論文

Design analyses were conducted on the hypersonic integrated control experimental aircraft (HIMICO) that flies at Mach 5. Aerodynamic performance of the aircraft was obtained by hypersonic wind tunnel experiments and CFD analyses. Structural feasibility in a high-temperature flight environment was confirmed by fluid-structure coupling simulation and structural analyses. The control law of hypersonic air intake was confirmed by hypersonic wind tunnel test. The control law of hydrogen ram combustor was confirmed by direct connect combustion experiments. The aerodynamic force measurements and flow field measurements for propulsion wind tunnel tests with Mach 5 flight simulating condition were prepared. Mechanical environment tests of onboard equipment were conducted simulating the onboard environment of sounding rocket.

In this study, the blast-wave formation and mushroom-cloud development associated with the Nagasaki atomic bombing were numerically investigated to construct initial conditions for subsequent meteorological simulations of “black rain.” To evaluate the three-dimensional transport of hazardous materials within the cloud, two distinct material tracers were treated as passive scalars in an Eulerian framework and solved together with the compressible airflow: bomb-derived radioactive material, assumed to represent plutonium released from the bomb, and ground-derived dust, representing non-radioactive smoke and dust associated with urban fires and disturbed surface materials. The computation was carried out in three stages: a one-dimensional theoretical blast-wave calculation, a high-resolution three-dimensional simulation of the early blast-wave/fireball interaction, and a large-domain three-dimensional simulation of the subsequent mushroom-cloud rise. The numerical results showed that sufficient grid resolution is required in both the blast-wave and mushroom-cloud stages to reproduce the cloud-rise behavior accurately. In the blast-wave stage, coarse grids smeared the reflected-shock/fireball interaction and led to an underestimation of the later cloud-top height. The background atmospheric profile also affected the buoyant rise, and a dry adiabatic atmosphere showed better agreement with historical nuclear-test data than the International Standard Atmosphere. The three-dimensional blast-wave simulation clarified that the reflected shock wave deforms the fireball and induces a strong axial updraft, which provides the initial flow structure for the subsequent mushroom-cloud rise. The mushroom-cloud simulation reproduced the buoyant ascent of the hot fireball, the horizontal spreading of the cloud head, and the development of a vortex-ring structure. The potential-temperature perturbation field visualized the hot core and updraft, while the relative-humidity field indicated possible condensate-cloud formation in the upper part of the cloud. The two solid-phase components exhibited distinct transport behaviors depending on their initial locations. Ground-derived dust was entrained mainly through the central updraft, whereas bomb-derived radioactive material was preferentially incorporated into the vortex-ring region around the cloud head. These results provide the three-dimensional initial distribution of hazardous materials for meteorological simulations of black rain. (Figure presented.)

In this study, a three-dimensional conjugate heat transfer (CHT) analysis was conducted for the entire hypersonic experimental vehicle (HIMICO), including its internal thermal protection system (TPS), to evaluate the thermal response characteristics and thermal integrity of the vehicle under the wind tunnel test conditions at Ramjet Engine Test Facility in JAXA. The results indicated that, after 30 s of exposure to the high-temperature flow, the temperatures of certain regions of the Ti-6Al-4V structural components and the Cork insulation material exceeded their allowable operating limits, suggesting the potential for degradation of material strength and deterioration of thermal insulation performance. Comparison with wind tunnel test data showed good agreement in the surface temperature distributions at both the nose cone tip and the wing region, thereby demonstrating the validity of the present analysis approach. Furthermore, the results revealed that small geometric features, such as bolt holes, can significantly affect localized temperature rises and should therefore be considered in detailed thermal design evaluations.

