Details of a Researcher - Matsuo, Akiko

Papers - Matsuo, Akiko

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Experimental Research for Clustering with the Coupled Cylindrical Rotating Detonation Engine

Sakata R., Inada M., Itouyama N., Matsuoka K., Kasahara J., Kawasaki A., Matsuo A., Funaki I.

AIAA SciTech Forum and Exposition, 2024 2024

It is possible to achieve a high flow rate and high thrust by clustering small cylindrical rotating detonation engines while reducing the axial length of the engines. In this study, a coupled cylindrical rotating detonation engine with two cylindrical RDEs (both combustors had a combustor inner diameter of 23 mm and an axial length of 30 mm) placed next to each other was fabricated and tested for clustering. The objective of the experiment was to achieve two-engine synchronized initiation with a single igniter. Experiments were conducted on the inner wall of the combustor with different connecting hole diameters d and wall heights h to evaluate the ignition mechanism, combustion mode, and propulsion performance. The propellants were gaseous ethylene and oxygen, and experiments were conducted under constant conditions of mass flow rate (the range of 40 ± 2 g/s), equivalent ratio (the range of 1 ± 0.1), and back pressure (approximately 10 kPa). The experimental results showed that when the two combustion chambers were completely separated by a wall, ignition occurred with a time delay in the chamber without an igniter. However, when a large hole (10 mm diameter) was placed in the wall separating the two combustion chambers, the flame propagated through the hole and synchronous initiation was successful. Synchronous initiation was also successful when the wall height was lowered (7 mm height), and high performance was obtained under both conditions. The flame was visualized by a high-speed camera during ignition and combustion, and the difference in combustion mode was confirmed. Frequency analysis was performed from the self-illuminated images taken by the high-speed camera to classify the

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Combustion Characteristics of Reflective Shuttling Detonation Combustor with Different Equivalence Ratio

Miyashita M., Matsuo A., Shima E., Kawasaki A., Itouyama N., Matsuoka K., Kasahara J.

AIAA SciTech Forum and Exposition, 2024 2024

A three-dimensional numerical analysis was performed on a reflective shuttling detonation combustor to investigate the propagation phenomena of detonation in a fuel-rich condition, with different equivalence ratio. The computational domain was a rectangular region where the hydrogen-air premixed gas flows in from the bottom injector with changing equivalence ratio from 1.0 to 2.0. As a result, continuous propagation of detonation was observed in each case, and the same trend of the propagation mode was observed. In addition, A decrease in the temperature value during combustion and an increase in the propagation velocity were observed as the equivalent ratio increased. Both trends were consistent with the trends of CJ temperature and CJ velocity with increasing equivalent ratio.

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A Bread Board Model Testing for In-Space Flight Demonstration of a Liquid-Propellant Detonation Engine System

Kawasaki A., Nakata K., Sato T., Sawada S., Kudo Y., Suzuki Y., Itouyama N., Matsuoka K., Matsuyama K., Kasahara J., Nakata D., Namera M., Eguchi H., Uchiumi M., Matsuo A., Funaki I., Nakamura S., Higashino K., Hirashima H.

AIAA SciTech Forum and Exposition, 2024 2024

In-space flight demonstration of a rotating detonation engine system using liquid nitrous oxygen and liquid ethanol as the propellant (DES-2) is planned in FY2024 by a collaborative research group of Nagoya University, Shizuoka University, Keio University, the Muroran Institute of Technology, and Institute of Space and Astronautical Science at the Japan Aerospace Exploration Agency (ISAS/JAXA). In this manuscript, a ground combustion test of the bread board model (BBM) of the DES-2, which has been conducted in May 2023, is briefly described. In the testing, the DES-2 BBM was operated in stand-alone mode and achieved detonation combustion as planned.

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Lagrangian dispersion and averaging behind a two-dimensional gaseous detonation front

Watanabe H., Matsuo A., Chinnayya A., Itouyama N., Kawasaki A., Matsuoka K., Kasahara J.

