Papers - Matsuo, Akiko

Three-dimensional Shock Wave Interactions in Sabot Separation Launched in Hypersonic Regime

Kasahara, H., and Matsuo, A.

Transactions of JSASS (Transactions of the Japan Society for Aeronautical and Space Sciences)  65 ( 4 ) 147 - 159 2022

Research paper (scientific journal), Joint Work, Accepted,  ISSN  05493811

　View Summary

A sabot is used in a supersonic accelerator to reduce the in-bore balloting in the tube. Sabot petals are separated from the projectile after muzzle discharges; however, this sabot separation induces flight deviation caused by the shock wave interaction with the projectile. In this study, we investigate the three-dimensional (3D) shock wave interaction between the projectile and the sabot based on a previous experimental configuration by using 3D computational fluid dynamics coupled with rigid-body dynamics. We observed three types of shock wave interaction on the projectile: namely, shock wave oscillation, recompression shock wave attenuation, and shock wave sweep. The shock waves from the projectile and the sabot oscillate around the projectile tip because the gap between the projectile and the sabot is too small for the ambient air to pass through. This oscillation induces a significant aerodynamic perturbation on the projectile and the sabot. Multiple shock wave reflections appear between the projectile and the sabot petals after the oscillation ends. These multiple shock wave reflections generate a complex flow field between the projectile and the sabot, and an unsteady aerodynamic perturbation on the sabot.

• Access to Document (DOI)

Flight Demonstration of Detonation Engine System Using Sounding Rocket S-520-31: System Design

Kawasaki A., Matsuyama K., Matsuoka K., Watanabe H., Itouyama N., Goto K., Ishihara K., Buyakofu V., Noda T., Kasahara J., Matsuo A., Funaki I., Nakata D., Uchiumi M., Habu H., Takeuchi S., Arakawa S., Masuda J., Maehara K., Nakao T., Yamada K.

AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022 (AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022)     1 - 12 2022

　View Summary

A research group at Nagoya University, in collaboration with Keio University, the Muroran Institute of Technology, and Institute of Space and Astronautical Science at the Japan Aerospace Exploration Agency (ISAS/JAXA), has successfully demonstrated world’s first operation of Detonation Engine System (DES) in space using the sounding rocket S-520-31, which lifted off from the JAXA Uchinoura Space Center, Japan, at 5:30 a.m. (Japan Standard Time) on July 27, 2021. The DES was equipped with two different types of detonation rocket engines, namely a rotating detonation engine (RDE) and a pulse detonation engine (PDE), which generated 500-N-class axial thrust and 1-Nm-class rolling torque, respectively. The RDE and PDE were tested in space during the ballistic flight of the rocket. In this manuscript, the system-level design of the DES was disclosed. Key point are as follows: the DES consists of an RDE and a PDE, and propellant storage/feed, avionics, instrumentation, and structure subsystems; the overall dimensions are 480 mm in the diameter × 1810 mm in the length. The dry mass was 175 kg, while gross (wet) mass was 181 kg including propellant and pneumatic gases. Commercial off-the-shelf (COTS) components are widely utilized after carefulevaluation and testing, and with redundant architecture.

• Access to Document (DOI)

Flight Demonstration of Detonation Engine System Using Sounding Rocket S-520-31: Performance of Rotating Detonation Engine

Goto K., Matsuoka K., Matsuyama K., Kawasaki A., Watanabe H., Itouyama N., Ishihara K., Buyakofu V., Noda T., Kasahara J., Matsuo A., Funaki I., Nakata D., Uchiumi M., Habu H., Takeuchi S., Arakawa S., Masuda J., Maehara K., Nakao T., Yamada K.

AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022 (AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022)   2022

　View Summary

To create a new flyable propulsion system, Detonation Engine System (DES) has developed which can be stowed in the sounding rocket S-520-31. This paper focused on the first flight demonstration of RDE in space environment, which is integrated into DES, using S-520-31. The flight result was compared with the ground test data to validate its performance. In the flight experiment, the stable combustion of the RDE with aerospike nozzle was observed by onboard digital and analog cameras. With time-averaged mass flow of 182 ± 11 g/s and equivalence ratio of 1.2 ± 0.2, the RDE generated the time-averaged-thrust of 518 N, and the specific impulse achieved 290 ± 18 s, which is almost identical to the ideal value of constant pressure combustion. Due to the RDE combustion, the angular velocity increased by 5.7 deg/s for total, and the time-averaged torque during 6 s of operation was 0.26 N∙m. The high frequency sampling data identified the detonation frequency during record time as 20 kHz in the flight, which was confirmed in the DES ground test through the high frequency sampling data analysis and high-speed video imaging.

• Access to Document (DOI)

Flight Demonstration of Detonation Engine System Using Sounding Rocket S-520-31: Performance of Pulse Detonation Engine

Buyakofu V., Matsuoka K., Matsuyama K., Goto K., Kawasaki A., Watanabe H., Itouyama N., Ishihara K., Noda T., Kasahara J., Matsuo A., Funaki I., Nakata D., Uchiumi M., Habu H., Takeuchi S., Arakawa S., Masuda J., Maehara K., Nakao T., Yamada K.

AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022 (AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022)     1 - 14 2022

　View Summary

A research team of Nagoya University, Keio University, Muroran Institute of Technology, and Institute of Space and Astronautical Science (ISAS) at Japan Aerospace Exploration Agency (JAXA) successfully conducted a space demonstration of the detonation engine system (DES) using the S-520-31 sounding rocket of JAXA/ISAS. This paper presents the results of an S-shaped pulse detonation engine (PDE) flight test. The aim of the flight is to confirm that the PDE and its system operate in space, to confirm the reproducibility of the PDE cycle, and to despin the sounding around its axis. The PDE and its system successfully operated for 14 cycles in space. The experimental plateau pressure was 80 ± 3 % of the calculated plateau pressure, suggesting that detonation occurred in 14 cycles. Pressure profile of the 14 cycles was highly reproduced as well. The average angular acceleration change of PDE operation, mixture supply, and oxygen supply were determined using gaussian distribution. The results suggested that the PDE despinned the sounding rocket by the thrust from detonation combustion, which was in good agreement with a quasi-steady-state model with an accuracy of 101 ± 15%.

• Access to Document (DOI)

Flight Demonstration of Detonation Engine System Using Sounding Rocket S-520-31: History from Development to Flight

Itouyama N., Matsuyama K., Matsuoka K., Kawasaki A., Watanabe H., Goto K., Ishihara K., Buyakofu V., Noda T., Kasahara J., Matsuo A., Funaki I., Nakata D., Uchiumi M., Habu H., Takeuchi S., Arakawa S., Masuda J., Maehara K., Nakao T., Yamada K.

AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022 (AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022)     1 - 9 2022

　View Summary

Detonation engines exert various characteristics and high propulsion performance depending on their operation modes. Our group has conducted the demonstration of the integrated RDE/PDE system (Detonation Engine System, DES) in outer-space environment using a sounding rocket named S-520-31 operated by Japan Aerospace Exploration Agency. This paper introduces the development history of DES from main integrity evaluations of preflight model to the outer-space demonstration of flight model.

• Access to Document (DOI)

Flight Demonstration of Detonation Engine System Using Sounding Rocket S-520-31: Flight Path and Attitude

Watanabe H., Matsuyama K., Matsuoka K., Kawasaki A., Itouyama N., Goto K., Ishihara K., Buyakofu V., Noda T., Ito S., Kasahara J., Matsuo A., Funaki I., Nakata D., Uchiumi M., Habu H., Takeuchi S., Arakawa S., Masuda J., Maehara K., Nakao T., Yamada K.

AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022 (AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022)   2022

• Access to Document (DOI)

Experimental Clarification on Detonation Phenomena of Liquid Ethanol Rotating Detonation Combustor

Yoneyama K., Ishihara K., Ito S., Watanabe H., Itouyama N., Kawasaki A., Matsuoka K., Kasahara J., Matsuo A., Funaki I.

AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022 (AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022)   2022

　View Summary

Rotating detonation combustion was successfully tested for liquid ethanol and gaseous oxygen heterogeneous mixture in a cylindrical combustor for the first time. The rotating detonation combustor (RDC) we operated has a diameter of 20 mm and variable lengths of 60, 210 mm. Pure ethanol and industrial ethanol were selected as liquid fuels. Under the conditions of ethanol and oxygen mass flow rate of 26-40 g/s, equivalence ratio of 0.4-1.7, and backpressure of 10-17 kPa, we confirmed detonation combustion and deflagration combustion. In detonation combustion, as the ethanol manifold supply temperature increases, the detonation propagation speeds, and luminance of combustion were enhanced. It is possible to say enhanced evaporation behavior resulted in stable detonation and showed high peaks in luminance value. Utilizing the control surface method to evaluate experimental thrust, estimated thrust showed good agreement with experimental thrust. For liquid-fueled cylindrical RDCs, the control surface methods can also be used to evaluate thrust.

• Access to Document (DOI)

Numerical Investigation of Rotating Detonation Engine with Injection from the Combustor Side Wall

Sada T., Matsuo A., Shima E., Watanabe H., Kawasaki A., Matsuoka K., Kasahara J.

AIAA AVIATION 2022 Forum (AIAA AVIATION 2022 Forum)   2022

　View Summary

Numerical analysis of cylindrical rotating detonation engine (RDE) with injectors on the combustor side wall is conducted using the stoichiometric H2-Air premixed gas. The combustor is 70 mm in length and 20 mm in diameter. The results are compared to the conventional injector configuration, which has injectors on the combustor bottom, and the combustion features with side injection are investigated. The instantaneous pressure and temperature fields show the characteristics of propagating waves are different depending on the injector configuration. From the spatiotemporal average profiles, the combustion occurs more rapidly in the case of injection from the bottom. However, the case of injection from the side wall also shows the consumption of fuel. Compared to the theoretical values, the specific impulse is 106 % from the bottom and 103 % from the side. In this combustor length, both injection conditions can provide similar thrust performance to the theoretical values.

• Access to Document (DOI)

Experimental Research on Thrust Performance of Rotating Detonation Engine with Liquid Ethanol and Gaseous Oxygen

Sato T., Ishihara K., Yoneyama K., Ito S., Itouyama N., Watanabe H., Kawasaki A., Matsuoka K., Kasahara J., Matsuo A., Funaki I.

AIAA AVIATION 2022 Forum (AIAA AVIATION 2022 Forum)   2022

　View Summary

Rotating detonation combustor (RDC) is one of the combustors using detonation waves, which are hypersonic combustion waves, and is expected to simplify the system and improve thermal efficiency due to their hypersonic combustion and compression performance by shock waves. Most of these studies use gas propellants, and liquid propellants are rarely used. Since liquid propellants are used in many combustors such as aircraft engines, it is important to evaluate the performance of RDC with liquid propellants. In this study, a cylindrical RDC, which is an RDC without inner cylinder, with a liquid ethanol and gaseous oxygen was tested at a mass flow rate of 31.3 ± 3.2 g/s, an equivalent ratio of 0.46-1.42 ± 0.12, a back pressure of 13.2 ± 0.9 kPa, and fuel injector with 24×φ0.2 or 6×φ0.4 to evaluate the performance and visualize the inside of the combustion chamber. As a result, when fuel injector was 24×φ0.2, detonation waves were observed, and high propagation velocity and high thrust performance were achieved. From the internal self-luminous and CH* radicals visualization from side wall, a circumferential DDT (deflagration to detonation transition) was observed. In addition, it was found that the detonation wave lifted about 2-3 mm from the combustor bottom, the main combustion region was occurred at 20 mm from the bottom, and the combustion region, including the main combustion region, requires about 50-70 mm from the bottom, which is correlated with the internal pressure, brightness distributions, and the image of acrylic damage. When fuel injector was 6×φ0.4, transition of detonation wave was not observed. The deflagration wave lifted about 10 mm, main combustion occurred up to 45 mm, and partial combustion continued to more downstream compared to detonation combustion. Acrylic damage began to occur downstream from the area with the highest pressure and brightness.

