研究者詳細 - 松尾　亜紀子

論文 - 松尾　亜紀子

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Lagrangian characterization of induction and reaction timescales in a cellular gaseous detonation

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

Physics of Fluids 37 （ 2 ） 2025年02月

ISSN 10706631

A Lagrangian approach was proposed to analyze induction and reaction times in the cellular gaseous detonation. Two-dimensional simulations in an argon-diluted and non-diluted hydrogen-based mixtures were performed with detailed chemistry, along particle trajectories. The distribution of the induction and reaction times inside the cell was significantly different between the Eulerian and the Lagrangian perspectives, the latter showing non-monotonic behavior. Preferential thermodynamic paths laid along the Rankine-Hugoniot curve and behind transverse waves (TW). All particles were ignited within half and one cell cycle for the diluted and non-diluted mixture, respectively. The ignition mechanisms were not only one-dimensional, but also multi-dimensional, with ignition behind the TW being the most important, and collision of TW and triple points being secondary. A new topology inside the cell could be drawn, from the intersection of the ignition front with TW. TW appeared as phase waves in the ( x , t ) diagram. Comparison of H2O mass fraction between local and equilibrium values indicated that a local chemical disequilibrium remained (superequilibrium), due to TW. Equating the mean sonic plane with thermochemical equilibrium in the non-diluted case is not completely accurate. Furthermore, the characteristic time scales for chemical and hydrodynamic phenomena were compared. The diffusive phenomenon did not make any contribution in the mixtures tested. In comparison with the Zel'dovich-von Neumann-Döring model, a shorter average induction time was found in the non-diluted mixture, which is not in line with the results from previous Favre approaches. The average reaction time was also shorter in both mixtures.

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Numerical Investigation of Detonation-assisted Fuel Injection Method in Scramjet Engines with Rotating Detonation Combustor

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

AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2025 2025年

This study conducted three-dimensional numerical analyses for the interaction of the exhaust flow of a Reflective Shuttling Detonation Combustor (RSDC) and supersonic air-flow to investigate the new concept of detonation combustion as an injection system for scramjet engines. The computational targets of RSDC are the region where the hydrogen-air premixed gas flows from three rows of injectors set at the bottom in fuel-rich conditions. The additional fuel for the supersonic turbulent combustion in the scramjet combustor was added from the side wall of RSDC. The reflective wall distance, the height, and the width of RSDC were set to 45.0 mm, 53 mm, and 5 mm. The supersonic air-flow area simulated scramjet combustor is the region where the air flows at Mach number 3, and the RSDC is related to the main flow area as an injector. The governing equations used in this study were a three-dimensional compressible Navier-Stokes equation and a chemical conservation law of 9 species. The mixed scale model for LES was also used as the turbulence model. As a result, turbulent combustion was observed in the supersonic flow area by mixing the supersonic flow and the exhaust flow of RSDC, including additional fuel. The dynamic pressure ratio of the exhaust flow was 23% greater than that of the single-hole injection (previous injection system of scramjet engine) and 13% greater than that of deflagration combustion. The penetration height was improved, and the turbulent combustion was continuously achieved without an ignition system or a cavity due to the detonation combustion inside the RSDC.

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Synchronized Initiation of Two Cylindrical Rotating Detonation Engines

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

Journal of Propulsion and Power 40 （ 6 ） 849 - 858 2024年11月

ISSN 07484658

A coupled cylindrical rotating detonation engine (RDE) 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 tested for rocket clustering application. 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 combustors with different connecting-hole diameters and wall heights to evaluate the ignition delay time, combustion mode, and propulsion performance. The propellants were gaseous ethylene and oxygen, and experiments were conducted under constant conditions of mass flow rate (40 ± 2g∕s), equivalence ratio (1.0±0.1), and backpressure (approximately 10 kPa). When the two combustion chambers were completely separated by a wall, ignition occurred with a time delay of 260 ms in the chamber without an igniter. However, when a large hole (10-mm diameter) was placed in the wall separating the two combustion chambers, synchronous initiation was successful. Synchronous initiation was also successful when the wall height was lowered (7-mm height). Under both conditions, the same level of specific impulse was achieved as for RDEs operating at the same mass flux.

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Wave-converging pressure increase in curved cylindrical rotating detonation combustors

Oda Y., Sawada S., Itouyama N., Matsuoka K., Kasahara J., Kawasaki A., Matsuo A., Funaki I.

Proceedings of the Combustion Institute （Proceedings of the Combustion Institute） 40 （ 1-4 ） 2024年08月

共著, 査読有り, ISSN 15407489

The internal combustion structure of a curved cylindrical rotating detonation combustor (RDC) was experimentally investigated through optical observation and pressure measurements. Experiments were conducted under low back pressure conditions using two combustion chambers having, in both cases, an internal diameter of 17.5 mm, a curvature of radius of 38.1 mm, and an outlet angle of 90 deg, with gaseous C2H4 and O2 as the propellants. One was made of SUS and the other of resin. In addition to vibrations and bottom center pressure, pressure distribution on the sidewalls was measured for the SUS combustor, and high-speed camera observations of self-luminescence from the radial direction were conducted for the resin combustor. These measurements were compared by varying the equivalence ratio. The frequency analysis results obtained from vibrations and self-luminescence indicated that the strongly-coupled detonation mode exhibited a higher peak frequency, suggesting that the detonation waves may have different propagation speeds or rotational positions. In terms of pressure distribution and self-luminescence, only the strongly-coupled detonation mode inside near the bottom surface exhibited high pressure and brightness values. This suggested the potential for converging pressure by employing rotating detonation waves and a curved tube.

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Visualization and thrust measurement of rotating detonation engine with various channel expansion angles

Nakajima K., Sawada T., Matsuoka K., Itouyama N., Kasahara J., Kawasaki A., Matsuo A.

Proceedings of the Combustion Institute 40 （ 1-4 ） 105577 2024年07月

研究論文（国際会議プロシーディングス）, 共著, 査読有り, ISSN 15407489

Although a rotating detonation engine (RDE) is expected to achieve a pressure gain (PG) at which the total pressure of the product at the engine exit exceeds the total pressure of the supplied oxidizer, many non-ideal phenomena occurring in the RDE hinder PG achievement. This study focused on the channel expansion angle near the injector outlet, which is considered to affect the detonation wave structure and the PG performance. To this end, thrust and self-luminescence were conducted by varying the channel expansion angle near the axially injected oxidizer injector outlet; that is 90° 30°and 6° Results indicate that thrust was not affected by the expansion angle. However, the corresponding propagation velocities of the detonation waves were 2160, 2025, and 1755 m/s, indicating that the propagation velocity decreased as the expansion angle decreased. The PG performance, which is the ratio of the total pressure at the combustor exit to the total pressure of the oxidizer plenum, was 57, 57, and 49 %. At the expansion angle of 6°, the PG performance decreased due to a significant increase in the total pressure of the oxidizer plenum. Visualization of the self-luminescence showed that as the expansion angle decreased, the burned gas backflow became more significant, confirming unstable propagation.

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Combustion Structure of a Cylindrical Rotating Detonation Engine with Liquid Ethanol and Nitrous Oxide

Sato T., Nakata K., Ishihara K., Itouyama N., Matsuoka K., Kasahara J., Kawasaki A., Nakata D., Eguchi H., Uchiumi M., Matsuo A., Funaki I.

Combustion and Flame （Combustion and Flame） 264 2024年06月

ISSN 00102180

Liquid propellants are commonly used in rocket engines due to their high energy density. In this study, we focused on the use of liquid ethanol and liquid nitrous oxide as propellants in a cylindrical rotating detonation engine (RDE) for practical applications, and examined the effects of various conditions on the bipropellant RDE performance. We varied the vapor quality of nitrous oxide in a mixture of liquid and gaseous nitrous oxide due to the flash boiling. In addition, we varied the liquid ethanol temperature, the injector stiffness, and momentum angle of the propellant. We then analyzed the effects of these changes on the propagation mode of the RDE under atmospheric pressure conditions. The results showed that a rotating detonation wave was observed under the combination of high vapor quality of nitrous oxide, high injector stiffness, high liquid ethanol temperature, and high momentum angle. The propagation velocity was obtained as 73∼83 % of the Chapman-Jouguet velocity. Long-duration combustion tests were also conducted, and the internal combustion structure was confirmed from the erosion of the carbon-carbon (C/C) composite material. The detonation wave was considered to propagate near the bottom of the RDE with liquid propellants from the erosion heights of the C/C composite. The erosion height was confirmed to be consistent with the fill height calculated using the sound speed of the nitrous oxide. In addition, characteristic exhaust velocity efficiencies of more than 85 % were achieved in all combustion tests.

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Combustion Structure of a Cylindrical Rotating Detonation Engine with Liquid Ethanol and Nitrous Oxide

Sato, T., Nakata, K. Ishihara, K., Itouyama, N., Matsuoka, K., Kasahara, J., Kawasaki, A., Nakata, D., Eguchi, H., Uchiumi, M., Matsuo, A., and Funaki, I.

Combustion and Flame 264 2024年06月

研究論文（学術雑誌）, 共著, 査読有り

Synchronized Initiation of the Two Cylindrical Rotating Detonation Engines

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

Journal of Propulsion and Power 2024年06月

研究論文（学術雑誌）, 共著, 査読有り

The circumferential force on a cylindrical rotating detonation engine

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

Proceedings of the Combustion Institute 40 （ 1-4 ） 105490 2024年06月

研究論文（国際会議プロシーディングス）, 共著, 査読有り, ISSN 15407489

Rotating detonation engines (RDEs) utilize single or multiple detonation waves continuously traveling at velocities ranging from 1 to 3 km/s. These waves induce a strong swirl near the inlet surface, generating circumferential shear stress on the surface of the combustion chamber. This stress works as torque in the axial direction. This characteristic is promising for turbine applications. However, for long-term operation, such applications require reduction in the thermal damage risk. While annular RDEs, which have an inner wall surface, involve the thermally independent center body, cylindrical RDEs have no center body and more suitable. On the other hand, the contact surface exposed to the swirling flow is smaller. This study evaluated the differences in combustion mode and circumferential phenomena between annular and cylindrical RDEs through the measurement of thrust and torque. The results showed that the torque generated by the cylindrical RDE was as large (around 0.1 Nm) as that generated by the annular RDE under mass flow rate of approximately 0.05 kg/s. This was due to the stronger swirl derived from the 16 % faster detonation waves. In addition, the presence of strong detonation combustion around the outer wall of the cylindrical RDE was indicated through the control surface method. When a fuel‒oxidizer was supplied to the strong detonation combustion region, the detonation waves reached velocities around the Chapman–Jouguet value, with the maximum torque in this study being 0.127 Nm. From these results, the effect of the propagation mode of detonation waves on the circumferential flow and force was evaluated. The results indicated that the increase of detonation wave velocity enhanced the circumferential flow and force. The tendency to increase became large as the number of the detonation waves increased.

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Nitrous Oxide/Ethanol Cylindrical Rotating Detonation Engine for Sounding Rocket Space Flight

Ishihara K., Sato T., Kimura T., Nakajima K., Nakata K., Itouyama N., Kawasaki A., Matsuoka K., Matsuyama K., Kasahara J., Eguchi H., Nakata D., Uchiumi M., Matsuo A., Funaki I.

Journal of Spacecraft and Rockets 62 （ 1 ） 44 - 54 2024年05月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00224650

There are few experimental studies on rotating detonation engines (RDEs) with liquid propellants. This study reveals the static thrust performance of a cylindrical RDE with ethanol and liquid nitrous oxide as propellants under atmospheric pressure. This RDE had an inner diameter of 40 mm, a maximum combustor length of 230 mm, a nozzle contraction ratio of 1.7, and a nozzle expansion ratio of 9.1. Nineteen experiments were conducted at total mass flow rates of184–210 g∕s, mixtureratios of 3.6–5.9, andcombustionpressures of 0.35–0.46 MPa, resulting in a maximum detonation velocity of 1840 m∕s (approximately 80% of the theoretical detonation velocity, 2251 m∕s), maximum thrust at sea level of 294 N, and maximum specific impulse at sea level of 148 s. In addition, the maximum characteristic exhaust velocity, c*,was1716 m∕s, which was 99% of the theoretical value. The characteristic length of the combustion chamber at this time was 0.15 m. Since conventional rocket combustion requires 1.57 m to achieve the same c* efficiency, this study shows that detonation combustion can reduce the combustor size by 88%.

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Effect of Channel Expansion Angle near Injector Outlet on a Rotating Detonation Engine Performance

Nakajima, K., Matsuoka, K., Itouyama, N., Kasahara, J., Kawasaki, A., and Matsuo. A.

ShockWaves（ICDERS　Special Issue） 2024年05月

研究論文（学術雑誌）, 共著, 査読有り

Effect of channel expansion angle near injector outlet on a rotating detonation engine performance

Nakajima K., Matsuoka K., Itouyama N., Kasahara J., Kawasaki A., Matsuo A.

