西山 和孝

The study on the post- HINODE Solar Observation Mission has been started by members in the Solar physics community. One candidate of the mission targets on the observation of the high latitude region of the Sun, which requires the injection of the space observatory (spacecraft) into the orbit largely inclined with the ecliptic plane. Reported in this paper are the trajectory design results for this orbit transfer, which contains a sequential application of the Electric Propulsion Delta-V Earth Gravity Assist (EDVEGA) procedure.

We have been developing several types of flight sensors for spacecraft surface contamination. Solar-cell type sensors were developed for the M-V-2 launch vehicle and the Hayabusa probe to measure contamination caused by solid spin motors and ion engines, respectively. Lunar-A probe launch by the M-V-2 rocket was canceled, but the Hayabusa's sensors are providing a strange long term degradation trend independent from ion engine activities. Another type of sensors using quartz crystal microbalances (QCM) were developed for the M-V-5 launch vehicle. The QCMs did not show clear contaminant deposition at the event of spin motor firings, but detected some depositions at nose fairing opening. Recently, new compact QCMs for spacecraft surface contamination measurements and material erosion measurements has been under development. Some flight programs using the QCMs are under discussion.

A 20-cm diameter xenon ion thruster with electron cyclotron resonance (ECR) discharge generates 30 mN of thrust at a total electric power consumption of 1 kW for spacecraft propulsion by ejecting 1.1 keV ion beam. By optimizing the discharge chamber length, magnetic field and propellant flow injection, ion beam currents of 500 mA at a microwave power of 100 W had been obtained at a frequency of 4.25 GHz with SmCo magnets arranged on a flat discharge chamber. The performance was highly dependent on the propellant injection method that affects electron-heating process. Two-dimensional microwave E-field distributions inside the discharge chamber were experimentally investigated for the best and the worst injector layouts. Microwave power absorption coefficient was estimated using the E-field distributions with and without plasma discharge. The coefficient decreases as the microwave power decreases and electron density gets close to an ECR cutoff density in all cases. The worst injector layout showed larger reflection and smaller absorption coefficient even at small beam currents. In the best configuration, microwave reflection was sufficiently smaller than 10% and 70-90% of the microwave power launched into the discharge chamber was absorbed by plasma electrons. © 2008 Elsevier B.V. All rights reserved.

DECi-hertz Interferometer Gravitational wave Observatory (DECIGO) is the future Japanese space gravitational wave antenna. The goal of DECIGO is to detect gravitational waves from various kinds of sources mainly between 0.1 Hz and 10 Hz and thus to open a new window of observation for gravitational wave astronomy. DECIGO will consist of three drag-free spacecraft, 1000 km apart from each other, whose relative displacements are measured by a Fabry - Perot Michelson interferometer. We plan to launch DECIGO pathfinder first to demonstrate the technologies required to realize DECIGO and, if possible, to detect gravitational waves from our galaxy or nearby galaxies. © 2008 IOP Publishing Ltd.

© 2008 IOP Publishing Ltd. DECIGO pathfinder (DPF) is a milestone satellite mission for DECIGO (DECi-hertz Interferometer Gravitational wave Observatory) which is a future space gravitational wave antenna. DECIGO is expected to provide us fruitful insights into the universe, in particular about dark energy, a formation mechanism of supermassive black holes, and the inflation of the universe. Since DECIGO will be an extremely large mission which will formed by three drag-free spacecraft with 1000m separation, it is significant to gain the technical feasibility of DECIGO before its planned launch in 2024. Thus, we are planning to launch two milestone missions: DPF and pre-DECIGO. The conceptual design and current status of the first milestone mission, DPF, are reviewed in this article.