• Access to Document (DOI)

Cylindrical Rotating Detonation Engine with Propellant Injection Cooling

Goto, K., Ota, K., Kawasaki, A., Itouyama, N., Watanabe, H., Matsuoka, K., Kasahara, J., Matsuo, A., Funaki, I., and Kawashima, H.

Journal of Propulsion and Power (Journal of Propulsion and Power)  38 ( 3 ) 410 - 420 2021.11

Research paper (scientific journal), Joint Work, Accepted,  ISSN  07484658

　View Summary

A cylindrical rotating detonation engine (24 mm diameter) with multiple injector holes on the combustor side wall for both propellant injection and cooling by injector gas flow was designed and tested. As each set of injectors created its own detonations, the synchronized detonation waves propagated along the three injector rows. From the axial view, those detonation combustion regions stood off from the injector surface as the non-well-mixed propellant existed between the wall and the combustion zone. This combustion region tended to broaden in lateral and radial directions as the mass flow rate increased by pushing its region with the injected propellant. Wall heat flux and heat balance were evaluated by a one-dimensional unsteady heat transfer model with propellant injection cooling. Even when the flow rate was doubled, the increase in the wall heat flux was only 18–25%. This heat trend and the image of standoff self-chemiluminescence from the injector surface implied that a non-well-mixed unburned propellant acted as a heat-reduction layer to ease heat load into the combustor. Measurements and thermal analysis verified the flow structure near the injector and heat-exchange mechanism due to the propellant gas flow, which has a potential for thermal steady operation.

• Access to Document (DOI)

Torque Around the Axial Direction on Rotating Detonation Engines

Sawada, S., Goto, K., Ishihara, K., Kawasaki, A., Matsuoka, K., Kasahara, J., Matsuo, A, and Funaki, I.

Journal of Propulsion and Power (Journal of Propulsion and Power)  38 ( 1 ) 59 - 70 2021.09

Research paper (scientific journal), Joint Work, Accepted,  ISSN  07484658

　View Summary

A rotating detonation engine (RDE) generates a continuous thrust with one or more rotating detonation waves. Because of the velocity on the order of kilometers/second, the reaction zone is relatively small. Therefore, the RDE realizes a short combustion chamber length. However, the detonation waves induce an azimuthal motion of propellant, resulting in torque around the thrust axis. Because the motion does not contribute to the thrust, the torque is important in terms of performance loss. Herein, we conducted combustion tests with a six-axis force sensor to simultaneously measure 0.149 0.009 Nm torque and 48.1 0.9 N thrust. A comparison of detonation waves captured by high-speed camera revealed that the torque followed the direction and was offset when the waves existed in both of two directions simultaneously, which indicates the possibility of controlling the torque. Under a mass flow rate at 87 9 g∕s and an equivalence ratio at 1.43 0.28, when the azimuthal component of shear force was 8.8 0.6% of the thrust, 0.77 0.10% of the total kinetic energy of the exit flow was distributed to the azimuthal component of velocity and did not contribute to the thrust. We therefore concluded that the effect of the azimuthal motion on the RDE’s performance was small.

• Access to Document (DOI)

Coupling Simulation of Thermal Response to Exhaust Gas at External Nozzle

Fujii, K., Matsuo, A., Oki, J., Taguchi, H., Chiga, T., and Ikeda, Y.

