AIAA 2009-5121
45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
2 - 5 August 2009, Denver, Colorado
Full-scale Firing Tests of Variable Flow Ducted Rocket
Engines employing GAP Solid Fuel Gas Generator
Hisahiro Nakayama1, Yoshiyuki Ikegami2, and Akihiko Yoshida3
Technical Research and Development Institute, Ministry of Defense
Kenji Koori4, Kiyoyuki Watanabe5, and Hidenori Tokunaga6
Aerospace Company, Kawasaki Heavy Industries, Ltd.
and
Haruo Shimizu7, and Shigeyuki Kanaizumi8
IHI Aerospace Co., Ltd.
Ram combustion performance of a full-scale, flightlike variable flow ducted rocket
engine was demonstrated through static firing tests. As the first step toward the
demonstration, self-ignition phenomena of fuel-air mixture in the transition phase and
ram combustion characteristics were studied by preliminary firing tests using a
heavyweight variable flow ducted rocket engine. The results of the firing tests and of a
CFD analysis of the flow field in the ram combustion chamber revealed that self-ignition
was considerably influenced by air-to-fuel ratio distribution of the recirculating flow
within the forward dome of the chamber. Subsequently, a flightlike variable flow ducted
rocket engine was designed based on the results of the preliminary firing tests and CFD
analysis. Excellent self-ignition in the transition phase, stable ram combustion and thrust
controllability during the ramjet phase were successfully demonstrated with the engine.
Nomenclature
Ae
F
FD
IM
m cap
m in
Pc
pe
Pt0
p0
va
ve
hd
hpr
=
=
=
=
=
=
=
=
=
=
=
=
=
=
2
nozzle exit area [m ]
net thrust [N]
thrust measured by load cell [N]
inlet margin [-]
flow rate of air captured by intake [kg/s]
flow rate of air induced into ram combustion chamber [kg/s]
total pressure at the end of ram combustion chamber [Pa]
static pressure at the end of nozzle [Pa]
total pressure of free stream [Pa]
static pressure of ambient [Pa]
speed of air flowing into intake [m/s]
speed of exhaust gas at the end of nozzle [m/s]
pressure loss in ram combustion chamber [-]
pressure recovery efficiency of intake [-]
1
Research Engineer, Air Systems Research Center, 1-2-10, Sakae, Tachikawa, Tokyo 190-8533, JAPAN.
Chief, Gifu Test Center, Naka, Kakamigahara, Gifu 504-0000, JAPAN.
3
Director, Advanced Defense Technology Center, 1-2-24, Ikejiri, Setagaya, Tokyo 154-0001, JAPAN.
4
Former Manager, Guidance System Design Section, 1, Kawasaki, Kakamigahara, Gifu 504-8710, JAPAN.
5
Assistant Manager, Guidance System Design Section, 1, Kawasaki, Kakamigahara, Gifu 504-8710, JAPAN.
6
Senior Staff Officer, Guidance System Design Section, 1, Kawasaki, Kakamigahara, Gifu 504-8710, JAPAN.
7
Manager, Defense Systems Quality Assurance Group, 900, Fujiki, Tomioka, Gunma 370-2398, JAPAN.
8
Engineer, Electronics Technologies Office, 900, Fujiki, Tomioka, Gunma 370-2398, JAPAN.
1
American Institute of Aeronautics and Astronautics
2
Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
D
I. Introduction
ucted rocket engine (DRE) is a kind of ramjet engine which uses a fuel-rich solid propellant. DRE has high
specific impulse since the engine does not carry oxidizer like conventional rocket engines, but uses air as an
oxidizer. As the engine has no compressor and turbine, its structure is very simple. Therefore, DRE is an
attractive candidate for tactical missile propulsion. Especially, variable flow ducted rocket engine (VFDRE) is
capable of providing a medium range air-to-air missile with high average flight speed and minimum time to target
by sustained propulsion, continuous optimization of intercept trajectory by thrust control, and superior intercept
maneuverability by high flight speed just before an intercept. Therefore, research and development of VFDREs
are carried out in several countries in order to examine the feasibility of VFDREs 1-6.
Typical operation sequence of a DRE in air-to-air missile application is shown in Fig. 1. A missile is
accelerated by the solid rocket booster up to more than Mach 2.0 after being launched from its carrier airplane.