In order to adapt to a wide variety of the space flights, such as small geosynchronous satellites and deep space explorers, feasibility study of performance enhancement options for the microwave discharge ion thruster μ10 is underway. Authors are considering the following five options: 1. Lower insertion loss DC blocks; 2. Direct monopole antenna insertion to the discharge chamber without using a circular waveguide part; 3. Optimization of gas injector layout which was originally located deep in the waveguide; 4. Additional magnet rings aiming ion loss reduction to the side wall of the discharge chamber; 5. New ion optics consists of a thinner screen grid and a smaller-hole accelerator grid. Not all but most of them have already been tested and reported in this article. The original models for Hayabusa asteroid explorer generated 8 mN at maximum. Larger thrust generation was impossible even if propellant flow rates and microwave powers were increased. It turned out to be feasible to increase the maximum thrust to a range of 10 - 11 mN with above mentioned options by supplying more flow rates and/or more microwave powers. © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

The "μ20" ion thruster, which successor to a microwave discharge ion thruster "μ10" mounted on "Hayabusa" asteroid explorer, is under research and development. A substantial improvement in performance of the "μ20" thruster has been achieved by equipping with unique magnet geometry and gas injector layout. The next goal has been an improvement in the propellant utilization efficiency. In order to increase the propellant utilization efficiency, new accelerator grid that aperture diameter were much reduced so aggressively from the conventional accelerator grid was designed and fabricated. Excessive accelerator grid impingement currents, however, were expected. In order to reduce accelerator current and decide the optimum aperture diameter distribution, this grid was ion machined for over 1,000hours. As a result of the ion machining process, accelerator current was like it had been before. The propellant utilization efficiency increased from 66.7% to 82.4%.

An ion engine with 10,000sec-class high specific impulse is expected as primary propulsion for interplanetary space mission with extremely high delta-V, and needs high voltage for ion acceleration over 10kV. The microwave discharge ion source without solid electrodes can supply sophisticated technologies and simple composition on electrical isolation for such a high voltage. New electro-static grid made of carbon-carbon composite material was fabricated based on the numerical simulation of “igx” code. The “μ10HIsp” ion engine combining the ECR ion source, microwave discharge neutralizer, DC blocks, propellant isolators, carbon-carbon composite grid and so on generated successfully a plasma beam with 10,000sec specific impulse using 15kV acceleration voltage.

Radiated electric field emissions from the prototype model of the ion engine system of the asteroid explorer Hayabusa (MUSES-C) were measured in approximate accordance to MIL-STD-461C. The typical noise level exceeded the narrowband specification at frequencies less than 5 MHz. The microwave discharge neutralizer generates broadband noise and narrowband oscillations that have a fundamental frequency of about 160 kHz and are accompanied by its harmonics up to the fifth. Leakage of 4.25 GHz microwaves for plasma production and its second harmonic were 65 dB and 35 dB above specifications, respectively. The X-band receiver onboard Hayabusa measured the noise from the ion engine system at the uplink frequency of 7.16 GHz through a horn antenna. This susceptibility test showed that the microwave discharge ion thruster is unlikely to interfere with deep space microwave communication.

The μ10 cathode-less electron cyclotron resonance ion engines made the Hayabusa spacecraft rendezvous with the asteroid Itokawa in 2005. Though the spacecraft was seriously damaged after the successful soft-landing and lift-off, the xenon cold gas jets from the ion engines rescued the Hayabusa. New attitude stabilization method using a single reaction wheel, the ion beam jets, and the solar pressure was established and enabled the homeward journey aiming the Earth return on 2010. The total accumulated operational time of the ion engines reaches 28,000 hours at the end of May 2007.

This paper describe that the instrumental improvement for application of the ECR Ion thruster to upper atmosphere observation system. It is developed the study of remote sensing method of neutral particles in the upper atmosphere using artificial energetic neutral atoms (ENAs) in the laboratory. It is generally known that atomic oxygen is dominant neutral gas in the upper atmosphere 200km and 700km altitude. And we need to estimate the atomic oxygen density by krypton ion beam in the laboratory. We studied the atomic oxygen source for ground experimental demonstration.

The cathode-less microwave discharge ion engine generates plasmas in an ion source and a neutralizer in the method of the collision ionization by electron impact ionization, in which thermal electrons are energized by microwave electronic field without emanating from a cathode. The plasma generation mechanism on the microwave discharge ion source was modeled and formulated based on the global O-dimensional Brophy's models conserving particles and energy into the discharge and out of the plasma in the form of charged particles to the walls, beam and plasma radiation. In addition, the ion current of the wall loss in the ion source was measured in the laboratory so that the total amount of the ion production current and the ratio of the ion current contributing to the ion beam are determined.

