AIAA 2009-5031
45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
2 - 5 August 2009, Denver, Colorado
Yoshihiro YAMANO
1 , Yoshiyuki IKEGAMI
2
, Hisahiro NAKAYAMA
1
, Eishu KIMURA
3
, Junichi SATO
1
,
Yuichi OTABE 1 , Akihiko YOSHIDA 4
Technical Research and Development Institute, Ministry of Defense
We launched full scale experimental vehicles with variable-flow ducted rocket engine to demonstrate its performance in various flight conditions. Observed transition time from burnout of an integral rocket booster to onset of ramjet combustion were satisfactorily short. Stable ramjet combustion in various angle of attack and sideslip angle, and thrust control capability of the engine were successfully demonstrated. Flight control of the vehicles, that have axially asymmetric configuration, via path angle and acceleration was demonstrated, as well.
P c
P t0
P max pr d
= total pressure at ram combustor exit
= total pressure of airflow
= maximum pressure through ram combustion phase
= total pressure recovery rate
= pressure loss in ram combustor
D
UCTED rocket engine (DRE) is one of the promising ramjet propulsion systems for the next generation missiles because DRE enables supersonic flight with higher specific impulse than conventional solid rocket motors. One of the major technical difficulties of DRE is sensitivity of stable ramjet combustion to incoming air mass flow through inlets, which is affected by flight conditions such as altitude, velocity, incident angle to inlets, and so on. To overcome this problem, Variable Flow Ducted Rocket Engine (VFDRE) was proposed and has been investigated by various researchers.
Among others, TRDI also has been investigating the VFDRE from 1990’s. Though we have conducted numerous wind tunnel tests [1] and RAMJET tests [2], it is essential to obtain the aerodynamic stability of the vehicle and the combustion stability of the ram burner in actual flight to demonstrate the feasibility of VFDRE. This report provides results of ground launch test of the VFDRE experimental vehicle. Figure 1 shows an experimental vehicle accelerated by an auxiliary booster.
Configuration of the VFDRE experimental vehicle with auxiliary booster is shown in figure 2. The vehicle has two rectangular, fixed geometry supersonic inlets underneath the body, four control fins, a gas generator chamber, and a ram combustor. Initially, the ram combustor is equipped with jettisonable nozzle and is filled with solid propellant, to form an integral rocket booster (IRB).
In the fore-body, electrical devices, such as control unit, inertial device, telemetry, are equipped. A destruction device for safety and a battery unit are equipped there, as well. The coordinate system of the VFDRE is also shown in figure 2.
DRE, as a variant of ramjet engines, requires sufficient mass of air flow into the ram combustor to start ramjet operation. Therefore, when the DRE is launched from an air plane, the velocity of the airplane is used so as to decrease the amount of booster propellant. However, under the condition of ground launch, IRB does not have sufficient volume for
1
2
Research Engineer, Air Systems Research Center, 1-2-10 Sakae-cho, Tachikawa, Tokyo 190-8533, JAPAN.
3
Chief, Gifu Test Center, Naka, Kakamigahara, Gifu 504-0000, JAPAN.
Deputy Head, Department of Guided Weapon Systems Development, 5-1 Ichigayahommura-cho, Shinjuku, Tokyo
162-8830, JAPAN.
4 Director, Advanced Defense Technology Center, 1-2-24 Ikejiri, Setagaya, Tokyo 154-0001, JAPAN.
1
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Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

entire booster propellant because the experimental vehicle was designed mainly for air launched missile. To solve the problem, an auxiliary booster with fixed fins was used.
Glycidyl Azide Polymer (GAP), major ingredient of the gas generator propellant, is a self-decomposing material. The decomposition process of GAP is maintained by heat generated by N
3
decomposition in the molecular structure, so that no additional oxygen is required. The fuel gas generated is mixed with air in ram combustor and start to burn spontaneously.
One of key challenges in this project was to ensure ignition and sustain of ramjet combustion. To achieve this objective, the fuel flow rate had to be adjusted according to the inlet air flow. In our engine, the heat-resistant rotating fuel control valve was inserted between gas generator and ram combustor, and was critically controlled to adjust the fuel flow rate. This ground launch test was also scheduled to verify the control sequence of the fuel control valve.
As mentioned above, ramjet combustor performance is influenced by incoming air flow conditions. To evaluate stability of ramjet combustor, we used inlet margin (IM), which is the margin between pressure recovered by the inlets and total pressure in ram combustor, and it characterizes available capacity of inlet air flow. Figure 3 shows the definition of
IM. The experimental VFDRE vehicle was designed to work optimally when IM over 10% is maintained. In order to maintain the designed IM, angle of attack and sideslip angle were limited within -5~+5 degree in our experimental vehicle.
