................................................................................... 10 

6. Interactions with NASA Glenn Research Centerís PDE Research Program........................ 10 

7. References ............................................................................................................................. 10 

 Appendix: PDE-Ejector Hardware........................................................................................ 11 
NASA/CRó2003-212191 1 

1. Introduction 

Pulse detonation engine (PDE) technology is currently receiving a great deal of 

consideration because of its potential for high efficiency combustion performance with reduced 

hardware complexity [1-2]. Although most of the current research efforts are focused on 

resolving the barriers to a pure PDE system, there are equally worthy options that involve the 

incorporation of a PDE as a component in hybrid engine concepts. The present NASA GRC 

funded three-year research project focuses on the study of a PDE driven ejector that would be 

applicable to a hybrid Pulse Detonation/Turbofan Engine. The ejector would be intended for 

applications that would replace the high-pressure compressor and high-pressure turbine sections 

of the core of a high bypass turbofan engine. The objective of the present study is to characterize 

the thrust augmentation achievable with a PDE-ejector and thus provide critical experimental 

data from which to assess the performance enhancements possible with this technology. 

The major potential advantages of the PDE-ejector as envisioned here include reduced 

costs due to the elimination of expensive compressor and turbine components with resulting 

reduced engine weight, along with improved specific fuel consumption and specific power 

inherent in the incorporation of a PDE component. There is currently insufficient data on the 

performance of such an ejector system to allow further development. Some of the key issues can 

be resolved if suitable measurements of the thrust augmentation were available over a suitable 

parameter space. Thus, at the present time the hybrid PD/Turbofan engine is at a Technology 

Readiness Level (TRL) of 2 and the present research intends to increase PDE-ejectors to TRL 3 

through the provision of critical experimental data that further defines the operational potential 

of this novel technology. 

The current research project on thrust augmentation experiments (NAG3-2657) was 

officially awarded to Penn State on January 30, 2002. The progress made on the project during 

the previous eight months in described in this report. 

2. Key Accomplishments in FY02 

The following tasks have been completed during the first eight months of the project. 

• A thrust stand for measuring average thrust for multi-cycle operation of a detonation tube has 

been designed and implemented. 

• The average thrust for an existing 1.3 in. diameter detonation tube has been measured and 

analyzed for ethylene as the fuel and a mixture of oxygen and nitrogen as the oxidizer. 
NASA/CRó2003-212191 2 

• Based on these initial measurements, a 2.25 in. diameter detonation tube has been designed 

and fabricated as the driver for the PDE-ejector thrust measurement experiments. 

The modular design of the system will allow experimentation to assess the effects of ejector 

geometry (diameter, length and shape) on thrust augmentation. The ejector section for the 

experiment has been designed and fabrication is near completion. Experiments will 

commence in the very near future. 

2.1. Thrust Stand Description

A thrust stand for measuring average thrust for multi-cycle operation of a detonation tube 

has been designed and implemented. A schematic of the system is shown in Fig. 1. The thrust 

stand features a solid table top on which the detonation tube is mounted on frictionless rails. 

A thrust block with a spring (k=14.4. lbf/in.) is attached to the injector assembly end of the 

detonation tube. Calibration of the assembly was conducted by measuring the displacement of 

the assembly for a range of forces. A force gage was used to supply a known force to the 

system, whereas a linear voltage displacement transducer (LVDT) was used to measure the 

displacement. The calibration curve for the system for the 1.3 in. diameter detonation tube 

(described in the next section) is shown in Fig. 2. The calibration results show that due to 

PDE

Frictionless rails

Engine supports

Thrust springs

Thrust block

Fig. 1. Schematic of thrust stand setup. 
NASA/CRó2003-212191 3 

frictional forces, and stiffness of the flex supply lines and ignition system, the system is not 

completely described by Hookeís law. For forces lower than about 5 lbf, the system does not 

move. However, for forces greater than 5 lbf, the measured displacement allows the applied 

force to be measured. 

The theoretical response of the system was also modeled. For the modeling, the system 

was considered to be frictionless. The damping of the system was unknown, but for the 

modeling, the damping response was assumed to be under-damped. Response outputs were 

simulated for a range of damping conditions. Matlab-Simulink was used to model the following 

displacement equation for a 10 Hz periodic force, F(t) with an average of 6 lbf, a mass, m, of 

100 lbm (system mass) and a spring constant, k, of 14.4 lbf/in. 

kx

dt

dx

c

dt

xd mF  +



 



 +













 = 2

2

The applied force, F(t), was constructed to mimic the time history of expected 

detonations for 10 Hz operation. The time response results for the calculations are shown in 

Fig. 3 for a range of damping coefficients, c, from 0.85 to 1.75. In each of these individual 

graphs the abscissa represents time in seconds, whereas the ordinate represents the calculated 

force response (i.e. kx(t)). The results show that steady state displacement (and therefore, thrust) 

is achieved within 2 s of operation irrespective of the actual damping coefficient. Comparison 

Fig. 2. Calibration showing applied force versus voltage measured for LVDT transducer. 
NASA/CRó2003-212191 4 

with detonation tube firing results (presented later) indicates that the damping coefficient of the 

system is about 1.5 lbf-s/in. 

