研究者詳細 - 松尾　亜紀子

論文 - 松尾　亜紀子

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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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