西山 和孝

Ionic liquid electrospray thrusters represent an alternative propulsion method for spacecraft to conventional plasma propulsion because they do not require plasma generation, which significantly increases the thrust efficiency. The porous emitter thruster has the advantages of simple propellant feeding and multi-site emissions, which miniaturize the thruster size and increase thrust. However, the multi-scale nature, that is, nano- to micrometer-sized menisci on the millimeter-size porous needle tip, makes modeling multi-site emissions difficult, and direct observation is also challenging. This paper proposes a simple model for multi-site emissions, which assumes that the ionic conductivity or ion transport in the porous media determines the ion-emission current. The conductivity was evaluated by comparing the experimental and numerical data based on the model. The results suggest that the ionic conductivity of the porous emitter is suppressed by the ion–pore wall friction stress. Additionally, the model indicates that the emission area expansion on the porous emitter creates the unique curve shape of the current vs voltage characteristics for multi-site emissions.

An understanding of the degradation mechanism of a microwave discharge cathode is the key to extending the lifetime of microwave ion thruster systems. This study investigates the effect of nozzle contamination by sputtered Ag-polytetrafluoroethylene (PTFE) on microwave discharge cathode performance. The current–voltage characteristics were measured for nominal and contaminated (by PTFE spray with 0.2 µm thick or tape with 0.15 mm thick) cathodes. The contamination thickness and area on the nozzle were varied to investigate the characteristic differences. It was confirmed that the anode voltage increased by 20 V or more in the case of the contaminated cathode. The anode voltage was measured for the sputter-contaminated cathode to evaluate the effect of contamination under more realistic conditions. After 630 h of sputter-contamination operation, it is estimated that sputtered particles were deposited to a thickness of 77 µm at most, and the anode voltage increased by 8 V. The results show that the downstream surface of the nozzle is critical for maintaining cathode performance. The insulating coating formed by the sputtered PTFE may interfere with ion absorption and degrade electron emission capability. A theoretical model based on the extended Brophy model supports these results. This study provides important information for the use of PTFE-based materials around ion thrusters.

Hayabusa2 is an asteroid sample return mission by Japan Aerospace Exploration Agency. After the first touchdown to the asteroid Ryugu, the nadir-viewing wide-angle optical navigation camera (ONC-W1) of Hayabusa2 experienced significant degradation in sensitivity that are likely caused by dust adhesion to its optical system. Images of the asteroid and target markers were darkened compared to those captured before the touchdown. Several large dark spots were also observed in images after the touchdown. Subsequently, some of them disappeared after Small Carry-on Impactor (SCI) separation, which caused largest impulse to the spacecraft. The dust removal operation using the electric propulsion system of Hayabusa2 was also attempted. We also evaluated the effects of the optical degradation to the image-based navigation, particularly for the second touchdown. This chapter describes the degradation, attempts for dust removal, and impact to image-based navigation.

DESTINY+ is an upcoming JAXA Epsilon medium-class mission to fly by the Geminids meteor shower parent body (3200) Phaethon. It will be the world's first spacecraft to escape from a near-geostationary transfer orbit into deep space using a low-thrust propulsion system. In doing so, DESTINY+ will demonstrate a number of technologies that include a highly efficient ion engine system, lightweight solar array panels, and advanced asteroid flyby observation instruments. These demonstrations will pave the way for JAXA's envisioned low-cost, high-frequency space exploration plans. Following the Phaethon flyby observation, DESTINY+ will visit additional asteroids as its extended mission. The mission design is divided into three phases: a spiral-shaped apogee-raising phase, a multi-lunar-flyby phase to escape Earth, and an interplanetary and asteroids flyby phase. The main challenges include the optimization of the many-revolution low-thrust spiral phase under operational constraints; the design of a multi-lunar-flyby sequence in a multi-body environment; and the design of multiple asteroid flybys connected via Earth gravity assists. This paper shows a novel, practical approach to tackle these complex problems, and presents feasible solutions found within the mass budget and mission constraints. Among them, the baseline solution is shown and discussed in depth; DESTINY+ will spend two years raising its apogee with ion engines, followed by four lunar gravity assists, and a flyby of asteroids (3200) Phaethon and (155140) 2005 UD. Finally, the flight operations plan for the spiral phase and the asteroid flyby phase are presented in detail.