Generally, integrated rocket booster (IRB) is a very useful booster for DREs. The booster propellant is cast
integrally within the ram combustion chamber of the DRE in order to increase volumetric efficiency of
propulsion section. Just after the IRB burned out, the port covers of the DRE are fragmented and the IRB nozzle
is ejected by detonator. Then the air captured by the intake is induced into the ram combustion chamber.
Subsequently, the solid fuel gas generator is ignited to generate hot fuel gas. The fuel gas mixes with the air and
self-ignition of the fuel-air mixture occurs in the ram combustion chamber. The period from the port cover
fragmentation to the beginning of ram combustion is defined as the transition phase. Prompt termination of the
transition phase is required for the DRE in order to minimize velocity loss during this phase. Therefore, as well as
stable ram combustion, the DRE requires prompt self-ignition of fuel-air mixture.
In general, DREs in air-to-air missile application have two intakes on the lower portion of their body as shown
in Fig. 2 in order to keep geometric adaptability between missiles and carrier airplanes. These intakes are
positioned adjacent to the body to provide missiles with good aerodynamic characteristics. Such a geometry of
DRE as this causes a distortion to the internal flow of DRE, therefore development of a recirculating flow within
the forward dome of the ram combustion chamber becomes insufficient. There is a possibility that the nonfullydeveloped recirculating flow causes insufficient fuel-air mixing and unstable ram combustion since the
recirculating flow within the forward dome plays an important role in fuel-air mixing and flame holding. Prompt
self-ignition of fuel-air mixture as well as stable ram combustion under such constraints is required for DRE
development.
The objective of this study is to demonstrate the technical feasibility of VFDREs for tactical missile
applications using full-scale VFDRE demonstrators by static firing test: prompt self-ignition of fuel-air mixture,
stable ram combustion and thrust controllability. This paper describes the results of the VFDRE demonstration.
II. Approach
The VFDRE program in Technical Research and Development Institute (TRDI) has been conducted, targeting
advanced propulsion for tactical missile application. TRDI experimentally designed and manufactured two types
of VFDREs, which were the heavyweight and the flightlike lightweight demonstrators, with Kawasaki Heavy
Industries and IHI Aerospace. These VFDREs were used for the performance demonstration in this study. This
study consists of three types of firing tests.
1)
Preliminary Firing Test: Static firing tests were conducted using the heavyweight VFDRE demonstrator. A
CFD analysis of the flow field in the ram combustion chamber of the heavyweight VFDRE demonstrator
was also performed. Self-ignition phenomena of fuel-air mixture and ram combustion characteristics were
examined based on the results of the firing tests and CFD analysis.
2)
Validation Firing Test: A fuel supply pattern at the onset of ramjet phase, which is to achieve prompt selfignition of fuel-air mixture, was proposed. A static firing test was conducted using the heavyweight
VFDRE demonstrator in order to validate the fuel supply pattern at the onset of ramjet phase.
3)
Integrated Firing Test: The flightlike lightweight VFDRE demonstrator was designed based on the results
of Preliminary Firing Test and Validation Firing Test. Static firing tests were conducted using the flightlike
VFDRE demonstrator in order to examine the thrust performance of the flightlike VFDRE demonstrator.
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III. Experimental Details
A. Solid Fuel Gas Generator
Glycidyl Azide Polymer (GAP) was employed as a solid fuel gas generator (SFGG) in this study. Hot fuel gas is
generated by self-sustaining combustion of the SFGG and the gas flows into the ram combustion chamber. Some
amount of a metallic fuel is added into the SFGG to increase flame holding capability and heat of combustion in the
ram combustion chamber. The length of the SFGG grain is determined to give burning time of approximately 20 s.
Self-decomposition characteristics of GAP give good self-sustaining combustibility without oxidizer to the
SFGG. Furthermore, the pressure sensitivity of the SFGG burning rate is relatively high within the pressure range
tested in this study. These characteristics of the SFGG, good self-sustaining combustibility and high pressure
sensitivity, give wide throttleability to the fuel supply into the ram combustion chamber.
B. Fuel Flow Controller
A thermally resistant rotary control valve was used in the fuel flow controller. The cross section area of flow
passage from the SFGG chamber to the ram combustion chamber is altered by gate opening of the valve. The SFGG
chamber pressure is controlled by changing the cross section area. Change in the SFGG chamber pressure alters the
SFGG burning rate, and then the fuel flow rate can be controlled.
