月崎 竜童

© 2019 American Association for the Advancement of Science. All rights reserved. The Hayabusa2 spacecraft arrived at the near-Earth carbonaceous asteroid 162173 Ryugu in 2018.We present Hayabusa2 observations of Ryugu's shape, mass, and geomorphology. Ryugu has an oblate "spinning top" shape, with a prominent circular equatorial ridge. Its bulk density, 1.19 ± 0.02 grams per cubic centimeter, indicates a high-porosity (>50%) interior. Large surface boulders suggest a rubble-pile structure. Surface slope analysis shows Ryugu's shape may have been produced from having once spun at twice the current rate. Coupled with the observed global material homogeneity, this suggests that Ryugu was reshaped by centrifugally induced deformation during a period of rapid rotation. From these remote-sensing investigations, we identified a suitable sample collection site on the equatorial ridge.

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.

© 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 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.

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)(P-3(2))6p (2)[3](0) 5/2 to Xe II 5p(4)(P-3(2))6s (2)[2] 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 mu N 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.

In order to improve the thrust force of the μ 10 microwave discharge ion thruster, its ion source which is biased at a screen voltage is separated into three parts; 1.waveguide, magnet, and a grid holder, 2.magnet spacers, and 3.a screen grid by ceramic rings. Connection of a high voltage power supply is selectable among each component independently, and maximum beam currents are measured in all six cases in addition to the nominal setup. Except for one case, the maximum thrust force is increased from the HAYABUSA's 8mN and HAYABUSA2's 10mN. The highest thrust force 11.2mN is recorded at Isp 3150sec when the magnet spacers are not connected to the power supply. At this time, the potential of the isolated magnet spacers is increased from 1500V to 1525V. The increase of the potential of the spacers regulates collisions of ions, resulting in increase of the ion currents. This study firstly demonstrates the effect of the segmentation of the discharge chamber of microwave ion thrusters and achieves the improvement of 4% of the total efficiency and of 40% of the thrust force since the HAYABUSA's flight model.

<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>

<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>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>

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.

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．

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

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]

In order to reveal the physical processes taking place within the "mu 10" microwave discharge ion thruster, 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. In this study, the axial density profiles of excited neutral xenon were successfully measured under ion beam acceleration by using a novel laser absorption spectroscopy system. The target of the measurement was metastable Xe I 5p(5)((2)P(3/2)(0))6s[(3)/(2)](2)(0) which absorbed a wavelength of 823.16 nm. Signals from laser absorption spectroscopy that swept a single-mode optical fiber probe along the line of sight were differentiated and converted into axial number densities of the metastable neutral particles in the plasma source. These measurements revealed a 10(18) m(-3) order of metastable neutral particles situated in the waveguide, which caused two different modes during the operation of the mu 10 thruster. This paper reports a novel spectroscopic measurement system with axial resolution for microwave plasma sources utilizing optical fiber probes. (C) 2011 American Institute of Physics. [doi:10.1063/1.3665954]

Based on the success of the Japanese asteroid explorer Hayabusa, the ECR ion thruster &mu;10 will be installed in Hayabusa's successor, Hayabusa-2, and is scheduled to be commercialized for use in geostationary satellites within the next three years. To increase the thrust force of the &mu;10 as much as possible without major design changes, luminescence measurements were conducted using an optical fiber probe. The probe gave an internal view of the &mu;10, and it was discovered that there was plasma in the waveguide. As the plasma, the density of which is higher than the cut-off density, interferes with the transmittance of microwaves, the propellant injection location was changed. In addition to the change in propellant injection location, the grid system was also refined. These improvements increased the thrust force from 8.0 mN to 10.1 mN with a decrease in specific impulse by 40 sec from 3200 sec to 3160 sec.

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 &micro;10 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.