Shock Waves （Shock Waves） 2024年

ISSN 09381287

A rotating detonation engine (RDE) is expected to achieve a pressure gain (PG) in which the total pressure of the product at the engine exit exceeds the total pressure of the supplied oxidizer. However, many non-ideal phenomena exist in the RDE, which hinder the achievement of a PG. This study focused on the channel expansion angle near injector outlet which is considered to impact the structure of detonation wave and the PG performance. Combustion tests were conducted by varying the channel expansion angle near the axially injected oxidizer injector outlet to 90∘ and 30∘. As a result, the thrust was not affected by the expansion angle, but the propagation velocity of detonation wave at an expansion angle of 90∘ was approximately 5% higher than that at an expansion angle of 30∘. A comparison utilizing the pressure increase ratio obtained from the fluctuating pressure in the combustion chamber suggested that the Mach number of the detonation wave was higher for an expansion angle of 90∘. As a result of evaluating the PG performance utilizing the equivalent available pressure obtained from thrust measurements, it was not confirmed to differ with changes in the expansion angle.

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The Rotating Detonation Engines with The Helical Combustion Chambers

Sawada S., Itouyama N., Matsuoka K., Kasahara J., Braun J., Paniagua G., Kawasaki A., Watanabe H., Higashino K., Matsuo A., Funaki I.

AIAA SciTech Forum and Exposition, 2024 2024年

Rotating detonation engines (RDEs) utilize a detonation wave circling at velocities on the order of kilometers inside a combustion chamber. In this study, the rotation direction was controlled by strategically positioning the ignition source at a specific location within a helical combustion chamber featuring sinusoidal circle cross section. Twenty-five combustion tests were conducted experimentally, involving two types of geometries with different helical directions. When the ignition was placed 30.3–45.7 mm far from the inlet surface, the direction of the rotating detonation wave was decided by the helical direction in all cases. By assuming that the direction of the rotating detonation wave was decided with a probability of 0.5, the probability of the 25 tests was 3×10-8, which was well beyond the 0.05 significance level. The direction of the rotating detonation wave was thus clearly depended on the helical direction. Moreover, the helical geometry is promising for the integration of combustion and torque extraction, thanks to the efficient extraction of torque primarily derived from the circumferential pressure difference on the helical crests. The torque generated was 0.040– 0.068, achieved with a mass flow rate of approximately 0.029 kg/s. The helical direction also controlled the direction of the torque.

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Overview for the research and development of Rotating Detonation Engine Systems using a liquid-liquid propellant combination

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

AIAA SciTech Forum and Exposition, 2024 2024年

Rotating detonation engines (RDEs) exert various characteristics and high propulsion performance depending on their operation mode. Our group has already demonstrated the RDE equipped with gas ethylene and gas oxygen in an outer-space environment using a sounding rocket. For the next step of R&D of detonation engine systems, we are planning to perform the in-space demonstration of detonation engine with liquid propellants to clarify its milestones on the achievement of space application of detonation engines as same as the conventional chemical propulsion systems, especially the propellant combination of ethanol and liquid nitrous oxide. This paper introduces the overview of our group’s measures on the in-space demonstration of the RDE system in the last few years.

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Operating Characteristics of Cylindrical Rotating Detonation Engines Using Liquid Ethanol and Liquid Nitrous Oxide

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

AIAA SciTech Forum and Exposition, 2024 2024年

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 supersonic combustion and compression performance by shock waves. Liquid propellants are commonly used in rocket engines due to their high energy density. However, most of these studies use gas propellants, so it is important to investigate the effect of the combustor length, mixture ratio, and propellant temperature on bipropellant RDE operation. In this study, cylindrical RDEs with liquid ethanol and liquid nitrous oxide were tested under atmospheric pressure conditions. The rotating detonation wave was observed in the cylindrical RDE with the liquid ethanol. About the effect on propagation mode, detonation mode was not observed when the mixture ratio was low, and the range of the mixture ratio in which the detonation mode was observed decreased as the combustor length decreased. When the ethanol liquid temperature was low, the detonation mode was not observed because the atomization performance and combustibility became lower. In addition, the rotating propagation mode was observed when the vapor quality of nitrous oxide was higher than about 9% because the propellant velocity was increased at the higher vapor quality of nitrous oxide, and this improved the atomization. Moreover, the throat pressure was close to the critical pressure for isentropic flow with a smaller ethanol injector diameter. The results of the experimental characteristic exhaust velocity in all combustion tests were in close match with the ideal value for the constant pressure mode. However, when the combustor length was 90 mm, the values in deflagration mode were much lower than the ideal value. Therefore, these results indicated that the combustion region was longer in the deflagration mode and with large ethanol injector diameter than in the detonation mode and small diameter, respectively.

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Numerical Investigation on Throatless Diverging Rotating Detonation Engines

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

AIAA SciTech Forum and Exposition, 2024 2024年

In this study, the flow characteristics and propulsive performance of throatless diverging rotating detonation engines (RDEs) are investigated. Seven diverging angles, 0, 1, 2, 3, 4, 5, and 6 deg were examined with constant length 70 mm and bottom diameter 20 mm using stoichiometric hydrogen-air premixed gas as propellant. The temperature and pressure in the combustor decreased as the diverging angle increased, and the continuous detonation propagation could not be observed in the case of 6 deg. Supersonic exhaust was achieved for all conditions with α > 0 deg, and the exit Mach number increased with larger angle and was 1.71 in 5 deg channel. The expansion ratio increased, but the combustion efficiency decreased with larger angle, so the thrust and specific impulse did not monotonically increase with increasing diverging angle. The specific impulse reached a maximum around 3 deg in this geometry.

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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年

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年

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年

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年08月

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., Kawasaki A., Watanabe H., Itouyama N., Kasahara J., Matsuo A.

Acta Astronautica （Acta Astronautica） 207 219 - 226 2023年06月

研究論文（学術雑誌）, 共著, 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年05月

研究論文（学術雑誌）, 共著, 査読有り

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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., Funaki I.

Transactions of the Japan Society for Aeronautical and Space Sciences （Transactions of the Japan Society for Aeronautical and Space Sciences） 66 （ 2 ） 46 - 58 2023年03月

研究論文（学術雑誌）, 査読有り, 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., Higashino K.

Journal of Propulsion and Power （Journal of Propulsion and Power） 39 （ 3 ） 297 - 307 2023年02月

研究論文（学術雑誌）, 共著, 査読有り, 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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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., Maehara K., Nakao T., Yamada K.

Journal of Spacecraft and Rockets （Journal of Spacecraft and Rockets） 60 （ 1 ） 181 - 189 2023年01月

研究論文（学術雑誌）, 共著, 査読有り, 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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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., Yamada K.

Journal of Spacecraft and Rockets （Journal of Spacecraft and Rockets） 60 （ 1 ） 273 - 285 2023年01月

研究論文（学術雑誌）, 共著, 査読有り, 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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The blast mitigation mechanism of a single water droplet layer and improvement of the blast mitigation effect using multilayers in a confined geometry

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

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

研究論文（学術雑誌）, 共著, 責任著者, 査読有り, 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年

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年

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., Matsuo A.

Process Safety and Environmental Protection （Process Safety and Environmental Protection） 160 491 - 501 2022年04月

研究論文（学術雑誌）, 共著, 責任著者, 査読有り, 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年02月

研究論文（学術雑誌）, 共著, 責任著者, 査読有り

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., Paniagua G.

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

研究論文（学術雑誌）, 共著, 査読有り, 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., Paniagua G.

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

研究論文（学術雑誌）, 共著, 責任著者, 査読有り, 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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Flight Demonstration of Detonation Engine System Using Sounding Rocket S-520-31: Performance of Rotating Detonation Engine

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

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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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年

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%.

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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年

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.

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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年

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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年

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.

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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年

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

Kasahara H., Matsuo A.

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

研究論文（学術雑誌）, 共著, 査読有り, 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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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年

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.

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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年

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.

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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., Kawashima H.

Journal of Propulsion and Power （Journal of Propulsion and Power） 38 （ 3 ） 410 - 420 2021年11月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 07484658

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.

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Torque Around Axial Direction on Rotating Detonation Engines

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

Journal of Propulsion and Power （Journal of Propulsion and Power） 38 （ 1 ） 59 - 70 2021年09月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 07484658

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.

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Coupling Simulation of Thermal Response to Exhaust Gas at External Nozzle

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

Journal of Spacecraft and Rockets （Journal of Spacecraft and Rockets） 59 （ 1 ） 260 - 270 2021年09月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00224650

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.

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Experimental Study on Truncated Conical Rotating Detonation Engines with Diverging Flows

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

AIAA Propulsion and Energy Forum, 2021 （AIAA Propulsion and Energy Forum, 2021） 2021年07月

共著

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.

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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., Funaki I.

AIAA Propulsion and Energy Forum, 2021 （AIAA Propulsion and Energy Forum, 2021） 2021年07月

共著

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.

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Three-dimensional numerical investigation of hypersonic projectile launched by railgun on transitional ballistics

Kasahara H., Matsuo A.

Journal of Spacecraft and Rockets （Journal of Spacecraft and Rockets） 58 （ 4 ） 1 - 17 2021年06月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00224650

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.

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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., and Kawashima, H.

Shock Waves 2021年01月

研究論文（学術雑誌）, 共著, 査読有り

Thrust validation of rotating detonation engine system by moving rocket sled test

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

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

研究論文（学術雑誌）, 共著, 査読有り, ISSN 07484658

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.

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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年

ISSN 9781624106095

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） 2021年

ISSN 09381287

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.

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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年

ISSN 9781624106095

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.

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

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

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00224650

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.

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Mach 4 propulsion wind tunnel test of high-mach integrated control experimental aircraft (Himico)

Taguchi H., Hirotani T., Takahashi H., Oki J., Sato T., Tezuka A., Tsue M., Tsuchiya T., Nakaya S., Matsuo A.

Accelerating Space Commerce, Exploration, and New Discovery conference, ASCEND 2021 （Accelerating Space Commerce, Exploration, and New Discovery conference, ASCEND 2021） 2021年

Mach 4 propulsion wind tunnel experiment of High Mach Integrated Control Experimental Aircraft (HIMICO) has been conducted in order to evaluate the installed performance of hypersonic engine, which is operated by hydrogen fuel. Surface temperature of the airframe, total pressure recovery of hypersonic air intake, temperature field around the exhaust nozzle, and longitudinal aerodynamic forces are measured in the experiment. The effect of the flow field around the airframe to the hypersonic engine performance and the effect of the engine exhaust to the aerodynamic forces are discussed based on the experimental results.

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Numerical analysis on behavior of dilute water droplets in detonation

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

Proceedings of The Combustion Institute 2020年10月

研究論文（学術雑誌）, 共著, 査読有り

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Optical Measurement of Fluid Motion in Semi-Valveless Pulse Detonation Combustor with High-Frequency Operation

Kubota A., Matsuoka K., Kawasaki A., Kasahara J., Watanabe H., Matsuo A., Endo T.

Combustion Science and Technology （Combustion Science and Technology） 192 （ 2 ） 197 - 212 2020年02月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00102202

© 2018, © 2018 Taylor & Francis Group, LLC. The purge layer of a semi-valveless pulse detonation cycle (PDC) needs to be minimized for operating at a gas-dynamic upper frequency limit. Therefore, it is essential to better understand the process of burned gas backflow for minimizing the purge layer thickness. The flow field of the semi-valveless PDC was visualized to illustrate the movement of burned gas. A combustor of length of 95 mm with a 10-mm-square cross section was used. Supercritical ethylene and oxygen gas were used as fuel and oxidizer, respectively, and the operation frequency was 604 Hz. The unsteady refilling process of the detonable mixture was modeled by an isentropic flow. In addition, the detailed burned gas blowdown process with deflagration-to-detonation transition (DDT) and the backflow were captured. It was shown that the retonation wave generated by the DDT process was the primary trigger of the burned gas backflow. When the duration required for the DDT process was sufficiently shorter than that of the burned gas blowdown process, it was found the latter could be reproduced with approximately 90% accuracy by one-dimensional numerical analysis without the DDT process.

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Cylindrical rotating detonation engine cooling by means of propellant injection

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

AIAA Propulsion and Energy 2020 Forum （AIAA Propulsion and Energy 2020 Forum） 1 - 9 2020年

ISSN 9781624106026

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. An engine cooling concept for cylindrical rotating detonation engine which had an injector surface on the combustor side wall has been tested and demonstrated. 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 in the testing conditions ranging from 31 to 59 g/s in this study. Combustion tests for 4.0 ~ 4.9 s were successfully done, and all injector side wall temperature increases were suppressed compared to that of combustor base plate, which had no cooling structure. This could be due to the cooling effect by the heat exchange of propellant injection. In the 4.9 s combustion test with 31 g/s, all thermocouples inserted in the combustor side wall which had the propellant injector surface showed a temperature decreasing 2.5 s after ignition even though the combustion was continuing, and implied the combustion mode shift.

