船木 一幸

A coupled cylindrical rotating detonation engine (RDE) with two cylindrical RDEs (both combustors had a combustor inner diameter of 23 mm and an axial length of 30 mm) placed next to each other was tested for rocket clustering application. The objective of the experiment was to achieve two-engine synchronized initiation with a single igniter. Experiments were conducted on the inner wall of the combustors with different connecting-hole diameters and wall heights to evaluate the ignition delay time, combustion mode, and propulsion performance. The propellants were gaseous ethylene and oxygen, and experiments were conducted under constant conditions of mass flow rate ([Formula: see text]), equivalence ratio ([Formula: see text]), and backpressure (approximately 10 kPa). When the two combustion chambers were completely separated by a wall, ignition occurred with a time delay of 260 ms in the chamber without an igniter. However, when a large hole ([Formula: see text] diameter) was placed in the wall separating the two combustion chambers, synchronous initiation was successful. Synchronous initiation was also successful when the wall height was lowered (7-mm height). Under both conditions, the same level of specific impulse was achieved as for RDEs operating at the same mass flux.

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 of [Formula: see text], mixture ratios of 3.6–5.9, and combustion pressures of 0.35–0.46 MPa, resulting in a maximum detonation velocity of [Formula: see text] (approximately 80% of the theoretical detonation velocity, [Formula: see text]), 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, [Formula: see text], was [Formula: see text], 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 [Formula: see text] efficiency, this study shows that detonation combustion can reduce the combustor size by 88%.

Rotating detonation engines (RDEs) have been actively researched around the world for application to next-generation 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 [Formula: see text] 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 [Formula: see text]: 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 [Formula: see text]. This study shows that detonation combustion can significantly reduce engine weight while maintaining thrust performance.

Abstract This paper presents a control algorithm for ensuring the stable operation of a 6 kW Hall thruster. The Japan Aerospace Exploration Agency is developing a “wide-range” power processing unit (PPU) to power a 6 kW Hall thruster system, a candidate for all-electric satellites. The PPU provides discharge oscillation, power, and discharge current control. The increments in the control loop are fixed, so the PPU digital controller does not need to calculate them, and the algorithm’s computational complexity is minimal. Results of integration tests on the 6 kW Hall thruster and the PPU breadboard model showed that the three control functions run correctly. The total controlled PPU power was adjusted in the range of 3.45–3.55 kW under constant power control with a target of 3.5 kW. Oscillations of ± 0.05 kW around the target power are acceptable. 70% of the peaks in the discharge current acquired for 100 ms exceeded the limit of 1.5A. Discharge oscillation control varied the coil magnetic field and reduced them to 0%. The PPU successfully controlled the coil current, reducing the discharge current from 13.7 A to 13 A. The propulsion efficiency increased from 59% to 60.5%.

2021年7月27日早朝5：30，JAXA内之浦宇宙空間観測所からデトネーションエンジンシステムを搭載した観測ロケットS-520-31号機が打ち上げられた．高度約200kmにてメタン–酸素推進剤による回転デトネーションエンジン（RDE）の6秒間作動およびパルスデトネーションエンジン（PDE）の2Hz作動を実施した．取得されたフライトデータから，RDE作動で時間平均推力518N，比推力290±18sおよび速度増速量8.0m/sを達成した．PDE作動では1サイクル当たりの圧力時間積分値が5％以内の高精度での繰り返しインパルス生成およびロケット機軸周りのスピンレート減少が確認された．本結果は，地上燃焼試験データとよく一致し，宇宙空間でのデトネーションエンジン作動が実証された．デトネーション波の判定に用いた圧力・加速度センサの高速サンプリングデータおよびRDEプルーム撮影用のデジタルカメラ画像は，JAXA/ISASで開発された再突入データ回収システムRATSにて回収することに成功した．

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.

<p>The ion energy angle distribution and its relationship to plasma parameters for spot and plume modes are elucidated for a LaB₆ hollow cathode with a radiative heater. Measurements were conducted using a retarding potential analyzer (RPA) and a single Langmuir probe. The ion energy distribution function (IEDF) characteristics showed different tendencies in the current density and mass flow-rate dependence under different plasma modes. The IEDF peak potential for the spot mode varied from 16 to 23 V with increasing current density, and the IEDF peak potential for the plume mode varied from 16 to 32 V with decreasing mass flow rate. Considering angle dependency of ion energy, when the observation angle was changed from the radial direction to the axial direction, the IEDF peak potential increased from 29 to 40 V for the plume mode (10 A, 10 sccm) and increased slightly from 16 to 18 V for the spot mode (20 A, 30 sccm). The probe measurement analysis revealed that the IEDF peak energies are the same as, or exceed, the plasma potential and have a qualitative correlation with the electron temperature spatial distribution.</p>

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.

<p>In order to improve ion engine's performance, the neutralization performance with two field emission cathodes in an ion engine is investigated. The neutralization performance is evaluated by the potential difference between cathode and ground, using a 100μN class ion thruster developed at Kyushu University and two 50×50mm² field emission cathodes with carbon nanotube emitter. The potential difference between cathode and ground is not only determined by cathode electron supply capacity and the position of the cathodes but also foot print of the neutralizers, it would be due to the space charge limitation. That is, the potential difference between cathode and ground is improved with increase in total emission current, and that with a single field emission cathode at emission current of 6mA is -20V, on the contrary, that with two field emission cathodes is -12V. </p>

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.

Multi-pole magnetosphere has complex magnetic fields and current distribution. We focused on the structure for the thrust performance improvement of the magnetic sail. Magnetic sail is a propulsion system using the solar winds and artificial magnetic fields on the coil mounted on spacecrafts. We made the multipole coil, it generates complex magnetic fields has multiple magnetic pole. We conducted experimental simulation of multipole magnetic sail using the multipole coil we made and the large size plasma plume generator consisted of three MPD arcjets. Then, we measured the magnetic fields deviation rate and the current distribution using probe method. Hence, we identified the magnetopause size as 340 mm. Moreover, we observed two stronger current regions, ring current region at 150 mm from the coil center and magnetopause current region at 270 mm from the coil center. Further, we evaluated the relationship of magnetospheric structure generated by a multi-pole coil and induced current.

© 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.

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 C2H4-O-2 and H-2-O-2 as propellants. Images with a strong self-luminescence 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. (c) 2020 The Combustion Institute. Published by Elsevier Inc. All rights reserved.