國中 均

An understanding of the plasma physics inside a microwave discharge cathode is key to extending the lifetime of microwave ion thruster systems. However, probes can only measure the plume region due to their low spatial resolution and electromagnetic disturbance. In this study, we develop a microwave discharge-based cathode with a small optical window in the discharge chamber that provides visual access to the cathode interior. The cathode has the same anode currents as those of a flight model in the diode mode (anode voltage error is within 7%). Laser-induced fluorescence spectroscopy is applied to the cathode. The axial and radial ion velocity distribution functions (IVDFs) in the plume region and the axial IVDFs inside the cathode are measured. The measured functions, which represent the number density of Xe II (P-3(2))6p[3](5/2), are compared to a previously reported number density of Xe II measured by an electrostatic probe in the plume region. The functions exhibit multimodal characteristics. Theoretical models based on the measured current oscillation support these characteristics.

This paper presents the design methodology and performance testing of an additively manufactured resistojet operating on hydrogen as propellant. Additive manufacturing allows to produce complex monolithic resistors, resulting in reliable high efficiency thrusters. The concept, to be used in combination with advanced cryogenic storage technologies, is proposed for short time and high specific impulse orbit transfers. The simplified two-dimensional thermal design approach adopted is discussed, and its application to the engineering of the resistor is shown for both Inconel 718 and tungsten. The paper reports the performance testing of the proof-of-concept version of the thruster, manufactured in Inconel 718. Experiments on hydrogen show near ideal performance, demonstrating a peak thermal efficiency of 96%. The thruster proves the validity of the design methodology proposed, and the feasibility of the approach to develop monolithic additively manufactured hydrogen resistojets as main propulsion units.

© 2019 IAA Japan's second asteroid explorer Hayabusa2 was successfully launched on Dec 3, 2014, to return a sample from asteroid 162173 Ryugu by 2020. Four xenon ion thrusters based on electron cyclotron resonance discharge propelled the spacecraft for 547 h during its first year in space. Hayabusa2 completed an Earth gravity assist on Dec 3, 2015, followed by 798 and 2593 h of ion thruster operation, called the first and second transfer phases of delta-v, respectively. The third transfer phase of delta-v was conducted from Jan 10, 2018, to Jun 6, 2018, in which the final 2475-h ion thruster operation was executed before the rendezvous with Ryugu. The cumulative operating times for the four ion thrusters are 6,450, 11, 5,193, and 6418 h. This paper summarizes the 6515-h powered flight by the ion engine system, which produced 1015 m/s delta-v, in terms of thruster performance change, roll torques generated by various combinations of ion thrusters, and spacecraft surface erosion history measured by two quartz crystal microbalances located near the thrusters. In parallel with the space flight operation, an engineering model of the microwave discharge neutralizer has been under long-duration testing on the ground since 2012. It has accumulated 55,170 h of diode-mode operation as of Mar 15, 2019.

The microwave cathode was developed as a neutralizer for the microwave ion thrusters of the Japanese asteroid explorers Hayabusa and Hayabusa2. Since it emits hundreds of mA of electron current, ion currents collect at the wall of the cathode, which causes fatal destruction due to sputtering. In an effort to reduce the sputtering voltage, this study investigates the effect of the strength of the magnetic field at the nozzle on the anode voltage. Firstly, a magnetic field is applied at the nozzle by a coli. Using the coil, decreasing the magnetic field intensity increases the electron density at the exit of the nozzle. It is presumed that the applied magnetic field facilitates the detachment of magnetic lines by the electrons inside the microwave cathode, resulting in a reduction of the anode voltage. By weakening the nozzle magnetic field, trapped electrons are reduced and the transportability to the outside is improved. Secondly, to realize the same magnetic field intensity achieved in the first experiment without any additional power consumption, the author proposes the use of a magnetic shield. The magnetic shield reduces the anode voltage from 37 V to 32 V at 180 mA, the nominal current of the flight model. Since the sputtering rate exponentially increases with the anode voltage, reducing the anode voltage through these techniques is effective in increasing the lifetime of the cathode.