Journal of Spacecraft and Rockets (AIAA)  59 ( 1 ) 260 - 270 2021.09

Research paper (scientific journal), Joint Work, Accepted,  ISSN  00224650

　View Summary

This paper examines thermal response behind the high-Mach integrated control (HIMICO) experiment’s engine. In this experimental aircraft, instruments must be protected from heat load due to exhaust gas; therefore, a coupling calculation between the fluid and the wall is conducted to confirm the performance of HIMICO’s thermal protection system (TPS). First, the validity of the coupling calculation is confirmed through comparison with aerodynamic heating on a hollow cylinder. The present calculation result can reproduce the surface temperature distribution on the cylinder better than previous work has managed because we consider the turbulence effect. Second, one-dimensional heat-transfer analysis is conducted on the external nozzle, and the appropriateness of the calculation result is confirmed through comparison with the ramjet-engine experiment. Finally, a coupling calculation between the fluid and the wall is conducted to investigate the local temperature distribution. The calculation result indicates that the temperature increase easily meets design requirements and that TPS performance is sufficient.

• Access to Document (DOI)

Experimental Study on Truncated Conical Rotating Detonation Engine with Diverging Flows

Nakata, K, Ota, K., Ito, S., Ishihara, K., Goto, K., Itouyama, N., Watanabe, H., Kawasaki, A., Matsuoka, K., Kasahara, J., Matsuo, A., and Funaki, I.

AIAA Propulsion and Energy 2021 Forum (AIAA Propulsion and Energy Forum, 2021)   2021.07

Joint Work

　View Summary

This study focused on acceleration of propellant to supersonic speed with a compact engine. Increase in the exhaust velocity of a rocket engine leads to improvement of the thrust performance. Several combustion tests were conducted for a small rotating detonation engine (RDE) whose channel is truncated conical shaped (diverging angle (formula presented)), under low back pressure conditions. In these tests, gaseous (formula presented) were used as the propellant, and the mass flow rate were ranged from 56 to 123 g/s. The pressure ratio between the maximum value in the engine and at the exit was approximately 0.16, which was confirmed to be significantly below the critical value for a sonic flow. Namely, the exhaust flow was supersonic, even though there was no convent section in the engine. In a range of propellant mass flow rate, specific impulses were approximately 110% compared to those in a cylindrical RDE with a uniform cross-section combustor.

• Access to Document (DOI)

Study of Cylindrical Rotating Detonation Engine with Propellant Injection Cooling System

Ota, K., Goto, K., Itouyama, N., Watanabe, H., Kawasaki, A., Matsuoka, K., Kasahara, J., Matsuo, A., and Funaki, I.

AIAA Propulsion and Energy 2021 Forum (AIAA Propulsion and Energy Forum, 2021)   2021.07

Joint Work

　View Summary

A cylindrical RDE with a propellant injection cooling system with an injector on the entire wall surface of the combustor has been tested and demonstrated. In this study, in order to search the optimum flow balance between the side and the bottom, the experiment was conducted under a total test condition of 30 g/s, and various balance condition. Under all conditions, a stable single rotating detonation was observed. In the 5.0-second combustion test, the temperature rise in the cooling area stopped and became steady state at a maximum of 850 K. This may be due to the heat exchange and film cooling effect of propellant injection. In addition, the propulsion performance was 94% of the theoretical value calculated assuming that the gas flow reached the speed of sound at the outlet of the combustor. It had been demonstrated that sufficient cooling and propulsion performance can be obtained by injecting 76% from the side wall and 24% from the bottom wall.

• Access to Document (DOI)

Three-Dimensional Numerical Investigation of Hypersonic Projectile Launched by Railgun on Transitional Ballistics

Kasahara, H., and Matsuo, A.