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Prediction of pressure loss in injector for rotating detonation engines using single-element simulations

Suzuki T., Matsuo A., Daimon Y., Kawashima H., Kawasaki A., Matsuoka K., Kasahara J.

AIAA Propulsion and Energy 2020 Forum （AIAA Propulsion and Energy 2020 Forum） 1 - 10 2020年

ISSN 9781624106026

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. This paper newly proposes a prediction method for estimation of pressure loss in injector for rotating detonation engines with lower cost calculation. This method is supposed to be used to design new injector shapes. Although the conventional full-scale simulation contains plenum, all injectors, and combustion chamber, the computational domain in the prediction method is limited to single-element of injector, which includes single injector and plenum. The effects of combustion chamber including the detonation propagation are given by the time-evolving boundary conditions at injector outlet. The time-evolving boundary conditions are prepared by the full-scale simulation or theoretical calculation of one cycle in RDE, and mass flow rate is determined for single-element simulations. The single-element simulation consists of two phases. First, a non-reactive steady flow is calculated for new injector shape, which is used as an initial condition. Then, time-evolving profile is periodically applied to injector outlet as a boundary condition. Eventually, plenum pressure converges to appropriate value. In this paper, the single-element simulations are carried out for the rectangle injector and the chamfered injector for validation of the new method. The error of pressure loss prediction is less than 0.6%, which is 50 times faster, in comparison with full-scale simulation.

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Numerical analysis of the mean structure of gaseous detonation with dilute water spray

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

Journal of Fluid Mechanics （Journal of Fluid Mechanics） 887 1 - 40 2020年

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00221120

© 2020 The Author(s). Published by Cambridge University Press. Two-dimensional (2-D) numerical simulations based on the Eulerian-Lagrangian method that take droplet break-up into account are conducted to clarify the mean structure of gaseous detonation laden with a dilute water spray. The premixed mixture is a slightly diluted stoichiometric hydrogen-oxygen mixture at low pressure. The simulated results are analysed via 2-D flow fields and statistical Favre spatiotemporal averaging techniques. Gaseous detonation with water droplets (WD) propagates stably with a velocity decrease compared with the dry Chapman-Jouguet speed. The mean structure of gaseous detonation with dilute water spray shares a similar structure as the one without water spray. However, the hydrodynamic thickness is changed due to the interaction with water spray. Overall interphase exchanges (mass, momentum and energy) that take place within the hydrodynamic thickness induce a decrease of the detonation velocity and lower the level of fluctuations downstream of the mean leading shock wave. Droplet break-up occurs downstream of the induction zone and in our case, the water vapour from the evaporation of water spray does not affect the reactivity of gaseous detonation. The laminar master equation for gaseous detonation laden with inert WD shows that the hydrodynamic thickness should rely on the gaseous sound speed, and works well as the working mixture is weakly unstable and its cellular structure is regular. The droplet flow regimes and break-up modes have also been determined. The characteristic lengths of detonation and interphase exchanges have been ordered under the present simulation conditions and have been shown to be intimately intertwined.

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Pressure and visualization measurements on pulsed combustion thrustor

Asahara M., Kasahara J., Matsuoka K., Kawasaki. A., Matsuo A., Funaki I.

AIAA Scitech 2020 Forum （AIAA Scitech 2020 Forum） 1 PartF 2020年

ISSN 9781624105951

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Pulsed combustor, it has features such as taking out work without a compressor and mixing promotion by a complicated flow field. The purpose of my research is to evaluate the thrust of a pulse combustor that substitutes this check valve with an ejector mechanism without moving parts, and to clarify the pressure-gain of pulsed combustion and grasp the combustion state with pressure and visualization measurement. Although the total pressure of exhaust gas was lower than the oxygen plenum pressure in all shots, these were almost equal in case in which the pulse combustion was stably continued. It was also found that the expansion wave and the compression wave reciprocate in the combustion section, and that the behavior of the pressure wave and the flame were synchronized. A pulse combustion cycle due to this pressure wave reciprocation was estimated, and the sound velocity obtained from this was in agreement at the order level with theoretical value.

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Mode classification of combustion and propulsive performance of reflective shuttling detonation combustor

Yamaguchi M., Matsuoka K., Kawasaki A., Kasahara J., Watanabe H., Matsuo A.

AIAA Scitech 2020 Forum （AIAA Scitech 2020 Forum） 1 PartF 2020年

ISSN 9781624105951

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Detonation combustors, such as a rotating detonation combustor and a pulse detonation combustor, have advantages in its higher thermal efficiency and possibility to configure smaller scale combustor than that of conventional internal combustion systems. In a previous study, we proposed a new detonation combustor named a reflective shuttling detonation combustor (RSDC) in which detonation waves propagate in opposite directions repeating reflection at side walls. In the present study, a rectangular combustor (53 × 45 × 5 mm) with non-premixed triplet injectors was used to clarify the effect of equivalence ratios and mass flow rates on combustion modes and propulsive performance. As a result, both detonation and deflagration modes were observed. These modes were classified into four types (Single, Double, Single strong single weak, and Deflagration modes) based on CH* images captured by a high-speed camera and a band-pass filter whose peak value is 430 nm. For mass flow rates and equivalence ratios, it is suggested that a normalized fill height h/λ, which varies depending on these parameters, affects wave number transition as a rotating detonation combustor. For propulsive performance, static pressure measured at the bottom of the combustion chamber was normalized with the theoretical value of a conventional isobaric combustor. The normalized pressure for detonation modes were lower than that of deflagration modes. This might be attributed to higher dynamic pressure caused by the wave propagation and/or insufficient combustion.

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Combustion pressure distributions and thrust performances in small cylindrical rotating detonation engines

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

AIAA Scitech 2020 Forum （AIAA Scitech 2020 Forum） 1 PartF 2020年

ISSN 9781624105951

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The internal flow and wave structure of a nozzle-less, cylindrical rotating detonation engine (RDE) was investigated in this research. For a C2H4–O2 mixture, pressure distributions and chemiluminescence inside the chamber were obtained by combustion experiments. Pressure distributions suggest that combustion region is finished near 20 mm from bottom of the cylindrical RDE, and Mach number distributions obtained by Rayleigh flow theory also reveal flow reaches the sonic speed at the exit of the cylindrical RDE. Chemiluminescence images taken from the side of the cylindrical RDE show that strong luminance area ends at approximately 15-20 mm, which also means that combustion in the RDE finishes around that point. Moreover, a forward-tilting detonation wave which stably rotated at 1414 m/s was observed in the images, and it extended to the downstream of burned gas. From these results of strong luminescence at bottom and shock wave extending to the exit, the flow and wave structure inside cylindrical RDEs are proposed.

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Investigation of combustion modes and pressure of reflective shuttling detonation combustor

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

Proceedings of the Combustion Institute （Proceedings of the Combustion Institute） 38 （ 3 ） 3615 - 3622 2020年

研究論文（学術雑誌）, 共著, 査読有り, ISSN 15407489

© 2020 The Combustion Institute. Detonation combustors are considered promising alternatives to conventional combustors because they offer high thermal efficiency and fast combustion. However, especially for the rotating detonation combustor, the theoretical propulsive performance has not been confirmed in experimental studies because the highly unsteady flow field hinders the measurements process. To understand the involved phenomena in more detail, a reflective shuttling detonation combustor (RSDC) with a rectangular combustion chamber was developed. The interior of the chamber can easily be visualized owing to its two-dimensional quality. Utilizing the RSDC, several combustion tests with gaseous ethylene and oxygen were conducted for different values of mass flow rates and equivalence ratios. Combustion modes from the tests were classified into four types based on the fast Fourier transform (FFT) analysis of the luminous intensity of the CH* self-luminescence images captured by a high-speed camera and a band pass filter. Simultaneously, the theoretical total pressure of a conventional isobaric combustor was compared to the static pressure measured at the bottom of the RSDC chamber. For the detonation modes, the ratio between experimentally measured static pressure and the theoretical pressure varied depending on the location in the chamber owing to the distribution of the time-averaged static pressure. Furthermore, the pressure ratio of the detonation modes was up to 18% lower than that of the deflagration mode potentially owing to the flow velocity induced by the detonation waves.

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Experimental study of internal flow structures in cylindrical rotating detonation engines

Yokoo R., Goto K., Kasahara J., Athmanathan V., Braun J., Paniagua G., Meyer T., Kawasaki A., Matsuoka K., Matsuo A., Funaki I.

Proceedings of the Combustion Institute （Proceedings of the Combustion Institute） 38 （ 3 ） 3759 - 3768 2020年

研究論文（学術雑誌）, 共著, 査読有り, ISSN 15407489

© 2020 Published by Elsevier Inc. The internal flow structures of detonation wave were experimentally analyzed in an optically accessible hollow rotating detonation combustor with multiple chamber lengths. The cylindrical RDC has a glass chamber wall, 20 mm in diameter, which allowed us to capture the combustion self-luminescence. A chamber 70 mm in length was first tested using C2 H4 -O2 and H2 -O2 as propellants. Images with a strong selfluminescence region near the bottom were obtained, confirming the small extent of the region where most of the heat release occurs as found in our previous research. Based on the visualization experiments, we tested RDCs with shorter chamber walls of 40 and 20 mm. The detonation wave was also observed in the shorter chambers, and its velocity was not affected by the difference in chamber length. Thrust performance was also maintained compared to the longer chamber, and the short cylindrical RDC had the same specific impulse tendency as the cylindrical (hollow) or annular 70-mm chamber RDC. Finally, we calculated the pressure distributions of various chamber lengths, and found they were also consistent with the measured pressure at the bottom and exit. We concluded that the short-chamber cylindrical RDC with equal length and diameter maintained thrust performance similar to the longer annular RDC, further expanding the potential of compact RDCs.

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Numerical investigation of thermal response on external nozzle in supersonic aircraft by exhaust gas

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

AIAA Propulsion and Energy 2020 Forum （AIAA Propulsion and Energy 2020 Forum） 1 - 11 2020年

ISSN 9781624106026

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. This paper focuses on thermal response in rear part of engine on HIMICO (High Mach Integrated Control experiment). Firstly, pyrolysis reaction of CFRP, which is used in HIMICO as thermal protection system, is modeled from previous experimental data. Comparison between numerical data and experiment data shows good agreement. Secondly, one-dimensional heat transfer analysis with obtained pyrolysis reaction model on the external nozzle is conducted. By comparing with ramjet engine experiment, calculation result is confirmed to be appropriate. Thirdly, in addition to the thermal response analysis, CFD analysis in the rear part of the engine is conducted. Calculation result indicates that ramjet engine combustion test overestimates heat load on external nozzle compared to actual flight condition because of high back pressure.

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Propulsion performance of cylindrical rotating detonation engine

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

AIAA Journal （AIAA Journal） 58 （ 12 ） 5107 - 5116 2019年11月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00011452

© 2019 by Nagoya University. This study evaluated the propulsion performance of a nozzleless, cylindrical rotating detonation engine (RDE). Using a C2 H4 –O2 mixture, the RDE was tested in a low-back-pressure environment at propellant mass flow rates of 8–45 g/s. In high-speed imaging of the self-luminescence within the combustor, rotating luminous regions were observed at mass flow rates above 22 g/s. Measured pressure distributions suggest that burned gas reached sonic velocity at the combustion chamber outlet. This paper proposes the structure of internal flow in the RDE and confirms that calculated pressure distribution based on the structure was close to the experimental distribution. This study also estimated the RDE’s thrust by pressure and momentum exchange and confirmed it by experimental measurement. Moreover, the theoretical thrust calculated under the assumption that exhaust is a sonic flow agreed with the load cell thrusts, suggesting that RDE combustion is perfectly completed inside the chamber. Specific impulses are 80–90% of specific impulses for ideal correct expanded flow for all mass flow rates, and its value was close to that of an annular RDE. In addition, RDE performance will increase by about 20% if the RDE is equipped with a divergent nozzle and the gas is correctly expanded to back pressure.

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Prediction model of the flow properties inside a tube during hydrogen leakage

Nagase Y., Sugiyama Y., Kubota S., Saburi T., Matsuo A.

