17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference
11 - 14 April 2011, San Francisco, California
AIAA 2011-2235
Recent experimental results obtained on Continuous
Detonation Wave Engine
Bruno LE NAOUR 1 , François FALEMPIN 2 and Flore MIQUEL 3
MBDA France
Due to its thermodynamic cycle, the Continuous Detonation Wave Engine has
theoretically a higher performance than classical iso-pressure combustion propulsion
concepts. CDWE can also be considered to reduce the environmental conditions generated
by other detonation engines such as PDE while reducing the importance of initiation issue
and simplifying some integration aspects. After performing some specific experimental and
numerical programs at LIH, MBDA designed a preliminary demonstrator of CDWE with
the support of the EADS Nursery. A first assessment of the tests in progress is presented in
the following article.
D
I. Introduction
UE to its thermodynamic cycle, the continuous detonation wave engine (CDWE) has theoretically a higher
performance than another classical propulsion concept using the combustion process (+ 20 to 25% in term of
thermal efficiency).
The use of rotating detonation wave can also be considered to reduce the environmental conditions generated by
pulsed detonation engine (PDE) such as vibrations, while reducing the importance of initiation issue and simplifying
some integration aspects.
Compared to a PDE, this design allows an easier operation in reduced pressure environment and an increase in
engine mass flow rate and thrust-to-weight ratio.
Such a concept has been studied since a long time, particularly at LIH (Lavrentiev Institute for Hydrodynamics)
in Novosibirsk. A specific experimental program has been performed by MBDA and Lavrentiev Institute ([1], [2])
to understand unsteady, three dimensional flow behind the detonation wave and to address some key points for the
feasibility of an operational rotating wave engine for space launcher:
• two-phase mixture detonation,
• operating limits (injection pressure for example),
• noise generated by a CDWE operating at several kHz,
• heat fluxes (intensity, areas) and cooling strategies,
• composite materials (Carbon / Silicon Carbide) compatibility,
• engine thrust vectoring capability.
• NOx production while using liquid hydrocarbon fuel/air mixture.
In order to better assess the feasibility of such a system, specific experiments have been performed to address
some key points like thrust vectoring, heat fluxes and material compatibility, operation in low pressure environment.
In parallel, a dedicated effort has been undertaken to develop an adapted numerical simulation tool.
1
R&D Engineer, MBDA-France, Rond-point Marcel Hanriot – Route d'Issoudun, 18020 BOURGES Cedex, France.
Head of Advanced Power Airframe – Chief Engineer LEA Hypersonic Programs, MBDA France, 1 Avenue
Réaumur, 92358 LE PLESSIS ROBINSON, France.
3
Engineer, MBDA-France, Rond-point Marcel Hanriot – Route d'Issoudun, 18020 BOURGES Cedex, France.
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Copyright © 2011 by MBDA France. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.
II. Experimental works on CDWE
Based on the previous studies and considering the growing interest shown around the world for the concept of
CDWE, a subscale demonstration engine has been designed with the support of the EADS Nursery. This small scale
demo (Figure 1) aims at replicating experimental works performed with Lavrentiev Institute and then extend the test
duration and test H2/CH4 mixture.
The annular chamber of the engine has inner and outer diameters of 80mm and 100mm. The length of the
chamber is 90mm. The thickness of the outer wall is 25mm.
The engine is mainly made of an alloy based on copper and zirconium. There is no cooling system. Thus, heatsink working becomes a limiting factor regarding the tests duration (a few seconds). Fuel and oxidizer are injected
separately in the axial direction, through 570 injection holes (three-jet injectors). The injection manifolds are
separated in four sectors in order to be able to change the local flow rate and investigate the thrust vectoring effect
[3]. Presently, tests are performed without nozzle at the end of the chamber.
Ignition is performed tangentially to the annular chamber. It is initiated by the use of an explosive wire device.
The amount of energy released by the wire explosion is about 10J.
Fuel and oxidizer are injected from two tanks able to sustain a maximum pressure of 80bar. Both receivers'
volumes are 20dm3. The equivalence ratio of the injected mixture is controlled by the supply pressure. As receivers
are emptying, equivalence ratio and specific flow rate are decreasing. This allows us to explore a large quantity of
operating points. Fuel can be either hydrogen or a mixture of hydrogen and methane. The maximum duration of test
is fixed to 5s. Products of combustion/detonation are exhausted into air at atmospheric pressure.
Figure 1 - Small scale CDWE preliminary demonstrator
A. Instrumentation of the Engine
The instrumentation of the test-bench is by now limited to a small number of sensors. The essential ones are the
pressure and temperature transducers installed on the feeding lines in order to control the mass flow rate injected.
As far as the demonstrator is concerned, average static pressure in the engine is measured by the use of three
transducers placed along the chamber. Evolution of the temperature of the outer wall is recorded during the tests.
Two fast pressure transducers (PCB 113B03) are also used in order to detect detonation waves. The maximum
acquisition frequency of those fast pressure transducers is 85 kHz. This frequency is too low to acquire with
precision the detonation fronts, but high enough to perform a frequency analysis of the signal and then to determine
if we are in a detonation or deflagration process.