C. VFDRE Demonstrator
The schematics of the VFDRE demonstrators are shown in Fig. 3. The heavyweight VFDRE demonstrator was
used in Preliminary Firing Test and Validation Firing Test, and the flightlike VFDRE demonstrator in Integrated
Firing Test. These demonstrators are full-scale, approximately 8 inches in diameter. All the demonstrators have two
intakes at a 90-degree angle to each other on the lower portion of the body, considering geometric adaptability to
carrier airplanes. As the main concerns in this study are only the characteristics of self-ignition of fuel-air mixture
and the performance of ram combustion, tests of the IRB phase are not performed. The ram combustion sequence
is started by the injection of fuel gas from the SFGG chamber into the ram combustion chamber where air is
being induced. Therefore, there is no port cover, ejectable IRB nozzle and booster propellant grain in these
demonstrators.
In Preliminary Firing Test, the incident angle of the air flow into the ram combustion chamber is set to 60
degrees in order to develop a recirculating flow and to enhance fuel-air mixing within the forward dome of the ram
combustion chamber. A long-type air-duct is employed in the heavyweight VFDRE demonstrator to obtain the air
flow into the ram combustor at an angle of 60 degrees certainly. In Validation Firing Test, a short-type air-duct is
employed in the heavyweight VFDRE demonstrator to make the shape of the VFDRE demonstrator more flightlike.
However, the incident angle of the air flow into the ram combustion chamber is kept to 60 degrees so that the
condition of the recirculating flow within the forward dome of the ram combustion chamber becomes
approximately equivalent to that in the VFDRE demonstrator with long-type air-duct. In Integrated Firing Test, a
rectangular-shaped and two-ramp external compression supersonic intake is adopted and two intakes are positioned
adjacent to the body of the flightlike VFDRE demonstrator. The basic structure of the flightlike VFDRE
demonstrator is the same as that of the heavyweight VFDRE demonstrator, however the flightlike VFDRE
demonstrator consists of thin-walled components and its weight is much lighter.
D. Firing Test Configuration and Condition
Since thrust performance of DREs varies in response to flight speed and altitude like ramjet engines, it is
necessary to provide these VFDRE demonstrators with an air flow which is equivalent to the air flow in the actual
flight condition in order to examine their thrust performance by static firing test. Therefore, all the static firing tests
were performed in a ramjet test facility under simulated flight conditions. These static firing tests were conducted in
a connected-pipe form or in a quasi-free jet. The schematics of these static firing test configurations are shown in
Fig. 4. In connected-pipe test, a simulated air flow is provided directly into the ram combustion chamber of a
VFDRE demonstrator via pipe at subsonic speed. The flow rate and temperature of the simulated air flow is set to be
the same as those of the air flow induced into the ram combustion chamber via intake duct in the actual flight. In
quasi-free jet test, a VFDRE demonstrator is operated while the intake is mounted in the hot supersonic flows of
freely expanding nozzles. The flow speed, density and temperature of the simulated air flow are set to be the same as
those of the air flow in the actual flight. Therefore, net thrust, combustion performance and operating condition of
intake can be evaluated in quasi-free jet test. All the test conditions are Mach 2.3 in flight speed and 2000 m in
altitude. After the simulated air flow becomes steady at a target value, the SFGG is ignited to start the firing test.
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American Institute of Aeronautics and Astronautics
In Preliminary Firing Test and Validation Firing Test, it is not necessary to consider operating condition of
intake since the aim of these tests is to understand basic characteristics of self-ignition of fuel-air mixture and of
ram combustion. Therefore, the heavyweight VFDRE demonstrator has no intake and connected-pipe test was
employed in these firing tests. On the other hand, it is necessary to consider the operating condition of the intake of
the flightlike VFDRE demonstrator in Integrated Firing Test in order to understand the performance of the whole
flightlike VFDRE demonstrator including its intake. Therefore, quasi-free jet test was employed in this firing test
and the flightlike VFDRE demonstrator was operated under simulated flight conditions. The angle of attack and
sideslip angle formed by the intake and the air flow are 0 degree respectively.
