Details of a Researcher - Matsuo, Akiko

Papers - Matsuo, Akiko

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Space Flight of Liquid Rotating Detonation Engine Using Sounding Rocket S-520-34

Sato, T., Matsuoka, K., Itouyama, N., Yasui, M., Matsuyama, K, Ide, Y., Nakata, K., Suzuki, Y., Ishibashi, R., Suzuki, S., Kasahara, J., Kawasaki, A., Hirashima, H., Nakata, D., Eguchi, H., Uchiumi, M., Himeno, T., Yahata, Y., Matsuo, A., Funaki, I., Habu, H., Arakawa, S., Masuda, J., Kawahara, K., Usuki, T., Maehara, K., Shida, M., Nakao, T., and Yamada, K.

Journal of Spacecraft and Rockets 2025.11

Research paper (scientific journal), Joint Work, Accepted

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Internal Flow Structure in Rotating Detonation Engines with Non-constant Area Channels

Nakata, K., Kimura, T., Itouyama, N., Matsuoka, K., Kasahara, J., Kawasaki, A., Watanabe, H., Matsuo, A., Funaki, I., Higashino, K., Athmanathan, V., Meyer, T., Paniagua, G.

AIAA Journal 2025.11

Research paper (scientific journal), Joint Work, Accepted

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Numerical investigation of a detonation-assisted fuel injection system in supersonic crossflow

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

Combustion and Flame 281 2025.11

ISSN 00102180

A novel detonation-assisted fuel injection system was developed in this study to achieve highly efficient supersonic combustion in scramjet engines. In this configuration, a Rotating Detonation Combustor (RDC) with an annular shape was coaxially employed around the main fuel injector, where hydrogen was injected perpendicular to the supersonic airflow. The computational domain consisted of a three-dimensional rectangular region, into which a Mach 2.4 supersonic flow was introduced to simulate flight conditions corresponding to Mach 8.0 at an altitude of 30 km. The main fuel injector and the RDC were coaxially connected to the isothermal lower wall of the scramjet engine. In the RDC, a stoichiometric premixed H₂–O₂ mixture was supplied from the bottom to generate a detonation wave. The compressible Navier–Stokes equations were solved under unsteady conditions. As a result, a detonation wave propagated continuously within the RDC, even when connected to the combustor section exposed to the supersonic main stream. The detonation products, accelerated to supersonic speeds, were discharged together with the main fuel into the scramjet combustor. This configuration generated large-scale vortex structures in the main stream, leading to a combustion efficiency up to 1.9 times higher and a 56% reduction in combustor length. This enhancement was primarily attributed to the high-enthalpy detonation products containing reactive radicals, which assisted main fuel penetration, increasing the penetration height by approximately 85%. Furthermore, high-frequency pressure disturbances with helically distributed spatial patterns were observed on the lower wall of the scramjet combustor, which were considered to contribute to combustion enhancement. These results demonstrate the effectiveness of detonation-assisted injection and highlight its potential as a solution to the long-standing challenges of stable and efficient combustion in supersonic propulsion systems.

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Numerical evaluation of the effectiveness of detonation gas injection for enhancing combustion performance in scramjet engines

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

Proceedings of the Combustion Institute 41 ( 105941 ) 2025.10

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

In this study, we focus on a novel approach that utilizes high-enthalpy detonation products as an energy source to assist supersonic combustion. Three-dimensional numerical analyses were conducted to evaluate the thermal and chemical effects of burned gas injected into a supersonic flow to enhance combustion in scramjet engines. The computational domain consisted of a three-dimensional square cross-section simulating a scramjet combustor into which supersonic airflow entered at Mach 2.4, corresponding to flight conditions at Mach 8.0. A main fuel injector and a burned-gas injection port were coaxially mounted on the isothermal lower wall of the combustor. To independently assess thermal and chemical effects, two types of burned-gas injection were considered: high-temperature H₂O without radicals (thermal assistance only) and burned gas containing radicals (thermal and chemical assistance). The results demonstrated that burned-gas injection provided a preheating effect due to its higher temperature than the main fuel, and radical injection further increased the heat release rate at comparable temperature. These effects were corroborated by the spatial distribution of the Damköhler number ( Da ₂), which revealed a clear expansion of reaction-dominated regions ( Da ₂ > 1) in the injection cases. Consequently, streamwise fuel consumption was accelerated; in the most enhanced case, the same level of fuel consumption was achieved 75.3 % earlier along the combustor length compared with the baseline. Furthermore, when comparing the burned-gas injection cases, it was suggested that only thermal and chemical assistance could shorten the required combustor length by 26.0 %. The fuel-consumption time scale decreased approximately linearly with increasing enthalpy input, with additional reductions observed in radical-containing cases. These findings demonstrate the effectiveness of this novel detonation-assisted injection strategy in promoting supersonic combustion and underscore its potential for future scramjet applications.

