國中 均

平成27年度宇宙輸送シンポジウム（2016年1月14日-15日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000050126レポート番号: STEP-2015-062

平成27年度宇宙輸送シンポジウム（2016年1月14日-15日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000050096レポート番号: STEP-2015-032

平成27年度宇宙輸送シンポジウム（2016年1月14日-15日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000050084レポート番号: STEP-2015-020

平成26年度宇宙輸送シンポジウム（2015年1月15日-16日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000036068レポート番号: STEP-2014-010

平成26年度宇宙輸送シンポジウム（2015年1月15日-16日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000036121レポート番号: STEP-2014-064

平成26年度宇宙輸送シンポジウム（2015年1月15日-16日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000036106レポート番号: STEP-2014-049

平成26年度宇宙輸送シンポジウム（2015年1月15日-16日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000036077レポート番号: STEP-2014-019

© 2015 American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Hayabusa2 is the second asteroid sample return mission by JAXA. The ion engine system (IES) for Hayabusa2 is based on that developed for Hayabusa with modifications necessary to improve durability, to slightly increase thrust, and to reflect on lessons learned from Hayabusa mission. Hayabusa2 will rendezvous with a near-earth asteroid 1999 JU3 and will take samples from its surfaces. More scientific instruments than Hayabusa including an impactor to make a crater and landers are on board thanks to the thrust enhancement of the IES. After 2.5 years of short development time, the spacecraft was launched from Tanegashima Space Center in Kagoshima Prefecture on-board an H-IIA rocket on December 3, 2014. The IES was quickly checked out in orbit and cruise operation by ion propulsion for the electric delta-V Earth gravity assist (EDVEGA) was almost completed in the first week of June 2014. Accumulated operating times for four ion thrusters are 594 h, 11 h, 54 h and 592 h, respectively.

第47回月・惑星シンポジウム (2014年8月4日-6日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり資料番号: SA6000033027

High power electric propulsion system is strongly required for future orbital space transportation. MPD (Magneto-Plasma-Dynamic) thrusters and DC (Direct Current) arcjets with hydrogen as a propellant are promising candidates for the missions because of their high performance and adaptability to high power operation. However, to use hydrogen for long term orbital missions, its storage in orbit is crucial issue to be considered. Firstly, we proposed a hydrogen storage and feed system for electric thrusters by applying our technologies derived from the liquid hydrogen launch vehicles. Secondly, we present R&amp D activities of hydrogen MPD thruster and DC arcjet, especially focusing on the improvement of their performance and durability. Then, development strategy of hydrogen electric thrusters is also discussed. Finally, advantages of hydrogen electric thruster were shown compared with conventional xenon thrusters through mission analyses of lunar orbit insertion and GTO-GEO transportation.

In order to improve the thrust force of the ECR Ion Thruster μ10, two kinds of experiments were conducted. At the first experiment, a quartz plate was set at a waveguide to prevent the plasma from situating in the waveguide. Though it was introduced to improve the transmittance of microwave from the waveguide to a discharge chamber, the beam current did not increase. Secondly, a design of a spacer which locates between magnet rings was changed in two ways, different thick spacers and an anode spacer to improve the ion production cost. In addition to a 7-mm spacer (original) 4, 10, and 13-mm spacers were manufactured to eliminate the possibility of producing ions which cannot be accelerated by a screen grid. The anode spacer had a role to decrease ions which collide with the spacer. As a result, the higher spacer did not improve the beam current, however, the anode spacer improve the beam current by 5%.