IV. Results and Discussion
A. Preliminary Firing Test
In this firing test, the SFGG chamber pressure was controlled in order to examine the operating condition of the
fuel flow controller. The history of the SFGG chamber pressure is shown in Fig. 5. The zero point on the horizontal
axis corresponds to the onset of SFGG ignition. The command of the control pattern is indicated by dashed line and
the measured pressure by solid line in this figure. The SFGG chamber pressure and gate opening of the rotary
control valve of the fuel flow controller are monitored and the fuel flow controller is operated so that the SFGG
chamber pressure is able to follow up the command. The measured SFGG chamber pressure is in good agreement
with the command. This result shows that the fuel flow controller in this study works well as expected.
The histories of the pressure and gas temperature in the ram combustion chamber are shown in Fig. 6. The
pressure and gas temperature were measured at the forward dome of the ram combustion chamber. The zero point on
the horizontal axis corresponds to the onset of SFGG ignition. The measured pressure is indicated by solid line, the
predicted pressure by dashed line, and the measured gas temperature by chain line in this figure. The predicted
pressure is obtained from the fuel flow rate on the assumption that the fuel gas burns in the ram combustion chamber
at a certain degree of combustion efficiency. The fuel flow rate can be estimated from the SFGG chamber pressure.
The measured and predicted pressures in the ram combustion chamber are non-dimensionalized by being divided by
the maximum predicted pressure in the ram combustion chamber. The test result shows that the measured pressure is
much lower than predicted and the gas temperature is relatively low until 11.6 s from the onset of SFGG ignition.
Since the fuel flow controller worked well as expected, it is concluded that self-ignition of fuel-air mixture could not
occur until 11.6 s. Although the self-ignition delay of 1 s or less is a target in this study in order to minimize the
velocity loss during the transition phase, the target is not attained in this firing test. However, since the pressure in
the ram combustion chamber agrees well with the predicted pressure and the gas temperature is sufficiently high
after 11.6 s passed, it is considered that fuel-air mixture burned stably in the ram combustion chamber as expected
until the SFGG burnout from the point of 11.6 s.
The gas temperature in the ram combustion chamber decreases by 300 K temporally from 15 s to 21 s during the
ram combustion. From the result of a CFD analysis of the flow field in the ram combustion chamber, as described
hereinafter, it is considered that the fuel-air mixture within the forward dome of the ram combustion chamber is fuel
rich from the beginning to the end in this firing test. The increment of the SFGG chamber pressure during this period
increases the fuel flow rate, and consequently the fuel-air mixture within the forward dome becomes more fuel rich.
It is considered that the gas temperature drop is caused by such an excessive fuel rich condition within the forward
dome.
B. CFD Analysis of Internal Flow of Ram Combustion Chamber
In order to investigate the cause of the long self-ignition delay, a CFD analysis of the flow field in the ram
combustion chamber was performed. The computational domain of the CFD analysis is shown in Fig. 7. The
computational domain consists of 450 thousands of structured grids. A wall function is introduced in order to
restrain the growth of the number of structured grids. The air and fuel gas are considered as perfect gas and k-ε
model is introduced as turbulence model. The flow field in the ram combustion chamber is solved by 3-dimensional
Navier-Stokes analysis assuming that the flow is a compressible steady flow.
As will be appreciated from Fig. 5, since the fuel flow rate is changing continuously, the flow field in the ram
combustion chamber is not completely in a steady state. However, since the timescale of the residence time of
combustion gas in the ram combustion chamber is millisecond and sufficiently shorter than the timescale of change
in the state of the ram combustion chamber, the flow field can be considered to be in a steady state in this analysis.
Air-to-fuel ratio (A/F) distribution is obtained by solving the fuel-air mixing state in each individual grid. The
fuel flow rate at 0.5 s was the largest but self-ignition did not occur in Preliminary Firing Test. By contrast, self4
American Institute of Aeronautics and Astronautics
ignition occurred at 11.6 s even though the fuel flow rate was the smallest. Therefore, the analysis was made for the
conditions at 0.5 s and at 11.6 s in order to investigate the relation between A/F distribution in the ram combustion
chamber and the long self-ignition delay. No chemical reaction between air and fuel gas was taken into account in
this analysis because self-ignition did not occur until 11.6 s from the onset of SFGG ignition. The condition of this
analysis is shown in table 1.
The air flow trajectories at 0.5 s and at 11.6 s are shown in Fig. 8. 400 trajectories of air are selected in a random
manner and plotted in each figures. Lager amount of air flows into the forward dome of the ram combustion
chamber and a recirculating air flow is developed better at 11.6 s than at 0.5 s.