2.2. Thrust Measurements for 1.3 in. Diameter Detonation Tube

An existing detonation tube with an internal diameter of 1.3 in. was used for the initial 

average thrust measurements. A picture of the detonation tube, the injector used for the setup 

and a schematic showing high-frequency pressure transducer locations are shown in Fig. 4. 

For these experiments, ethylene (C2H4) was used as the fuel and a mixture of oxygen and 

nitrogen (O2+0.5N2) was used as the oxidizer. The equivalence ratio was targeted to be 1.1. 

High speed Valvetek valves were used to control the injection timing. For the results discussed 

next, 30-40 cycles at 10 Hz operation constituted a single firing. The fill time for each cycle was 

set to be between 70-80 ms. 

Pressure measurements made near the head end of the detonation tube (Port 1, Fig. 4(c)) 

using a high-frequency pressure transducer are shown in Fig. 5. The top graph shows the peak 

pressures realized for individual detonations. The pressure traces for two consecutive detonation 

events during the firing are magnified in the bottom section of the figure. These results are 

typical of multi-cycle detonation tube operation. 

Fig. 3. Calculated thrust response results. For each trace, the abscissa represents time (s), 

whereas the ordinate represents force response (lbf). 
NASA/CRó2003-212191 5 

The corresponding thrust response from the displacement measurements is shown in 

Fig. 6. Results are shown here for two firings of 35 and 40 cycles, respectively. As mentioned 

earlier, the system does not respond to forces less than about 5 lbf due to static friction and 

preload from flex lines used for supplying propellants, and ignition system wiring. However, the 

results show that the average thrust for the multi-cycle firing is reached within 2 s (i.e. 20 cycles) 

of the start of the firing. Based on these average thrust measurements, the mixture based Isp for 

the firing is about 130 s. Cooper et al. [3] carried out single-shot PDE thrust measurements for 

the same propellant combination (i.e. C2H4, O2 and N2) using a pendulum mechanism. Their Isp

results for C2H4/(O2+0.33N2) and C2H4/(02+0.57N2) mixtures at an equivalence ratio of one are 

162 and 150 s, respectively. The current results of Isp=130 s for C2H4/(02+0.5N2) at an 

equivalence ratio of 1.1 agrees qualitatively with their results. 

The thrust measurements described here for the 1.3 in. diameter detonation tube 

demonstrates that steady state thrust can be accurately measured for multi-cycle operation of a 

PDE or PDE-ejector system. In the next sub-section, the PDE-ejector system that has been 

designed for this project is described. 

TOP VIEW

30°

1.7mm

2.8mm

2.0mm

CROSS SECTION

ÿ35.2mm

 (a) Image of detonation tube (b) Impinging jet injector 

1 2 3 4 5 6 7 8 9

Igniter

1.57m

Transducer Port

(c) Detonation tube schematic. 

Fig. 4. Detonation tube used for average thrust measurements. 
NASA/CRó2003-212191 6 

2.3. PDE-Ejector Description

The existing 1.3 in. diameter detonation tube cannot be used for the PDE-ejector 

experiments since the flanges and pressure ports would interfere with the ejector part of the 

setup. Based on these considerations, design of a completely new PDE-ejector system was 

Time [s]

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5

Relative Pressure [psi]

0

100

200

300

400

500

600

Time [s]

0.878 0.879 0.880 0.881 0.882 0.883 0.884 0.978 0.979 0.980 0.981 0.982 0.983 0.984

Relative Pressure [psi]

0

100

200

300

400

500

100 ms

Fig. 5. Pressure measurements near injector end of detonation tube. 

Fig. 6. Thrust measurement results for two firings of 35 and 40 cycles, respectively. 
NASA/CRó2003-212191 7 

undertaken. A larger detonation tube with an inner diameter of 2.25 in. was chosen for this 

geometry to increase the expected average thrust level (~18 lbf) at the same operating conditions 

as described earlier (viz. 10 Hz operation with C2H4/(02+0.5N2) propellants at 1.1 equivalence 

ratio). The higher thrust level will provide greater accuracy in the thrust measurement. 

