西山 和孝

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.

© 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.

© Copyright 2017 by the International Astronautical Federation (IAF). All rights reserved. DESTINY+ (Demonstration and Experiment of Space Technology for Interplanetary Voyage, Phaethon Flyby and Dust Science) is a candidate of ISAS Epsilon class small program. The mission of DESTINY+ is to validate key technologies for our future deep space exploration. DESTINY+ will demonstrate the high performance electric propelled vehicle technology and execute the flyby exploration of asteroid 3200 Phaethon. DESTINY+ starts its voyage from a low elliptic orbit, spirals up the orbits, fly-by the Moon, escapes from the Earth, and depart for the asteroid 3200 Phaethon. It will detect and analyze interplanetary and interstellar dust particles during deep space cruise. This paper will introduce an overview of the DESTINY+ mission.

<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>

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.

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 km) 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 N2 beam. It is suggested that the hyperthermal N2 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 N2 pressure inside the air breathing ion engine. An ion beam current of 16 mAat an acceleration voltage of 200 V provided a thrust of 0.13 mN for both hyperthermal N2 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 O2 molecules inside the air breathing ion engine.

A 20-cm diameter electron cyclotron resonance xenon ion thruster for space propulsion is under development that generates 500 mA of ion beam current at a microwave discharge power of 100 W. It does not have any moving mechanical parts for microwave impedance matching. Extracted ion currents and reflected microwave powers were experimentally investigated around a nominal frequency of 4.25 GHz for different flow rates. Optimized frequency tuning within 0.6% of the nominal frequency minimized the microwave reflection and maximized the ion current at each flow rate between 0.39 and 1.27 mg/s. However, constant frequency operation at 4.266 GHz is recognized as the best strategy because it provided with fare performance in wide range of flow rate and almost minimum reflection during tentative stop of beam extraction after high voltage breakdowns.

The &mu;10 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&middot;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.

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.

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.

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.

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.

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.

Radiated electric field emissions from the prototype model of the Ion Engine System (IES) of the MUSES-C mission were measured in accordance to MIL-STD-461 E. The average noise level exceeded the narrowband specification at frequencies less than 5 MHz. The microwave discharge neutralizer generates a broadband noise and narrowband oscillations which have a fundamental frequency of about 160 kHz and are accompanied by its harmonics up to the 5th. The leakage of the 4.25 GHz microwave for plasma production and its second harmonic were 65 dB and 35 dB above specification, respectively. The X-band receiver onboard the MUSES-C measured the noise from the IES at the up-link frequency of 7.2 GHz through a horn antenna. This susceptibility test proved that the microwave discharge ion thruster will never interfere the deep space microwave communication.

Noise and oscillatory behavior of a plasma column produced in front of the microwave discharge neutralizer developed for MUSES-C mission were experimentally investigated. Radiated electric field emissions were measured following to MIL-STD-461 E. The average noise level exceeded the narrowband specification by 30 dB&mu;V/m at frequencies less than 5 MHz. Noise in electron emission current was also measured by using a current probe and a spectrum analyzer, and was compared with the noise of a hollow cathode. The microwave discharge neutralizer generates a broadband noise and oscillations which have a fundamental frequency of about 160 kHz and are accompanied by its harmonics up to the 5th. Considering the dependence on the diameter of the plasma column, they are probably the radial oscillation modes of ion acoustic waves. Although the hollow cathode shows nearly the same noise level at frequencies less than 1 MHz, intense oscillation exists in the 1-10 MHz range, which is generated by the keeper plasma.

An electron cyclotron resonance (ECR) ion source was designed and its performance was studied in comparison with a DC ion source. Ion production cost of 340 V/ion for ECR ion source and 189 V/ion for DC ion source were obtained at the pressure of 0.5 m Torr and the discharge power of 50 W. A luminosity distribution of the Ar-II line was imaged through an interferometer filter and a luminous arch hill was observed near the ECR region, where the Plasma was resonantly generated. Ion wall loss measurement showed that the ion production cost became worse due to the excessive ion wall loss near the cusped magnets, where the plasma confinement was inferior to that of the DC ion source.