The A/F distributions of the recirculating flow within the forward dome of the ram combustion chamber are
shown in Fig. 9. The A/F value within the forward dome at 0.5 s is considerably low due to high fuel supply from
the SFGG and the weak recirculating air flow. Consequently, it is considered that the fuel-air mixture within the
forward dome is too fuel-rich to ignite spontaneously at 0.5 s. On the other hand, the fuel-rich condition within
the forward dome is improved at 11.6 s due to decreased fuel supply from the SFGG and the well-developed
recirculating air flow. From the result of this CFD analysis, it is concluded that the A/F distribution condition
within the forward dome is one of the important factors which control self-ignition delay. From this result, it is
considered that an improvement of the fuel-rich condition within the forward dome is needed to achieve prompt selfignition of fuel-air mixture.
C. Validation Firing Test
Based on the results of Preliminary Firing Test and CFD analysis, the fuel supply pattern at the onset of ramjet
phase was improved to achieve prompt self-ignition. A firing test to validate the improved fuel supply pattern at the
onset of ramjet phase was performed using the heavyweight demonstrator with short-type air-duct. Since the aim of
this firing test is to validate the improved fuel supply pattern at the onset of ramjet phase, long burning time of the
SFGG is not required. Therefore, a shorter SFGG of which the length is approximately one-third of the normal
SFGG is used in this firing test. The improved control pattern of the SFGG chamber pressure is shown in Fig. 10.
The zero point on the horizontal axis corresponds to the onset of SFGG ignition. The command of the control
pattern is indicated by dashed line and the measured pressure by solid line in this figure. First, the SFGG chamber
pressure is raised to increase the gas temperature in the ram combustion chamber by supplying large amount of hot
fuel gas, and then lowered to make an appropriate A/F distribution condition for self-ignition of fuel-air mixture
within the forward dome by decreasing fuel gas supply from the SFGG chamber. This process is repeated three
times to ensure self-ignition of fuel-air mixture.
The histories of the pressure and gas temperature in the ram combustion chamber are shown in Fig. 11. The zero
point on the horizontal axis corresponds to the onset of SFGG ignition. The measured pressure is indicated by solid
line, the predicted pressure by dashed line, and the measured gas temperature by chain line in this figure. These
pressures are non-dimensionalized by being divided by the maximum predicted pressure in the ram combustion
chamber. By the first SFGG chamber pressure rise, the gas temperature in the ram combustion chamber is raised as
expected, however the measured pressure in the ram combustion chamber does not agree with the predicted pressure.
This result shows that the first SFGG chamber pressure spike did not cause self-ignition of fuel-air mixture in the
ram combustion chamber. At 0.9 s, the beginning of the second SFGG chamber pressure rise, the pressure and gas
temperature in the ram combustion chamber rise substantially. Therefore, it is considered that self-ignition of fuelair mixture occurred at this moment. The target self-ignition delay of 1 s or less in this study is attained in this firing
test. It is concluded that the improved fuel supply pattern at the onset of ramjet phase is proved to be effective.
Furthermore, since the measured pressure in the ram combustion chamber agrees well with the predicted pressure
after 0.9 s passed, it is also considered that fuel-air mixture burned stably in the ram combustion chamber as
expected until the SFGG burnout from the point of 0.9 s.
The gas temperature in the ram combustion chamber increases temporally by 500 K after 6 s passed. The SFGG
burns out approximately at 6 s and the fuel supply decreases around the SFGG burnout. During the process of
decrease in fuel supply, the A/F value of the recirculating flow within the forward dome of the ram combustion
chamber decreases and stoichiometric mixture ratio is probably attained there. It is considered that the gas
temperature rise is caused by the improvement of such a fuel rich condition within the forward dome.
D. Integrated Firing Test
The target self-ignition delay of 1 s or less in this study was attained and an fuel supply pattern at the onset of
ramjet phase for prompt self-ignition of fuel-air mixture was established in Validation Firing Test. Stable ram
combustion was also attained using the short-type air-duct which is necessary to a flightlike VFDRE. Based on these
results, a flightlike lightweight VFDRE demonstrator was designed. The thrust performance of the flightlike
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American Institute of Aeronautics and Astronautics
VFDRE demonstrator was examined in Integrated Firing Test. The control pattern of the SFGG chamber pressure in
this firing test is shown in Fig 12. The zero point on the horizontal axis corresponds to the onset of SFGG ignition.
The command of the control pattern is indicated by dashed line and the measured pressure by solid line in this figure.