Abstract In electron cyclotron resonance (ECR) thrusters, the plasma mode transition is a critical phenomenon because it determines the maximum thrust performance. In ECR ion thrusters, ionization generally occurs in the magnetic confinement region, where electrons are continuously heated by ECR and confined by magnetic mirrors. However, as the flow rate increases, ionization is also observed outside the magnetic confinement region, and this induces the plasma mode transition. In our previous work, two-photon absorption laser-induced fluorescence (TALIF) analysis revealed that the stepwise ionization from the metastable state plays an important role in the ionization process. However, the distribution of the stepwise ionization has not yet been revealed because of the long lifetime of the metastable state. In this study, this distribution was investigated using one experimental and two numerical approaches. First, TALIF was applied to two types of gas injection with clear differences in thrust performance and ground-state neutral density distribution. In the first simulation, the metastable state particle simulation was used to estimate the excitation rate distribution. In the second study, simulations of the electric field of microwaves were used to estimate the contribution of the stepwise ionization to the plasma density. The experimental and numerical results revealed that the stepwise ionization spreads outside the magnetic confinement region because of the diffusion of metastable particles, and this spread induces the plasma mode transition, explaining the difference between the two types of gas injection.

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 article was originally published online on 9 November 2021 with an error in Eqs. (13) and (14). The equations are shown correctly below. All online and printed versions of the article were corrected on 10 November 2021.

In this study, a novel single-piece thin multi-layer tungsten resistive heater was successfully fabricated using additive manufacturing and tested as an electrothermal thruster. The heater has 12 resistive layers, with each layer having a thickness and height of 0.15 and 81 mm, respectively, and can provide high heating efficiency. A single-piece or monolithic heater was manufactured via additive manufacturing technique, which drastically improved its reliability and decreased its manufacturing cost. In the heating and thrust measurement tests that used nitrogen gas as a propellant, the heater reached a gas temperature of ∼2000 K at a 140-A heater current without experiencing any failure. The tungsten-heater resistance linearly increased with an increase in temperature due to the temperature dependence of tungsten's resistivity. The specific impulse and thrust increased with the heater temperature in accordance with the theoretical prediction. Even including a voltage drop due to a contact resistance, the achieved heater efficiency reached 63% at a 100-A heater current even without a thermal insulation around the thruster. The heater efficiency decreased with an increase in the heater temperature due to heat loss to the surroundings. The heat-loss analysis indicated that both thermal conduction and radiation heat losses were crucial for improving the heater performance at a high-temperature operation of over 2000 K.

JAXA’s asteroid explorer Hayabusa2 completed its operation near the asteroid 162173 Ryugu, which started in June 2018, and carried out a maneuver away from the asteroid on November 13, 2019. In the outbound operation, the total delta-v performed by its ion propulsion was about 1,015 m/s, the space powered flight time reached 6,515 hours, 24 kg of propellant xenon was consumed, and 42 kg remained. On the return trip, 2,400 hours of operation was carried out in two parts, from December 2019 to February 2020 and from May to August 2020. Trajectory correction maneuver TCM-0 was carried out with one ion thruster from September 15 to 17, 2020, which was the last operation of the ion engine system, followed by several TCMs by chemical propulsion. The capsule returned to Earth on December 6, 2020. The total delta-v in the round trip was about 1.3 km/s, and the powered flight time was 9,398 hours. After consuming 31 kg of propellant xenon, 35 kg remained, a series of close flyby with an L-type asteroid 2001 CC21 in 2026 and rendezvous with a fast rotator asteroid 1998 KY26 in 2031 has been proposed as an extended mission of Hayabusa2 and its ion engine were restarted on January 5, 2021. The cumulative operating times for the four ion thrusters are 6,996, 2,880, 9,220, and 8,941 hours, respectively. 12,632-hour powered flight by the ion engine system produced about 1.7 km/s delta-v. An engineering model of Hayabusa2 neutralizer has been subjected to ground durability tests since the summer of 2012 prior to launch. 75,277 hours have passed by the end of September 2021, and it is still operating without failure and testing is ongoing.

