Published papers

© 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.

© 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.

© 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.

© 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.

© 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.

© 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.

© 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.

© 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.

© 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.

© 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.

© 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.

© 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.

© 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.

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.

© 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.

© 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.

© 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.

© 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.

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.

© 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.