The SFGG chamber pressure is controlled in a step pattern and a ramp pattern in order to examine the ram
combustion stability of the flightlike VFDRE demonstrator.
The history of the pressure in the ram combustion chamber is shown in Fig. 13. The zero point on the horizontal
axis corresponds to the onset of SFGG ignition. The measured pressure is indicated by solid line and the predicted
pressure by dashed line in this figure. These pressures are non-dimensionalized by being divided by the maximum
predicted pressure. The pressure in the ram combustion chamber increases immediately just after the beginning of
fuel supply from the SFGG chamber. Prompt self-ignition of fuel-air mixture is achieved as in Validation Firing
Test. In an actual flight, the fuel flow rate needs to be controlled in various patterns by changing the SFGG chamber
pressure in response to changes in flight speed and altitude. The ram combustion responds rapidly and smoothly to
changes in the SFGG chamber pressure that is in a step pattern and a ramp pattern in this firing test. From this result,
it is considered that the flightlike VFDRE demonstrator has sufficiently practicable throttleability of the fuel flow
rate for actual use. Although stable ram combustion is successfully demonstrated, the measured pressure is a little bit
higher than predicted. From this result, it is assumed that ram combustion efficiency is slightly higher than expected.
Inlet margin (IM)7 is a parameter that is used for the evaluation of intake performance and condition. IM is
defined by the following equation.
ìï
üï
Pc
IM = í1 ý ´ 100
ïî h pr (1 - h d )Pt 0 ïþ
(1)
IM is defined by pressure recovery efficiency of intake, pressure loss in ram combustion chamber, and total pressure
at the end of ram combustor. Pressure recovery efficiency is a performance parameter of intake, pressure loss in ram
combustion chamber is determined by the condition of internal flow and the inner geometry of ram combustion
chamber, and total pressure at the end of ram combustor represents the condition of ram combustion. Therefore, IM
is not only a performance parameter of intake but also an important performance parameter that represents the
operating condition of the whole DRE.
IM is physically a difference of total pressure of the air flowing into ram combustion chamber and total pressure
in ram combustion chamber. As the total pressure of the air is greater than the total pressure in ram combustion
chamber in stable ram combustion condition, IM is normally greater than 0. The value of IM changes during DRE
operation due to changes in air and/or fuel flow rate, operating condition of intake, and ram combustion condition.
Significant decreases in IM caused by excessive rise in ram combustion pressure, etc. makes the air flow into
ram combustion chamber unstable. As IM approaches to 0%, the stability of DRE operation is degraded due to
insufficient and unstable air flow into ram combustion chamber. Then, the minimum value of IM needs to be an
appropriate positive value under all the possible flight conditions in order to make DRE operation stable. However,
if the minimum value of IM is set at an excessive level in DRE design, appropriate thrust margin7 is not attained and
missile flight performance is degraded. Therefore, the flightlike VFDRE demonstrator is designed to make the
minimum IM approximately 10% under the conditions of 0-degree angle of attack and sideslip angle and also of the
maximum expected operating pressure of the ram combustion chamber in order to keep appropriate thrust margin.
The history of the IM is shown in Fig. 14. The zero point on the horizontal axis corresponds to the onset of
SFGG ignition. The measured IM is indicated by solid line and the predicted IM by dashed line in this figure. The
pressure recovery efficiency of the intake used for the IM calculation is obtained from wind-tunnel tests of the intake
conducted prior to Integrated Firing Test. The pressure loss in the ram combustion chamber is estimated using data
from the open literatures8,9. Although the measured IM is a little bit smaller than predicted, since the measured
minimum IM is nearly equal to 10% as designed and roughly agrees with the predicted IM in this firing test, it is
considered that the whole flightlike VFDRE demonstrator including its intake is in good operating condition.
As Integrated Firing Test is conducted in a quasi-free jet, no pipe is connected to the flightlike VFDRE
demonstrator. Therefore, net thrust is easily obtained from the load cell output. The net thrust is obtained from the
following equation in this study7.
F = FD + m in ve - m cap va + ( pe - p0 )Ae
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American Institute of Aeronautics and Astronautics
(2)
The history of the net thrust is shown in Fig. 15. The zero point on the horizontal axis corresponds to the onset of
SFGG ignition. The net thrust is non-dimensionalized by being divided by the maximum net thrust. The net thrust
responds rapidly and smoothly to changes in the SFGG chamber pressure that is in a step pattern and a ramp pattern.