One of the most important goals in aerospace engineering applications is the creation of new “flyable” systems. In a flight demonstration using the sounding rocket S-520-34, we show the first successful operation of a bipropellant cylindrical rotating detonation engine using liquid ethanol and liquid nitrous oxide, Detonation Engine System 2 (DES2), in a space environment. From the pressure and temperature histories, the combustion was finished before all propellants were consumed because nitrogen was supplied earlier than the ideal depletion time due to spin stabilization of the sounding rocket. Therefore, the combustion pressure decreased from the nitrogen-supply start time. The short-time Fourier transform result indicated that the deflagration mode, two-wave mode, and singlewave mode occurred in sequence. This was attributed to the locally lower liquid temperatures, wall temperature, and mixture ratio at ignition near the wall, where the rotating detonation wave propagated. A comparison of the filling mass and consumption indicated that the mass flow rate estimated using control surface theory reflects an actual phenomenon. As for the propulsive performance, the experimental characteristic exhaust velocity was almost the same as the ideal value. Moreover, a specific impulse efficiency of more than 90% was achieved throughout the rotating detonation engine operation.

Liquid propellants are widely used in rocket engines, and investigating the operational characteristics of RDEs using liquid propellants is essential for their practical application. In this study, combustion experiments were conducted on a single-annular RDE using liquid propane and liquid nitrous oxide as propellants, and its operational characteristics were evaluated. As a result, successful operation in the detonation mode was achieved with an operating frequency of 15–16 kHz. However, a gradual decrease in the propane mass flow rate was observed during combustion. This behavior was attributed to the possible occurrence of vapor lock in the propane injector feed lines caused by heat generated during combustion. To examine this possibility, the injector feed line wall temperature was calculated using the heat conduction equation under the assumptions of a one-dimensional semi-infinite solid, a constant incoming heat flux, and single-phase liquid flow in the injector. The results showed that for an incoming heat flux of q=0.2 MW/m2, the wall temperature near the injector outlet exceeds the boiling temperature of propane, indicating the potential for bubble formation at the injector outlet.

Flow inside a rotating detonation engine (RDE), which is studied for use in aircraft and rocket propulsion, is affected by the geometry of the injector component. Typically, RDEs are cylindrical in nature and have a circular injection pattern. However, although hollow RDEs are now being studied, the injector pattern has remained the fixed-radius, circular injector pattern. In order to understand the impact of uniform injection across the entire cross-sectional area of a combustor inlet, the current research sought to conceptualize, design, and test a bi-propellant RDE with a symmetric injector pattern, and elucidate detonation propagation. In this investigation, different injector patterns concepts were studied, and a symmetric, square, lattice pattern injector was designed and fitted to a square combustion chamber. Combustion tests were conducted with mass flowrates between 30-38 g/s, equivalence ratios between 1.0-1.6, and from back pressures of approximately 5 kPa. Experimental data revealed unidirectional detonation wave motion, reflected waves coexisting with the detonation wave under stable operation, non-uniform pressure near the combustor inlet, increased wave velocity compared to a previous hollow engine tested at Nagoya University, potentially improved reactant mixing in the chamber, and a 94° phase shift when decomposing the detonation wave motion into rectangular coordinates.

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

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

© 2020 The Combustion Institute Two-dimensional numerical simulations are conducted based on the Eulerian-Lagrangian method to model a gaseous detonation laden with monodispersed water droplets. The premixed mixture is a slightly diluted stoichiometric hydrogen oxygen mixture at low pressure. The outcome of the interactions of the droplet breakup with the cellular instabilities and the non-uniform flow behind the leading front is analyzed. The simulation results are also analyzed using instantaneous flow fields and Favre average profiles for water droplets. Breakup occurs mainly near the detonation front. The mean final diameter of the water droplets at the end of the breakup process is the same regardless of the initial strength of the leading shock or whether it is lower or greater than the Chapman-Jouguet value. The polydispersity comes from local phenomena behind the leading shock, such as forward jets coming from triple point collisions, transverse waves and a combination of both. The total breakup time is longer than that estimated from post-shock conditions and the present finding is in line with the previous experimental results on spray detonation.

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.