Journal of Fluid Mechanics (Journal of Fluid Mechanics) 968 2023.08

ISSN 00221120

Two-dimensional numerical simulations with the particle tracking method were conducted to analyse the dispersion behind the detonation front and its mean structure. The mixtures were 2H-O-7Ar and 2H-O of increased irregularity in ambient conditions. The detonation could be described as a two-scale phenomenon, especially for the unstable case. The first scale is related to the main heat release zone, and the second where some classical laws of turbulence remain relevant. The dispersion of the particles was promoted by the fluctuations of the leading shock and its curvature, the presence of the reaction front, and to a lesser extent transverse waves, jets and vortex motion. Indeed, the dispersion and the relative dispersion could be scaled using the reduced activation energy and the parameter, respectively, suggesting that the main mechanism driving the dispersion came from the one-dimensional leading shock fluctuations and heat release. The dispersion within the induction time scale was closely related to the cellular structure, particles accumulating along the trajectory of the triple points. Then, after a transient where the fading transverse waves and the vortical motions coming from jets and slip lines were present, the relative dispersion relaxed towards a Richardson-Obukhov regime, especially for the unstable case. Two new Lagrangian Favre average procedures for the gaseous detonation in the instantaneous shock frame were proposed and the mean profiles were compared with those from Eulerian procedure. The characteristic lengths for the detonation were similar, meaning that the Eulerian procedure gave the mean structure with a reasonable accuracy.

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Impact of mixture mass flux on hydrodynamic blockage ratio and Mach number of rotating detonation combustor

Noda, T., Matsuoka, K., Goto, K., Kawasa, A., Watanabe, H., Itouyama, N., Kasahara, J., and Matsuo, A.

Acta Astronautica (Acta Astronautica) 207 219 - 226 2023.06

Research paper (scientific journal), Joint Work, ISSN 00945765

To analyze non-ideal phenomena, such as burned gas backflow and non-detonation combustion, which affect the rotating detonation wave Mach number, simultaneous self-luminous visualization, time-averaged static pressure, fluctuating pressure, and thrust measurements with gaseous ethylene and oxygen were performed. Consequently, by doubling the number density of the fuel injectors, the hydrodynamic blockage ratio at the oxidizer inlet increased approximately 1.7-fold under the same oxidizer inlet area conditions. This may be attributed to the increase in the detonation propagation Mach number owing to the enhanced mixing of fuel and oxidizer. The relationship between the parasitic combustion fraction in front of the rotating detonation wave and the Mach number was also investigated by using a distributed heat release model. Consequently, it was suggested that experimental Mach number decreased from approximately 4.1 to 2.8 with increase in a mixture mass flux, and the theoretical detonation wave propagation Mach number was 7.3.

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Numerical investigation of the effects of injector configuration on flow structures in annular and cylindrical rotating detonation combustors

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

Science and Technology of Energetic Materials 84 ( 2 ) 17 - 23 2023.05

Research paper (scientific journal), Joint Work, Accepted

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Visualization and Performance Evaluation of a Liquid-Ethanol Cylindrical Rotating Detonation Combustor

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

Transaction of the Japan Society for Aeronautical and Space Sciences (Transactions of the Japan Society for Aeronautical and Space Sciences) 66 ( 2 ) 46 - 58 2023.03

Research paper (scientific journal), Accepted, ISSN 05493811

Rotating detonation combustors (RDCs) are among the combustors that use supersonic combustion waves known as detonation waves, and are expected to simplify engine systems and improve thermal efficiency due to their supersonic combustion and compression performance using shock waves. Research is also being actively conducted worldwide on a cylindrical RDC; a RDC without an inner cylinder, which is expected to simplify and downsize the combustor. However, most of the research was performed using gas propellants, and liquid propellants were rarely used. Since liquid propellants are used in many combustors, it is important to evaluate the performance of RDCs with liquid propellants. In this study, a cylindrical RDC with a liquid ethanol-gas oxygen mixture was constructed and tested at a flow rate of 31.5 ± 5.0 g/s, an equivalence ratio of 0.46-1.39, and a back pressure of 14.5 ± 2.5 kPa. The thrust was shown to depend strongly on the combustor bottom pressure history. In addition, the start-up process of the cylindrical RDC with liquid fuel was clarified by self-luminous and CH+ radical visualizations. It was found that the detonation wavefront propagated at a distance of 2-3mm from the combustor bottom, and the main combustion region was 10-15mm in height.