© 2019 Author(s). Magnetoplasmadynamic (MPD) thrusters are operated in a quasisteady state with about 1.0 ms pulse created by a pulse forming network (PFN). However, there is still no precedent to verify the operation time quantitatively. The nonsteady region of the pulse can lead to an error of the thrust performance against that of steady state operation. In addition, the propellant gas outside the discharge chamber can be consumed since the exhaust velocity exceeds the estimated velocity. This paper shows the first step in quantitative evaluation of the quasisteadiness of an MPD thruster operation. First, we developed a new power supply that outputs a flat-topped and less nonsteady region pulse with a variable pulse width. Compared with that of a PFN, the nonsteady region "tr + tf" decreased from 0.532 to 0.110 ms. By implementing the circuit shorter and adjusting the gate resistance, the surge voltage in the experiment was suppressed to 309 V, which is less than 2% error of that in the PSIM simulation, 305 V. Second, we operated an MPD thruster using the new power supply for discharge and the external magnetic field. As a result, we obtained operation time characteristics by sweeping the operation time from 0.3 to 5.0 ms. The current waveforms are in a range of 620 ± 70 A. We confirmed the consistency of the thrust, 0.32 ± 0.03 N. From the correlation between the input energy and the impulse, it is possible to discuss quasisteadiness of the MPD thruster operation using the determination coefficient and the offset.

© 2019 IAA To improve the performance of the 10-cm-class microwave discharge ion thruster μ10 for use in future deep space exploration missions planned by the Japan Aerospace Exploration Agency (JAXA), a new discharge chamber was designed, and its performance was tested. The maximum beam current in the new discharge chamber geometry was 16% higher than that in the original geometry, which was used in the Hayabusa 2 space explorer, under the same discharge power. To investigate the reason for this performance improvement, the multi-charged ion ratio in the plume, the beam current density profiles, and the ion current in the discharge chamber were measured by probes. It was found that the multi-charged ion efficiency and the beam divergence efficiency in the redesigned configuration were not significantly different from those in the Hayabusa 2 configuration. This shows that the increase in the ion beam current enhances the thrust. In addition, it was confirmed that the total ion current inside the new discharge chamber is higher than that in the Hayabusa 2 configuration. The ion extraction efficiency, however, was lower than that in the Hayabusa 2 configuration. This suggests that the increase in the total ion current per unit of incident microwave power is the cause of the performance improvement. In the redesigned configuration, the thrust is 12.0 mN, the specific impulse is 3122 s, the discharge loss is 162 W/A, and the propulsion efficiency is 39.6% at the peak performance point.

This paper presents the first laboratory-based study to measure the azimuthal velocities of ions in the beam of a gridded ion thruster. Through the operation of gridded ion thrusters in space, it has been confirmed that these thrusters cause an unexpected roll torque about the ion beam axis. To reveal the physical mechanism that produces this torque, laser-induced fluorescence spectroscopy has been applied to a microwave ion thruster that was installed in Japanese asteroid probes. This technique can be used to measure the azimuthal velocity by estimating the Doppler shift of the Xe II 5p 4()6p 5/2 to Xe II 5p 4()6s 3/2 transition at 834.659 nm. The measurement was conducted without a neutralizer cathode to avoid the possibility of the cathode affecting the trajectory of the ion beam. The measured velocity functions are the sum of the spectra of the high velocity beam ions and those of charge exchange ions. By deconvolving these spectra, the azimuthal velocities were successfully measured and were found to range from -700 to 620 m s-1 with an accuracy of 25%. The measured azimuthal velocity profile was accurately reproduced by the simulated velocity profile obtained using a model, which includes the effects of the maximum possible misalignment of the accelerator grid with respect to the screen grid and the Lorentz force produced by the magnetic field leaked from the discharge chamber. A roll torque of 0.5 ±0.1 μN m about the thrust axis was calculated from the velocity profile, which is lower than that reported in flight data, but additional mechanisms are suggested to explain this discrepancy.