Journal of Loss Prevention in the Process Industries （Journal of Loss Prevention in the Process Industries） 62 2019年11月

査読有り, ISSN 09504230

© 2019 Elsevier Ltd We numerically investigated high-pressure hydrogen leakage from transportation facilities, focusing on the steady mass flow rate and pressure distribution in a tube during the leakage. We studied steady leakage from a square opening in a square duct as well as leakage from a ruptured cylindrical tube with unsteady closure of a cutoff valve from fully open. A prediction model for the mass flow rate and pressure distribution inside the tube was proposed; such a model would help prevent physical hazards during an accident. We considered changes in the physical quantities according to the fluid dynamics occurring inside the tube. The flow properties were divided into two phases: (i) the unsteady expansion wave propagating inside a tube filled with hydrogen and (ii) the acceleration of hydrogen due to the reduction in the cross-sectional area between the tube and the leakage opening. To close the prediction model, we introduced contraction coefficient models depending on how the hydrogen leakage occurred. The mass flow rate and pressure drop during the leakage estimated by our prediction model showed good agreement with numerical simulation results when the contraction coefficient model was appropriately chosen. This model is considered highly applicable to the construction condition of pressure sensors, the operating conditions of a valve, and the prediction of mass flow rate during an accident.

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Research and development of rotating detonation engine system for the sounding rocket flight experiment S520-31

Kasahara J., Kawasaki A., Matsuoka K., Matsuo A., Funaki I., Nakata D., Uchiumi M.

AIP Conference Proceedings （AIP Conference Proceedings） 2121 2019年07月

ISSN 9780735418615

© 2019 Author(s). A detonation is a combustion wave that propagates at supersonic speed (2∼3 km/s) in a combustible mixture. There are many fundamental studies of detonation waves and detonation engine systems. The detonation cycle has a higher thermal efficiency than a conventional constant-pressure combustion cycle. Therefore, it is expected that a high-efficiency propulsion system can be realized using detonation waves.A rotating detonation engine (RDE) uses continuous detonation propagating at a bottom in an annular combustor. As detonation waves propagate at a supersonic speed only in the bottom region of the RDEs, the combustor can be shortened. However, the combustor needs cooling system due to high heat flux to the combustor wall. In this experimental study, we performed combustion tests of RDE system using gaseous ethylene and oxygen as the propellant. This RDE system performance will also be demonstrated in space environment by the sounding rocket. We measured the combustor pressure, temperatures, heat flus, mass flow rate and thrust. The RDE system used in this study is shown in Figure 1. We performed the long-duration rotating detonation engine combustion tests for at sea level condition. The stable trust histories were obtained.

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Semi-valveless pulse detonation cycle at a kilohertz-scale operating frequency

Matsuoka K., Taki H., Kawasaki A., Kasahara J., Watanabe H., Matsuo A., Endo T.

Combustion and Flame （Combustion and Flame） 205 434 - 440 2019年07月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00102180

© 2019 The Combustion Institute A high operating frequency of a pulse detonation engine is required to increase the thrust-to-engine weight ratio or thrust density. The semi-valveless pulse detonation cycle (PDC) proposed by Matsuoka et al. (2017) can achieve a high operating frequency exceeding several kilohertz. For achieving a higher operating frequency close to the upper limit of gas dynamics, it is necessary to minimize the process in which the buffer layer is applied to avoid self-ignition of the detonable mixture. Experiments were conducted, and a one-dimensional numerical model was developed to investigate the minimum thickness of the buffer layer and the required duration for the stable PDC operation. Ethylene was used as a fuel and pure oxygen as an oxidizer. The total length of two combustors with an inner diameter of 10 mm was 40 and 80 mm. Therefore, the thickness of the buffer layer of approximately 20 mm was suggested for the stable PDC operation. This result indicated that 10% of the duration of one PDC was required to prevent self-ignition (SI). In the failed PDC, the early and late SI were confirmed. Interestingly, high-frequency PDC operation with a short combustor can suppress late SI and results in a higher success rate with the same thickness of the buffer layer. Furthermore, a stable PDC operation of a 1916 Hz with a combustor with a total length of 40 mm was demonstrated.

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Numerical investigation of the interaction between a shock wave and a particle cloud curtain using a CFD–DEM model

Sugiyama Y., Ando H., Shimura K., Matsuo A.

Shock Waves （Shock Waves） 29 （ 4 ） 499 - 510 2019年05月

ISSN 09381287

© 2018, Springer-Verlag GmbH Germany, part of Springer Nature. A two-dimensional numerical simulation of the interaction between a shock wave and a particle cloud curtain (PCC) in a shock tube was conducted to develop the numerical method and to understand how the particle layer mitigates the shock wave. In the present study, computational fluid dynamics/the discrete element method in conjunction with drag force and convective heat transfer models were used to separately solve the continuum fluid and particle dynamics. The applicability of the method to the gas flow and particles was validated through comparison with gas–particle shock-tube experiments, in which the PCC was generated by free fall, and particles initially had a gradient of its volume fraction and falling velocity in height. When the incident shock wave interacted with the PCC, it was reflected from and transmitted through the PCC. The transmitted shock wave had a curved front because the initial gradient in the volume fraction of particles locally changed the interaction between the shock wave and the particles. We calculated the effects of the drag force and heat transfer in mitigating the strength of the transmitted shock wave. The propagation of the transmitted and reflected shock waves and the motion of the PCC induced by the gas flow behind the shock wave agreed well with previous experimental data. After the interaction between the gas flow and the PCC, drag force and heat transfer were activated by the gradients in pressure, velocity, and temperature between them, and the gas flow lost momentum and energy, which weakened the transmitted shock wave. At the same time, the PCC gained momentum and energy and was dispersed. The contact forces between two particles affected the local dispersion of the PCC.

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Propulsive performance and heating environment of rotating detonation engine with various nozzles

Goto K., Nishimura J., Kawasaki A., Matsuoka K., Kasahara J., Matsuo A., Funaki I., Nakata D., Uchiumi M., Higashino K.

Journal of Propulsion and Power （Journal of Propulsion and Power） 35 （ 1 ） 213 - 223 2019年01月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 07484658

Geometric throats are commonly applied to rocket combustors to increase pressure and specific impulse. This paper presents the results from thrust measurements of an ethylene/gas-oxygen rotating detonation engine with various throat geometries in a vacuum chamber to simulate varied backpressure conditions in a range of 1.1–104 kPa. For the throatless case, the detonation channel area was regarded to be equivalent the throat area, and three throat-contraction ratios were tested: 1, 2.5, and 8. Results revealed that combustor pressure was approximately proportional to equivalent throat mass flux for all test cases. Specific impulse was measured for a wide range of pressure ratios, defined as the ratio of the combustor pressure to the backpressure in the vacuum chamber. The rotating detonation engine could achieve almost the same level of optimum specific impulse for each backpressure, whether or not flow was squeezed by a geometric throat. In addition, heat-flux measurements using heat-resistant material are summarized. Temporally and spatially averaged heat flux in the engine were roughly proportional to channel mass flux. Heat-resistant material wall compatibility with two injector shapes of doublet and triplet injection is also discussed.

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Supersonic combustion induced by reflective shuttling shock wave in fan-shaped two-dimensional combustor

Yamaguchi M., Matsuoka K., Kawasaki A., Kasahara J., Watanabe H., Matsuo A.

Proceedings of the Combustion Institute （Proceedings of the Combustion Institute） 37 （ 3 ） 3741 - 3747 2019年01月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 15407489

© 2018 The Combustion Institute. As a novel detonation combustor that differs from a pulse and a rotating detonation engine, a reflective shuttling detonation combustor (RSDC), in which detonation waves shuttle repeatedly, was proposed. In a fan-shaped two-dimensional combustor, detonation waves propagate, repeating attenuation and re-ignition by a shock reflection at the side wall. In the demonstration experiment, chemiluminescence visualization and pressure measurement with ethylene-oxygen mixture were conducted at the same time. As the result, a single shuttling wave coupled with pressure rise was observed in the combustor. The tangential velocity of the wave was 1526 ± 12 m/s and approximately 60% of the estimated Chapman-Jouguet velocity of 2513 m/s. The ratio of pressure in front of the wave to one behind the primary wave or the reflected wave was in good agreement with one-dimensional shock theory, and it was suggested that the rapid reaction behind the reflected shock wave sustained the continuous propagation of the shock wave.

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Critical condition of inner cylinder radius for sustaining rotating detonation waves in rotating detonation engine thruster

Kawasaki A., Inakawa T., Kasahara J., Goto K., Matsuoka K., Matsuo A., Funaki I.

Proceedings of the Combustion Institute （Proceedings of the Combustion Institute） 37 （ 3 ） 3461 - 3469 2019年01月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 15407489

© 2018 The Combustion Institute. Published by Elsevier Inc. All rights reserved. We describe the critical condition necessary for the inner cylinder radius of a rotating detonation engine (RDE) used for in-space rocket propulsion to sustain adequate thruster performance. Using gaseous C2H4 and O2 as the propellant, we measured thrust and impulse of the RDE experimentally, varying in the inner cylinder radius ri from 31 mm (typical annular configuration) to 0 (no-inner-cylinder configuration), while keeping the outer cylinder radius (ro = 39 mm) and propellant injector position (rinj = 35 mm) constant. In the experiments, we also performed high-speed imaging of self-luminescence in the combustion chamber and engine plume. In the case of relatively large inner cylinder radii (ri = 23 and 31 mm), rotating detonation waves in the combustion chamber attached to the inner cylinder surface, whereas for relatively small inner cylinder radii (ri = 0, 9, and 15 mm), rotating detonation waves were observed to detach from the inner cylinder surface. In these small inner radii cases, strong chemical luminescence was observed in the plume, probably due to the existence of soot. On the other hand, for cases where ri = 15, 23, and 31 mm, the specific impulses were greater than 80% of the ideal value at correct expansion. Meanwhile, for cases ri = 0 and 9 mm, the specific impulses were below 80% of the ideal expansion value. This was considered to be due to the imperfect detonation combustion (deflagration combustion) observed in small inner cylinder radius cases. Our results suggest that in our experimental conditions, ri = 15 mm was close to the critical condition for sustaining rotating detonation in a suitable state for efficient thrust generation. This condition in the inner cylinder radius corresponds to a condition in the reduced unburned layer height of 4.5-6.5.

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Using an extended CFD-DEM for the two-dimensional simulation of shock-induced layered coal-dust combustion in a narrow channel

Shimura K., Matsuo A.

Proceedings of the Combustion Institute （Proceedings of the Combustion Institute） 37 （ 3 ） 3677 - 3684 2019年01月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 15407489

© 2018 Elsevier Ltd. We used the computational fluid dynamics-discrete element method (CFD-DEM) model of compressible flow to numerically investigate the flame structure during shock-wave-induced layered coal-dust combustion, which poses a significant risk in coal mines. This represents the first attempt to apply a CFD-DEM model to compressible reactive gas-particle flow. The Eulerian-Lagrangian model for compressible gas-particle flow was extended to simulate shock-particle interactions in a reactive flow field. The particle interactions were predicted by the DEM, and source terms for homogeneous and heterogeneous reactions were included in the governing equations. The calculations were validated for inert particle dispersion from shock-particle interactions and flame propagation velocity in layered coal-dust combustion. The results were consistent with those from previous experiments. Furthermore, the flame structure in layered coal-dust combustion was revealed using the CFD-DEM approach. The simulation of the layered coal-dust combustion indicated that the shock wave was initially generated by gas detonation in a narrow channel with coal-dust particles at the bottom. The predicted propagation mechanism during layered coal-dust combustion was consistent with that reported in a previous numerical study based on the Eulerian-Eulerian approach (Hoium et al., 2015). Moreover, the flame comprised a leading shock wave and diffusion flame; the coal-dust particles were dispersed and heated by the shock wave and combustion products, respectively. The diffusion flame structure propagated at 350-500 m/s, resulting in devolatilization behind the reaction front. However, the CFD-DEM results indicated that the particle dispersion heights were higher than those predicted by the Eulerian-Eulerian approach (despite similarities in the inert particle dispersion results of these methods), attenuating the compression wave from the reaction front and slowing the leading shock wave.

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Numerical investigation on propagation behavior of gaseous detonation in water spray

Watanabe H., Matsuo A., Matsuoka K., Kawasaki A., Kasahara J.

Proceedings of the Combustion Institute （Proceedings of the Combustion Institute） 37 （ 3 ） 3617 - 3626 2019年01月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 15407489

© 2018 Elsevier Ltd. A two-dimensional (2D) numerical simulation is conducted to clarify the propagation behavior of gaseous detonation in a water spray and its structure. The computational target refers to the experiment conducted by G. Jarsalé et al., and C2H4-air gaseous detonation propagates where the water droplets (WDs) are sprayed. The parameters used are the C2H4-air equivalence ratio and WD mass fraction. The flow field, Favre-averaged one-dimensional profile, and cellular structure are revealed in 2D simulations. Stable propagation of gaseous detonation is observed in the water spray, and the decrease in velocity relative to the Chapman-Jouguet velocity without WDs is as much as 3.2%. Adding WDs changes the cellular pattern, especially for leaner mixtures. The weak triple point decays, and the cell width increases because of the longer induction length due to decreased velocity. The WD presence changes the detonation flow field substantially, and evaporation occurs primarily at 10 mm behind the shock wave. The high-evaporation region propagates at the detonation speed, and the compression wave formed when the detonation reflects from the two-phase medium propagates backward. Furthermore, WD evaporation suppresses the velocity, vorticity, and temperature fluctuations. Rapid evaporation with WDs leads to lower hydrodynamic thickness than that without WDs or in the Zel'dovich-von Neumann-Döring model.