Journal of Spacecraft and Rockets (Journal of Spacecraft and Rockets)  58 ( 4 ) 1 - 17 2021.06

Research paper (scientific journal), Joint Work, Accepted,  ISSN  00224650

　View Summary

Supersonic accelerators use sabots to suppress bore balloting. However, sabot separation induces flight deviation in projectiles owing to their interactions with fluids and shock waves. A seamless calculation of the acceleration phase in the tube and the full separation process using 3-D computational fluid dynamics coupled with rigid body dynamics is presented in this study to investigate the shock-wave interactions around a railgun-launched projectile. The railgun acceleration induced a normal shock wave propagating ahead of the projectile; when the projectile reached the tube end, a substantial expansion at the muzzle generated two shock waves, namely, a spherical precursor shock wave and a Mach disk, outside the tube. Aerodynamic forces on the projectile/sabot decreased to almost zero just after the tube exit, with the ambient fluid around the tube being faster than the projectile by the substantial expansion. After exiting, the projectile penetrated the precursor shock wave at t 0.15 ms; this increased the aerodynamic forces acting on the projectile, initiated sabot separation, and generated multiple shock-wave interactions, which induce unsteady aerodynamic loads on the projectile.

• Access to Document (DOI)

Thrust Validation of Rotating Detonation Engine System by Moving Rocket Sled Test

Goto, K., Kato, Y., Ishihara, K., Nishimura, J., Matsuoka, K., Kasahara, J., Matsuo, A., Funaki, I., Nakata, D., Higashino, K., and Tanatsugu, N.

Journal of Propulsion and Power (Journal of Propulsion and Power)  Vol. 37 ( No. 3 ) 419 - 425 2021

Research paper (scientific journal), Joint Work, Accepted,  ISSN  07484658

　View Summary

This Paper presents the results of a rotating detonation engine rocket sled test demonstration on a 127 mm gauge rail at 100 m long high-speed running track test equipment to evaluate the thrust performance of a stand-alone rotating detonation engine propulsion system. The developed sled test demonstrator contained all onboard subsystems, including engine, gas cylinders, valves, controller, and a data logger. The sled test demonstrator successfully generated thrust, and three or four detonation waves of 1460 to 1800 m∕s were observed during the 2.0 s burn time. The measured specific impulse in the sled test was 144 s as estimated from the acceleration of the demonstrator and the mass decrease in the onboard cylinders. The test results were compared with the data from a laboratory-stand firing test using a load cell, and the specific impulses of the sled test were found to be comparable to the results of the stand firing test.

• Access to Document (DOI)

Experimental measurement of torque and force on a rotating detonation engine with six-axis force sensor

Sawada S., Kawasaki A., Matsuoka K., Kasahara J., Matsuo A., Funaki I.

AIAA Scitech 2021 Forum (AIAA Scitech 2021 Forum)     1 - 15 2021

ISSN  9781624106095

　View Summary

Rotating detonation Engine (RDE) is highly expected for the future propulsion systems due to its more compact structure than conventional internal combustion engines. It is because detonation waves circling in the order of km/s compress propellant instead of the mechanical complex compressor. Of interest on RDE, the torque around the z axis is important for practical use of the system. Due to the detonation waves, fluid inside the RDE produces friction on the chamber wall, which causes force and torque aside from thrust on RDE. In this study, we measured the torque by introducing the 6-axis force sensor which output the torque and axial force simultaneously. And we observed several modes, some of which were dominated positive or negative propagation duration, and others of which were contained both of propagation. From the results, we clarified the torque closely connected to the propagation of detonation waves in terms of the direction and strength. Moreover, we evaluated the effect on thrust performance of RDE. And we concluded that the contribution to RDE performance loss was effectively zero.

• Access to Document (DOI)

Propulsive performane of cylindrical rotating detonation engine with propellant injection cooing

Goto K., Ota K., Kawasaki A., Watanabe H., Itouyama N., Matsuoka K., Kasahara J., Matsuo A., Funaki I.