Thrust is estimated thanks to three strain gauges installed between the engine and the structure of the test-bench.
B. First test series and first results
A first test campaign has been performed on the demonstrator with gaseous hydrogen (GH2) and gaseous
oxygen (GO2). The aim was to make sure of the good performance of the engine compared to the one already tested
at LIH, to check the repeatability of the combustion/detonation process under the same initial conditions of injection
as well as to ensure of the reliability measurement devices. This paper describes the results obtained during one of
the tests performed during this campaign.
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For this test, both GH2 and GO2 pressures were set to ensure an initial equivalence ratio of ~1.6 and a specific
mass flow rate of ~40kg/s/m² injected in the combustion chamber. One second after ignition, and as tanks were
emptying, those values were found to be 1.2 for equivalence ratio, and 33kg/s/m². Those tests condition are chosen
to be within detonation the stability domain experimentally established by the LIH (Figure 2).
Stability domain of a rotating detonation wave (Source : LIH)
2
1,8
1,6
Equivalence ratio
1,4
1,2
1
DETONATION
0,8
0,6
Limit for a 1bar
counter-pressure
0,4
0,2
COMBUSTION
0
10
15
20
25
30
35
40
45
50
55
60
65
70
Specific mass flow rate (kg/s.m²)
Figure 2 : Stability domain established by LIH. The dotted curve
is the stability frontier between detonation and combustion
process. The continuous curve describes the test conditions
applied for the test presented in this paper.
Engine is ignited at t=+0,6s after opening the fuel and oxidiser tanks in order to ensure a good feeding of the
lines. Due to the fact that fast pressure transducers were not cooled, test was stopped after one second of
detonation/combustion.
Fast pressure data are filtered with a band-pass Butterworth filter before being analysed. Figure 4 shows the data
obtained from transducer n°1 after filtering. Ignition shock can be observed at t=+0,6s.
Figure 3 : demonstrator after ignition
Figure 4 : Fast pressure transducers data after
Butterworth filtering.
Spectrum of those data is shown on Figure 5. One can notice a first peak at ~22kHz which stays constant during
the first 700ms before translating to a frequency of ~30kHz. A fast periodic phenomenon is then detected by the
sensors. If directly converted in term of speed, the velocity of this phenomenon would be 6200m/s and up to
8460m/s. This is clearly higher than the theoretical Chapman-Jouguet velocities and indicates that there is more than
one detonation wave running within the chamber.
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Figure 5 : Fourier transform of the fast pressure filtered data,
as a function of time.
An assumption is made on the number of waves we effectively get during this test : two waves are running
between t=+0,6s and t=1.3s. Then three waves are running until the end of the test. This leads to the detonation
waves velocities shown on Figure 6 which are found to be around 3% of the theoretical Chapman-Jouguet velocities
calculated for the same condition of injection.
3500
Velocity (m/s)
3000
2500
2000
1500
Two detonation
waves
1000
Three detonation
waves
500
0
0,7
0,8
0,9
1
1,1
1,2
1,3
1,4
1,5
Time (s)
Measured velocities
Chapman-Jouguet velocities
Figure 6 : comparison between measured detonation wave velocities and theoritical
Chapman-Jouguet velocities.
LIH shown that number of waves is a function of the detonation cells size. In order to check if the results
obtained are in good agreement with the empirical laws they established [2], we estimated the cells size of the
detonation waves for some mixture condition during the test (Matignon et al. [4]). By using equations 1 and 2
(where λ is the cell size, h is the height of detonation wave, and n the number of waves), we can determine the
interval of wave number that we should expect for the applied test conditions.
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h  (12  5) * 
 * Dmean
72
n*h
eq.1
eq.2
LIH laws give us that the number of waves should be between two and four waves during the whole test (Figure
7). The assumption we made is going from two to three waves, which is in good agreement with their results and
allows us to confirm that we effectively get the estimated number of waves running within the combustion chamber
during about 870ms.
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Number of waves
4
3
2
1
0,7
0,8
0,9
1
1,1
1,2
1,3
1,4
1,5
Time (s)
LIH lower limit
LIH upper limit
Estimated number of waves during test
Figure 7 : comparison between the number of waves
calculated with the LIH laws and the estimated ones.
As far as force transducers are concerned, they were used to obtain an estimation of the thrust and specific
impulse of the chamber. Filtered data of the total recorded on-ground thrust is shown on Figure 8. Data were also
expressed in term of thrust in vacuum, that is to say taking into account the atmospheric counter pressure, in order to
be compared with calculation performed with the 0D CDWRE code developed in ICARE laboratory in Orleans,
France [5].
Figure 8 : filtered signal of the total thrust measured
Maximum thrust measured during test was 54,7N on ground, that is to say 338,5N of equivalent thrust in
vacuum. The corresponding vacuum specific impulse is 341,7s. For the same condition of injection, CDWRE code
gives us a thrust of 330N and a specific impulse of 310s. The calculating results are in good agreement with
measurements and allow us to consider the CDWRE code as reliable tool for pre-dimensioning.