Plasma mode-transition and hysteresis have been reported in several electron cyclotron resonance (ECR) plasma sources and have also been observed in ECR ion thrusters. From observation outside the thruster, the optical emission is significantly changed after plasma mode-transition, which indicates the electron heating, excitation and ionization processes are also changed. Therefore, to investigate these processes quantitatively, the ground-state neutral density and spontaneous emission by electron impact excitation are directly measured using two-photon absorption laser-induced fluorescence spectroscopy. This measurement system is applied to two types of thrusters. In one thruster, both mode-transition and hysteresis were observed, while in another, no hysteresis was observed due to the partial prevention of ECR heating. The experimental results indicate that the spontaneous emission sharply increases by the mode-transition and hysteresis. The tendency for this increase was relatively small with the partial prevention of ECR heating; however, the mode-transition was not deleted. Analysis of the excitation and ionization process revealed that the increase of indirect (stepwise) excitation and ionization from metastable particles can contributes to the sharp increase of the spontaneous emission. In addition, quantitative estimation of the collisional and ECR heating indicated that collisional heating cannot be neglected after the mode-transition, which indicates that collisional heating can contribute to the sharp increase due to the increase of electron temperature. The indirect ionization and collisional heating could thus be the main cause of the mode-transition and hysteresis.

This paper reports measurements of the xenon ground state and excited state densities inside a mu 10 microwave ion thruster. This thruster exhibits a 40% thrust enhancement upon changing from the waveguide to the discharge chamber propellant injection mode. In the present work, the associated mechanism was quantitatively evaluated using two-photon laser induced fluorescence (TALIF) spectroscopy to monitor the thruster waveguide. The 834.7 nm emission from excited state xenon was investigated with a 224.3 nm dye laser to excite the Xe I 5p61 S0 6pMODIFIER LETTER PRIME [3/2]2 state, compared with the emission without the laser. The resulting data confirm that the neutral density exhibits a linear relationship with the propellant flow rate in the cold gas and ionized state, while the ion acceleration decreases the neutral density by the same order of magnitude as the propellant utilization efficiency is changed. As the propellant flow rate increases, the collisions of neutrals that generate excited states occur in the waveguide and, when this process plateaus, the ground state emission suddenly increases. Propellant injection from the discharge chamber is evidently effective at suppressing collisions with electrons in the waveguide that generate excited states and that potentially interfere with microwave propagation.

© 2021 IAA 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.

© 2020 IAA A hollow cathode is an efficient electron source in the self-heating mode utilizing the discharge power. However, in sub-ampere currents, it needs keeper power to maintain the thermionic electron discharge, which could decrease the thrust efficiency. To address this problem, we propose using a microwave cathode, which is based on the flight model of a microwave ion thruster neutralizer cathode, as an alternative to a hollow cathode. First, we redesigned the magnetic field of a microwave cathode discharge chamber and tested it in the diode mode configuration. The electron emission current is doubled compared to the original performance. Next, we coupled the improved microwave cathode with a 200-W class Hall thruster and compared the characteristics and performance with a hollow cathode. We confirmed that the magnetic field polarity affects the ignition characteristics. We measured the thrust by an inverted thrust stand, the ion energy distribution functions by a retarding potential analyzer, and the beam profiles by an ion collector. The thrust and thrust efficiency are equivalent for both types of cathode. The specific impulse is 10% higher in the case of the microwave cathode. Since the potential difference between the microwave cathode and ground rapidly increased at currents above 600 mA, this could be taken to be the trade-off point against the hollow cathode.

