船木 一幸

平成30年度宇宙科学に関する室内実験シンポジウム (2019年2月28日-3月1日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)相模原キャンパス), 相模原市, 神奈川県資料番号: SA6000139049

平成30年度宇宙科学に関する室内実験シンポジウム (2019年2月28日-3月1日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)相模原キャンパス), 相模原市, 神奈川県資料番号: SA6000139036

平成30年度宇宙科学に関する室内実験シンポジウム (2019年2月28日-3月1日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)相模原キャンパス), 相模原市, 神奈川県資料番号: SA6000139035

平成29年度宇宙科学に関する室内実験シンポジウム (2018年2月26日-27日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)相模原キャンパス), 相模原市, 神奈川県資料番号: SA6000123021

<p>スペース重力波アンテナDECIGOに搭載されるスラスタには，推力雑音が0.1µN程度以下[0.1～10Hz帯]という性能が想定されている．そこで本研究ではスラスタの性能評価に向け，微小な推力雑音を測定する装置として捻じれ振り子を用いたスラストスタンドの開発を行った．またこれに伴いプロトタイプの装置の感度評価を行った．本講演では装置開発の現状と今後の展望を報告する．</p>

Study of a Rotating Detonation Engine (RDE) has been carried out in many research institutions. The RDE has the advantage of high efficiency, simple structure and short combustor length. Toward the practical use for rocket engines, the evaluation of thrust efficiency at vacuum condition is necessary. In addition, it’s necessary to find a cooling method to endure the high heat load by detonation combustion. In this study, ethylene - oxygen mixtures are used as propellant, and we applied C/C composite, which is a heat resistant material, to a RDE and carried out a long-duration combustion test at sea level. As a result, we succeeded in the long-duration rotating detonation engine combustion demonstration of 6- 10 s at sea level for the first time in the world as rocket engine use, and the maximum thrust and the maximum specific impulse were achieved 301 N and 144 s.

A rotating detonation turbine engine has simpler structure and higher thermal efficiency than a conventional gas turbine engine. In this study, we designed a rotating detonation turbine engine with single stage centrifugal compressor, combustion chamber, and single stage radial flow turbine which are placed on one side of rotor disk. We performed cold flow experiments and combustion experiments. As results of cold flow test, compressor can supply air 52 g/s. In combustion experiments, we supply pressured oxidizer and fuel from outside of engine. As a results, combustion propagation speed was 600 - 1300 m/s. this is about 25 – 45 % of Chapman – Jouget detonation velocity. In addition, rotational speed rises by 160 rpm during combustion.

The importance of treatment of the cathode boundary was investigated intended for future self-consistent simulation of cathode coupling effects. JAXA fully kinetic particle code was applied to the discharge simulation of a SPT-100-like Hall thruster (discharge voltage 300 V and mass flow rate 5 mg/s operation condition), and four different cases of cathode boundary treatment were tested. Agreement of thrust and discharge current within 10 % error compared to the measurement were achieved without using any anomalous diffusion models in the baseline case treating the cathode boundary empirically. The preliminary simulation results of the cases treating the cathode boundary suggested that the electron injection conditions (quasi-neutrality, energy, density) can have significant impact on the simulation result of the plasma properties and electron flow.

Theoretically, a Rotating Detonation Engine (RDE) has higher thermal efficiency than a conventional gas turbine engine. In addition, RDE has a potential to shroten the combustuion chamber volume and makes it possible to be a simple strucuture. In order to practicalize RDEs, the optimum propellant injection configurations is necessary for steady combustion and high engine performance because it has a significant role for filling sufficient mixing layers in the combustion chamber. However, CFD analysis are usually done with premixed propellant, whereas experimental investigations are limited in terms of injector configulations. In this work, the slight difference of propagation states of detonation waves due to an increase in the number of fuel holes was not observed. Also, we shows a transiton state of detonation waves with a narrow oxidizer slit, and discuss the phenomenon.

Rotating detonation engines (RDEs) can get thrust or work by keeping detonation waves propagating in circumferential direction. Our objective is to apply rotating detonation engines (RDEs) to aircraft engines since we expect that the aircraft such as airplanes is to fly longer range by more simple structure. In this study, we designed a rotating detonation turbine engine (RDTE) with a single stage centrifugal compressor and a single stage radial flow turbine. The outer diameter of RDTE is about 200 mm. The revolution speed is 110000 rpm and rated air mass flow rate was 0.610 kg/s. We made two-parallel-plane rotating detonation combustor. We performed hot test with ethylene and oxygen. We analyzed several combustion modes with streak pictures and pressure time history. We observed the number of detonation waves, quenching and re-ignition phenomena by streak pictures.

A rotating detonation engine is an engine using continuous detonation in an annular combustor, and make thrust. Because detonation wave propagates in supersonic and in very small region, the combustor can be shortened. Programs in the combustor are high pressure loss when propellant is injected and necessity of cooling due to high heat flux. However, the physics of detonation combustion in an annular combustor is still not clear. Especially combustion efficiency, influence of injector shape of a rotating detonation combustor are not reported a lot. This paper reports measurement result of combustor stagnation pressure and influence of injector on characteristic velocity efficiency. Then convergent-divergent nozzle was used to increase thrust efficiency. In addition we measured temperature of the combustor and estimated heat flux.