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Effect of reflection wall distance and supply conditions on reflective shuttling detonation combustor

Nagaoka T., Matsuoka K., Takahashi Y., Watanabe H., Itouyama N., Kawasaki A., Kasahara J., Matsuo A.

Combustion and Flame 276 2025.06

ISSN 00102180

In a reflective shuttling detonation combustor (RSDC), detonation waves propagate between the two reflection walls of a thin combustion chamber. As the chamber is two-dimensional and does not have any curvature, optical and pressure measurements can be conducted simultaneously. In this study, pressure measurements as well as self-luminous and schlieren visualizations were performed using a chamber with a reflection wall distance of 90 mm with ethylene and oxygen. Consequently, the detonation wave number increased from 2 to 5 following the total mass flux, indicating that the wave number depended on the reflection wall distance because the maximum wave number increased with increasing distance. In addition, although the reflection wall distance was varied, the mode transition, in which the wave number increased from 1 to 2, 2 to 3, and 3 to 4, occurred at similar values to the critical value at which the maximum mixture fill height was non-dimensionalized by the cell size. The dimensionless detonation wave propagation distance, that is, the reflection wall distance divided by the wave number and maximum mixture fill height, was 3.0 ± 0.3, which is almost the same as that of the previous 45 mm combustor. Additionally, when the ratio of the reflective wall distance to the wave number is the same, the dimensionless quantities match, except in a few cases. Comparing the model results with the schlieren results, different trends were caused by non-ideal phenomena arising from the supply conditions, such as the equivalence ratio and the relationship between the plenum pressure and combustion pressure.

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Direction Control of Rotating Detonation Waves with Helical Combustion Chambers

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

Journal of Propulsion and Power 41 ( 2 ) 164 - 177 2025.03

ISSN 07484658

This paper explores the innovative direction control of rotating detonation waves in rotating detonation engines (RDEs) by adjusting the ignition location and employing helical combustors with a sinusoidal cross section. In our experimental setup, we conducted 25 combustion tests using two distinct combustor geometries, each featuring different helical profile directions. The following conclusion drawn from the results: when the ignition was positioned 30.6–46.0 mm from the inlet, the detonation wave direction was invariably influenced by the helical direction. This correlation was statistically significant, with an occurrence probability (assuming a random direction probability of 0.5) being 3 × 10−8, far exceeding the 0.05 significance level. Furthermore, the helical combustors generated a measurable torque due to the pressure differentials created by shock waves within the combustor. This torque, recorded between 0.040 and 0.077 N ⋅ m at a mass flow rate of 28.5–29.0 g/s, indicates the potential of power extraction from the combustor. Notably, the torque direction was also controllable via the helical direction. This study presents a significant advancement in propulsion technology, demonstrating a novel method to control detonation wave direction and torque generation in RDEs through helical combustor design, paving the way for more efficient and controllable propulsion systems.

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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 ) 2025.02

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

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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Experimental Investigation of Gaseous Film Cooling for Cylindrical Rotating Detonation Engines

Chan A., Nakata K., Itouyama N., Matsuoka K., Kasahara J., Kawasaki A., Matsuo A., Higashino K., Funaki I.