The electron cyclotron resonance ion engine has long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizes Four μ 10s, each generating a thrust of 8 mN, specific impulse of 3200 s, and consuming 350 W of electric power, propelled the Hayabusa asteroid explorer launched on May 2003. After vacuum exposure and several baking runs to reduce residual gas, the ion engine system established continuous acceleration. Electric propelled delta-V Earth gravity assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit In 2005, Hayabusa, using solar electric propulsion, managed to successfully cover the solar distance between 0.86 and 1.7 AU. It rendezvoused with, landed on, and lifted off from the asteroid Itokawa. During the 2-year flight, the ion engine system generated a delta-V of 1400 m/s while consuming 22 kg of xenon propellent and operating for 25,800 h. Copyright © 2007 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Microwave discharge ion engines "μ10"are dedicated to the main propulsion on the HAYABUSA asteroid explorer. In a development program, various tests and assessments were conducted on the ion engines and the spacecraft. They include endurance tests, an electromagnetic interference susceptibility test, an interference test between the plasma and communication microwave, a beam exhaust test on the spacecraft, assessments on the plasma interference with a solar array, and so on. The spacecraft was launched in deep space by the M-V rocket in May 2003. After vacuum exposure and several runs of baking for reduction of residual gas, the ion engine system established continuous acceleration of the spacecraft toward the asteroid ITOKAWA. The spacecraft passed through a perihelion of 0.86 astronomical unit (AU) in February 2004 and an aphelion of 1.7 AU in February 2005, becoming the first solar electric propulsion system to travel this far toward and away from the Sun. The HAYABUSA succeeded in rendezvousing with the target asteroid in September 2005. © 2006 IEEE.

This paper presents the development status of a 20-cm diameter microwave discharge ion thruster which generates 25 ∼ 30 mN of thrust with an electric power of 1 kW. By optimizing the discharge chamber length, magnetic field and propellant flow injection, ion currents of up to 530 mA at a net microwave power of 100 W had been obtained at a frequency of 4.25 GHz with a coaxial cable to circular waveguide transformer. Almost the same performance has been achieved with a new antenna directly inserted into the discharge chamber, which removes the need for a long circular waveguide. Higher frequencies up to 5.8 GHz and stronger magnets have been tested for performance improvement and turned out to be very promising. © 2005 Elsevier B.V. All rights reserved.

Sounding rockets and satellites have been dedicated to measuring "natural" energetic neutral atoms (ENAs) in order to investigate ring current ions and precipitating ions. We studied the remote sensing method on the neutral particles in the upper atmosphere using "artificial" ENA in the laboratory. The measurement scheme is described below. A space-borne plasma generator emits beam ions, which experience charge exchange collision with neighboring neutral particles to transform into artificial ENAs. Once ENAs are produced, they draw ballistic trajectories holding information, such as density and composition. In the upper atmosphere between 200 km and 700 km altitude, the primary constituent is atomic oxygen, to which krypton ion beams are sensitive. Measurement of artificial ENA at a remote region reveals global distribution of the neutral particles with high resolutions in time and space. One of the feasibility studies demonstrated the density measurement of O2 out of the mixture of O2 and N2 using a xenon ion beam in the laboratory. Copyright 2006 by the American Geophysical Union.

マイクロ波放電型イオンエンジンの開発に伴い、そのイオン源の解析理論の構築が必要となっている。本研究ではその第1段階として、現在研究開発が進められているμ20のイオン源に対してプローブ計測を行った。μ20のイオン源には、特徴である放電室内の磁場形状や推進剤の供給方法から、周方向の分布が存在する事が予想される。今回のプローブ計測により、内側の磁石列間において周方向の分布が確認された。電子の受けるΔBドリフトの方向に沿って、電子温度の増加とプラズマ密度の減少が見られた。また、推進剤の供給位置を変えた場合に電子温度・プラズマ密度に違いがあることがわかった。これらの結果より、放電室形状を最適にすることでμ20の性能向上が期待できる。資料番号: AA0049213012レポート番号: JAXA-SP-05-020