A pulse detonation thruster (PD thruster) is a kind of reaction control system for controlling the attitude of a spacecraft using high-precision short-pulse thrust by intermittent detonation wave. Using a liquid-purge method proposed by Matsuoka et al. (Combustion Science and Technology, Vol. 187, No. 5, pp. 747-764, 2015), a 3N-calss PD thruster demonstrator was constructed in which the total length of combustor was 491.5 mm, inner diameter was 10 mm, and the total weight was 1241 g. In the liquid-purge method (LIP method) residual burned gas in combustor is cooled and purged by the latent heat of injected liquid droplet. This purge technique can realize a pulse detonation cycle using only fuel and oxidizer. Therefore it is possible to operate the pulse detonation cycle at high frequency without a purge process and a purge gas system. The thrust measurement experiment using the PD thruster at an operation frequency of 50 Hz was carried out and the thruster performance was evaluated. In a condition that injection duration was 6 ms, a time-averaged thrust of 2.3 ± 0.1 N, a propellant-based specific impulse of 113 ± 11 s, and a thrust-to-engine weight ratio of 0.19 were obtained. This thrust-to-engine weight ratio was the same level of a conventional 1N-calss monopropellant thruster. Furthermore the impulse in each pulse detonation cycle was 20.8 ± 2.6 mNs and the impulse repeatability was ± 12.3%. Moreover, the centroid time that is defined as the delay from the input signal to the thrust was 7.9 ± 0.1 ms. Therefore it was confirmed that a PD thruster has high thruster performance, namely, the high impulse repeatability and the high thrust response.

After the successful launch on the world first spacecraft, Sputnik 1 by the former Soviet Union in 1957, 58 years has passed. In 1960, Pioneer 5 of the United States escaped the Earth's gravity at the first time, and since then many interplanetary explorers had set to sail interplanetary. However, even in the present day, interplanetary voyages are not still easy. First, interplanetary missions require large amounts of delta-V, and second, the opportunity to get to the destination opens only every synodic period with the destination celestial body. For example, the synodic period with Mars is about 2 years, which means the opportunity to get to Mars opens every 2 years. For such circumstances, this paper proposes a new type of low-thrust orbit design method, "Interplanetary Parking Method" that realizes "anytime" launch of deep-space explorers. The proposed interplanetary parking method enables to make an Earth return orbit with an arbitrary time-of-flight connecting to the minimum energy transfer orbit to a destination. While the time-of-flight of the transfer orbit is fixed, the Earth return orbit with the arbitrary time-of-flight virtually eliminates the severe launch window constraint in interplanetary missions. As application of the proposed method, the paper demonstrates dual launch trajectory design of explorers to different destinations i.e., Mars and Venus. The proposed method will widen the scope of opportunity for interplanetary missions.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. A laboratory model of a hollow cathode with a radiative heater was built and tested for discharge currents of 10-50 A. Plasma ignitions and also discharge operations were successfully demonstrated even if the heater was not wrapped with an insulator, which will lead to robustness against heater breakdown. The voltage-current characteristic indicates that mode transition occurred between 30 and 40 A for a mass flow rate of 30 sccm. To investigate the flow field, a Hybrid-PIC simulation was conducted for a mass flow rate of 30 sccm and a discharge current of 30 A. The result shows acceptable distributions of the electron density, electric potential, and electron temperature inside and outside the cathode. The keeper’s floating voltage was close to the experimental data, but was slightly higher than experimental data. Changing the parameters of the anomalous resistivity can adjust the keeper’s floating voltage, but it has strong impact on the plasma properties.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In this paper, we present external discharge plasma thruster (XPT), a prototype of erosion free low power Hall thruster which produces and sustains plasma discharge completely outside a cavity. Details of this novel Hall thruster design and preliminary test results are presented. The thrust and the anode specific impulse ranged from 0.5 to 17.4 mN, and from 108 to 1240 sec respectively at anode potentials of 100-150-200-250 V with anode mass flow rates of 0.48-0.95-1.43 mg/s. The anode efficiency ranged from 2.4 to 25.6 % at discharge powers from 11 to 412 W. Preliminary experiment results suggest that XPT has similar discharge characteristics with SPT and TAL. Performance of XPT is comparable to SPT and TAL at the same power level, and very stable operation (Δ<0.2) is possible over wide range of operational conditions.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Magnetoplasmadynamic (MPD) thruster is a candidate of next generation electric propulsion system for a mission that high thrust is required. We are developing the high power Self-field MPD arcjet of several hundred kW class using a numerical and thermal design tool. We are testing quasi-steady MPD thrusters as the validation tool for numerical tool. In order to validate the numerical simulation or the predict of sheath effect, the electrode temperature measurement is necessary flowing to a great impact on the numerical simulation. This paper reports the cathode temperature measurement of the quasi-steady state MPD thruster with argon propellant in 160-1940 kW range and the cathode temperature using the 2-color pyrometer that we recently developed. The cathode tip temperature is near 3000 K above 13 kA. We reveals the temperature distribution heating only the cathode tip peculiar to the quasi-steady state experiment in lower the theoretical critical current. In addition, the cathode temperature change is clarified of 1 ms quasi-steady operation.