論文

A research group at Nagoya University, in collaboration with Keio University, the Muroran Institute of Technology, and Institute of Space and Astronautical Science at the Japan Aerospace Exploration Agency (ISAS/JAXA), has successfully demonstrated world’s first operation of Detonation Engine System (DES) in space using the sounding rocket S-520-31, which lifted off from the JAXA Uchinoura Space Center, Japan, at 5:30 a.m. (Japan Standard Time) on July 27, 2021. The DES was equipped with two different types of detonation rocket engines, namely a rotating detonation engine (RDE) and a pulse detonation engine (PDE), which generated 500-N-class axial thrust and 1-Nm-class rolling torque, respectively. The RDE and PDE were tested in space during the ballistic flight of the rocket. In this manuscript, the system-level design of the DES was disclosed. Key point are as follows: the DES consists of an RDE and a PDE, and propellant storage/feed, avionics, instrumentation, and structure subsystems; the overall dimensions are 480 mm in the diameter × 1810 mm in the length. The dry mass was 175 kg, while gross (wet) mass was 181 kg including propellant and pneumatic gases. Commercial off-the-shelf (COTS) components are widely utilized after carefulevaluation and testing, and with redundant architecture.

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.

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.

A detonation engine system is successfully demonstrated for the first time in space using sounding rocket S-520-31 of the Japan Aerospace Exploration Agency/Institute of Space Astronautical Science. Detailed flight results of an S-shaped pulse detonation engine (PDE) installed in the rocket are presented herein. The flight is conducted to confirm that the PDE and its system operate at scheduled sequences in space, confirm the reproducibility of the PDE cycle, and despin the rocket around its axis. It is confirmed that the PDE operated successfully for 14 cycles in space. The experimental plateau pressure of 2.0 0.1 MPais 80 3% of the calculated plateau pressure, which suggests that detonation occurred in 14 cycles. The pressure profiles of the cycles are similar, and the pressure integrals are 2.0 0.1 kN⋅ s∕m2, confirming the excellent reproducibility of the PDE cycle. A probability statistical approach assuming a Gaussian distribution is applied to determine the average angular acceleration difference between processes of the PDE operation, mixture supply, and oxygen supply. The results suggested that the PDE despun the rocket via the thrust produced via detonation combustion, which is consistent with a quasi-steady-state model with an accuracy of 101 15%.

Rotating detonation engines (RDEs) have been actively researched around the world for application to nextgeneration aerospace propulsion systems because detonation combustion has theoretically higher thermal efficiency than conventional combustion. Moreover, because cylindrical RDEs have simpler combustors, further miniaturization of conventional combustors is expected. Therefore, in this study, with the aim of applying RDEs to space propulsion systems, a cylindrical RDE with a converging-diverging nozzle was manufactured; the combustor length Lc was changed to 0, 10, 30, 50, and 200 mm; and the thrust performance and combustion mode with the different combustor lengths were compared. As a result, four combustion modes were confirmed. Detonation combustion occurred with a combustor length of Lc = 0 mm: that is, a converging rotating detonation engine. The thrust performance of this engine was 94 to 100% of the theoretical rocket thrust performance, which is equivalent to the thrust performance of conventional rocket combustion generated at Lc = 200 mm. This study shows that detonation combustion can significantly reduce engine weight while maintaining thrust performance.

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%.

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.

Detonation engines exert various characteristics and high propulsion performance depending on their operation mode. Our group has demonstrated the integrated Detonation Engine System (DES) in outer-space environment using the sounding rocket named S-520-31in Japan Aerospace Exploration Agency. The system was operated by the detonation of premixed gas composing of methane gas and oxygen gas. As the next step, we are planning to perform the in-space demonstration of detonation engine with liquid propellants in order to clarify its milestones on the achievement of space application of detonation engines as same as the conventional chemical propulsion systems. This paper introduces the overview of our group’s measures about the in-space demonstration of detonation engine, especially for rotating detonation engine, in the last few years.

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

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

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

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.

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.

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.

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.

Rotating detonation combustion was successfully tested for liquid ethanol and gaseous oxygen heterogeneous mixture in a cylindrical combustor for the first time. The rotating detonation combustor (RDC) we operated has a diameter of 20 mm and variable lengths of 60, 210 mm. Pure ethanol and industrial ethanol were selected as liquid fuels. Under the conditions of ethanol and oxygen mass flow rate of 26-40 g/s, equivalence ratio of 0.4-1.7, and backpressure of 10-17 kPa, we confirmed detonation combustion and deflagration combustion. In detonation combustion, as the ethanol manifold supply temperature increases, the detonation propagation speeds, and luminance of combustion were enhanced. It is possible to say enhanced evaporation behavior resulted in stable detonation and showed high peaks in luminance value. Utilizing the control surface method to evaluate experimental thrust, estimated thrust showed good agreement with experimental thrust. For liquid-fueled cylindrical RDCs, the control surface methods can also be used to evaluate thrust.

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.

The inner flow containing the ignition process of a throatless and diverging (constant diverging angle α=5 deg) rotating detonation engine (RDE) was experimentally investigated by both optical observation and pressure measurement. An RDE that had an acrylic engine wall with a length of 70 mm and an inlet diameter of 20 mm was tested under a low back pressure condition using gaseous C2H4-O2 as the propellants. A high-speed camera captured the ignition process, of which there were two types: deflagration-to-detonation transition and direct initiation. In addition to the visualizations of the inner flow, the pressure values in an RDE with an acrylic wall were compared with those of an RDE having a stainless steel wall with 8 pressure ports in the axial direction. The comparison revealed that the RDE exhaust flow was supersonic. A strong self-luminescence and CH* luminosity area near the injector surface during combustion suggested that the short completion of detonation combustion enabled the thermal choking of the subsonic propellant gases even when the channel was diverging. Moreover, the axial position of the erosion in the acrylic wall indicated that the heat release region was near the injector surface. Finally, a theoretical flow model for the RDE was compared to the experimental data, and the results suggested that the inner flow in the RDE can be reasonably described as a quasi-one-dimensional steady flow.

In-space flight demonstration of a rotating detonation engine system using liquid nitrous oxygen and liquid ethanol as the propellant (DES-2) is planned in FY2024 by a collaborative research group of Nagoya University, Shizuoka University, Keio University, the Muroran Institute of Technology, and Institute of Space and Astronautical Science at the Japan Aerospace Exploration Agency (ISAS/JAXA). In this manuscript, a ground combustion test of the bread board model (BBM) of the DES-2, which has been conducted in May 2023, is briefly described. In the testing, the DES-2 BBM was operated in stand-alone mode and achieved detonation combustion as planned.

A three-dimensional numerical analysis was performed on a reflective shuttling detonation combustor to investigate the propagation phenomena of detonation in a fuel-rich condition, with different equivalence ratio. The computational domain was a rectangular region where the hydrogen-air premixed gas flows in from the bottom injector with changing equivalence ratio from 1.0 to 2.0. As a result, continuous propagation of detonation was observed in each case, and the same trend of the propagation mode was observed. In addition, A decrease in the temperature value during combustion and an increase in the propagation velocity were observed as the equivalent ratio increased. Both trends were consistent with the trends of CJ temperature and CJ velocity with increasing equivalent ratio.