We report the experimental and simulated azimuthal ion velocities of a gridded ion thruster, which generates a roll torque around the thrust axis. Laser-induced fluorescence spectroscopy was applied to two microwave ion thrusters with opposite magnetic polarities. A comparison of the measured results revealed a net misalignment of the grid optics and showed that the ions are continuously accelerated from inside the discharge chamber towards a direction downstream of the grid optics. To investigate the effect of the electromagnetic field, the authors conducted a two-dimensional particle-in-cell Monte Carlo collision (2D-PIC-MCC) numerical simulation. The numerical simulation agrees with the measurements and reveals that the ions are azimuthally accelerated by a gradient B drift, curvature drift, E x B drift and the Lorentz force. The reproduced roll torque is 3.1 +/- 2.3 mu Nm and arises due to the mechanical tolerance of the grid optics. The roll torque shows good agreement with the result observed in the space operation. Therefore, the roll torque can be predicted by using our experiment and simulation.

The microwave power absorption efficiency of the mu 10 ECR ion thruster, utilized in the Japanese asteroid explorers Hayabusa and Hayabusa2, is investigated in order to allow performance measurement and provide information for its improvement. A model detailing the local electron behavior in a real ECR plasma discharge, based the magnetic field characteristics, is presented. Three methods to evaluate the microwave power absorption efficiency are proposed: an estimation based on the chamber geometry and magnetic field characteristics, a measuremen based on performance parameters and a measurement performed with Langmuir probes. The equations used for each method are analytically derived. The local electron behavior model is confirmed with a Langmuir probe experiment. Measurement of the microwave power absorption efficiency is performed with the two independent methods proposed. Results from the two experiments show good agreement with each other and with the theory. Finally, a diffusion model explaining the different electron temperature distributions observed in the chamber is proposed. The model and experiments clarify the physics behind previously observed performance variations and give valuable hints for future chamber improvement.

© 2018 Author(s). This paper presents calibration devices and methods for the measurement of electric thruster performance parameters using a seesaw-type thrust stand to measure the mass loss of solid propellant in a vacuum. In previous studies, impact hammers and electrostatic combs have been manufactured for the calibration of the thrust and impulse using seesaw-type thrust stands. However, these conventional devices rely on self-calibration, which means that the input delivered by the device in unknown, and must undergo a calibration process themselves. In this paper, the manufactured calibration devices successfully reproduced known impulses, thrusts, and mass losses in a vacuum. By reproducing known inputs based on known masses, the proposed calibration devices can omit the conventionally required self-calibration process. The calibration results showed linear relations between outputs and known inputs and agreed with the theoretical values to within an error of 10%. Additionally, the uncertainties of all known inputs were less than 1.5%. On the basis of these results, the average thrust, impulse, and mass loss were measured using a calibrated thrust stand for the first time. The cumulative impulses obtained from the measured impulse and average thrust agreed with each other to within an error of 5%. The error of the measured mass loss per 1000 shots with respect to the actual mass loss measured using an electronic balance ranged from 1% to 17%.

The Japan Aerospace Exploration Agency launched the asteroid sample return spacecraft "Hayabusa2" on December 3, 2014. Hayabusa2 will reach the C-type asteroid 1999 JU3 in 2018, and return back to the Earth in 2020. Sample collections from three sites, four surface rovers deployment and a 4 MJ-class kinetic impact crater generation are planned in the 1.5 years of the asteroid-proximity operation. The mission objective of Hayabusa2 has three aspects, science, engineering and exploration, all of which would be expanded by the successful round-trip journey. This paper describes the outline of the Hayabusa2 mission and the current flight status after the seven month of the interplanetary cruise. (C) 2016 IAA. Published by Elsevier Ltd. All rights reserved.