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Thrust Performance of Converging Rotating Detonation Engine Compared with Steady Rocket Engine

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

Journal of Propulsion and Power (Journal of Propulsion and Power) 39 ( 3 ) 297 - 307 2023.02

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

Rotating detonation engines (RDEs) have been actively researched around the world for application to nextgeneration aerospace propulsion systems because detonation combustion has theoretically higher thermal efficiency than conventional combustion. Moreover, because cylindrical RDEs have simpler combustors, further miniaturization of conventional combustors is expected. Therefore, in this study, with the aim of applying RDEs to space propulsion systems, a cylindrical RDE with a converging-diverging nozzle was manufactured; the combustor length Lc was changed to 0, 10, 30, 50, and 200 mm; and the thrust performance and combustion mode with the different combustor lengths were compared. As a result, four combustion modes were confirmed. Detonation combustion occurred with a combustor length of Lc = 0 mm: that is, a converging rotating detonation engine. The thrust performance of this engine was 94 to 100% of the theoretical rocket thrust performance, which is equivalent to the thrust performance of conventional rocket combustion generated at Lc = 200 mm. This study shows that detonation combustion can significantly reduce engine weight while maintaining thrust performance.

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Space Flight Demonstration of Rotating Detonation Engine Using Sounding Rocket S-520-31

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., and Yamada, K.

Journal of Spacecraft and Rockets (Journal of Spacecraft and Rockets) 60 ( 1 ) 273 - 285 2023.01

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

To create a new flyable detonation propulsion system, a detonation engine system (DES) that can be stowed in sounding rocket S-520-31 has been developed. This paper focused on the first flight demonstration in the space environment of a DES-integrated rotating detonation engine (RDE) using S-520-31. The flight result was compared with ground-test data to validate its performance. In the flight experiment, the stable combustion of the annulus RDE with a plug-shaped inner nozzle was observed by onboard digital and analog cameras. With a time-averaged mass flow of 182 11 g∕s and an equivalence ratio of 1.2 0.2, the RDE generated a time-averaged thrust of 518 N and a specific impulse of 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 in total, and the time-averaged torque from the rotational component of the exhaust during 6 s of operation was 0.26 N ⋅ m. The high-frequency sampling data identified the detonation frequency during the recorded time as 20 kHz in the flight, which was confirmed by the DES ground test through high-frequency sampling data analysis and high-speed video imaging.

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Flight Demonstration of Pulse Detonation Engine Using Sounding Rocket S-520-31 in Space

Buyakofu, V., 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 (Journal of Spacecraft and Rockets) 60 ( 1 ) 181 - 189 2023.01

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

A detonation engine system is successfully demonstrated for the first time in space using sounding rocket S-520-31 of the Japan Aerospace Exploration Agency/Institute of Space Astronautical Science. Detailed flight results of an S-shaped pulse detonation engine (PDE) installed in the rocket are presented herein. The flight is conducted to confirm that the PDE and its system operate at scheduled sequences in space, confirm the reproducibility of the PDE cycle, and despin the rocket around its axis. It is confirmed that the PDE operated successfully for 14 cycles in space. The experimental plateau pressure of 2.0 0.1 MPais 80 3% of the calculated plateau pressure, which suggests that detonation occurred in 14 cycles. The pressure profiles of the cycles are similar, and the pressure integrals are 2.0 0.1 kN⋅ s∕m2, confirming the excellent reproducibility of the PDE cycle. A probability statistical approach assuming a Gaussian distribution is applied to determine the average angular acceleration difference between processes of the PDE operation, mixture supply, and oxygen supply. The results suggested that the PDE despun the rocket via the thrust produced via detonation combustion, which is consistent with a quasi-steady-state model with an accuracy of 101 15%.

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The blast mitigation mechanism of a single water droplet layer and improvement of the blast mitigation effect using multilayers in a confined geometry  

Shibue, K., Sugiyama, Y., and Matsuo, A.