© 2020 IAA The authors redesigned the discharge chamber for the μ10 microwave discharge ion thruster to improve its thrust performance and succeeded in enhancing the maximum beam current and thrust efficiency. However, it was found that the ion current ratio extracted from the discharge chamber with the redesigned configuration was lower than that obtained with the original configuration. To investigate the relationship between ion extraction and the magnetic field geometry, the ion loss current distribution in these two types of discharge chamber were measured by electrostatic probes. Using planar probes with a guard ring, the ion current that flowed into the wall was measured without disturbing the ion beam current. The results show that ionization occurs mainly near the upstream magnet. In addition, the ion flux on the sidewall in the redesigned discharge chamber is about 1.5-2 times larger than that in the original discharge chamber. This suggests that the distance between the edge of the plasmaproduction region and the chamber wall with consideration of the Larmor radius of ions is an important parameter in discharge chamber design. In addition, although the ion beam current showed a tendency to saturate at high microwave power, the ion loss to each part in the discharge chamber increased in proportion to input microwave power. The decrease in the extracted ion ratio in the redesigned discharge chamber is considered to be caused by a decrease in the electrostatic ion transparency of the screen grid. Therefore, in a well-tuned microwave discharge ion thruster, it is difficult to improve the thrust efficiency by increasing the discharge power. A design that suppresses the wall loss of ions is thus important.

© 2020 IAA In space operation of the microwave discharge ion thruster μ10 on the asteroid explorers Hayabusa and Hayabusa2, the propellant utilization efficiency deteriorated much more than in the ground endurance test. In this study, a fault tree analysis and experimental simulations of space operation were performed, focusing on the grid-derived internal carbon contamination. It was found that the performance deterioration due to the waveguide contamination matched that in Hayabusa2 qualitatively and quantitatively. Based on the experimental verification, the future performance is experimentally predicted.

© 2019 Elsevier Ltd This paper reports the design, fabrication, and testing of a reversible thermal panel breadboard model (RTP-BBM). RTP is a flexible, re-deployable radiator that autonomously controls the temperature of a heat source. It promotes heat dissipation by deploying the radiator surface when the heat source is at a high temperature. Conversely, in a cold case, heat dissipation is conserved by stowing the radiator surface. Herein, deployment/stowing and thermal vacuum tests were conducted herein to evaluate the validity of the design, and model correlations were conducted via thermal analysis. The RTP-BBM comprises high thermal conductivity graphite sheets as the flexible fin, and shape-memory alloys (SMA) as a temperature sensitive passive actuator. The deployment/stowing test was conducted in a thermal constant bath, confirming that the fin was deployed and stowed according to the SMA temperature. However, temperature hysteresis of up to +60 °C was confirmed between heating and cooling cycles. In the thermal vacuum test, power step and power cycle tests were conducted. Results showed that the fin deployed and stowed according to the temperature of the onboard equipment while autonomously regulating the temperature. Additionally, the thermal analysis model correlated with the experimental results, showing good agreement within ±6 °C.

<p>The authors investigate the discharge chamber of the microwave ion thruster μ10 by using kinetic particle simulation. First, to investigate the plasma phenomena qualitatively, we conduct a particle-in-cell (PIC) simulation model. The simulation results indicate that the distribution of ion density is ring-shaped. To verify the simulation result with the experimental result, the simulation result is compared with the optical emission distribution. In low propellant flow rates, the distribution of ion density agrees with the optical emission distribution. However, in high propellant flow rates, the optical emission distribution is different from simulation results in the waveguide due to the excited neutral particles. In the thruster, the performance strongly depends on the location of injecting the propellant. Hence, to develop the plasma simulation for quantitative comparison with the experiment, the distribution of the neutral density is evaluated by using direct Monte Carlo simulation (DSMC). The results show the neutral density in the waveguide increases corresponding to the ratio of waveguide injection, which indicates that the density is one of the most important parameters for quantitative evaluation with the experiment.</p>