<p>Conventionally, neutralizers in ion thruster systems do not generate thrust force. Hence, the power consumption of a neutralizer negatively affects the thrust efficiency of the ion thruster system. Therefore, in this paper, a negative ion source that generates thrust force as well as neutralizes the positive ion beam was newly developed using fullerene as a propellant so as to realize a more efficient ion thruster system. To develop the negative ion source, two measurements were conducted. The first measurement was an E &times; B probe to identify the species of positive and negative ions. The second measurement was a magnetically filtered Faraday probe to measure quantitatively the negative ion currents. Based on the measurements, it is concluded that the negative current is not carried by electrons but by negatively charged fullerenes. Finally, the negative ion source was successfully coupled with a positive ion source. To the best of our knowledge, this is the first paper to report the demonstration of an ion thruster using a negative ion source instead of a cathode.</p>

<p>As a part of a Japanese collaborative research and development project on practical use of a high power anode layer type Hall thruster, a 5 kW class anode layer Hall thruster (RAIJIN94) has been developed and the thrust performance has been evaluated. The thrust was measured in the ion engine endurance test facility at ISAS/JAXA using a pendulum thrust stand developed at the University of Tokyo. The thrust performance at 3 kW operation was measured (xenon anode mass flow rate of 9.8 mg/s and xenon cathode mass flow rate of 0.5 mg/s); the thrust, specific impulse, and thrust efficiency were found to be 160 mN, 1600 sec and 0.42, respectively. The thrust performance depends on magnetic field configuration, that is, the strength of the magnetic field and the ratio of trim coil to inner/outer coil.</p>

<p>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 increase thrust by 20%, 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 will be on board thanks to the thrust enhancement of the IES. An improved neutralizer with stronger magnetic field for longer life has been under endurance test in diode mode since August 2012 and has accumulated the operational hours of 25600 h ( > mission requirement: 14000 h) by July 2015. The IES flight model was developed within 2.5 years. The spacecraft was launched from Tanegashima Space Center in Kagoshima Prefecture on-board an H-IIA launch vehicle on December 3, 2014. </p>

宇宙航空研究開発機構・宇宙科学研究所・電気推進研究室が、米欧ロとは技術的に一線を画して研究開発したマイクロ波放電式イオンエンジンは、「はやぶさ」小惑星探査機の主推進として採用され、地球〜小惑星間宇宙往復航海を世界に先駆けて実現した。高効率・省電力でプラズマを生成しながら１台当たり２年間にも及ぶ耐久性を宇宙で実証した。宇宙活動と同時並行で行われた地上におけるさらなる研究開発は、光ファイバーを用いた新たな探針法によりイオン源内部現象を解明し、性能向上をもたらした。改良されたイオンエンジンは、「はやぶさ２」小惑星探査機において、新たな小惑星に向けてその能力を今まさに発揮中である。本稿では、従前の電極を用いる直流放電式システムと比較しながら、電子サイクロトロン共鳴型イオン源の高い性能と耐久性を解説する。

小惑星探査機「はやぶさ2」による小惑星サンプル採取量の推定は,小惑星到着後,はやぶさ2のサンプリング地点を決定する上で非常に重要となる.はやぶさ2のサンプリング機構は,様々な小惑星表面状態に対応できるよう,小惑星表面に弾丸を衝突させ,舞い上がった粒子を採取する仕組みとなっている.そのため,衝突クレーター形成時の放出粒子の挙動を知ることは,小惑星サンプル採取量や採取過程を明らかにすることにつながる.小惑星には様々な表面状態が存在すると想定されるが,はやぶさ2のサンプリング候補地点として,粉体層からなるレゴリスが有力視されている.そこで本稿では,はやぶさ2のサンプリング機構を模擬し,小惑星レゴリスからのサンプル採取時の放出粒子の挙動や採取量を実験的に明らかにした.