The microwave discharge ion engine “μ10” has a long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. A single μ10 generates a thrust of 8 mN, specific impulse of 3,200 seconds, and consuming 350 W of electric power. Four μ10s propelled the “Hayabusa” asteroid explorer, launched on May 2003, combining a new orbit change scheme “Delta-V Earth Gravity Assist”. In 2005, Hayabusa, using solar electric propulsion, managed to successfully cover the distance between 0.86 AU and 1.7 AU from Sun in the solar system, as well as rendezvous with, land on, and lift off from the asteroid Itokawa. During the 2-year flight, the ion engine system generated a delta-V of 1,400 m/s while consuming 22 kg of xenon propellant and operating for 25,800 hours. Hayabusa solved failures of reaction wheels and chemical thrusters by means of xenon cold gas jet from the ion engine system and aims to return to Earth in 2010.

The electron cyclotron resonance ion engine, "μ10,"" has a long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four μ10, each generating a thrust of 8 mN, specific impulse of 3,200 seconds, and consuming 350 W of electric power, propel the "HAYABUSA" asteroid explorer that was launched on May 2003. After vacuum exposure and several runs of bailing to reduce residual gas, the ion engine system established continuous acceleration. Delta-V Earth Gravity Assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit. In 2005, HAYABUSA, using solar electric propulsion, managed to successfully cover the distance between 0.86 AU and 1.7 AU in the solar system, as well as rendezvous with, land on, and lift off from the asteroid Itokawa. During the 3-year flight, the ion engine system generated a delta-V of 1,400 m/s while consuming 22 kg of xenon propellant and operating for 25,900 hours.

A new electron cyclotron resonance (ECR) ion source "μ6" has been developed. It is a scaled down model of the μ-series microwave discharge ion thrusters. The ECR condition at 4.25 GHz is obtained in a multi-cusp arrangement using a set of two SmCo magnet rings on a base plate of a cylindrical ionization chamber with a plasma volume 6 cm in diameter and 2 cm in length. The microwave is launched with a quarter-wavelength straight electric antenna and an N-type connector. The ionizer length and chamber diameter were experimentally optimized. The xenon ion current to a negatively biased metal grid has been measured to estimate ion beam current that could be extracted with ion optics. In order to achieve the same level of ion current density at the same discharge pressure range as larger thrusters μ10 and μ20, larger discharge power density was necessary. Although ion production cost of 1000-2000 W/A is much worse than larger thrusters, the μ6 operation in a thrust range of 1-3 mN is still attractive for micro-thrusting applications. Sudden increase of the ion current was observed when the pressure and the microwave power density were four times larger than those of larger thrusters. A current density of 6-10 mA/cm2 has been reached with an incident microwave power of 70-140 W and a chamber pressure of 0.06-0.13 Pa. This implies that the μ6 with properly designed ion optics would generate a 14 mN thrust at an ion production cost of 400 W/A and propellant utilization efficiency of 70%. The thrust density will be as large as that of electron-bombardment or radio-frequency ion thrusters. The most unique feature of the μ6 is its very wide operating range.

In order to adapt to a wide variety of space flights as well as to advance the technology of microwave discharge ion engines, the "μ20μ is under research and development. The μ20 is a 20-cm diameter class ion engine and was designed for 30 mN/kW thrust-to-power ratio. The grid assembly made of high stiffness carbon-carbon composite material was machined and passed the vibration test. Magnetic field and propellant injection method of the ion source has been optimized. The performance is highly dependent on the propellant injection method that affects electron-heating process. The ion source can generate 530 mA ion current while consuming 100 W of 4.25 GHz microwave power. Total system performance has been estimated with the use of individual performance data of the ion source and the neutralizer. Thrust-to-power ratio can be increased up to about 30 mN/kW with thinner screen grid and diffusing gas injection. Total power consumption was 1080 W at full throttle operation, under which the thrust was 32 mN and the specific impulse was 2540 s, respectively.

