國中 均

Some numerical wear tests are conducted for the carbon/carbon ion optics of a microwave ion thruster μ10 engineering model to evaluate the accuracy and precision of JAXA's ion optics code (JIEDI). Through comparisons with experiment, the JIEDI code showed good agreement with a real-time 18,000-hrs life test when incorporating the motion of eroded grid materials and a low-energy sputtering yield model for the energy below 300 V. Numerical error caused by the uncertainty of physical model is also studied and it is found that uncertainty in beam current and plasma parameters cause 10% or less error to estimate grid hole erosion profiles. The grid erosion profile is most sensitive to the uncertainty in sticking factor, which indicates what percentage of eroded grid material arriving at a grid surface will re-deposit onto the grid surface. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The Asteroid Explorer Hayabusa was launched by M-V rocket from Uchinoura Space Center, JAXA on May 9, 2003. Conquering several troubles during totally 7 years of orbital flight, it returned to the vicinity of the earth and completed the powered-flight by the ion thruster in the begging of 2010. After successive trajectory correction maneuvers for the reentry ,the mother spacecraft successfully released a small capsule with asteroid sample. The capsule has entered the earth atmosphere in the desert of the Australia on June 13, 2010, and successfully have been recovered by June 15. The present paper overviews the return operation of the Hayabusa mother spacecraft and reentry flight and recovery operation of the sample return capsule. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

The cathode-less electron cyclotron resonance ion engines, μ10, propelled the Hayabusa asteroid explorer, launched in May 2003, which is 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. It rendezvoused with the asteroid Itokawa after a two year deep space flight with a delta-V of 1.4 km/s, 22 kg of xenon propellant consumption and 25800 hours of total accumulated operational time of all the four ion engines added up. Though it succeeded in landing on the asteroid on November 2005, the spacecraft was seriously damaged. This delayed the Earth return in 2010 from the original plan in 2007. Reconstruction on the operational scheme using remaining functions and newly uploaded control logic made Hayabusa leave for Earth in April 2007. After a coasting period of 2008, the ion propulsion was reignited in February 2009. Although most of the neutralizers were degraded and unable to be used by fall of 2009, a combination of an ion source and its neighboring neutralizer has been successfully operated for the last 3230 hours including a series of final trajectory correction maneuvers. Before reentry, the total accumulated operational time reached 39637 hours consuming a total of 47 kg Xenon propellant. Total duration of powered spaceflight is 25590 hours which provided a delta-V of 2.2 km/s and a total impulse of 1 MN·s, approximately. Finally, the spacecraft returned to Earth. Its reentry capsule, which may contain samples from asteroid Itokawa, was retrieved from the Australian outback according to plan. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

Hayabusa Sample Return Capsule (SRC) has accomplished reentry flight directly from the interplanetary-transfer orbit with velocity of about 12 km/s and successfully recovered on June 14, 2010. The SRC is considered to deploy parachute at the altitude of 5 km as planned, while the chute-trigger timer has been carefully set to cover contingency cases due to malfunctions of triggering devises taking account of dispersions of orbital and atmospheric density and errors in the aerodynamic coefficients. All the components separated at the chute-deployment were searched-out based on the landing position of the instrument module.

In order to solve some problems on PPTs, especially on propellant feed, we achieved an entirely new idea of a PPT with powdered propellant. We roughly estimated the thrust performance of a powdered propellant PPT comparing with a traditional solid propellant PPT. And we obtained a reliable result that the former could show as high performance as the latter in some cases that the propellant feed was sufficiently uniform. Next we made a prototype model of propellant feed system for a powdered propellant PPT. We tested the uniformity of feed system and it was proved that this system provided propellant uniformly enough. From this fundamental test, we got a rigid design policy for developing a practical feed system. Copyright © 2009 by Takefumi Saito.

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.