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000016113レポート番号: STEP-2013-040

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000016097レポート番号: STEP-2013-024

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000016080レポート番号: STEP-2013-007

平成25年度宇宙輸送シンポジウム（2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所（JAXA)(ISAS)）, 相模原市, 神奈川県資料番号: SA6000016079レポート番号: STEP-2013-006

Hayabusa2 is an asteroid sample return mission. It is the follow-on mission of Hayabusa, which was the first spacecraft in the world that brought back the surface material of an asteroid to the earth. Hayabusa2 is now almost ready to launch in 2014. The target asteroid of Hayabusa2 is (162173) 1999 JU3, which is a C-type asteroid. The scientific purpose is to study not only the formation and evolution of the solar system but also the organic matter and water, which existed in the early stage of the solar system. Since we have the experiences and heritages of Hayabusa, we modified the spacecraft in a lot of parts. Thus the spacecraft becomes much more robust and reliable with some new technological challenges. After the launch in 2014, it will arrive at the asteroid in the summer of 2018, stay there for one and half years, and it will come back to the earth at the end of 2020.

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF資料番号: AA0061856113レポート番号: STEP-2012-030

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県マイクロ波放電式イオンスラスタの内部現象解明のためには、プラズマ生成に直結するスラスタ内部のマイクロ波電場測定が非常に重要である。本研究では電気光学素子ファイバプローブ（EOプローブ）をイオンスラスタの電場測定に適用することを提案した。EOプローブは金属を含まず微小なため、内部状態に擾乱を与えず非破壊で放電室に挿入することが可能である。本研究では（1）EOプローブを用いたビーム加速下のプラズマ中電場測定技術の確立と（2）未解明現象である高流量時の推力低下メカニズムおよび推進剤投入位置による性能変化メカニズムの解明を目標に研究を行った。電場測定結果は、FDTD解析結果と一致した。またプローブ挿入によるスラスタへの影響は8% 以下、加速プラズマ下でのプローブ感度変動は10% 以下であった。現象解明では中心軸上およびECR領域の電場分布を測定し考察を行った。その結果、導波管投入の際、高流量時に導波管内でマイクロ波の反射が起こり、電子加熱が行われるECR領域の電場強度が低下することがわかった。また推進剤を放電室より投入することは導波管内で反射が起きる流量をより高流量側へシフトさせる効果があり、その結果ECR領域での電場強度が最大になる流量がより高流量側へシフトし、推力増強につながったことがわかった。形態: カラー図版あり形態: PDF資料番号: AA0061856123レポート番号: STEP-2012-040

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF資料番号: AA0061856092レポート番号: STEP-2012-009

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF資料番号: AA0061856091レポート番号: STEP-2012-008

平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF資料番号: AA0061856104レポート番号: STEP-2012-021

Very low earth orbit satellites enable researchers to find out about aeronomy, accurate gravity and magnetic field mapping, and high-resolution earth surveillance. They orbit the earth at an altitude of lower than 250 km, where the effect of atmospheric drag cannot be discounted. In order to use this orbit, some kind of propulsion for drag make-up is required and propellant mass increase proportionally to the mission time. The Air Breathing Ion Engine (ABIE) is a new type of electric propulsion system which can be used to compensate the drag of a satellite. In the ABIE propulsion system, the low density atmosphere surrounding the satellite is taken in and used as the propellant for the Electron Cyclotron Resonance (ECR) ion engines to reduce the required propellant mass. Therefore ABIE is a promising propulsion system for aerodynamic drag free missions longer than two years. Feasibility and performance of the ABIE depend on the compression ratio and an air intake efficiency. Generally, pressure of a discharge camber is lower than a propellant tank pressure in propulsion system, and the propellant flows to the reaction chambcr from the tank. In the case of ABIE, a static pressure of atmosphere which corresponds to tank pressure is lower than the discharge chamber pressure. The air intake is the most important component to realize the ABIE. The temperature of the atmosphere is from 700K to HOOK at 200km, which is sufficiently low compared with the orbital velocity of 8km/s. Therefore, it can be said it is a uniform and well collimated supersonic flow parallel to the orbital direction. Moreover, the density is thin enough and it is a free molecular flow. The air intake consists of a collimator section and a reflector section. The collimator section will be composed of gaps between concentric cylinders. This part does not intercept the entering neutral particles, and they impact the reflection part on the downstream side directly. However, the backflow from the discharge chamber to the upstream side through the collimator section cannot easily leak out, because it is thermalized to the same level of temperature as the chamber walls and it has a velocity in a random direction. We simulate the relation between the ABIE and the rarefied atmosphere on such a super low earth orbit in a vacuum chamber. We verified the pressure rise inside the air intake. Copyright © (2012) by the International Astronautical Federation.