C. Long duration test
Fast pressure transducers could not be used more than one second due to the fact they were not cooled. After
removing them, it becomes possible to perform longer tests, up to five seconds, but the determining information that
those sensors allowed us to obtain on the detonation waves is lost. Nevertheless, it is still possible to compare the
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signals measured by the static pressure transducers within the detonation zone. Figure 9 shows the static pressure
signals obtained for a one second duration test and a five second duration test. Both are performed with the same
initial injection conditions.
1,1
Static pressure (bar)
1,05
1
0,95
0,9
0,85
0,8
0,75
0,7
0
0,5
1
1,5
2
2,5
3
3,5
4
4,5
5
5,5
Time (s)
1 second test
5 seconds test
Figure 9 : static pressure signals within the detonation zone
for a one second test and a five second test.
One can clearly notice the good repeatability of measurements under the same conditions of injection. Both
signals are identical until the end of the one second duration test. On the five seconds duration test, pressure is then
kept constant to 0.95bar until t=2.2s. According to those measurements, we can consider that detonation process
occurred during the longer duration test. Injection conditions allowed detonation waves to rotate within the engine
during about 1.6 seconds.
III. Large demonstration engine
In parallel of the tests on going, a larger scale engine is to be manufactured and tested in existing test facility.
The combustion chamber is 350mm (external inner diameter) and 280mm (internal inner diameter) and will be able
to operate with GH2/GO2 or GH2/LO2 or liquid hydrocarbon/air with the change of supply lines and injection wall.
This engine mock-up is modular and actively cooled.
As for the actual demonstration engine, the injection wall is divided in 8 sectors in order to be able to investigate
the thrust vectoring effect with a diverging nozzle or with a center core nozzle (aerospike). Moreover, the engine
will be equipped with a complete weighing system providing thrust vector components and corresponding moments.
Thanks to its modularity, the engine will be used, in first step as a non-flying workhorse allowing to address all
the key points such as:
• effect of injection configuration and conditions (2-phase mixing),
• stable operation domain and key parameters influencing it,
• effects of high speed tangential flow (skin friction and heat fluxes),
• thermal and mechanical strength of the combustion wall (fuel-cooled structure, high frequency mechanical
shocks),
• effect of dissymmetric injection on thrust vectoring when including a full nozzle,
• generated environment (vibration and acoustics),
• …
In a second step, the modularity will allow to progressively replace all the engine components by flight-worthy
ones in order to finally obtain a flight-worthy demonstrator which will be tested to really access the achievable
performance when taking into account all the technology issues.
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IV. Conclusion
Preliminary tests of the CDWE have been realized with a mixture of GH2-GO2. The objective of these tests was
to check the whole performances of the mock-up and to ensure of its reliability to perform repeatable firings before
using a mixture of GCH4+CH2 with GO2. The use of fast pressure transducers allowed us to estimate the number of
waves running within the chamber during nearly one second after ignition. The number of waves and their velocities
were found to be in good agreement with Chapman-Jouguet conditions on one side and with LIH empirical laws on
the other side. This validates the know-how acquired during the test campaigns performed at LIH in Novosibirsk.
Thrust has been measured and specific impulse calculated. Values are within the same range than the one
estimated by calculation at LIH for such a mixture and with the CDWRE code developed by ICARE.
In parallel, efforts are focused on the development of a larger modular and actively cooled demonstration engine
which should be available for tests by the end of year 2012.
Acknowledgments
Authors particularly wish to thank Dr Sergey Zhdan and Dr Fedor Bykovskyi from LIH for their key
contribution in all experimental works performed on CDWE, as well as Dr Dmitry Davidenko and Yohann Eude
from ICARE Laboratory for their outstanding contribution to numerical simulation activities.
References
1
Daniau, E., Falempin, F., Zhdan, S., Bykovskyi, F.A., Detonation Wave Propulsion Systems : First Step Toward
Operational Engines, ISABE 2005.
2
Daniau, E., Falempin, F., Bykovskyi, F.A., Zhdan, S., Pulsed and rotating Detonation Propulsion System : First Step
Toward Operational Engines, AIAA 2005-3233.
3
Falempin, F., Le Naour, B., R&T Efforts on Pulsed and Continuous Detonation Wave Engines, AIAA-2009-7284.
4
Matignon, C., Etude de la detonation de deux mélanges stoechiométriques (CH4/H2/O2/N2 et CH4/C2H6/O2/N2) –
Influence de la proportion relative des deux combustibles et de la temperature initiale élevée, Thèse de l'Université de Poitiers,
December 2000.
5
Davidenko, D.M., Eude, Y., Gökalp, I., Ideal Performance of a Continuous Detonation Wave Rocket Engine Fed with
Hydrogen-Oxygen Mixture, ICPCD2010, Saint-Petersburg, October 2010.
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