In order to adapt to a wide variety of space flights as well as to advance the technology of microwave discharge ion engines, the "μ20" is under research and development. The μ20 is a 20-cm diameter class ion engine and was designed for 30 mN/kW thrust-to-power ratio. The grid assembly made of high stiffness carbon-carbon composite material was machined and passed the vibration test. Magnetic field and propellant injection method of the ion source has been optimized. The performance is highly dependent on the propellant injection method that affects electron-heating process. The ion source can generate 530 mA ion current while consuming 100 W of 4.25 GHz microwave power. Total system performance has been estimated with the use of individual performance data of the ion source and the neutraliser. Thrust-to-power ratio can be increased up to about 30 mN/kW with thinner screen grid and diffusing gas injection. Total power consumption was 1080 W at full throttle operation, under which the thrust was 32 mN and the specific impulse was 2540 s, respectively.

資料番号: AA0049206001レポート番号: JAXA-SP-05-001E

Four units of the microwave discharge ion engine "μ10", which generates 8mN thrust and 3,200sec specific impulse with 350W electric power, are devoted to the main propulsion on the "HAYABUSA" asteroid explorer. The spacecraft was input in the deep space by M-V rocket on May 9, 2003. After vacuum exposure and several runs of baking for reduction of residual gas the ion engine system established the continuous acceleration of the spacecraft. In the first year the spacecraft changed its eccentricity of the orbit by the maneuver of the on engines m a one-year Earth-synchronous orbit. On May 19, 2004 just one year later after the launch, the spacecraft steered its course by means of the Earth swing-by and got on the transfer orbit toward the target asteroid, even on which the ion engines accelerated continuously. The spacecraft passed by the aphelion 1.7 astronautical unit on February 18, 2005 so that the microwave discharge ion engine μ10s became the electric propulsion to arrive at the furthest space from Sun. On the beginning of May 2005, the ion engine system reaches the total number of space operational time 23,000hour during two years, and the spacecraft approaches the asteroid at 400,000km remaining distance. The HAYABUSA spacecraft is about to rendezvous with the asteroid "Itokawa" on September 2005. Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A completely new solar electric propulsion concept, the Air Breathing Ion Engine (ABIE), has been proposed for spacecraft drag makeup at very low altitudes, ranging from 150 to 200 km. ABIE scoops up neutral atoms and molecules traveling at an orbital velocity of approximately 8 km/s, ionizes them by means of an electron cyclotron resonance plasma source that is efficient in a wide range of low gas pressures, and accelerates the ionized air particles electrostatically to exhaust velocities larger than 100 km/s. The key technology of this thruster is the design of a propellant inlet which allows the incoming flow to enter the discharge chamber, yet it prevents the thermalized gas from escaping upstream. In this system, an air-breathing-type neutralizer may also be employed, in which case the need to carry on-board xenon propellant is eliminated and results in gains in payload mass if the mission duration is longer than 2 years. This technology should give researchers access to a part of the atmosphere that is currently very difficult to measure and is thus called the "ignorosphere." Promising applications other than aeronomy include science missions involving accurate gravity and magnetic field mapping, and high-resolution Earth surveillance.

In recent years, concern has been raised about the interference phenomena (chemical reaction, sputtering etc.) between ionospheric plasma and high-voltage space systems such as the International Space Station and so on. In order to solve the physical mechanisms of these phenomena and establish the prevention technology, it is important to accumulate experimental data based on ground simulation tests. Therefore we have developed a cathode-less microwave discharge oxygen ion source for an ionospheric plasma simulator. The oxygen ion source for ionospheric plasma simulation demands plasma density more than 10¹²m^{- 3} and high durability (our target is more than 70hours). This ion source is capable of generating oxygen plasma of the order of 10¹⁴m^{- 3} for more than 100hours with 25W microwave power and 3sccm oxygen gas. In order to investigate the operational performance of the oxygen ion source, we conduct various measurement tests such as plasma diagnosis using a single probe, ionic mass spectroscopy by a Quadrupole Mass Spectrometer (QMS), atomic oxygen flux measurement using a silver coated QCM (Quartz Crystal deposition Monitor) and 70hours operating test etc. This paper reports the experimental results of each measurement tests and assesses performance as an ion source for ionospheric plasma simulator.