A rotating detonation engine (RDE) is expected to achieve a pressure gain (PG) in which the total pressure of the product at the engine exit exceeds the total pressure of the supplied oxidizer. However, many non-ideal phenomena exist in the RDE, which hinder the achievement of a PG. This study focused on the channel expansion angle near injector outlet which is considered to impact the structure of detonation wave and the PG performance. Combustion tests were conducted by varying the channel expansion angle near the axially injected oxidizer injector outlet to 90∘ and 30∘. As a result, the thrust was not affected by the expansion angle, but the propagation velocity of detonation wave at an expansion angle of 90∘ was approximately 5% higher than that at an expansion angle of 30∘. A comparison utilizing the pressure increase ratio obtained from the fluctuating pressure in the combustion chamber suggested that the Mach number of the detonation wave was higher for an expansion angle of 90∘. As a result of evaluating the PG performance utilizing the equivalent available pressure obtained from thrust measurements, it was not confirmed to differ with changes in the expansion angle.

© 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.

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.

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.

Converging–diverging nozzles are common in rocket engine systems to increase the exhaust velocity and improve thrust performance. In this study, we focused on the acceleration of subsonic burned gas without a structural throat via detonation to realize a simple and compact engine. We developed and tested a rotating detonation engine (RDE) without a throat and with a diverging channel (constant diverging angle α 5 deg). Gaseous C2H4 and O2 were used as the propellants, and the mass flow rate ranged from 62 to 134 g∕s in the combustion tests under low back-pressure conditions. We measured pressure and thrust, as well as high-speed imaging of self-luminescence of the combustion and imaging of the exhaust plume. The pressure at the exit was less than one-fifth of the maximum pressure in the RDE, significantly below the value for a sonic flow. The results suggested that the exhaust flow was supersonic, with values up to Mach 1.7, without the need of a converging section within the engine. In addition to the estimated Mach number from the measured pressure, the exhaust plume images coherently indicated the existence of supersonic exhaust.

Detonation engines exert various characteristics and high propulsion performance depending on their operation modes. Our group has conducted the demonstration of the integrated RDE/PDE system (Detonation Engine System, DES) in outer-space environment using a sounding rocket named S-520-31 operated by Japan Aerospace Exploration Agency. This paper introduces the development history of DES from main integrity evaluations of preflight model to the outer-space demonstration of flight model.

A research team of Nagoya University, Keio University, Muroran Institute of Technology, and Institute of Space and Astronautical Science (ISAS) at Japan Aerospace Exploration Agency (JAXA) successfully conducted a space demonstration of the detonation engine system (DES) using the S-520-31 sounding rocket of JAXA/ISAS. This paper presents the results of an S-shaped pulse detonation engine (PDE) flight test. The aim of the flight is to confirm that the PDE and its system operate in space, to confirm the reproducibility of the PDE cycle, and to despin the sounding around its axis. The PDE and its system successfully operated for 14 cycles in space. The experimental plateau pressure was 80 ± 3 % of the calculated plateau pressure, suggesting that detonation occurred in 14 cycles. Pressure profile of the 14 cycles was highly reproduced as well. The average angular acceleration change of PDE operation, mixture supply, and oxygen supply were determined using gaussian distribution. The results suggested that the PDE despinned the sounding rocket by the thrust from detonation combustion, which was in good agreement with a quasi-steady-state model with an accuracy of 101 ± 15%.

To create a new flyable propulsion system, Detonation Engine System (DES) has developed which can be stowed in the sounding rocket S-520-31. This paper focused on the first flight demonstration of RDE in space environment, which is integrated into DES, using S-520-31. The flight result was compared with the ground test data to validate its performance. In the flight experiment, the stable combustion of the RDE with aerospike nozzle was observed by onboard digital and analog cameras. With time-averaged mass flow of 182 ± 11 g/s and equivalence ratio of 1.2 ± 0.2, the RDE generated the time-averaged-thrust of 518 N, and the specific impulse achieved 290 ± 18 s, which is almost identical to the ideal value of constant pressure combustion. Due to the RDE combustion, the angular velocity increased by 5.7 deg/s for total, and the time-averaged torque during 6 s of operation was 0.26 N∙m. The high frequency sampling data identified the detonation frequency during record time as 20 kHz in the flight, which was confirmed in the DES ground test through the high frequency sampling data analysis and high-speed video imaging.