AIAA Scitech 2021 Forum (AIAA Scitech 2021 Forum)     1 - 7 2021

ISSN  9781624106095

　View Summary

An engine cooling concept for cylindrical rotating detonation engine which had an injector surface on the combustor side wall has been tested with three combustor length of 21, 30, and 6 mm. Thrust measurement of the cylindrical RDE (24-mm-diameter) was conducted with monitoring K-type thermocouples inserted in combustor wall. Single rotating detonation wave was observed with the combustor length of 30 and 69 mm in this study. Cooling effect due to the propellant injection was confirmed as the nearly saturated temperature response in the combustor side wall. when the chamber length is more than 30 mm, the specific impulse maintained more than 80% of theoretical value assuming sonic condition at the chamber exit. The result indicated that modest combustor length as an efficient thruster exists in the range of 30 to 69 mm.

Investigation into the effective injector area of a rotating detonation engine with impact of backflow

Goto K., Yokoo R., Kawasaki A., Matsuoka K., Kasahara J., Matsuo A., Funaki I., Kawashima H.

Shock Waves (Shock Waves)  31 ( 7 ) 753 - 762 2021

ISSN  09381287

　View Summary

For rotating detonation engines, the high-pressure region behind the detonation causes backflow into the plenum, making it difficult to properly design injectors to achieve the target pressure balance due to blockage of a part of the injector area during engine operation. In this paper, we present the pressure and thrust measurement of a rotating detonation engine with two different triplet injectors (fuel injector diameters of 0.8 mm and 1.0 mm) using gaseous methane, gaseous ethylene, and gaseous oxygen. The detonation wave propagation velocity with the fuel injector diameter of 0.8 mm was approximately 200 m/s higher than that with the fuel injector diameter of 1.0 mm. Combustor pressures and specific impulses were almost identical for both fuel injector diameters in this study. For our evaluation of the extent to which the available injector area can be utilized during engine operation, the effective injector area ratio was defined as the ratio of the plenum pressure during burn time to the pre-ignition value. Regardless of fuel species and fuel injector orifice diameter, the effective injector area ratio decreased proportionally with the ratio of combustor pressure to pre-ignition plenum pressure. This result implies that the pressure balance between the upstream plenum pressure and the combustor pressure can be roughly determined taking the effect of backflow into consideration.

• Access to Document (DOI)

Development of an S-Shaped Pulse Detonation Engine for a Sounding Rocket

Valentin Buyakofu, Matsuoka, K., Matsuyama, K., Kawasaki, A., Watanabe, H., Itouyama, N., Goto, K., Ishihara, K., Noda, T., Kasahara, J., Matsuo, A., Funaki, I., Nakata, D., Uchiumi, M., Habu, H., Takeuchi, S., Arakawa, S., Masuda, J., and Maehara, K.

Journal of Spacecraft and Rockets (AIAA)  59 ( 3 ) 850 - 860 2021

Research paper (scientific journal), Joint Work, Accepted,  ISSN  00224650

　View Summary

This paper presents the results of an S-shaped pulse detonation engine (PDE) ground firing test in the form of a detonation engine system. The world’s first technology demonstration of PDE in space using a sounding rocket is planned, and the aim is to control the rocket spin rate in the axial direction using pulsed detonation. The PDE operation at full sequence was successful. The despin rate change of the rocket between continuous oxygen supply and successful PDE operation is expected to be 0.95 deg ∕s per run. This change in despin rate can be measured by an onboard gyro sensor, making the system flyable. The test results were compared with data from thrust measurement tests conducted in a laboratory, the results of which confirmed the thrust generation under an ambient pressure of 0.5 0.1 kPa. The average thrust values in the thrust measurement experiments showed good agreement of 101 3% with a quasi-steady-state model introduced to predict the PDE thrust. These results demonstrate the feasibility of the newly developed PDE and its system as the world’s first technology demonstration of detonation propulsion in space.

• Access to Document (DOI)