The electron cyclotron resonance microwave discharge ion thrustery mu 10, installed in the Japanese asteroid explorer Hayabusa, experienced trouble during a round-trip voyage between Earth and an asteroid due to degradation of the neutralizer. The neutralizer is a critical component limiting the thruster lifetime. The mechanism of the performance degradation is investigated here to find out methods to extend its lifetime for next-generation mu 10 thrusters. Fault tree analysis during on-the-spot inspections of the neutralizer, previously used for 20,000 h endurance testing on the ground and in experimental simulations, leads to the following proposed degradation mechanism. The internal-surface materials are sputtered by doubly ionized xenon and accumulate on other surfaces, forming thin films containing iron. Next, thermal cycles result in the peeloff of the thin films due to the difference between the thermal expansions of the films and the base plates: Finally, thin films containing iron (that is, ferromagnetic flakes) are magnetically attracted to the tips of the magnetic circuit. These flakes inhibit plasma production, resulting in a performance degradation of the neutralizer. A magnetic circuit covered with molybdenum as a sputtering-resistant material is shown to be effective in prolonging the life of the neutralizer.

Two optical fiber measurement techniques are used in this paper to reveal the physical mechanism of the enhancement of the thrust force of the mu 10 electron cyclotron resonance ion thruster. The beam current of the mu 10 thruster was increased in previous studies by changing the propellant injection method. In this study, to observe the difference in plasma distributions, optical fiber probes were inserted into the thruster under beam acceleration. The first measurement was laser absorption spectroscopy. By traversing the optical fiber, the number densities of Xe I 5p(5)(P-2(3/2)0)6s[3/2](2)(0) 1 were obtained along the center axis. The second measurement was an electric-optic element probe measurement conducted to measure the intensities of the microwave electric field. Both measurements suggest that there is plasma in the waveguide in the conventional model of the thruster. This phenomenon is possibly caused by the leakage of electrons from the electron cyclotron resonance region to the waveguide. As a result, this paper concludes that the suppression of plasma in the waveguide is a very important measure to improve the performance of microwave thrusters.

Next generation arcjets should have light-weight design and prolonged lifetime. For the former topic, it is shown that the radiator mass can be drastically reduced by the effective use of propellant as a coolant at the lower temperature region on the radiator. Resulting thruster weight of 2.0 kg including the radiator is possible for 15 kWe arcjet. For the latter topic, replaceable cathode system is proposed and some key issues are mentioned.

The Small Demonstration Satellite-4 (SDS-4) of JAXA launched on May 18, 2012 (JST) is equipped with a Japan's first quartz crystal microbalance (QCM) for spacecraft surface contamination monitoring. The QCM was installed on one of the satellite outer surface and occasionally observed gradual frequency decrease (=contamination) under the ground clean room environment for about a year. The QCM frequencies just before and after the launch by the H-IIA Launch Vehicle No. 21 (H-IIA F21) were almost the same, which indicated good cleanness inside the H-IIA's payload fairing. The frequency rapidly increased to the initial level during the first week after the launch probably due to removal or erosion of contaminants on the crystal surface by attack of atoms and ions in the orbit at an altitude of about 700 km. Contamination was never dominant during seventeen months of the space operation. Long term trend of the QCM frequency seems to be affected by the upper atmosphere density changing with the F10.7 solar radio flux.