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Solar electric propulsion is a key enabling technology that has improved the efficiency of space transport. With specific impulses that are typically ten times higher than the chemical counterpart, electric motors allow a considerable saving in propellant mass at the expense of longer times of flight. However, the length of the transfer process and the specific operational needs require to develop a different operational concept for the navigation and orbit control that can be sustained during the different phases of the mission. In this paper, a trade-off is performed among several operational concepts and solutions for multi-revolutions SEP transfers with application to the DESTINY+ mission. The GTO-to-Moon low-thrust transfer is first computed in a high-fidelity model with a tangential thrust strategy and later optimized with a five-order Legendre-Gauss-Lobatto collocation method. The impact of eclipses, radiation, thrust outages and misfires, and orbit tracking is analyzed in detailed and included in the transcript optimal problem as algebraic constraints where possible. Numerical results show that the driving factors for the optimal trajectory are related to the operations of the spacecraft rather than the final mass or time of flight.

© 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.

© 2020 The Japan Society for Aeronautical and Space Sciences To investigate the neutral xenon density distribution of electric thrusters such as ion and Hall thrusters, two-photon absorption laser-induced fluorescence (TALIF) spectroscopy was applied to a microwave cathode. First, the background pressure of the vacuum chamber was measured by TALIF. In the present measurements, the ground state was excited by a 224.29 nm laser, and 834.68 nm fluorescence was detected. The first measurement confirmed that the fluorescence intensity linearly increases with respect to the ground state number density. Based on this result, the density of neutral ground-state xenon was measured at the exit of the nozzle of the microwave cathode. The variation in the density with the microwave power was successfully measured at xenon flow rates of 0.029 and 0.098 mg/s. The measured densities varied from 2.3 © 1019 to 8.4 © 1019 m3 with a maximum error of «20% due to the plasma fluorescence.

© 2019 IAA 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 coil. 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 Elsevier Ltd This paper reports the first study to measure xenon neutral ground state particle density of microwave cathode by two-photon laser induced fluorescence spectroscopy (TALIF). Xenon is commonly used as a propellant in electric propulsion like Hall thrusters, ion thrusters, and their cathodes. For electric propulsion, information about neutral particles is important such as the ionization degree and the charged exchange collisions (CEX). The measurement target is XeI 5p61S06p[3/2]2, which absorbs at a wavelength of 224.29 nm and emits fluorescence of 834.7 nm. The measurement system was demonstrated for three cases: cold gas, without electron extraction, with electron extraction. From three cases, the measurement system can detect a neutral ground state particle density of 1019m−3 order without and with a plasma. In a cold gas, the neutral ground state particle density is (8.4±0.4)×1019m−3 at 0.098 mg/s. Without electron extraction, the neutral ground density decreases by ionization and excitation With electron extraction, the density varied from 0.6 to 2.3 times compared to without electron extraction depending on anode voltage.

© 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 Author(s). This paper reports the numerical investigation of plasma properties for the microwave discharge ion thruster μ10. The model consists of a particle in cell simulation and a Monte Carlo collision simulation. The results indicate that the plasma density and the electron temperature in the confined region are larger than those in other regions and are qualitatively consistent with probe measurements. Moreover, we traced the trajectories of charged particles to investigate the plasma generation and transport. The electron trajectories indicate that electrons are strongly confined by the mirror magnetic field and the sheath, which indicates that the confinement depends on the electron energy. As a result, the electron energy distribution function is a combination of two Maxwellian distributions. Although the hot electrons account for 3.4% of all electrons, they account for 50.1% of the ionization and can generate plasma with an excitation loss of 1/3 of that of cold electrons. The ion trajectories indicate that they are affected by the magnetic field. To investigate the effect of the magnetic field on the transport, we evaluate the ion and electron current percentage toward the wall and compare with the wall surface percentage. The ion and electron current ratios differ because of diffusion with respect to the magnetic field. The ion current percentage is larger than the surface area percentage in the grid, which indicates that ions are transported to the grid more efficiently due to the magnetic field. Therefore, the effect on ions by the magnetic field is one of the most important criteria for microwave discharge ion thrusters.