Thrust throttle range of 25-100% is achieved in this firing test. Thrust controllability, which is the most significant
advantage of VFDREs, is demonstrated successfully with the flightlike VFDRE demonstrator.
V. Conclusion
In order to investigate ram combustion characteristics of VFDREs and demonstrate thrust performance of
VFDREs, full-scale VFDRE demonstrators, which were the heavyweight and the flightlike, were developed and
evaluated through static firing tests in a ramjet test facility under a simulated flight condition. The results were as
follows:
1)
Self-ignition of fuel-air mixture in the ram combustion chamber was considerably influenced by the air-tofuel ratio distribution of the recirculating flow within the forward dome of the chamber. Prompt selfignition of fuel-air mixture could be achieved by improving the fuel-rich condition within the forward
dome of the chamber.
2)
Excellent self-ignition in the transition phase, stable ram combustion and thrust controllability during the
ramjet phase were successfully demonstrated using the flightlike lightweight VFDRE demonstrator with a
rectangular-shaped and two-ramp external compression supersonic intake.
References
1
Besser, H., “History of Ducted Rocket Development at Bayern-Chemie,” AIAA-2008-5261, Proceedings of 44th
AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Hartford, CT, 2008.
2
Besser, H., Weinreich, H., and Kurth, G., “Fit for Mission-Design Tailoring Aspects of Throttleable Ducted Rocket
Propulsion Systems,” AIAA-2008-5262, Proceedings of 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit,
Hartford, CT, 2008.
3
Hewitt, P. W., “Status of Ramjet Programs in the United States,” AIAA-2008-5265, Proceedings of 44th
AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Hartford, CT, 2008.
4
Qin Fei, He Guoqiang, Li Jiang, and Liu Peijin, “Integrated Flow Field Calculation of Combustor and Inlet in Variable Flow
Ducted Rocket,” AIAA-2008-5172, Proceedings of 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit,
Hartford, CT, 2008.
5
Komai, I., and Frederick, R. A., “An Analytical Assessment for the Temperature Sensitivity of Variable-flow Gas
Generators,” AIAA-1994-3196, Proceedings of 30th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit,
Indianapolis, IN, 1994.
6
Bhat, V. K., and Singh, H., “Propellants for Variable Flow Ducted Ramjet (VFDR) Propulsion,” AIAA-1997-2977,
Proceedings of 33rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Seattle, WA, 1997.
7
Kubota, N., and Kuwahara, T., Ramjet, Published by Nikkan Kogyo Shimbun Ltd., 1996.
8
Experimental and Analytical Methods for the Determination of Connected-Pipe Ramjet and Ducted Rocket Internal
Performance, AGARD-AR-323, Propulsion and Energetic Panel, Working Group 22, 1994.
9
Mahoney, J. J., Inlets for Supersonic Missiles, AIAA Education Series, Published by AIAA, 1991.
Fuel
Gas
Air
At 0.5 s
At 11.6 s
Table 1. Condition of CFD analysis
Flow Rate [kg/s]
Total Pressure [MPa]
Total Temp. [K]
9.3
0.55
566
1.0
11.0
1770
0.3
2.0
1770
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American Institute of Aeronautics and Astronautics
Remarks
No self-ignition
Self-ignition
IRB Propellant
IRB Phase
Intake
Fuel Flow Controller
Port Cover
1. IRB Propellant
Combustion
Solid Fuel
Gas Generator
SFGG Chamber
Rocket to Ramjet Transition
Ramjet Phase
4. Air Induction Start
IRB Nozzle
Ram Combustion Chamber
3. Port Cover Fragmentation
5. SFGG Ignition and Combustion
2. IRB Nozzle
Ejection
6. Ram Combustion Start
Figure 1. Typical operation sequence of DRE in air-to-air missile application
Figure 2. Typical airframe geometry of tactical missile equipped with DRE