The microwave discharge ion engine generates plasmas of the main ion source as well as the neutralizer using 4GHz microwave without discharge electrodes and hollow cathodes, so that long life and durability against oxygen and air are expected. MUSES-C “HAYABUSA” spacecraft installing four microwave discharge ion engines was launched into deep space by M-V rocket on May 9, 2003. After vacuum exposure and several runs of baking for reduction of residual gas the ion engine system established the continuous acceleration of the spacecraft toward an asteroid. The Doppler shift measurement of the communication microwave revealed the performance of ion engines, which is 8mN thrust force for a single unit with 3,200sec specific impulse at 23mN/kW thrust power ratio. At the beginning of December 2003 the accumulated operational time exceeded 7,000 hours and units.

資料番号: AA0046978005レポート番号: JAXA-SP-03-001

The microwave discharge ion engine generates plasmas of both the main ion source and the neutralizer using 4GHz microwave without discharge electrodes and hollow cathodes, so that long life and durability against oxygen and air are expected. The MUSES-C "HAYABUSA" asteroid explorer installing four microwave discharge ion engines "μ10"s was input in the deep space by M-V rocket No.5 on May 9, 2003. After vacuum exposure and several runs of baking for reduction of residual gas the ion engine system established the continuous acceleration of the spacecraft toward the asteroid "ITOKAWA". The Doppler shift measurement of the communication microwave revealed the performance of ion engines, which is 8mN thrust force for a single unit with 3,200sec specific impulse at 23mN/kW thrust power ratio. At the end of March 2004 the ion engine system accumulated the operational time 10,600 hour & unit and consumed 12kg Propellant, and then input the spacecraft on the orbit to encounter Earth. HAYABUSA will execute the Earth swing-by on May 19 and arrive at the asteroid in 2005 and return to Earth again in 2007. © 2004 by the American Institute of Aeronautics and Astronautics, Inc.

The MUSES-C spacecraft was launched on May 9, 2003 and has been propelled by microwave discharge ion engines. The mission duration is four years in total and most of it is the cruising phase driven by ion propulsion. After one-year acceleration it succeeded the Earth swing-by on May 19, 2004 and still on the way to the asteroid "ITOKAWA". Active thrust vector control contributes to keep the reaction wheels within an appropriate rotation rate. The hydrazine thruster firing tunes off the ion engines before it and restart them again after it. Depending on the operation of ion engines the heater power is adjusted. These operations are autonomously controlled on the spacecraft without supervisions from Earth. A ground operation system that integrates all the process required for ion engine operations such as orbit synthesis, orbit determination, operation scheduling, command generation and telemetry data analysis has been developed. Automation of on-board and ground operations saves the operation cost and time, and makes the operation stable and reliable.

A completely new solar electric propulsion concept, the Air Breathing Ion Engine (ABIE), has been proposed for spacecraft drag makeup at very low altitudes, ranging from 150 to 200 km. ABIE scoops up neutral atoms and molecules traveling at an orbital velocity of approximately 8 km/s, ionizes them by means of an electron cyclotron resonance plasma source that is efficient in a wide range of low gas pressures, and accelerates the ionized air particles electrostatically to exhaust velocities larger than 30 km/s. The key technology of this thruster is the design of a propellant inlet which allows the incoming flow to enter the discharge chamber, yet it prevents the thermalized gas from escaping upstream. In this system, an air-breathing-type neutralizer may also be employed, in which case the need to carry on-board xenon propellant is eliminated and results in gains in payload mass of approximately 200 kg per mission-year, as estimated for a spacecraft cross section of 1.5 m2 orbiting at an altitude of 150 km. This technology should give researchers access to a part of the atmosphere that is currently very difficult to measure and is thus called the "ignorosphere." Promising applications other than aeronomy include science missions involving accurate gravity and magnetic field mapping, and high-resolution Earth surveillance. Copyright © 2003 by the International Aeronautical Federation. All rights reserved.