Hayabusa2 is an asteroid exploration mission to return surface samples of a near-Earth C-type asteroid (162173) 1999 JU(3). Because asteroids are the evolved remnants of planetesimals that were the building blocks of planets, detailed observation by a spacecraft and analysis of the returned samples will provide direct evidence regarding planet formation and the dynamic evolution of the solar system. Moreover, C-type asteroids are expected to preserve the most pristine materials in the solar system, a mixture of minerals, ice, and organic matter that interact with each other. Space missions are the only way to obtain such pristine materials with geologic context and without terrestrial contamination. Hayabusa2 will launch off in 2014, arrive at 1999 JU(3) in mid-2018, and fully investigate and sample the asteroid at three different locations during its 18-month stay. The concept and design of the Hayabusa2 sampler are basically the same as that on-board Hayabusa, and impact sampling with a 5-g Ta bullet will be made at three locations of the asteroid. The sample container has three separate chambers inside to store samples obtained at different locations separately. The spacecraft will return to Earth with samples in December 2020. Returned samples will be investigated by state-of-the-art analytical techniques in 2020 to understand the evolutionary history of the solar system from 4.56 Gyr ago to the present by combining results from laboratory examinations of the returned samples with remote-sensing datasets and comparing all results of observations of meteorites, interplanetary dust particles, and future returned samples.

The microwave ion thruster &mu;10's ion beam current saturated at a large mass flow rate when propellant gas was injected from a waveguide inlet and it was improved by additional propellant inlets to a discharge chamber. In order to understand the mechanism of these phenomena, it is important to measure distributions of the microwave electric field inside the discharge chamber, which is directly related to plasma production. In this study, we applied an electro-optic (EO) probe to measuring the microwave electric field. The probe contains no metal and can be accessed in the discharge chamber with less disruption to the microwave distribution. We measured electric-field profiles along the centerline and in the ECR area of &mu;10 with the EO probe. Consequently, this paper revealed that when the propellant was injected from the waveguide inlet, microwave was reflected in the waveguide at large mass flow rate, which disturbed a propagation of microwave to the ECR area. It also revealed that when the propellant was injected from the discharge chamber inlet, the mass flow rate where the microwave reflection occurred shifted to larger rate, which resulted in the increase of the beam current．

The basic properties of an air breathing ion engine, which uses upper atmospheric gases as a propellant, were experimentally investigated. The N-2 environment in a sub-low Earth orbit (altitude of 140-200 kin) was simulated by a laser detonation beam source, which has been previously used in studies on atomic oxygen-induced material degradation. The basic properties of the air breathing ion engine were studied using a hyperthermal N-2 beam. It is suggested that the hyperthermal N-2 molecules thermalized by scattering at the reflector surface in the air breathing ion engine. The efficiency of the collimator was experimentally investigated and the collimator was found to maintain the N-2 pressure inside the air breathing ion engine. An ion beam current of 16 mA at an acceleration voltage of 200 V provided a thrust of 0.13 mN for both hyperthermal N-2 and atomic oxygen beams. The maximum ion beam current was found to be limited by the space-charge effect. The experimental results strongly indicated the recombination of atomic oxygen into O-2 molecules inside the air breathing ion engine.

宇宙工学は、宇宙への往来の実現を目指し、技術を切磋琢磨してきた。その成果の端的な例は、「はやぶさ」にて実現された地球〜小惑星間往復航行（２００３年〜２０１０年）である。それにより、科学や技術分野を越えて、より大きな世界観を得ることができた。次の新しい知見を得るために、科学的な意義はもちろんのこと、「宇宙を自在に往来する独自能力の維持発展」と「人類の活動領域の宇宙への拡大」という宇宙工学・宇宙探査に跨る目標を担い、「はやぶさ２」小惑星探査ミッションが開発中である。

In order to understand the internal phenomena in a microwave discharge ion thruster, it is important to measure the distribution of the microwave electric field inside the discharge chamber, which is directly related to the plasma production. In this study, we proposed a novel method of measuring a microwave electric field with an electro-optic (EO) probe based on the Pockels effect. The probe, including a cooling system, contains no metal and can be accessed in the discharge chamber with less disruption to the microwave distribution. This method enables measurement of the electric field profile under ion beam acceleration. We first verified the measurement with the EO probe by a comparison with a finite-difference time domain numerical simulation of the microwave electric field in atmosphere. Second, we showed that the deviations of the reflected microwave power and the beam current were less than 8% due to inserting the EO probe into the ion thruster under ion beam acceleration. Finally, we successfully demonstrated the measurement of the electric-field profile in the ion thruster under ion beam acceleration. These measurements show that the electric field distribution in the thruster dramatically changes in the ion thruster under ion beam acceleration as the propellant mass flow rate increases. These results indicate that this new method using an EO probe can provide a useful guide for improving the propulsion of microwave discharge ion thrusters. (C) 2012 American Institute of Physics. [http://dx.doi.org/10.1063/1.4770116]