In some cases, however, we had to exceed over this limit on purpose in order to observe phenomena in the transient fuel flow change.
Data acquired from the ground launch test was used to evaluate ramjet transition characteristics, ramjet combustion stability against maneuver, thrust controllability, and the attitude control system of the VFDRE experimental vehicle in flight. Data from direct-connect flow tests (also known as “connected pipe tests”) and hardware in the loop (HWIL) simulations previously conducted was also used as reference in order to investigate experimental vehicle performance.
The launch test was conducted in a branch of Air System Research Center in Niijima Island which is a missile range of
TRDI. Test equipment and materials for sequence control, measurements and safety monitoring were set up. The schematic view of the launch test is shown in Figure 4. The data acquired on the vehicle such as acceleration, angular rate, velocity, and combustion data were telemetrically received and recorded.
Two flight patterns were planned and are shown in Figure 5. The purpose of the 1st flight pattern is to collect data in the straight flight path, as a nominal case of this launch test. Data in the turning flight path with angle of attack change were also collected. In the 2nd flight pattern, data when the airframe rolls and turns to change sideslip angle were collected.
The vehicle rolls through 90 degrees during ramjet phase in the 2nd flight pattern.
A.
Transition
Transition performance from boost phase to ram combustion phase is very important for our VFDRE. Figure 6 schematically shows pressure trace in the ram combustor from burnout of the IRB to onset of steady ramjet operation.
Transition time was defined as the time from IRB nozzle ejection to the onset of ram combustion, which is defined as the time when the pressure in the ram combustor exceeds the threshold after the first pressure increase shown in figure 6. Here, the first pressure increase is caused by IRB sliver combustion.
Figure 7 is pressure trace of the ram combustor measured during the 2nd flight. Transition time during the 1st flight and the 2nd flight was 0.17 second and 0.15 second, respectively. Transition time was expected to be less than 0.25 second by the system design requirement. Therefore, transition to ram combustion phase satisfied the requirement. Besides, transition times obtained by direct-connect flow tests with the same airflow conditions were 0.18 second and 0.22 second, respectively. While the test conditions were not exactly the same as the flight test, no significant difference was observed.
B.
Ram Combustion
Figure 8 and figure 9 show the ram combustion stability during the two flight patterns described previously.
Both of the flights included the turning maneuver to make variation of the angle of attack or side slip angle to observe the influence on ram combustion stability. The variation of angle of attack or side slip angle caused by maneuvers is shown in the top of the figures. The time histories of the angle of attack and sideslip angle are approximately agreed with the predicted calculation under some disturbance (only shows the maximum and minimum value). To confirm the continuation of ram combustion, estimated pressure without ram combustion is also shown in bottom of the figures (described as Estimated Pressure (Unburned)). In the middle of the figure, IM shows the inlets’ condition. If IM is more than 0%, the inlets act as super critical condition, that is to say the inlets are in normal operating range. If IM is less than 0%, the inlets have no margin for incoming air called sub critical condition. In this condition, ram combustion is said to be unstable and has the risk of extinguishment.
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In Figure 8, from 7.5 second to 19 second, the vehicle flew without maneuver, but as the fuel flow was controlled to reduce after 15.3 second in order to investigate throttling capability of the experimental vehicle. Though the pressure slightly decreased at the time, it remained stable through straight flight. As illustrated in figure 8, inlet margin was mostly maintained value greater than 10% as planned until 19 seconds.
The pressure went down after 19 second because the inlet air flow decreased as the angle of attack was decreased to -8 degree, which is well below the optimum operation limit. At that time IM rapidly decreased and became negative, which indicates inlets were operated at sub-critical condition. Ram combustion seemed to be sustained after 10G turn from 19 second, since the ram combustor pressure is higher than the estimated unburned value.
Figure 9 shows the 2nd flight data. In the 2nd flight, the vehicle rolls through 90 degrees during the ram combustion phase, and then subsequent dive took place so as to investigate effect of sideslip maneuver.
All the time through the ram combustion phase, sideslip angle was controlled within ±5 degree designed limit. IM was kept positive through the ram combustion phase and ram combustor pressure was well above the value, which is estimated pressure without ram combustion. Thus, stable operation with rolling and turning maneuvers was demonstrated.
C.
Thrust Controllability
Gas generator chamber pressure, rotation angle of fuel control valve, fuel mass flow rate and acceleration during the
2nd flight are shown in figure 10. The gas generator chamber pressure is increased as the valve rotation angle decreased
(i.e., “closed”), and then the fuel flow rate into the ram combustor is increased. As a result, the acceleration in body axis direction (X b
), i.e., thrust, is also increased.