In order to reveal the physical processes taking place within the ECR ion thruster "μ 10", internal plasma diagnosis is indispensable. However, the ability of metallic probes to access microwave plasmas biased at a high voltage is limited from the standpoints of the disturbance created in the electric field and electrical isolation. This paper firstly attempts to measure the electronic excitation temperature using an optical fiber with little disturbance. Secondly, coupling with this result, it goes on to successfully measure the axial density profiles of excited neutral xenon under ion beam acceleration by using a novel laser absorption spectroscopy system. The targets of this measurement were metastable Xe I 823.16 nm and non-metastable Xe I 828.16 nm. As a consequence, these two measurements confirmed the existence of electrons at the anti-node of microwaves in the waveguide, which was caused by a propellant injection from the waveguide. Therefore this paper concludes that it is important to avoid the attenuation of microwave propagation to the discharge chamber by the electrons in the waveguide. © 2012 by Ryudo Tsukizaki.

In order to reveal the internal phenomenon theoretically within the electron cyclotron resonance ion thruster μ 10, internal microwave electric field measurement is very important because it is closely related to plasma producing mechanism. We have established a technology of electric field measurement with an optical fiber sensor which uses an Electro-Optic crystal (EO probe). This technology enables electric field measurement in plasma source under beam acceleration without disturbing microwave electric field. In this study, first, validity of electric field measurement using the EO probe in the atmosphere was demonstrated by comparing experimental results with FDTD simulation. Then, we measured axial electric field distribution in the accelerated plasma. This experiments indicated that electric field distribution in the μ10 thruster was related to its beam current. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The JIEDI (JAXA's Ion Engine Development Initiative) tool has been developed as a numerical tool for the lifetime qualification of ion thruster's ion optics with high precision and accuracy. The numerical wear test from beginning-of-life to end-of-life (EOL) by the JIEDI tool was conducted for the ion optics of HAYABUSA's microwave ion thruster (μ10 engineering model (EM)). It becomes clear that the erosion profiles for high ion beam current hole are most important to estimate the EOL of μ10EM ion optics. The ion optics of μ10EM ion thruster encounters its EOL by structural failure of the decelerator grid, which is mainly caused by sputtering of ions and neutrals scattered by elastic collisions. The estimated lifetime of the ion optics is 545 khours at the longest. Through the numerical wear test for μ10EM ion optics, the process for lifetime estimation of ion optics by the JIEDI tool was demonstrated. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The cathode-less electron cyclotron resonance ion engines, μ10, propelled the Hayabusa asteroid explorer, launched in May 2003, which is focused on demonstrating the technology necessary for a sample return from an asteroid, using electric propulsion, optical navigation, material sampling in a zero gravity field, and direct re-entry from a heliocentric orbit. It rendezvoused with the asteroid Itokawa after a two-year deep space flight using the ion engines. Though it succeeded in landing on the asteroid on November 2005, the spacecraft was seriously damaged. This delayed Earth return in 2010 from the original plan in 2007. Reconstruction on the operational scheme using thrust vector control of ion engines, Xe cold gas jets and solar pressure torque made Hayabusa leave for Earth in April 2007. Although most of the neutralizers were degraded and unable to be used in fall of 2009, a combination of an ion source and its neighboring neutralizer has kept the orbit maneuver to Earth including a series of final trajectory correction maneuvers. Finally, the spacecraft decayed in atmosphere and only the reentry capsule was retrieved from the Australian outback on June 14th, 2010. For the round trip space odyssey between Earth and the asteroid, the ion engines served the total accumulated operational time 39,637 hour·unit, the powered spaceflight in 25,590 hours, delta-V of 2.2 km/s, total impulse of 1 MN·s and 47 kg Xenon propellant consumption. © 2011 IEEE.