International Journal of Multiphase Flow (International Journal of Multiphase Flow) 159 2023

Research paper (scientific journal), Joint Work, Corresponding author, Accepted, ISSN 03019322

The blast mitigation mechanism of a water droplet layer was numerically studied, considering convective heat transfer and quasi-steady drag. The propagation of a one-dimensional blast wave in a shock tube was modeled, leading to momentum and energy transfer between the water droplets and the air. Convective heat transfer absorbed part of the energy, thereby directly mitigating the blast wave. Although the quasi-steady drag transferred less energy than the convective heat transfer, it exhibited greater blast mitigation. Thus, a parametric study using characteristic values and describing the water droplet layer was conducted to examine the blast mitigation mechanisms involved in the quasi-steady drag, aside from the energy absorption effect. It was determined that each time the shock wave reached the air/water droplet interface, the quasi-steady drag divided the wave into transmitted and reflected waves. Because the one-dimensional blast wave helped to carry some of the energy, the division reduced the energy carried by the incident blast wave, resulting in its mitigation. Finally, the blast mitigation effect of multilayers was compared with that of a single layer. The result showed that increasing the number of layers synergistically improved the blast mitigation effect by dividing the shock wave multiple times.

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In-space Demonstration of Rotating Detonation Engines: from Gaseous Propellent to Liquid Propellant Applications

Itouyama N., Matsuyama K., Matsuoka K., Kawasaki A., Ishihara K., Ito S., Nakata K., Sato T., Kasahara J., Matsuo A., Funaki I.

AIAA SciTech Forum and Exposition, 2023 2023

Detonation engines exert various characteristics and high propulsion performance depending on their operation mode. Our group has demonstrated the integrated Detonation Engine System (DES) in outer-space environment using the sounding rocket named S-520-31in Japan Aerospace Exploration Agency. The system was operated by the detonation of premixed gas composing of methane gas and oxygen gas. As the next step, we are planning to perform the in-space demonstration of detonation engine with liquid propellants in order to clarify its milestones on the achievement of space application of detonation engines as same as the conventional chemical propulsion systems. This paper introduces the overview of our group’s measures about the in-space demonstration of detonation engine, especially for rotating detonation engine, in the last few years.

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Experimental Research on Internal Flow Structure of Cylindrical Rotating Detonation Engine Using Ethanol

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

AIAA SciTech Forum and Exposition, 2023 2023

Rotating detonation engine (RDE) is one of the combustors using detonation 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, and it is important to evaluate the performance of RDE with liquid propellants. In this study, a cylindrical RDE, which is an RDE without inner cylinder, with a liquid ethanol and gaseous oxygen was tested when mass flow rate ṁ was 16.5-62.0 g/s, equivalence ratio Φ was 0.42-1.46, and the back pressure pb was 13.0 ± 0.8 kPa. to evaluate operating characteristics, internal visualization, and thrust performance. As a result, Detonation was not observed when the equivalence ratio was greater than the stoichiometric value. Moreover, the combustor length decreased, detonation was not observed at any equivalence ratio when a total mass flux was 101±6 kg/(m2-s). Moreover, it is found that the equivalent ratio at ignition has a significant effect on successful ignition and combustion. Detonation was also observed when the total mass flux increased. These results suggest that combustor length, equivalent ratio, and total mass flux affect the propagation mode of a liquid RDE. Moreover, the propagation velocity is almost consistent with the acoustic frequency calculated by sound speed. From visualization results of different equivalence ratio, the brightness value distribution, pressure distribution, and composite image show that the main combustion area is almost the same. In thrust performance, a high specific impulse compared to the ideal specific impulse is achieved regardless of equivalence ratio and total mass flux. The thrust using control surface matched well with the experimental thrust.

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Numerical study of the effect on blast-wave mitigation of the quasi- steady drag force from a layer of water droplets sprayed into a confined geometry

Shibue, K., Sugiyama, Y., and Matsuo, A.

Process Safety and Environmental Protection (Process Safety and Environmental Protection) 160 491 - 501 2022.04

Research paper (scientific journal), Joint Work, Corresponding author, Accepted, ISSN 09575820

The blast-mitigation effect produced by water droplets sprayed into a confined geometry was studied. When a one-dimensional blast wave interacts with a layer of water droplets, blast mitigation occurs that can be expected to reduce the damage from an accidental explosion. Focus was on the blast-mitigation effect produced by the quasi-steady drag force between the shocked air and the stationary water droplets. Momentum transfer due to the drag mitigated the blast wave, and the maximum blast-wave mitigation occurred just behind the layer of droplets. The study parameters (the volume fraction of water droplets, the location and length of the region with sprayed water droplets, and the mass of the high explosive) demonstrated that higher momentum loss from the air resulted in greater mitigation. The momentum loss was evaluated using the volume fraction and length of the layer of water droplets and the quasi-steady drag force when the shock wave reached the layer. Relationships were formulated between the initial conditions, the momentum loss, and the blast-mitigation effect, which made it possible to evaluate quantitatively the mitigation effect on the blast wave caused by the water droplets.