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Numerical investigation of the interaction between a shock wave and a particle cloud curtain using a CFD-DEM model (accepted)

Sugiyama, Y., Ando, H., Shimura, K., and Matsuo, A.

Shock Waves 29 （ 4 ） 499 - 510 2019年

研究論文（学術雑誌）, 共著, 査読有り

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The effect of acceleration and exit velocity on hypersonic projectiles launched by a ground-based railgun

Kasahara H., Matsuo A.

AIAA Scitech 2019 Forum （AIAA Scitech 2019 Forum） 2019年

ISSN 9781624105784

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In this paper, the effects of projectile acceleration profile on interior and transition ballistics are investigated in order to design HVP (Hyper-Velocity projectile) for railgun. Our railgun facilities accelerate projectile to 600 – 2501 m/s in order to evaluate the effect of exit velocity on sabot separation. However, sabot melt prevents us from observing detail sabot separation and shock wave interactions. Thus, in this paper, CFD-RBD method is applied to calculate the projectile acceleration and sabot separation motion due to unsteady aerodynamic force. In the interior calculation, acceleration behavior is approximated by the cubic spline interpolation using our experimental result in Vexit = 658, 1268, 1875, 2048, 2501 m/s. Firstly, in the interior ballistic calculation, the projectile and sabot induce compression waves due to acceleration. In 0.6 ms, the single normal shock wave stands ahead of projectile. The precursor shock wave reaches the end of tube and propagates outside of tube. When the precursor shockwave propagates outside of tube, Mach disk stands just before the precursor shock wave due to expansion. Especially, the cases of Vexit ≥ 1875 m/s show similar pressure distribution and precursor shock wave propagation outside of tube. On the other hand, supersonic exit velocity, in Vexit = 1268 m/s, increases distance between precursor shock wave and Mach disk. Also, in Vexit = 658 m/s, although Mach disk is close to the muzzle, precursor shock wave propagates far away from Mach disk, which generates completely different pressure distribution. Secondary, in transition ballistic calculation, sabot separation motion does not depend on projectile exit velocity in our railgun configuration. From this result, the design of HVP need not to consider sabot trajectory variation even if exit velocity changes from supersonic (Vexit = 658 m/s) to hypersonic (Vexit = 2501 m/s) regime.

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Propulsion performance of inner-cylinder-less rotating detonation engine

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

AIAA Scitech 2019 Forum （AIAA Scitech 2019 Forum） 2019年

ISSN 9781624105784

© 2019 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. We evaluated the propulsion performance of a nozzle-less, inner-cylinder-less rotating detonation engine (RDE). For a C2H4–O2 mixture, the RDE was tested in a low-back-pressure environment at several propellant mass flow rates ranging from 8 to 45 g/s. In high-speed imaging of the self-luminescence within the combustor, rotating luminous regions were observed at a mass flow rates greater than 22 g/s. The specific impulse efficiency was greater than 80% for all the mass flow rate, and approximately 90% in some cases, which is comparable with ones in conventional RDEs having inner cylinders. By the control surface analysis, it was clarified that propellants injection from the injector holes and pressure on the bottom of the combustion chamber account for the thrust. It was also suggested that the design of the outer cylinder of the combustion chamber and the injector arrangement may affect the thrust performance. Axial Mach number distributions within the engine were calculated under an assumption of isentropic expansion. As a result, the burned gas reached a sonic or supersonic velocity at the outlet of the combustion chamber. Axial pressure distributions also suggested that the acceleration of the burned gas was completed at a far-upstream region within the combustion chamber.

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Experimental performance validation of a rotating detonation engine toward a flight demonstration

Goto K., Yokoo R., Kim J., Kawasaki A., Matsuoka K., Kasahara J., Matsuo A., Funaki I., Nakata D., Uchiumi M.

AIAA Scitech 2019 Forum （AIAA Scitech 2019 Forum） 2019年

ISSN 9781624105784

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Thrust measurements of rotating detonation engine of (1) ethylene / gas-oxygen and (2) methane / gas-oxygen with various throat geometries in a vacuum chamber to simulate different back-pressure conditions ranging from 1.1-104 kPa were conducted. For throatless rotating detonation engine, we defined equivalent throat area as the detonation channel area, and then tested four nozzle contraction ratios of 1, 1.5, 2.5, and 8. Engines could be successfully ignited by electric ignitors when initial pressure was high enough to have, at least, one detonation cell in RDE channel. We measured the combustor pressure and reveled that it was almost proportional to the throat mass flux regardless of contraction ratios and the propellant combinations. The specific impulse of methane / gas-oxygen case could achieve 84 ± 1% of ideal specific impulse at the optimum expansion for each back pressure.

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Numerical investigation on characteristic lengths for gaseous detonation with dilute water spray

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

AIAA Propulsion and Energy Forum and Exposition, 2019 （AIAA Propulsion and Energy Forum and Exposition, 2019） 2019年

ISSN 9781624105906

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Two-dimensional (2D) numerical simulations based on Eulerian-Lagrangian method are conducted to clarify the droplet behavior within its lifetime within the detonation cell. The simulation results are analyzed via 2D instantaneous flow fields and Favre spatiotemporal average technique, by applying the recycling block method. Gaseous detonation with dilute water droplets (WDs) propagates stably with a 4% velocity decrease compared to dry CJ velocity in the simulation conditions. From the instantaneous flow field analysis, the droplet breakup occurs primarily in jets, downstream of the transverse wave, nearby the collision of transverse waves, and the interaction between the transverse wave and the jets. The Favre average one-dimensional profiles by grouping WDs based on the initial shock strength that WDs experience reveal the droplet life inside the cellular structure. The mean equilibrium diameter after the breakup is not affected by the initial shock strength.

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An experimental study of in-space rotating detonation engine with cylindrical configuration

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

AIAA Propulsion and Energy Forum and Exposition, 2019 （AIAA Propulsion and Energy Forum and Exposition, 2019） 2019年

ISSN 9781624105906

© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Utilizing detonation in a combustor potentially leads to a compact and propellant-efficient rocket engine. A rotating detonation engine (RDE) is a form of such detonation engine. For a practical realization of the RDE, maturation of its thermal design is essential. In this study, a miniature cylindrical rotating detonation engine was tested in a low-pressure environment (~ 6 kPa) with a film-cooled wall. For the combustor, measurement of lateral wall temperature and high-speed imaging were conducted during combustion. As a result, it has been clarified that there exist an appropriate range in the coolant flow rate to maintain detonation combustion. Additionally, it has been also clarified that film cooling is effective even under detonation combustion.

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Optical Measurement of Fluid Motion in Semi-Valveless Pulse Detonation Combustor with High-Frequency Operation

Kubota, A., Matsuoka, K., Kawasaki, A., Kasahara, J., Watanabe, H., Matsuo, A., and Endo, T.

Combustion Science and Technology 192 （ 2 ） 197 - 212 2018年12月

研究論文（学術雑誌）, 共著, 査読有り

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Experimental investigation on detonation dynamics through a reactivity sink

Boulal S., Vidal P., Zitoun R., Matsumoto T., Matsuo A.

Combustion and Flame （Combustion and Flame） 196 11 - 25 2018年10月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 00102180

© 2018 The Combustion Institute This article reports on an experimental investigation into dynamical behaviours of detonation in non-uniform mixtures, generated from stoichiometric propane–oxygen, oxygen and ethane, with initial temperature and pressure 290 K and 20 kPa, respectively. Composition gradients are parallel to the direction of detonation propagation, with an equivalence ratio (ER) that first decreases from lean values and then increases to rich ones. Composition distributions are characterized according to the depth of the ER sink. Gradients are generated in a 50 × 50-mm2-square cross-section and a 665-mm total length chamber. The mixture components are injected separately in the pre-evacuated chamber in their order of decreasing density through porous plates at the chamber top-end to ensure planar filling of the chamber. Non-uniform distributions are then precisely controlled as a function of time by means of optical oxygen sensors. A Chapman–Jouguet (CJ) detonation is transmitted at the chamber bottom-end from a 3.6-m-long driver tube. Fast pressure transducers, sooted plates and Schlieren visualizations coupled with high-speed cameras are used to characterize the longitudinal velocity, cellular structure and transmission, failure and re-initiation mechanisms of the detonation front. Shallowest ER sinks produce the supercritical transmission mode of the CJ detonation with continuous adaptation of velocity and multicellular structure to local composition. Deepest sinks lead to the subcritical behaviour characterized by sudden detonation failure from shock-flame decoupling when ER decreases, and without detonation re-initiation when ER increases again. Intermediate sink depths generate critical behaviour with detonation re-initiat ion at chamber walls from expanding combustion kernels and reflected transverse Mach waves and then from SWACER retro-active mechanism. An elaboration of the failure criterion used in a previous study is found to well predict conditions for shock-flame decoupling.

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Two-dimensional explosion experiments examining the interaction between a blast wave and a sand hill

Sugiyama, Y., Izumo, M., Ando, H., and Matsuo, A.

Shock Waves （Shock Waves） 28 （ 3 ） 627 - 630 2018年05月

研究論文（学術雑誌）, ISSN 09381287

© 2018, Springer-Verlag GmbH Germany, part of Springer Nature. Two-dimensional explosion experiments were conducted to discuss the interaction between a blast wave and sand and show the mitigation effect of the sand on the blast wave. The explosive used was a detonating cord 1.0 m in length, which was initiated in a sand hill shaped like a triangular prism and whose cross section was an isosceles triangle with base angles of 30∘. Sand-hill heights of 30 and 60 mm were used as parameters to discuss the effect of sand mass upon blast-wave strength. The interaction of the blast wave with the sand/air interface causes multiple peaks in the blast wave, which are induced by successive transmissions at the interface. The increase in the sand mass further mitigates the blast parameters of peak overpressure and positive impulse. The results of this experiment can be utilized to validate the numerical method of solving the problem of interaction between a compressible fluid and a particle layer.

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CFD-DEMを用いた衝撃波を伴う固気二相流の数値解析

木村謙仁, 志村啓, 松尾亜紀子

混相流 32 （ 1 ） 65 - 72 2018年03月

研究論文（学術雑誌）, 共著, 査読有り

不足膨張条件下における外部ノズル推力予測の簡略化手法

御子柴稜, 松尾亜紀子, 富岡定毅, 藤川貴弘, 土屋武司, 磯野達志

日本航空宇宙学会論文集 66 （ 1 ） 15 - 23 2018年02月

研究論文（学術雑誌）, 共著, 査読有り

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Two-dimensional CFD–DEM simulation of vertical shock wave-induced dust lifting processes

Shimura, K., and Matsuo, A.

Shock Waves （Shock Waves） 28 （ 6 ） 1285 - 1297 2018年

研究論文（学術雑誌）, ISSN 09381287

© 2018, Springer-Verlag GmbH Germany, part of Springer Nature. This paper conducts a numerical study of particle dispersion by vertical shock waves using a combination of computational fluid dynamics and the discrete element method (DEM). The Magnus force is important for particle dispersion, and the DEM approach can consider the rotation of individual particles and provide a detailed analysis of particle–particle and particle–wall interactions. Simulations are conducted and the results are compared with those of previous experiments, showing that the models can capture the particle dispersion process and shock wave geometry inside the mixed gas–particle region. During the particle dispersion process, the contact force due to the particle–particle interactions causes the particles to move upward after the shock wave passes and then decreases rapidly due to the low collision frequency in the particle cloud. The Magnus force is initially lower and has little effect until the contact force decreases, when it becomes dominant and maintains the particles’ upward movement. It is mainly driven by the gas rotation above the dust layer, and the effect of particle rotation is relatively small in comparison. The gas velocity gradient above the dust layer in the particle dispersion region is caused by gas–particle interactions. The particle dispersion region becomes wider and thinner over time, meaning that the gas velocity gradient above the dust layer becomes small and hence that the Magnus force (which is driven by the gas rotation) also becomes small. In contrast, the downward drag force remains constant during the vertical motion because it is primarily affected by the velocity difference between the particles and the gas. This means that the drag and Magnus forces eventually balance, and the particles stop rising.

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The effect of shape on the aerodynamic and thermal performance of hypersonic projectiles launched by a ground-based railgun

Kasahara H., Matsuo A.