A schematic of the PDE-ejector design is shown in plate (a) of Fig. 7. A picture of the 

fabricated detonation tube (ID and OD of 2.25 in. and 2.75 in., respectively) is also shown in the 

same figure. Additional detailed drawings of the setup are included in the appendix. The tube is 

six feet in length, with 3 ports located at 3 in. intervals in the center part for pressure transducer 

instrumentation. Pressure transducers at these positions will be used to verify that detonations 

occur in the main tube. Propellants for the detonation tube will be introduced through the 

impinging jet injector shown in Fig. 8. Flow control will use four high-speed Valvetek valves, 

two for oxygen, one for ethylene and one for nitrogen. 

A metal stand welded to the detonation tube (Fig. 7(b)) provides rigidity to the system. 

The detonation tube is bolted at the base of the stand to two frictionless sliders for thrust 

(a) 

(b) 

Fig. 7. (a) PDE-ejector schematic, (b) Image of 2.25 in. diameter detonation tube. 
NASA/CRó2003-212191 8 

measurements as described earlier. The ejector shroud is a tube six feet in length, with a 

diameter larger than that of the detonation tube. Three different diameter tube ejectors will be 

tested. These 0.125 in. wall thickness tubes have outer diameters of 4 in., 5 in., and 6 in., 

respectively, and have ports in 6 in. intervals for high frequency pressure transducers. 

The PDE-ejector system has been designed such that the detonation tube exit can be varied 

within the ejector shroud. A clamp system has been designed to accommodate the different 

diameter ejector shrouds. The design also features the capability of lateral positioning of the 

ejector shroud, a feature that allows accuracy in concentric positioning of the detonation tube 

with respect to the ejector shroud. The detonation tube and the ejector shroud are connected with 

long bars that provide rigidity to the complete system. The thrust measurement device is a 

rectangular bar with two free moving supports having a spring wrapped around each support. 

All required hardware sections are currently in different stages of fabrication with 

complete installation expected within 1 month. The new detonation tube fabrication is now 

complete, and initial experiments will concentrate on characterizing the average thrust for PDE 

engine only configuration. These experiments will be completed before the ejector shroud 

hardware fabrication is complete. 

 (a) (b) 

Fig. 8. Injector features 16 fuel-centered impinging jet elements. For each element, fuel is 

introduced through the central hole, whereas the oxidizer is introduced through six angled hole 

that impinge and mix rapidly with the central fuel jet. (a) Impinging jet injector face, (b) Central 

fuel tubes for injector. 
NASA/CRó2003-212191 9 

3. Measure of Technical Performance 

The technical milestones for the first year of the project are included here (taken directly 

from the proposal). 

Year 1 (9/1/01 ñ 8/31/02) Cost :$100,000

Milestone 1: Determine the thrust for a stand alone 1.3-inch diameter detonation tube (2/28/02). 

Milestone 2: Design and fabricate two ejector ducts for each PDE tube (2/28/02). 

Milestone 3: Install and test 1.3-inch PDE in one of the two companion ejector ducts. Ducts will 

be instrumented with high-speed pressure transducers (5/31/02). 

Milestone 4: Complete thrust augmentation measurements for the 1.3-inch PDE-ejector tube in 

one of the two ducts. (8/31/02). 

Note that for the proposal, the expected start date was 9/1/01, whereas the actual start 

date was delayed by 5 months (1/30/02). In short, Milestone 3 needs to be completed by 

10/31/02. As discussed earlier, technical milestone 1 has been completed and milestones 2-3 are 

near completion (within 1 month). 

Table 1. Timeline and Milestones. 

Year 1 Year 2 Year 3

Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4

Measure thrust for det. tube (1.3 in.)

Des./Fab 2 ejector ducts

Install det. tube in 1st ejector duct

Measure thrust for det. tube/1st ejector

Install det. tube in 2nd ejector duct

Measure thrust for det. tube/2ndejector

Schlieren studies for straight nozzle in

det. tube/duct assembly

Des./Fab optically-accessible section

Des./Fab det. tube conv. & div. nozzles

Schlieren studies for converging nozzle

in det. tube/duct assembly

Schlieren studies for converging nozzle

in det. tube/duct assembly

Final Report
NASA/CRó2003-212191 10 

4. Proposed Schedule and Relative Progress 

The timeline and milestones for the project is shown in Table 1 (taken from original 

proposal). As is evident from the table and earlier discussion, the current progress is on track 

with the proposed schedule. 