This study evaluates the system performances of a 10-W-class miniature ion thruster designed for 50kg small spacecraft. The miniature ion thruster used here, using microwave discharge, was specially designed for low microwave power operation, as low as 1.0W. Thruster performance of this thruster (ion beam current, required microwave power, and required gas flow rate) was measured by the experiments. This experiment included a neutralizer and power and gas needed for its operation. Specifications of sub-components needed for a miniature ion thruster system was estimated based on commercially available or space qualified products. As a result, performance of the miniature ion thruster system was evaluated using those thruster performance and sub-component specifications. One of the results is that the miniature ion thruster system can generate 297μN thrust with 1100s specific impulse and ΔV of 300m/s for 50kg spacecraft by 15.6W total power consumption and 2.7–3.5kg total wet weight of the system.

The electron cyclotron resonance microwave discharge ion thruster μ10, which was installed in Hayabusa, has experienced an autonomous stop in the final phase of Hayabusa project by the degradation of the neutralizer. It was shown that the neutralizer is a critical element that limits the thruster lifetime. To understand the mechanism of performance degradation, we presupposed a degradation mechanism by studying the neutralizer which had gone through a 20,000-hour endurance test and whose degradation had already occurred. According to this mechanism, degradation is triggered off by divalent ion sputtering against yokes of inside neutralizer. The objective of this study is to evaluate degradation in quantitative form by measuring net ion current distribution of the neutralizer and find out the method to improve endurance of the neutralizer. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

DECIGO, a Deci-hertz Interferometer Gravitational Wave Observatory, is a space gravitational wave antenna. The purpose of DEC1GO is to observe gravitational waves at the frequency band mainly between 0.1-1.0 Hz, and to open a novel window of gravitational wave astronomy. DECIGO will consist of three spacecrafts flying in a triangular formation with a side length of 1,000 km. The position of each satellite with respect to its two counterparts has to be controlled to ensure sufficient accuracy of the scientific measurements. Therefore, the propulsion system which can satisfy stringent requirements for drag-free control is indispensable. To accomplish this control by some propulsion systems, these thrust must be controlled precisely to counteract non-gravitational forces such as residual aerodynamic drag or solar radiation pressure. DECIGO Pathfinder (DPF) is the precursor mission to DECIGO designed to validate the core technologies. One of the enabling technologies in DPF mission is the precise micro-propulsion system necessary to achieve the unique propulsion requirements. The objective of this study is to develop a miniature microwave discharge ion thruster for this micro-propulsion system. In drag-free control, the thrust must be actively-controllable with fast response (&gt 10 Hz). It is different from a conventional ion thruster system, where the thrust is constant for a long time. In this experiment, the thrust dynamic range was between 7-100 %, and the thrust noise was less than 0.02 μN/Hz1/2 in the frequency range of 0.1-1.0 Hz. In addition, the thrust control with fast response was realized by the feedback control of the ion beam current.©2012 by the International Astronautical Federation.