AIAA Aerospace Sciences Meeting, 2018 （AIAA Aerospace Sciences Meeting, 2018）（ 210059 ） 2018年

ISSN 9781624105241

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In this paper, the effect of hypersonic-projectile design upon its aerodynamic coefficient is investigated to optimize projectile shape for ground-based-railgun launch. Previous studies on railgun-launch systems have suggested that projectiles experience significant thermal and aerodynamic effects when exiting the launcher because of their hypersonic exit velocity and the high enthalpy flow in the standard atmospheric condition. When designing a hypersonic projectile, the aerodynamic coefficient and thermal protection are crucial for maximizing range and withstanding heat. Two conditions of hypersonic flight, such as free flight and transitional ballistics, are simulated herein. To investigate the projectile’s geometric effect in the steady-flight condition, projectiles designed based on theory and empirical observation are compared. The aerodynamic result suggests that sharper projectiles do not always have smaller drag coefficients in simulation and theory because of a tradeoff between maximum pressure at the tip and pressure distribution on the latter half of the front shape. While temperature shows a similar trend compared to pressure, the maximum-temperature point is affected by projectile bluntness. Therefore, to optimize hypersonic design for thermal effect, it is inappropriate to adopt excess thermal protection at the tip, and an accurate prediction of maximum-heat-flux point is important. In addition to free-flight simulation, a transitional-ballistics simulation is performed. The interior and transitional calculations indicate that the precursor shockwave and sabot affect the flow field around the projectile. The sabot-separation phase can be divided into four stages based on the characteristics of the flow field. In the first and second stages, before the projectile undergoes the precursor shockwave, its shape has little effect on drag or velocity. However, when the projectile interacts with the sabot in the third and fourth stages, geometric effects can be observed. The theoretically optimized projectile is found to keep the highest flight velocity from transitional ballistics to free-flight.

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The effect of exit velocity and material on the sabot separation launched by ground-based railgun

Kasahara H., Matsuo A.

2018 Applied Aerodynamics Conference （2018 Applied Aerodynamics Conference） 2018年

ISSN 9781624105593

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In this paper, the sabot separation process in super/hypersonic flight from 600 to 2501 m/s is numerically investigated by using CFD and RBD (Rigid Body Dynamics). Firstly, the motion of sabot separation is calculated for validating our in-house code. CFD-RBD code shows good agreements with Mehmet’s experiment. Secondary, the effects of exit velocity from muzzle are examined. While all sabot velocities differ, sabot separations indicate the same trajectory. The variation of sabot velocity can be eliminated by non-dimensionalization considering exit velocity. Finally, the effects of sabot material, which means mass density, are investigated with railgun acceleration condition. Sabot separation shows different motion completely depending on sabot material. However, non-dimensionalization eliminates the effect of sabot material on separation. Proposed non-dimensionalization method can represent the various sabot separation motions in unique one trajectory, considering exit velocity and sabot density in time term.

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Preliminary experiments on rotating detonation rocket engine for flight demonstration using sounding rocket

Goto K., Nishimura J., Higashi J., Taki H., Ukai T., Hayamizu Y., Kikuchi K., Yamada T., Watanabe S., Hotta K., Inakawa T., Kubota A., Yamaguchi M., Daicho T., Kawasaki A., Matsuoka K., Kasahara J., Matsuo A., Funaki I., Yasuda K., Mori K., Yagihashi H., Nakata D., Uchiumi M., Higashino K.

AIAA Aerospace Sciences Meeting, 2018 （AIAA Aerospace Sciences Meeting, 2018）（ 210059 ） 2018年

ISSN 9781624105241

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Rotating detonation engine (RDE) is a promising candidate for future upper stage motor because of its high theoretical efficiency and short combustor length. Technological demonstration in space will be necessary to evaluate the performance and flight dynamics of RDE as a space propulsion system. Toward a flight test, we built an open-structure, rocket RDE-powered vehicle to measure thrust to validate a feasibility of an onboard propulsion system. We carried out 4.4 s combustion test, and achieved over 73% of ideal specific impulse above 1.1 MPa of static pressure in the RDE combustor.

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Numerical investigation of the gas-particle flow in the shock tube using discrete particle and continuum model

Kimura A., Matsuo A.

2018 Fluid Dynamics Conference （2018 Fluid Dynamics Conference） 2018年

ISSN 9781624105531

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The numerical investigation of the gas-particle flow in the shock tube is conducted using the continuum model (Two-Fluid model) and the discrete particle model (CFD-DEM). From calculated results of both methods, particles are concentrated at downstream edge of particle bed and dispersed at upstream edge of particle bed with the passage of time. The incident shock wave is divided into the reflected shock and the transmitted shock because of interactions with particles. These flow field can be observed in the experiment conducted by Rogue et al., and Two-Fluid model and CFD-DEM can qualitatively reproduce the interaction phenomena between shock wave and particles. In the experiment, the pressure oscillations behind the reflected and the transmitted shock wave are observed as well as CFD-DEM, but no oscillation is seen in Two-Fluid model. Therefore, Two-Fluid model can qualitatively reproduce the interaction phenomena of shock wave and particles qualitatively, although the behavior of individual particle is neglected in the model.

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水中移動体の先端形状が飛しょう安定性に与える影響に関する数値解析

奥野航平, 松尾亜紀子

日本機械学会論文集 83 （ 856 ） 2017年12月

研究論文（学術雑誌）, 共著, 査読有り

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Validation of pulse-detonation operation in low-ambient-pressure environment

Matsuoka K., Takagi S., Kasahara J., Matsuo A., Funaki I.

Journal of Propulsion and Power （Journal of Propulsion and Power） 34 （ 1 ） 116 - 124 2017年06月

研究論文（学術雑誌）, 共著, 査読有り, ISSN 07484658

© Copyright 2017 by Nagoya University. A pulse-detonation thruster can generate a high-repeatability small impulse at a high operating frequency. To operate a pulse-detonation cycle in a vacuum environment without a purging material, a liquid-purge method proposed by Matsuoka et al. ("Development of a Liquid-Purge Method for High-Frequency Operation of Pulse Detonation Combustor," Combustion Science and Technology, Vol. 187, No. 5, 2015, pp. 747-764) and throat at the exit of the combustor were introduced. In the demonstration experiment, gaseous ethylene-liquid nitrous oxide, as detonable mixture, and a throat having an inner diameter of 3.6 mm (blockage ratio = 87%) were used. The measured cyclic flame-propagation speeds were 2005 ± 90 m/s and 92 ± 4% of the estimated Chapman-Jouguet detonation speed. Consequently, a 50 Hz pulse-detonation operation without a purging material in the ambientpressure range of 0.035-1.5 kPa was confirmed. A quasi-steady model, in which gases in the combustor are in the stationary state, was newly developed to investigate the operating characteristic of the pulse-detonation thruster. The experimental pressure history in the combustor during the burned-gas blowdown process was in good agreement with that of the model. Moreover, using the model, the thrust performance of the pulse-detonation thruster with a converging-diverging nozzle was investigated. It was found that the estimated specific impulse was comparable with that of the theoretical steady-state rocket engine.

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Supersonic Combustion Using Multiple Stinger-shaped Injectors

Kouchi, T., Tomioka, S., Hirano, K., Matsuo, A., and Masuya, G.

Transactions of JSASS 60 （ 1 ） 56 - 59 2017年01月

研究論文（学術雑誌）, 共著, 査読有り

Experimental Visualization of the Structure of Rotating Detonation Waves in a Disk-Shaped Combustor

Nakagami, S., Matsuoka, K., Kasahara, J., Kumazawa, Y., Fujii, J., Matsuo, A., and Funaki, I.

Journal of Propulsion and Power, Special Section on Pressure Gain Combustion 33 （ 1 ） 80 - 88 2017年01月

研究論文（学術雑誌）, 共著, 査読有り

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Development of High-Frequency Pulse Detonation Combustor Without Purging Material

Matsuoka, K., Muto, K., Kasahara, J., Watanabe, H., Matsuo, A., and Endo, T.

Journal of Propulsion and Power, Special Section on Pressure Gain Combustion 33 （ 1 ） 43 - 50 2017年01月

研究論文（学術雑誌）, 共著, 査読有り

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Investigation of Fluid Motion in Valveless Pulse Detonation Combustor with High-Frequency Operation

Matsuoka, K., Muto, K., Kasahara, J., Watanabe, H., Matsuo, A. and Endo, T.

Proceedings of the Combustion Institute 36 （ 2 ） 2641 - 2647 2017年

研究論文（学術雑誌）, 共著, 査読有り

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Numerical Investigation on Detonation Velocity in Rotating Detonation Engine Chamber

Fujii, J., Kumazawa, Y., Matsuo, A., Nakagami, S., Matsuoka, K., and Kasahara, K.

Proceedings of the Combustion Institute 36 （ 2 ） 2665 - 2672 2017年

研究論文（学術雑誌）, 共著, 査読有り

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Experimental Study of the Structure of Forward-Tilting Rotating Detonation Waves and Highly Maintained Combustion Chamber Pressure in a Disk-Shaped Combustor

Nakagami, S., Matsuoka, K., Kasahara, J., Matsuo, A., and Funaki, I.

Proceedings of the Combustion Institute 36 （ 2 ） 2673 - 2680 2017年

研究論文（学術雑誌）, 共著, 査読有り

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直径5㎜の飛翔体加速装置に対する砲内弾道シミュレーション

浅野啓, 松尾亜紀子

防衛技術ジャーナル, 2016年10月号（427号）（防衛技術協会）（ 427 ） 46 - 50 2016年10月

研究論文（学術雑誌）, 共著

Multi-frame visualization for detonation wave diffraction

Nagura, Y., Kasahara, J. , and Matsuo, A.

Shock Waves 26 （ 5 ） 645 - 656 2016年09月

研究論文（学術雑誌）, 共著, 査読有り

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初期間隙が飛翔体加速装置の発射薬燃焼に与える影響に関する数値解析

浅野啓, 松尾亜紀子

EXPLOSION 26 （ 2 ） 73 - 81 2016年08月

研究論文（学術雑誌）, 共著, 査読有り

Numerical simulation of blast wave mitigation achieved by water inside a subsurface magazine model

Sugiyama, Y., Matsuo, A. Nakayama, Y., Matsumura, T., Wakabayashi, K., and Homae, T.

Journal of Loss Prevention in the Process Industries 43 521 - 528 2016年07月

研究論文（学術雑誌）, 共著, 査読有り

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Flight Validation of a Rotary-Valved Four-Cylinder Pulse Detonation Rocket

Matsuoka, K., Morozumi, T., Takagi, S., Kasahara,J., Matsuo, A., and Funaki, I

Journal of Propulsion and Power 32 （ 2 ） 383 - 391 2016年

研究論文（学術雑誌）, 共著, 査読有り

Numerical analysis on acoustic coupling of spinning detonation in a square tube

Sugiyama, Y., and Matsuo, A.

Journal of Thermal Science and Technology 11 （ 1 ） 1 - 8 2016年

研究論文（学術雑誌）, 共著, 査読有り

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Thrust Performance of Rotary-Valved Four-Cylinder Pulse Detonation Rocket Engine

Matsuoka, K., Sakamoto, R., Morozumi, T., Kasahara, J., Matsuo, A., and Funaki, I.

Transactions of the Japan Society for Aeronautical and Space Sciences 58 （ 4 ） 193 - 203 2015年07月

研究論文（学術雑誌）, 共著, 査読有り

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Scale Effect of Spherical Projectiles for Stabilization of Oblique Detonation Waves

Maeda, S., Sumiya, S, Kasahara, J., and Matsuo, A.

Shock Waves 25 （ 2 ） 141 - 150 2015年03月

研究論文（学術雑誌）, 共著, 査読有り

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Numerical simulation on blast wave around the small-scale model of subsurface magazine

Sugiyama, Y., Tanaka, T., Matsuo, A., Homae, T., Wakabayashi, K., Matsumura, T., and Nakayama, Y.

Science and Technology of Energetic Materials 76 （ 5 ） 115 - 119 2015年

研究論文（学術雑誌）, 共著, 査読有り

Numerical Investigations on Detonation Propagation in a Two-Dimensional Curved Channel

Sugiyama, Y., Nakayama, Y., Matsuo, A., Nakayama, H., and Kasahara, J.

Combustion Science and Technology 186 1662 - 1679 2014年09月

研究論文（学術雑誌）, 共著, 査読有り

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Supersonic Combustion Using a Stinger-Shaped Fuel Injector

Kouchi,T., Masuya, G., Hirano, K., Matsuo, A., and Tomioka, S.

Journal of Propulsion and Power 29 （ 3 ） 639-647 2013年05月

研究論文（学術雑誌）, 共著, 査読有り

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Numerical Prediction of Interior Ballistics Performance of a Projectile Accelerator Using a Granular or Tubular Solid Propellant

Miura, H., Matsuo, A., and Nakamura, Y.

Propellants, Eplosives, Pyrotechnics 38 （ 2 ） 204–213 2013年04月

研究論文（学術雑誌）, 共著, 査読有り

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Numerical Investigation for Pressure Mitigation Effects of Dike on Blast Wave

Miura, H., Matsuo, A., and Tabuchi, G.