D.
Airframe Control
Flight path angle and acceleration are shown in figure 11. During straight flight, the airframe attitude was controlled by flight path angle. When the vehicle turned, the airframe was controlled by acceleration. As shown in the figure, the flight path angle agreed well with the prediction. Thus the angular variation was controlled as planned during the ram combustion phase by controlling the flight path angle. The acceleration in Z b
axis also agreed with the prediction. These results show that the attitude control worked well to keep the asymmetrical body stable as well as to maintain the positions of inlets against the air flow within the optimal position during flight.
The conclusions of this paper are as follows:
(1) Short transition time from the burnout of the integral booster to the onset of the ram combustion, below the design objective 0.25 seconds in flight, was achieved.
(2) Stable ram combustion of the VFDRE experimental vehicle during maneuver such as varied angle of attack, sideslip angle and roll angle was demonstrated.
(3) Thrust controllability of the VFDRE experimental vehicle was also demonstrated.
(4) The attitude control system of the VFDRE experimental vehicle worked well to keep the asymmetrical body stable, as well as maintain the positions of inlets against the air flow within the optimal angle during flight.
The authors wish to thank Kawasaki Heavy Industries, Ltd. Aerospace Company for their help in development of this VFDRE experimental vehicle.
1 Nakayama, H., “Full-scale Firing Tests of Variable Flow Ducted Rocket Engines employing GAP Solid Fuel
Gas Generator,” AIAA Paper 2009-5121, 2009.
2 Tokunaga, H., “Development of Wide-Range Supersonic Intake for Variable Flow Ducted Rocket Engine,”
AIAA Paper 2009-5223, 2009.
3
American Institute of Aeronautics and Astronautics
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Figure 1: An experimental vehicle accelerated by an auxiliary booster
・
・
・
Figure 2: Configuration of VFDRE experimental vehicle
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P
IM 1 100
η 1 η P
P η 1
P c
= total pressure at ram combustor exit
P t0
= total pressure of airflow pr d
= total pressure recovery rate
= pressure loss in ram combustor
Figure 3: Definition of inlet margin
η P
Patrol Helicopter
Experimental
Vehicle
Tracking
Radar
Camera
Launch Pad
Control Place
Search Radar
Figure 4: Schematic view of the launch test
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Patrol Boats

Figure 5: Flight patterns of ground launch test
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Transition Time
Influence of Sliver
Pmax
⊿ Pmax
10% of ⊿ Pmax
△ Onset of Ram Combustion
△
Ignition of Gas Generator
Time
:Fuel gas starts to flow into the combustor
△
Port Cover Open
IRB Nozzle Status
1
△ IRB Nozzle Separation
Figure 6: Definition of transition time
Transition Time
0.15
Ram Combustor Pressure
∆
P
P max max
10% of “ ∆ P max
”
0
7.2
IRB Nozzle Status (Separation)
7.4
7.6
Time [s]
7.8
Figure 7: Transition time during the 2nd flight
8.0
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Nose Up
10
5
0
-5
-10
-15
Nose Down
30
25
20
15
10
-5
-10
5
0
1
0
0 5
Ram Combustion Phase
Angle of Attack
Predicted
(with Disturbance)
10G Turn
Maximum
Minimum
Inlet Margin
Test Data
Estimated Pressure
(Unburned)
10 15 20 25 30
Time [s]
Figure 8: Ram combustor pressure stability during the 1st flight
35
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Nose Left
10
5
-15
Nose Right
50
40
0
-5
-10
30
20
10
0
1
Rolls through 90°
Side Slip Angle
Predicted
(with Disturbance)
5G Turn
Maximum
Minimum
Inlet Margin
Test Data
Estimated Pressure
(Unburned)
0
0 5
Ram Combustion Phase
10 15
Time[s]
20 25
Figure 9: Ram combustor pressure stability during the 2nd flight
30 35
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20
0
-20
-40
2.0
1.5
1.0
0.5
0
15
10
5
0
60
40
20
0
10
Gas Generator Chamber Pressure
Command
15
Valve Rotation Angle[deg.]
Command[deg.]
Acceleration in Body Axis
Time [s]
Figure 10: Thrust control of VFDRE
Mass Flow Rate
20
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Ascent
40
20
0
-20
-40
-60
-80
Descent
200
150
100
50
0
-50
0
Controlled by Flight Path Angle
Flight Path Angle
Predicted
(with Disturbance)
Maximum
Minimum
Nominal
Acceleration in Zb Axis
Nominal
Command
5 10 15 20
Time [s]
Controlled by Acceleration Rate in Body Axis
25 30 35
Figure 11: Flight path and acceleration in Z b direction
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