The Hayabusa spacecraft launched in 2003 is now completing its round trip to a near Earth asteroid Itokawa, to which it accessed in 2005. Hayabusa suffered from a fuel leak and eruption incidents after its successful descent, touch-down and lift-off at the end of November in 2005. Hayabusa lost two reaction wheels among three aboard and the chemical propulsion. The only means left for Hayabusa are the ion engines and xenon gas reserved for them as well as a single reaction wheel. Hayabusa project team devised the use of xenon gas for cold gas propulsion, and also developed the new attitude control strategy taking the advantage of solar radiation pressure. The spacecraft started the delta-V maneuver using the ion engines from 2007 and will continue it by next February. It is scheduled for Hayabusa to reenter into the atmosphere and land in middle of Australia in June of 2010.

Neutral particles play an important role in the physics of the upper atmosphere, which expands and shrinks depending on solar activity. We have been investigated that remote sensing of the upper atmosphere structure via artificial Energetic Neutral Atoms (ENAs) generated by the charge exchange (CEX) collision between upper atmosphere atoms and the artificial ion beam. In our previous study, a krypton ion has high CEX cross-section, which calculated by the Impact-parameter method theoretically, than other noble gas to atomic oxygen. We started to some fundamental experiments for the krypton ENAs detector. We have been investigated to solid particle detector as a simple ENA particle detector. Avalanche Photodiode (APD) which has a self-amplification effect was selected as a solid detector. As a result of the heavy particles (ion and ENA) irradiation experiments to APD, APD is sensitive to the heavy particles with 1∼2keV drift energy, and its multiplication ratio is about 50 ∼ 100 times. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

The total electrode fall voltage in a channel of Magnetoplasmadynamic thruster was measured by using "zero-limit approximating method", which is one of the classical methods widely used in welding society. 5ch parallel-plate MPD was newly designed and operated in quasi-steady mode changing its electrode gap from 4mm to 1mm. The intercept of the extrapolated line was 18V that was considered to be the total electrode fall. The electrode fall voltage did not depend on the discharge current. It is considered that most of the electrode fall lies on the cathode side because measured plasma potential was almost closed to the anode potential.

Institute of Space and Astronautical Science of Japan Aerospace Exploration Agency (ISAS/JAXA) successfully developed and operated the microwave discharge ion engines "μ10" onboard Hayabusa asteroid explorer. The μ10 ion engines feature the cathode-less plasma generation in both the ion generators and neutralizes and the electrostatic grid system made of carbon-carbon composite with the results of long life and high reliability in space. Based on the space achievements of μ10 ion engines with 8mN thrust, 3,000sec Isp and 350W consumption power as the smallest electric propulsion in space, the u10HIsp is under development for deep space missions to such as Jupiter and Mercury. The integrated test with the plasma generators, a propellant isolator, a microwave DC block and a high Isp grid system established the thrusting operation with 10,000sec Isp using 15kV acceleration voltage, 24mN thrust, 12mN/kW thrust power ratio and 56% total efficiency. © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

In this study, we are proposing a novel miniaturized ion engine system μ1. Recently microspacecraft and propulsion system to be installed there have attracted a lot of attentions. To accomplish the miniaturization of spacecraft component, multifunctionalization of devices is important. The ion engine system proposed here consists of multiple engines distributed on microspacecraft. This micro ion engine system gives various motions to spacecraft by selecting the engines to be fired. A novel idea was introduced to this system to reduce the devices for the neutralization of ion beam. That is to use single plasma source as both ion beam source and neutralizing electron source only by electrical connection. This ion engine system is released from the necessity of a number of neutralizers. In order to realize our micro-ion engine system, we have two major challenges: 1) verification of the switching operation of ion engine mode and neutralizer mode and 2) very low power operation of the plasma source. In this study, we have accomplished these challenges. First, using small ion engine driven by microwave power, our idea was successfully verified. Secondly, developing antenna design method for small chamber, low power operation was achieved.

In order to give pulsed plasma thrusters (PPTs) propellant feed system, we suggested application of powdered material into propellant of PPTs. We compared variable propulsive performances of a typical ablative PPT (APPT) with solid propellant and a powdered propellant PPT (PP-PPT). One of the main measurements of PPTs is impulse measurement and another is mass shot measurement. We conducted these two measurements at once in a vacuum using the thrust mass balance we had developed. As a result, we found that dispersion of powder reduces the specific impulse. However, we also found that such propellant loss could be prevented by rubbing powder thinly into the rough surface, such as porous material, when feeding powdered propellant. In this conference, we mainly report the results of the thrust performance measurements.