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

A new combustor has been developed for an experimental test campaign on film cooling for hollow-type Rotating Detonation Engines (RDEs). In support of these research efforts, a modular design has been employed with interchangeable injector manifolds such that a range of coolants, propellants, injection geometries, and operating conditions can be examined. The engine is instrumented with pressure transducers and thermocouples to evaluate performance and heat loads at various locations along the axial direction. High-speed imaging has also been used to study the wave stability, internal combustion structure, and engine ignition process. A parametric study with varying coolant mass flow rate between 5-25% was conducted to characterize the influence of the coolant layer on the engine’s operation. Phase 1 of the study focused on gaseous propellants, utilizing ethylene (C2H4) as the fuel, oxygen (O2) as the oxidizer, and nitrogen (N2) as an inert coolant. In Phase 2, the investigation shifted to a fuel-based film coolant. Thermocouple measurements revealed a significant reduction in heat flux along the length of the combustion chamber. The highest cooling effectiveness was observed immediately after coolant injection, extending to the region near the wave’s triple point. Stable wave propagation was also observed with a reasonable reduction in engine performance. These initial findings demonstrate the potential of axially injected gaseous film cooling as an effective means of thermal management for RDEs.

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Flight Demonstration of Detonation Engine System Using Sounding Rocket S-520-34: Performance of Rotating Detonation Engine Using Liquid Propellants

Sato T., Matsuoka K., Itouyama N., Yasui M., Matsuyama K., Ide Y., Nakata K., Suzuki Y., Ishibashi R., Suzuki S., Kasahara J., Kawasaki A., Hirashima H., Nakata D., Eguchi H., Takano T., Uchiumi M., Matsuo A., Funaki I., Habu H., Arakawa S., Masuda J., Kawahara K., Usuki T., Maehara K., Shida M., Nakao T., Yamada K., Himeno T., Yahata Y.

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

One of the most important aspects of aerospace engineering applications is to create new “flyable” systems. We plan to demonstrate Detonation Engine System2 (DES2) operation using liquid propellants in the space environment for the first time, using the sounding rocket S-520-34. A cylindrical rotating detonation engine was used with liquid ethanol as fuel and liquid nitrous oxide as oxidizer. The plume images were successfully acquired by the onboard digital and analog cameras, and the ignition was successful, which can be confirmed from the image. From pressure and temperature histories, nitrous oxide was considered to be supplied into the combustor with nitrogen from X+125.5 s. Therefore, the combustion pressure decreased due to decrease in total mass flow rate. From the Short-time Fourier transform (STFT) result, a peak value of 21.0 kHz (two-wave mode) was measured, and a peak value of 13.5 kHz (single wave mode) was measured. Therefore, the bipropellant RDE operation was successful. As for the propulsive performance, we measured the RDE axial thrust from the flyable system’s axial acceleration directly, and the thrust was in good agreement with the value calculated from the combustion pressure. In addition, the specific impulse efficiency of 82±7% was achieved.

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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 2024.11

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., and Funaki, I.

Proceedings of the Combustion Institute (Proceedings of the Combustion Institute) 40 ( 1-4 ) 2024.08

Joint Work, Accepted, 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., and Matsuo, A.

Proceedings of the Combustion Institute 40 ( 1-4 ) 105577 2024.07

Research paper (international conference proceedings), Joint Work, Accepted, 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 2024.06

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 2024.06

Research paper (scientific journal), Joint Work, Accepted

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 2024.06

Research paper (scientific journal), Joint Work, Accepted

The circumferentia l force on a cylindrical rotating detonation engine

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

Proceedings of the Combustion Institute 40 ( 1-4 ) 105490 2024.06

Research paper (international conference proceedings), Joint Work, Accepted, 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, R., Nakajima, K., Nakata, K., Itouyama, N., Kawasaki, A., Matsuoka, K., Matsuyama, K., Kasahara, J., Eguchi, H., Nakata, D., Uchiumi, M., Matsuo, A., and Funaki, I.

Journal of Spacecraft and Rockets 62 ( 1 ) 44 - 54 2024.05

Research paper (scientific journal), Joint Work, Accepted, 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., Matsuo A.

Shock Waves (Shock Waves) 34 ( 2 ) 215 - 224 2024

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

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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