Japanese deep space spacecraft Hayabusa is the first spacecraft using an electron cyclotron resonance (ECR) microwave discharge ion thruster as the primary propulsion system. A discharge cathode to generate primary electrons to generate plasma is not necessary for the microwave discharge ion thruster. Hence the cathode erosion is not able to occur, which is one of a typical breakdown of ion thruster. In fact, the total accumulated operation time of Hayabusa's four ion thrusters named "10" thruster almost reached 40,000 hours, setting the world record. However, μ10 experienced an autonomous stop in the final phase of Hayabusa project by the degradation of the neutralizer. It was shown that the neutralizer is a critical element that limited the thruster lifetime. Therefore it is necessary to enhance the endurance of neutralizer to make the lifetime of the spacecraft longer. To understand the mechanism of performance degradation, we presupposed a degradation mechanism by studying the neutralizer which had gone through a 20,000-hour endurance test and whose degradation had already occurred. According to this mechanism, degradation is triggered off by divalent ion sputtering against walls of inside neutralizer. To prevent the degradation, to decrease sputtering voltage which related to applied neutralizer voltage to emit electrons and to decrease numbers of the ions are effective. The objective of this study is to evaluate degradation in quantitative form by measuring current distribution of the neutralizer and find out the method to improve endurance of the neutralizer. When ECR neutralizer emits electrons, the same charge of xenon ions hit inside of the neutralizer as a counterpart. This hit is necessary to emits electrons but also it triggers off the performance degradation. Therefore measuring the charge of the xenon ions is proper for improving endurance of the neutralizer. This positive charge flows to the earth ground in laboratory experiment. We measured this current by insulating each parts of the neutralizer and investigated current distribution when running the neutralizer. We found out that there are parts which the ions hit dominantly and the parts which the electrons hit dominantly. This knowledge enable us to understand what part is intimately related to decreasing numbers of ions which sputter inside neutralizer and decreasing sputtering voltage. In this paper, we report the results of net ion current distribution measurement and floating the antenna is effective to suppress the neutralizer voltage.©2012 by the International Astronautical Federation.

The μ10 cathode-less electron cyclotron resonance ion engines, have propelled the Hayabusa asteroid explorer for seven years since its launch in May 2003. The spacecraft was focused on demonstrating the technology needed 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. The final stage of the return cruise and the subsequent trajectory correction maneuvers have been accomplished by using a special combined operation of neutralizer A and ion source B after the exhaustion of the other neutralizers' lives by the autumn of 2009. The total duration of the powered spaceflight was 25,590 h, which provided a delta-V of approximately 2.2 km/s and a total impulse of 1 MN·s. The degradation trends of the thruster performances have been investigated. It seems that the main cause of the degradation was the decrease in effective microwave power input to the discharge plasma induced by the increase in the transmission loss of the microwave feed system, and not due to the increase in the gas leakage through the accelerator grid apertures enlarged by erosion. Unintentional engine stop events have been summarized and analyzed. Most of them occurred due to the limit check errors of the backward microwave powers. Such errors can be decreased by carefully monitoring the trend change in microwave backward power as a function of xenon flow rate in future missions.

The microwave discharge ion engine μ10's thrust force was improved by additional propellant inlets to a discharge chamber. However, internal plasma diagnostics was not carried out while ion beam was extracted. In order to understand the effects of the new propellant inlets, we measured excitation temperatures and axial number density distributions of metastable Xe I 5p⁵(²P⁰_3/2)6s[³/₂]⁰₂ inside of μ10 by a line pair method and laser absorption spectroscopy respectively. Firstly, the measurement of excitation temperatures was operated in two positions of the probe tip: 0cm and 5cm from a screen grid. This measurement confirmed that the temperatures marked between 0.42 and 0.68eV. Secondly, the number density distribution measurements were realized by a novel laser absorption spectroscopy utilizing optical fibers. As a result, 10¹⁷m^-3 order of metastable neutral particles were measured by coupling with the excitation temperatures. Consequently, this paper will reveal that the propellant injection from a waveguide inlet increased the electron number density in the waveguide, which disturbed a propagation of microwave to the discharge chamber. It will also reveal that the propellant injection from the discharge chamber was effective to suppress the plasma production in the waveguide, which resulted in the increase of the thrust.