Hayabusa2 at the asteroid Ryugu Asteroids fall to Earth in the form of meteorites, but these provide little information about their origins. The Japanese mission Hayabusa2 is designed to collect samples directly from the surface of an asteroid and return them to Earth for laboratory analysis. Three papers in this issue describe the Hayabusa2 team's study of the near-Earth carbonaceous asteroid 162173 Ryugu, at which the spacecraft arrived in June 2018 (see the Perspective by Wurm). Watanabe et al. measured the asteroid's mass, shape, and density, showing that it is a “rubble pile” of loose rocks, formed into a spinning-top shape during a prior period of rapid spin. They also identified suitable landing sites for sample collection. Kitazato et al. used near-infrared spectroscopy to find ubiquitous hydrated minerals on the surface and compared Ryugu with known types of carbonaceous meteorite. Sugita et al. describe Ryugu's geological features and surface colors and combined results from all three papers to constrain the asteroid's formation process. Ryugu probably formed by reaccumulation of rubble ejected by impact from a larger asteroid. These results provide necessary context to understand the samples collected by Hayabusa2, which are expected to arrive on Earth in December 2020. Science , this issue p. 268 , p. 272 , p. eaaw0422 ; see also p. 230

© 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.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. A 3D printed Inconel resistojet is proposed as an option for short time and high fuel efficiency orbit transfers. The current thruster is presented as a proof-of-concept for high performance high temperature variants. Experiments on N2 propellant have been conducted, and the measured performance parameters are presented. Finally, the extra application of the 3D printed resistojet as part of a hybrid electro-chemical thruster is presented.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. An MPD (Magneto-Plasma-Dynamic) thruster requires the input of a large amount of power, 100 kW to 1 MW, and it is difficult to operate in a steady state. Therefore, MPD thrusters are operated in quasi-steady state with a pulse generated by a PFN (Pulse Forming Network). However, there are two ambiguities regarding the steady state of the discharge. First, the discharge time of a PFN, typically 0.5 to 1.0 ms, is insufficient to quantitatively verify discharge steady state. Second, the unsteady region at the end of a discharge can lead to error in the evaluation of the steady state of the entire discharge. In this presentation, we propose a pulsed power supply which generates more rectangular pulses that are several longer and fewer unsteady regions than those of a PFN to evaluate the quasi-steady state of an MPD thruster. In an operation test with a 1.0 ms discharge, the unsteady regions was reduced from 0.532 to 0.085 ms, and the flat-top region which should be evaluated was increased from 0.393 to 0.880 ms. Therefore, the ratio of the evaluation time to the discharge time was improved from 42% to 91%. Using the power supply, we operated an MPD thruster and obtained the discharge time characteristics of the discharge waveform and the thrust performance by sweeping the discharge time from 1.0 to 5.0 ms. As a result, the discharge current was maintained at 900 ± 100 A for each discharge time and the thrust was in the range of 0.3 ± 0.02 N, which is interpreted as being constant. These results confirm the quasi-steady state of the MPD thruster operation from 1.0 to 5.0 ms. In addition, we also investigated the influence of residual gas in a vacuum chamber on the quasi-steady state by sweeping gas supply timing.

平成30年度宇宙輸送シンポジウム（2019年1月17日-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, 神奈川県資料番号: SA6000136057レポート番号: STEP-2018-001

平成30年度宇宙輸送シンポジウム（2019年1月17日-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, 神奈川県資料番号: SA6000136059レポート番号: STEP-2018-003