5. Collaborations 

 Collaborations with Prof. Merkle of University of Tennessee have been on-going since 

the start of the program. The PDE-ejector analysis performed by Prof. Merkle will help in 

guiding the experiments for the full course of this three-year project. 

6. Interactions with NASA Glenn Research Centerís PDE Research Program 

The major interactions with the NASA Glenn PDE research program have resulted 

through the periodic review meetings that have been held at NASA Glenn Research Center and 

follow up discussions. In particular, interactions have occurred with Jack Wilson, Don Paxson 

and Rene Fernandez. We have interacted with Jack Wilson in the area of fast actuating valves to 

share experiences and supplier information as we are using different systems. Dan Paxson has 

been very helpful in the area of the design of our ejector configuration in terms of comparison 

with his modeling effort. In this respect, his work has impacted us more than us impacting work 

at Glenn Research Center. Finally, discussions and interactions with Rene Fernandez led to his 

participation in the conference (and Proceedings publication) of ìAdvances in Confined 

Detonationsî that was held last July in Moscow, Russia. 

7. References 

[1] Bratkovich, T., and Bussing, T., ìA Pulse Detonation Engine Performance Model,î AIAAñ

95ñ3155, 1995. 

[2] Santoro, R., Broda, J., Conrad, C., Woodward, R., Pal, S., and Lee, S.-Y., îMultidisciplinary 

Study of Pulse Detonation Engine Propulsion,î JANNAF, 36th CS/PSHS/APS Joint Meetings 

(1999). 

[3] Cooper, M., Jackson, S., Austin, J., Wintenberger, E., Shepherd, J. E., ìDirect Experimental 

Impulse Measurements for Detonations and Deflagrations,î" AIAAñ01ñ3812 (2001). 
NASA/CRó2003-212191 11 

APPENDIX: PDE-Ejector Hardware 

In this appendix, PDE-ejector hardware drawings are included. The design features a PDE tube 

(2.25 in. diameter) with a multi-element (16) impinging jet injector. Modular design allows 

different size ejector tubes to be used for the experiments. Design also allows ejector tube 

positioning flexibility. 

 

Fig. A.1. Assembly drawing of modular pulse detonation ejector system. 

Fig. A.2. Detonation tube (2.25 in. diameter) for PDE-ejector experiments (#1 on assembly 

drawing). 
NASA/CRó2003-212191 12 

Fig. A.3. Constant diameter ejector tubes with pressure transducer ports (#3 on assembly 

drawing). 

Fig. A.4. Connecting plate (#4 on assembly drawing). 
NASA/CRó2003-212191 13 

Fig. A.5. Adjustable stand for ejector tubes (#2 on assembly drawing). 

Fig. A.6. Details of adjustable stand (#2 on assembly drawing). 
NASA/CRó2003-212191 14 

Fig. A.7. Sample clamp (#2 part of assembly drawing).

Fig. A.8. PDE injector with 16 impinging elements. Each element consists on one centered fuel 

hole and size angled oxidizer holes. 
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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21

Thrust Augmentation Measurements Using a Pulse Detonation Engine Ejector

Pulse detonation engine; PDE; Ejector

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Pennsylvania State University

Propulsion Engineering Research Center and

Department of Mechanical and Nuclear Engineering

University Park, Pennsylvania 16802

Robert J. Santoro and Sibtosh Pal

Project Manager, Richard DeLoof, Aeropropulsion Research Program Office, NASA Glenn Research Center,

organization code 0142, 216–433–6632.

The present NASA GRC-funded three-year research project is focused on studying PDE driven ejectors applicable to a

hybrid Pulse Detonation/Turbofan Engine. The objective of the study is to characterize the PDE-ejector thrust augmentation. A PDE-ejector system has been designed to provide critical experimental data for assessing the performance

enhancements possible with this technology. Completed tasks include demonstration of a thrust stand for measuring

average thrust for detonation tube multi-cycle operation, and design of a 72-in.-long, 2.25-in.-diameter (ID) detonation

tube and modular ejector assembly. This assembly will allow testing of both straight and contoured ejector geometries.

Initial ejectors that have been fabricated are 72-in.-long-constant-diameter tubes (4-, 5-, and 6-in.-diameter) instrumented with high-frequency pressure transducers. The assembly has been designed such that the detonation tube exit can

be positioned at various locations within the ejector tube. PDE-ejector system experiments with gaseous ethylene/

nitrogen/oxygen propellants will commence in the very near future. The program benefits from collaborations with Prof.

Merkle of University of Tennessee whose PDE-ejector analysis helps guide the experiments. The present research effort

will increase the TRL of PDE-ejectors from its current level of 2 to a level of 3.