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Investigation of reflective shutting detonation cycle by Schlieren and Chemiluminescence Photography

Taguchi, T., Yamaguchi, M., Matsuoka, K., Kawasaki, A., Watanabe, H., Itouyama, N., Kasahara, J., and Matsuo, A.

Combustion and Flame 236 2022.02

Research paper (scientific journal), Joint Work, Corresponding author, Accepted

Experimental investigation of inner flow of a throatless diverging rotating detonation engine

Nakata, K., Ishihara, K., Goto, K., Itouyama, N., Watanabe, H., Kawasaki, A., Matsuoka, K., Kasahara, J., Matsuo, A., Funaki, I., Higashino, K., Braun, J., Meyer, T., and Paniagua, G.

Proceedings of The Combustion Institute (Proceedings of the Combustion Institute) 39 ( 3 ) 3073 - 3082 2022

Research paper (scientific journal), Joint Work, Accepted, ISSN 15407489

The inner flow containing the ignition process of a throatless and diverging (constant diverging angle α=5 deg) rotating detonation engine (RDE) was experimentally investigated by both optical observation and pressure measurement. An RDE that had an acrylic engine wall with a length of 70 mm and an inlet diameter of 20 mm was tested under a low back pressure condition using gaseous C2H4-O2 as the propellants. A high-speed camera captured the ignition process, of which there were two types: deflagration-to-detonation transition and direct initiation. In addition to the visualizations of the inner flow, the pressure values in an RDE with an acrylic wall were compared with those of an RDE having a stainless steel wall with 8 pressure ports in the axial direction. The comparison revealed that the RDE exhaust flow was supersonic. A strong self-luminescence and CH* luminosity area near the injector surface during combustion suggested that the short completion of detonation combustion enabled the thermal choking of the subsonic propellant gases even when the channel was diverging. Moreover, the axial position of the erosion in the acrylic wall indicated that the heat release region was near the injector surface. Finally, a theoretical flow model for the RDE was compared to the experimental data, and the results suggested that the inner flow in the RDE can be reasonably described as a quasi-one-dimensional steady flow.

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Supersonic Exhaust from a Rotating Detonation Engine with Throatless Diverging Channel

Nakata, K., Ota, K., Ito, S., Ishihara, K., Goto, K., Itouyama, N., Watanabe, H., Kawasaki, A., Matsuoka, K., Kasahara, J., Matsuo, A., Funaki, I., Higashino, K., Braun, J., Meyer, T., and Paniagua, G.

AIAA Journal (AIAA Journal) 60 ( 7 ) 4015 - 4023 2022

Research paper (scientific journal), Joint Work, Corresponding author, Accepted, ISSN 00011452

Converging–diverging nozzles are common in rocket engine systems to increase the exhaust velocity and improve thrust performance. In this study, we focused on the acceleration of subsonic burned gas without a structural throat via detonation to realize a simple and compact engine. We developed and tested a rotating detonation engine (RDE) without a throat and with a diverging channel (constant diverging angle α 5 deg). Gaseous C2H4 and O2 were used as the propellants, and the mass flow rate ranged from 62 to 134 g∕s in the combustion tests under low back-pressure conditions. We measured pressure and thrust, as well as high-speed imaging of self-luminescence of the combustion and imaging of the exhaust plume. The pressure at the exit was less than one-fifth of the maximum pressure in the RDE, significantly below the value for a sonic flow. The results suggested that the exhaust flow was supersonic, with values up to Mach 1.7, without the need of a converging section within the engine. In addition to the estimated Mach number from the measured pressure, the exhaust plume images coherently indicated the existence of supersonic exhaust.

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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

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.

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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

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.

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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

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.

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