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American Institute of Aeronautics and Astronautics
Injector
SFGG Chamber
Nozzle
Ram Combustion Chamber
Solid Fuel
Gas Generator
60 degrees
Fuel Flow
Controller
Long-type Air-duct
(in Preliminary Firing Test)
60 degrees
Short-type Air-duct
(in Validation Firing Test)
(a) Heavyweight VFDRE Demonstrator
SFGG Chamber
Nozzle
Injector
Solid Fuel
Gas Generator
Ram Combustion Chamber
(in Integrated Firing Test)
External Compression Supersonic Intake
(Rectangular, 2 Ramps)
(b) Flightlike Lightweight VFDRE Demonstrator
Figure 3. Schematics of VFDRE demonstrators
Connected-pipe
Air Source (High
Temperature and
Pressure)
Plenum
Chamber
Heavyweight VFDRE Demonstrator
(a) Connected-pipe Test
Nozzle
Intake
Air Source (High
Temperature and
Pressure)
Plenum
Chamber
Flightlike Lightweight VFDRE Demonstrator
(b) Quasi-free Jet Test
Figure 4. Schematics of static firing test configurations
9
American Institute of Aeronautics and Astronautics
SFGG Chamber Pressure [MPa]
14
12
: Command
: Measured
10
8
6
4
2
SFGG
Burnout
0
-5
0
5
10
15
20
25
30
Time [s]
2
2.5
Nondimensional Ram Combustion
Chamber Pressure [-]
: Predicted (Pressure)
: Measured (Pressure)
: Measured (Temperature)
2
1.5
1.5
1
0.5
1
0
0.5
-5
0
5
10
-0.5
SFGG
Burnout
Self-ignition
15
20
Gas Temperature in
Ram Combustion Chamber [x 103 K]
Figure 5. History of SFGG chamber pressure (Preliminary Firing Test)
-1
30
25
Time [s]
Figure 6. Histories of pressure and gas temperature in ram combustion chamber (Preliminary Firing Test)
Injector
Air-duct
Nozzle
Ram Combustion Chamber
Figure 7. Computational domain of CFD analysis
10
American Institute of Aeronautics and Astronautics
Forward Dome
Forward Dome
(b) At 11.6 s (Self-ignition)
(a) At 0.5 s (No Self-ignition)
Figure 8. Air-flow trajectories
5.0
5.0
View
View
2.5
2.5
0.0
0.0
Forward Dome
Forward Dome
2.5
iew
iew
2.5
V
5.0
V
5.0
0.0
0.0
(b) At 11.6 s (Self-ignition)
(a) At 0.5 s (No Self-ignition)
Figure 9. A/F distributions of recirculating flow within forward dome of ram combustion chamber
SFGG Chamber Pressure [MPa]
14
To increase gas
temperature in
ram combustion
chamber
12
10
To improve
fuel-rich condition
within forward dome
: Command
: Measured
8
6
4
2
SFGG
Burnout
0
-2
0
2
4
6
8
10
12
Time [s]
Figure 10. Improved control pattern of SFGG chamber pressure (Validation Firing Test)
11
American Institute of Aeronautics and Astronautics
2.5
Nondimensional Ram Combustion
Chamber Pressure [-]
: Predicted (Pressure)
: Measured (Pressure)
: Measured (Temperature)
2
1.5
1.5
1
0.5
1
0
0.5
-2
0
2
-0.5
SFGG
Burnout
Self-ignition
4
6
Gas Temperature in
Ram Combustion Chamber [x 103 K]
2
8
10
-1
12
Time [s]
Figure 11. Histories of pressure and gas temperature in ram combustion chamber (Validation Firing Test)
SFGG Chamber Pressure [MPa]
14
: Command
: Measured
12
10
8
6
4
2
SFGG
Burnout
0
-5
0
5
10
15
Time [s]
Figure 12. Control pattern of SFGG chamber pressure (Integrated Firing Test)
12
American Institute of Aeronautics and Astronautics
20
Nondimensional Ram Combustion
Chamber Pressure [-]
1.2
: Predicted
: Measured
1
0.8
0.6
SFGG
Burnout
0.4
-5
0
5
10
15
20
Time [s]
Figure 13. History of pressure in ram combustion chamber (Integrated Firing Test)
60
Inlet Margin [%]
50
: Predicted
: Measured
40
30
20
Designed
Minimum IM
10
0
SFGG
Burnout
-5
0
5
10
15
20
Time [s]
Figure 14. History of inlet margin (Integrated Firing Test)
Nondimensional Net Thrust [-]
1.5
1
SFGG
Burnout
0.5
0
Thrust Throttle Range:
25-100%
-5
0
5
10
15
Time [s]
Figure 15. History of net thrust (Integrated Firing Test)
13
American Institute of Aeronautics and Astronautics
20