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

We are developing a new type of thruster which uses powdered propellant. Propellant transport methods in most thrusters using gas or liquid propellant are some pipes and valves, but in this study, we suggest the propellant transport system using electrostatic adsorption of powder. This propellant transport with no high-pressure systems enables the structure of the thruster to be simpler. In addition, we selected pulsed electromagnetic acceleration as propellant acceleration process. We present the result of fundamental examination for electrostatic adsorption of powder and of some research about pulsed electromagnetic acceleration using powdered propellant.

This paper gives experimental information for the electrode configuration and material problem in a quasi-steady, self-field MagnetoPlasmaDynamic (MPD) thruster. Although strenuous efforts have been paid for a long time in this field, some ambiguous points still exist. In this report, the performance characteristics of 7 types of electrode (Straight, 3 types of flared and 3 types of converging-diverging configurations) were compared. The results showed that the detail scale parameters (e.g. exit or throat diameter) did not affect the thrust efficiency whereas rough classification affected to some extent. About the electrode material, we obtained the performance characteristics using Y 2O3-W and La2O3-W, which are nonradioactive rare-earth oxide tungsten. Both cathodes showed obviously lower operating voltage compared to conventional ThO2-W cathode.

Cathode-less ion engines are on the Hayabusa asteroid explorer, and Kaufman-type engines are on the Engineering Test Satellite-VIII at present. A 5kW-class Hall thruster and a 200mN-class ion engine are under development. PPTs and laser micro-thrusters are prepared for the propulsion system of microsatellites. This paper reports the recent activities on electric propulsion conducted in Japanese universities, industries, and JAXA.

It has been said that the performance of MagnetoPlasmaDynamic (MPD) thruster is strongly affected by its electrode geometries. A lot of analytical or experimental researches have been conducted, but general guidelines have not been obtained yet. Recently, some researchers pointed out the existence of optimum electrode geometry by quasi 1-D numerical calcuration. The optimum geometry had had a few characteristic features, but they were not confirmed experimentally yet. In this study, considering such features, actual thrust performances were obtained using 7 types of electrode configuration. However, it was found that the throat or the exit diameter hardly affected the propulsive efficiency.

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

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

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

It is well known that the performance of a MPD (MagnetoPlasmaDynamic) thruster is very affected by its geometries. Recently, some researcher pointed out the existence of "optimum electrode geometry" in their 1-dimensional computational studies. Optimum geometry had a few interesting feature, commonly based on converging-diverging geometry. But that has not been confirmed experimentally yet. We investigated the effects of each geometric parameter, using 6 type electrode configurations. Straight, 2 types of Flared and 3 types of converging-diverging anodes were tested at self-field, quasi-steady condition. The result was contrary to the foregoing expectation the thruster performance was found to be almost indifferent to the change of the each parameter in the case of argon propellant.

High specific impulse ion propulsion is of concern as primary propulsion system for various interplanetary explorers. A high specific impulse microwave discharge ion engine has been developed in Japan Aerospace Exploration Agency (JAXA) since 2003. The aim of this development is achieving a specific impulse of 10,000sec with an ion acceleration voltage of 15kV. New electro-static grid made of Carbon-Carbon composite material and various high voltage insulating components such as a DC block and a propellant isolator were developed. The authors fabricated a prototype model of the ion engine, and successfully operated it. The paper will describe the outline and the experimental result of demonstration test of the high specific impulse microwave discharge ion engine. © 2006 IEEE.

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

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

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

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

The performance of the MagnetoPlasmaDynamic (MPD) thruster is strongly affected its electrode geometries. A lot of analytical or experimental research had been conducted, but any remarkable guideline has not obtained yet. Recently, some researcher pointed out the existence of "optimum electrode geometry" from numerical approaches. Optimum geometry had a few interesting feature, but not confirmed experimentally yet. In this paper, considering these features, the effects of each geometric parameter are discussed through systematic experiments. Resultantly, it was found that the throat diameter in converging-diverging configuration is very critical parameter. On the other hand, the exit diameter did not affect thruster performance so much.