Journal of Loss Prevention in the Process Industries 26 （ 2 ） 329 - 337 2013年

研究論文（学術雑誌）, 共著, 査読有り

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Initiation and Sustaining Mechanisms of Stabilized Oblique Detonation Waves Around Projectiles

Maeda, S., Sumiya, S., Kasahara, J., Matsuo, A.

Proceedings of the Combustion Institute 34 1973-1980 2013年

研究論文（学術雑誌）, 共著, 査読有り

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Comprehensive visualization of detonation-diffraction structures and sizes in unstable and stable mixtures

Nagura, Y., Kasahara, J., and Matsuo, A.

Proceedings of the Combustion Institute 34 1949-1956 2013年

研究論文（学術雑誌）, 共著, 査読有り

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Front shock behavior of stable curved detonation waves in rectangularcross - section curved channels

Nakayama, H., Kasahara, J., Matsuo, A., and Funaki, I.

Proceedings of the Combustion Institute 34 1939 - 1947 2013年

研究論文（学術雑誌）, 共著

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Numerical investigation on the detonation regime with longitudinal pulsation in circular and square tubes

Sugiyama, Y., and Matsuo, A.

Combustion and Flame 159 （ 12 ） 3646-3651 2012年12月

研究論文（学術雑誌）, 共著, 査読有り

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Visualization study of ignition modes behind bifurcated-reflected shock waves

Yamashita, H., Kasahara, J., Sugiyama, Y., and Matsuo, A.

Combustion and Flame 159 （ 9 ） 2954-2966 2012年09月

研究論文（学術雑誌）, 共著, 査読有り

Oblique Detonation Wave Stability Around a Spherical Projectile by a High Time Resolution Optical Observation

Maeda. S., Kasahara. J., and Matsuo. A.

Combustion and Flame 159 （ 2 ） 887-896 2012年02月

研究論文（学術雑誌）, 共著, 査読有り

Stable detonation wave propagation in rectangular-cross-section curved channels

Nakayama. H., Moriya. T., Kasahara. J., Sasamoto. Y., Funaki. I. and Matsuo. A.

Combustion and Flame 159 （ 2 ） 859-869 2012年02月

研究論文（学術雑誌）, 共著, 査読有り

Optical and Thrust Measurement of a Pulse Detonation Combustor with a Coaxial Rotary Valve

Matsuoka. K., Esumi. M., Ken B. Ikeguchi, Kasahara. J. Funaki. I., and Matsuo. A.

Combustion and Flame 159 1321-1338 2012年01月

研究論文（学術雑誌）, 共著, 査読有り

Experimental Investigation of a Multi-Cycle Single-Tube Pulse Detonation Rocket Engine with a Coaxial Rotary Valve

Matsuoka, K., Esumi, M., Ken B. Ikeguchi, Kasahara, J., Funaki, I., and Matsuo, A.

Transaction of JSASS Aerospace Technology Japan 10 （ ists28 ） Te_1-4 2012年

研究論文（学術雑誌）, 共著, 査読有り

Unsteady Propagation Process of Oblique Detonation Waves Initiated by Hypersonic Spherical Projectiles

Maeda, S., Kasahara, J., and Matsuo, A.

Transaction of JSASS Aerospace Technology Japan 10 （ ists28 ） Pe_1-6 2012年

研究論文（学術雑誌）, 共著, 査読有り

Propagation of Curved Detonation Waves Stabilized in Annular Channels with a Rectangular Cross-section

Nakayama. H., Takahiro M., Kasahara. J., Matsuo A., Sasamoto, Y., and Funaki. I.

Transaction of JSASS Aerospace Technology Japan 10 （ ists28 ） Pe_7-14 2012年

研究論文（学術雑誌）, 共著, 査読有り

The Influence of Heat Transfer and Friction on the Impulse of a　Detonation Tube

Kawane, K., Shimada, S., Kasahara, J., and Matsuo, A.

Combustion and Flame 158 （ 10 ） 2023-2036 2011年10月

研究論文（学術雑誌）, 共著, 査読有り

不足膨張噴流による膜物体の大変形シミュレーション

那須野雄介, 松尾亜紀子

ながれ 30 （ 4 ） 325 - 335 2011年08月

研究論文（学術雑誌）, 共著, 査読有り

圧縮性流体と多粒子系膜モデルを用いた膜物体の連成計算手法

那須野雄介, 松尾亜紀子

ながれ 30 （ 3 ） 229 - 239 2011年07月

研究論文（学術雑誌）, 共著, 査読有り

Three-Dimensional Simulation of Pressure Fluctuation in a Granular Solid Propellant Chamber within an Ignition Stage

Miura, H., Matsuo, A., and Nakamura, Y.

Propellants, Eplosives, Pyrotechnics 36 （ 3 ） 259-267 2011年06月

研究論文（学術雑誌）, 共著, 査読有り

Three-Dimensional Simulation of Pressure Fluctuation in a Granular Solid Propellant Chamber within an Ignition Stage

Miura, H., Nakamura, Y., and Matsuo, A.

Propellants, Explosives, Pyrotechnics 36 （ 3 ） 259–267 2011年06月

研究論文（学術雑誌）, 共著, 査読有り

爆轟管同軸回転バルブを用いたパルスデトネーションエンジンの性能評価

中道達也, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

航空宇宙技術 10 107-112 2011年

研究論文（学術雑誌）, 共著, 査読有り

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Oblique Detonation Waves Stabilized in Rectangular-Cross-Section Bent Tubes

Kudo, K., Nagura, Y., Kasahara, J., Sasamoto, Y., and Matsuo, A.

Proceedings of the Combustion Institute 33 2319-2326 2011年

研究論文（研究会，シンポジウム資料等）, 共著, 査読有り

Visualization of the Non-Steady State Oblique Detonation Wave Phenomena Around Hypersonic Spherical Projectile

Maeda, S., Inada, R., Kasahara, J., and Matsuo, A.

Proceedings of the Combustion Institute 33 2343-2349 2011年

研究論文（学術雑誌）, 単著, 査読有り

Combustion Wave Propagation and Detonation Initiation in the Vicinity of Closed-Tube End Walls

Yageta, J., Shimada, S., Matsuoka, K., Kasahara, J., and Matsuo, A.

Proceedings of the Combustion Institute 33 2303-2310 2011年

研究論文（学術雑誌）, 共著, 査読有り

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On the Characteristics of Two-dimensional Double Cellular Detonations with Two　Successive Reactions Model

Sugiyama, Y., and Matsuo, A.

Proceedings of the Combustion Institute 33 2227-2233 2011年

研究論文（学術雑誌）, 共著, 査読有り

固体発射薬の初期温度条件が薬室内差圧変動に及ぼす影響に関する数値解析

三浦啓晶，松尾亜紀子，中村雄一

Science and Technology of Energetic Materials 71 （ 6 ） 151 - 157 2010年12月

研究論文（研究会，シンポジウム資料等）, 共著

Propagation Characteristics of Shock Waves Driven by Gaseous Detonation Waves

Kato, S., Hashimoto, S., Uemichi, A., Kasahara, J., and Matsuo, A.

Shock Waves 20 （ 6 ） 479-489 2010年12月

研究論文（学術雑誌）, 共著, 査読有り

Analysis on Thermal Efficiency of Non-Compressor Type Pulse Detonation Turbine Engines

Maeda, S., Kasahara, J., Matsuo, A., and Endo, T.

Transactions of the Japan Society for Aeronautical and Space Sciences 153 （ 181 ） 192-206 2010年11月

研究論文（学術雑誌）, 共著, 査読有り

爆発安全研究への期待

松尾亜紀子

RISS Newsletter : Safety & Sustainability 第8号 2010年09月

研究論文（大学，研究機関等紀要）, 単著

高エネルギーの放出が生む流体現象を解く

松尾亜紀子

火薬学会2010年度春季研究発表会 2010年05月

研究論文（学術雑誌）, 単著

Numerical analysis on behavior of dilute water droplets in detonation

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

Proceedings of the Combustion Institute （Proceedings of the Combustion Institute） 2010年

研究論文（国際会議プロシーディングス）, 共著, 査読有り, ISSN 15407489

© 2020 The Combustion Institute Two-dimensional numerical simulations are conducted based on the Eulerian-Lagrangian method to model a gaseous detonation laden with monodispersed water droplets. The premixed mixture is a slightly diluted stoichiometric hydrogen oxygen mixture at low pressure. The outcome of the interactions of the droplet breakup with the cellular instabilities and the non-uniform flow behind the leading front is analyzed. The simulation results are also analyzed using instantaneous flow fields and Favre average profiles for water droplets. Breakup occurs mainly near the detonation front. The mean final diameter of the water droplets at the end of the breakup process is the same regardless of the initial strength of the leading shock or whether it is lower or greater than the Chapman-Jouguet value. The polydispersity comes from local phenomena behind the leading shock, such as forward jets coming from triple point collisions, transverse waves and a combination of both. The total breakup time is longer than that estimated from post-shock conditions and the present finding is in line with the previous experimental results on spray detonation.

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Thrust Measurement of a Multicycle Partially Filled Pulse Detonation Rocket Engine

Kasahara, J., Hirano M., Matsuo, A., Daimon, Y., and Endo, T.

Journal of Propulsion and Power 25 （ 6 ） 1281-1290 2009年11月

研究論文（学術雑誌）, 共著, 査読有り

KHIにおける低公害航空エンジン燃焼器用燃料噴射弁の開発

小林正佳，緒方秀樹，小田剛生，堀川敦史，木下康裕，松尾亜紀子

微粒化（日本液体微粒化学会） 18 （ 61 ） 4-11 2009年03月

研究論文（学術雑誌）, 共著

Numerical Investigation on Propagation Mechanism of Spinning Detonation in a Circular Tube

Sugiyama, Y., and Matsuo, A.

Proceedings of the Combustion Institute 32 2331-2337 2009年

研究論文（学術雑誌）, 共著, 査読有り

Matched Pressure Injections into a Supersonic Crossflow Through Diamond-Shaped Orifices

Tomioka, S., Izumikawa, M., Kouchi, T., Masuya, G., Hirano,K.,and Matsuo, A.

Journal of Propulsion and Power （AIAA） 24 （ 3 ） 471-478 2008年05月

研究論文（学術雑誌）, 共著, 査読有り

AGARD gun条件における軸対称固気二相流砲内弾道計算と手法の有効性

三浦啓晶，松尾亜紀子

Science and Technology of Energetic Materials 69 （ 1 ） 8-15 2008年01月

研究論文（学術雑誌）, 共著, 査読有り

Numerical Analysis of Projectile-Launch Tube Wall Friction Effects on Projectile Acceleration in Single-Stage Gun

Miura, H., Matsuo, A.

Materials Science Forum 566 71-76 2008年01月

研究論文（学術雑誌）, 共著

Multi-Dimensional Simulation on Ignition Stage of Granular Solid Propellant Varying Primer Configuration

Miura, H., Matsuo, A., and Nakamura, Y.

International Journal of Energetic Materials and Chemical Propulsion 7 （ 6 ） 507-522 2008年

研究論文（学術雑誌）, 共著, 査読有り

建物内部におけるガス爆発シミュレーション

松尾亜紀子

計算工学 12 （ 2 ） 1555-1562 2007年04月

研究論文（学術雑誌）, 単著

飛翔体加速装置における固体発射薬の燃焼シミュレーション

松尾亜紀子，三浦啓晶

日本航空宇宙学会北部支部20周年記念講演会（日本航空宇宙学会） 2007年03月

研究論文（その他学術会議資料等）, 共著

Comment on "Numerical studies on specific impulse of partially filled pulse detonation rocket engines" - Reply (Reply by the Authors to W. H. Heiser)

Matsuo, A., Endo, T., and Kasahara, J.

Journal of Propulsion and Power 23 （ 1 ） 254 2007年01月

研究論文（学術雑誌）, 共著, 査読有り

Unsteady Features on One-dimensional Hydrogen-air Detonations

Y. Daimon and A. Matsuo

Physics of Fluids （American Institute of Physics） 19 （ 11 ） 2007年

研究論文（学術雑誌）, 共著, 査読有り

Numerical Analysis of Gas Explosion inside Two Rooms Connected by Ducts

Hashimoto, A., and Matsuo, A.

Journal of Loss Prevention in the Process Industries 20 （ 4-6 ） 455-461 2007年

研究論文（学術雑誌）, 共著, 査読有り, ISSN 0950-4230

Homogeneous-Dilution Model of Partially Fueled Simplified Pulse Detonation Engines

Takuma Endo; Tomoaki Yatsufusa; Shiro Taki; Akiko Matsuo; Kazuaki Inaba; Jiro Kasahara

Journal of Propulsion and Power （AIAA） 23 （ 5 ） 1033-1041 2007年

研究論文（学術雑誌）, 共著, 査読有り

Numerical Studies on Specific Impulse of Partially Filled Pulse Detonation Rocket Engines

Sato, S., Matsuo, A., Endo, T., Kasahara, J.

