論文 - 松尾　亜紀子

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

　概要を見る

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

　概要を見る

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

　概要を見る

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

　概要を見る

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

　概要を見る

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

　概要を見る

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

　概要を見る

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

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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

　概要を見る

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

　概要を見る

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

　概要を見る

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

　概要を見る

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

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.

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

　概要を見る

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

　概要を見る

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

　概要を見る

© 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

　概要を見る

© 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

　概要を見る

© 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

　概要を見る

© 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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