Journal of Propulsion and Power 22 （ 1 ） 64-69 2006年12月

研究論文（学術雑誌）, 共著, 査読有り

マッハ反射による煤膜模様の形成機構

因幡和晃、松尾亜紀子、Shepherd, J.E.

日本燃焼学会誌 48 （ 146 ） 372-378 2006年11月

研究論文（学術雑誌）, 共著, 査読有り

デトネーションの数値解析のための熱化学モデルの提案

渡邊和也，　松尾亜紀子

日本燃焼学会誌 48 （ 145 ） 281-292 2006年08月

研究論文（学術雑誌）, 共著, 査読有り

Numerical Study on Underwater Explosion Simulation Surrounded by an Iron Wall Using Smoothed Particle Hydrodynamics

Kobashi, W. and Matsuo, A.

Science and Technology of Energetic Materials 66 （ 6 ） 421-424 2005年11月

研究論文（学術雑誌）, 共著, 査読有り

パルスデトネーションエンジン研究の背景

松尾亜紀子

日本燃焼学会誌 47 （ 140 ） 81-83 2005年05月

研究論文（学術雑誌）, 単著

Numerical Investigation on Acoustic Coupling of Transverse Waves in Two-Dimensional H2-O2-Diluent Detonations

Inaba, K., Matsuo, A., and Tanaka, K.

Transactions of the Japan Society for Aeronautical and Space Sciences 47 （ 158 ） 249-255 2005年02月

研究論文（学術雑誌）, 共著, 査読有り

Pressure History at the Thrust Wall of a Simplified Pulse Detonation Engine

Endo, T., Kasahara, J., Matsuo, A., Inaba, K., Sato, S., and Fujiwara, T.

AIAA Journal 42 （ 9 ） 1921-1930 2004年09月

研究論文（学術雑誌）, 共著, 査読有り

超音速燃焼器燃料噴射孔形状の設計問題における最適化と設計情報の知識発見

佐々木康友，松尾亜紀子

日本航空宇宙学会論文集 52 （ 607 ） 371-376 2004年08月

研究論文（学術雑誌）, 単著, 査読有り

Analogy between Wedge-Induced Steady Oblique Detonation and One-Dimensional Piston-Supported Unsteady Detonation

Daimon, Y. and Matsuo, A.

Science and Technology of Energetic Materials 65 （ 4 ） 111-115 2004年08月

研究論文（学術雑誌）, 共著, 査読有り

Oscillation Transition in Decay Process of Overdriven Detonation

Inaba, K., Matsuo, A. and Tanaka, K.

Science and Technology of Energetic Materials 65 （ 4 ） 134-139 2004年08月

研究論文（学術雑誌）, 共著, 査読有り

Reflection and Diffraction Phenomena of Blast Wave Propagation in Nuclear Fuel Cycle Facility

A. Miura, A. Matsuo, T.Mizukaki, T. Shiraishi, G. Utsunomiya, K. Takayama and I. Nojiri

JSME International Journal, Series B 47 （ 2 ） 287-292 2004年05月

研究論文（学術雑誌）, 共著, 査読有り

Parallelization of the Compressible Flow with the Finite Rate Chemical Reaction for Real Time Visualization

Daimon, Y., and Matsuo, A.

Interim Ras Research Progress Report 2 73-76 2004年03月

研究論文（大学，研究機関等紀要）, 共著

Spread Behavior of Explosion in Closed Space

A. Miura, T. Mizukaki, T. Shiraishi, A. Matsuo, K. Takayama, and I. Nojiri

Journal of Loss Prevention in the Process Industries 17 81-86 2004年

研究論文（学術雑誌）, 共著, 査読有り

超音速燃焼器の自発着火に対する支配パラメータの検討

高瀬一生，松尾亜紀子，三谷徹

日本航空宇宙学会論文集 51 （ 592 ） 191-197 2003年05月

研究論文（学術雑誌）, 共著, 査読有り

多目的遺伝的アルゴリズムを用いた超音速燃焼器燃料噴射口形状の最適化

石部哲史，松尾亜紀子

日本航空宇宙学会論文集 51 （ 592 ） 169-174 2003年04月

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Detailed Features of One-Dimensional Detonations

Y. Daimon and A. Matsuo

Physics of Fluids 15 （ 1 ） 112-122 2003年01月

研究論文（学術雑誌）, 共著, 査読有り

Triple Flameにおける予混合火炎の火炎特性

廣田光智，松尾亜紀子，溝本雅彦

日本機械学会論文集 69 （ 677 ） 177-184 2003年01月

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Spared Behavior of Explosion in Closed Space

A. Miura, T. Mizukaki, T. Shiraishi, A. Matsuo, K. Takayama, and I. Nojiri

Journal de Physique Ⅳ （ Ⅳ ISHPMIE ） 231-237 2002年10月

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軸対称壁噴流中に形成されるEdge Flameの安定機構

鳥飼宏之，松尾亜紀子，植田利久，溝本雅彦

日本機械学会論文集 68 （ 671 ） 2150-2155 2002年07月

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Criticality for Stabilized Oblique Detonation Waves Around Spherical Bodies in Acetylene/Oxygen/Krypton Mixtures

J. Kasahara, T. Arai, S. Chiba, K. Takazawa, Y. Tanahashi, and A. Matsuo

Proceedings of Combustion Institute 29 2817-2824 2002年07月

研究論文（学術雑誌）, 共著, 査読有り

Experimental Investigations of Momentum and Heat Transfers in Pulse Detonation Engines

J. Kasahara, K. Takazawa, T. Arai, Y. Tanahashi, S. Chiba, and A. Matsuo

Proceedings of Combustion Institute 29 2847-2854 2002年07月

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超音速燃焼器内の自発着火領域に関する数値解析

高瀬一生，松尾亜紀子，富岡定毅，三谷徹

日本航空宇宙学会論文集 50 （ 580 ） 181-187 2002年05月

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よどみ流中に形成されるEdge Flameの消炎特性と火炎構造

鳥飼宏之，松尾亜紀子，植田利久，溝本雅彦

日本機械学会論文集 B編 68 （ 666 ） 320 - 328 2002年02月

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酸水素希釈二次元爆轟波の横波構造と第二燃焼限界

因幡和晃，松尾亜紀子，田中克己

火薬学会誌 62 （ 677 ） 269-275 2001年11月

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Triple Flameの燃焼速度から考察する火炎先端の相互干渉

廣田光智，松尾亜紀子，溝本雅彦

日本機械学会論文集 67 （ 661 ） 2353 - 2360 2001年09月

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平板に衝突する音速噴流の振動機構に関する数値解析

沖中亮太，松尾亜紀子

日本航空宇宙学会論文集 49 （ 572 ） 316-321 2001年09月

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軸対称衝突噴流場の壁噴流領域に形成されるEdge Flameの火炎特性

鳥飼宏之，松尾亜紀子，植田利久，溝本雅彦

日本機械学会論文集Ｂ編 67 （ 653 ） 234 - 240 2001年01月

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超音速流中の物体周りに誘起される燃焼の非定常挙動

松尾亜紀子

可視化情報学会誌 20 （ 79 ） 21-26 2000年10月

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Flow Features of Shock-Induced Combustion Around Cylindrical Projectiles

Y. Kamiyama, A. Matsuo

Proceedings of Combustion Institute 28 671-677 2000年07月

研究論文（学術雑誌）, 共著, 査読有り

濃度こう配を有する予混合気中を伝ぱする火炎先端領域の特性

廣田光智，松尾亜紀子，溝本雅彦

日本機械学会論文集Ｂ編 66 （ 646 ） 1557 - 1564 2000年06月

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Numerical Simulations and Planar Laser-Induced Fluorescence Imaging Results of Hypersonic Reactive Flows

K.Toshimitsu, A.MATSUO, Michel R. Kamel, Christopher I. Morris and Ronald K. Hanson

Journal of Propulsion and Power 16 （ 1 ） 16-21 2000年01月

研究論文（学術雑誌）, 共著, 査読有り

同軸流中に形成される層流逆拡散火炎の火炎特性

大輝祐一，松尾亜紀子，溝本雅彦

日本機械学会論文集，Ｂ編 66 （ 641 ） 263 - 270 2000年01月

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A Comprehensive Examination of the Structure, Extinction of Turbulent Nonpremixed Flames Formed in a Counterflow

A. Kitajima, T. Ueda, A. Matsuo, M. Mizomoto

Combustion and Flame 121 301-311 2000年

研究論文（学術雑誌）, 共著, 査読有り

Response of Flame Displacement Speeds to Oscillatory Stretch in Wall-Stagnating Flow

T. Hirasawa, T. Ueda, A. Matsuo, M. Mizomoto

Combustion and Flame 121 312-322 2000年

研究論文（学術雑誌）, 共著, 査読有り

対向噴流乱流拡散火炎の火炎挙動と流れ場の相互干渉

北島暁雄，植田利久，松尾亜紀子，溝本雅彦

日本機械学会論文集Ｂ編 65 （ 635 ） 2526 - 2533 1999年07月

研究論文（学術雑誌）, 共著, 査読有り

Numerical Prediction of Envelope Oscillation Phenomena of Shock-Induced Combustion

A.Matsuo

Transactions of the Japan Society for Aeronautical and Space Sciences Vol. 41 （ No.133 ） 140-143 1998年11月

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Prediction Method of Unsteady Combustion Around Hypersonic Projectile in Stoichiometric Hydrogen-Air

A.Matsuo and K.Fujii

AIAA Journal Vol. 36 （ No. 10 ） 1834-1841 1998年10月

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流れ方向に振動する壁面近傍の壁面よどみ流

平沢太郎，植田利久，松尾亜紀子，溝本雅彦

日本機械学会論文集，Ｂ編 64巻（ 624号） 2683 - 2689 1998年08月

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Effect of Oscillatory Stretch on the Flame Speed of Wall-Stagnating Premixed Flame

T. Hirasawa, T. Ueda, A. Matsuo, and M. Mizomoto

Proceedings of Combustion Institute 27 （ - ） 875-882 1998年07月

研究論文（学術雑誌）, 共著, 査読有り

対向噴流拡散火炎における空気流の周期的変動の影響

大澤英雄，北島暁雄，松尾亜紀子，植田利久，溝本雅彦

日本機械学会論文集，Ｂ編 64巻（ 620号） 1228 - 1233 1998年04月

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Numerical Investigation of the One-Dimensional Piston Supported Detonation Waves

松尾亜紀子，藤井孝藏

Journal of Energy Conversion and Management Vol. 38 （ No. 10-13 ） 1283 - 1295 1997年11月

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衝撃波が誘起する燃焼

松尾亜紀子

計算工学 2 （ 3 ） 164-168 1997年09月

研究論文（学術雑誌）, 単著

Detailed Mechanism of the Unsteady Combustion Around Hypersonic Projectiles

松尾亜紀子，藤井孝藏

AIAA Journal 34 （ 10 ） 2082-2089 1996年10月

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Experimental Investigation of the Flames Structure and Extinction of Turbulent CounterFlow Non-premixed Flames

北島暁雄，植田利久，松尾亜紀子，溝本雅彦

Proceedings of Combustion Institute 26 137 - 143 1996年07月

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Computational Study of Large-Disturbance Oscillations in Unsteady Supersonic Combustion Around Projectiles

松尾亜紀子，藤井孝藏

AIAA Journal 33 （ 10 ） 1828-1835 1995年10月

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Flow Features of Shock-Induced Combustion Around Projectile Traveling at Hypervelocities

松尾亜紀子，藤井孝藏，藤原俊隆

AIAA Journal 33 （ 6 ） 1056-1063 1995年06月

研究論文（学術雑誌）, 共著

Numerical Investigation of Oscillatory Instability in Shock-Induced Combustion Around a Blunt Body

松尾亜紀子，藤原俊隆

AIAA Journal 31 （ 10 ） 1835-1841 1993年10月

研究論文（学術雑誌）, 共著

Numerical Prediction of Mechanism on Oscillatory Instabilities in Shock-Induced Combustion

松尾亜紀子，藤原俊隆

Dynamic Aspects of Explosion Phenomena, Progress in Astronautics and Aeronautics 154 516-531 1993年06月

研究論文（学術雑誌）, 共著

Numerical Investigation of Standing Oblique Detonation Supported by Two-Dimensional Blunted Wedge

松尾亜紀子，藤原俊隆

Transactions of the Japan Society for Aeronautical and Space Sciences 36 （ 111 ） 47-56 1993年05月

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Oxyhydrogen Oblique Detonation Supported by Two-Dimensional Wedge

Fujiwara, T., and Matsuo, A.

Memoirs of the Faculty of Engineering （Nagoya University） 41 （ 2 ） 295-302 1990年03月

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