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資料番号: AA0047899058

The Institute of Space and Astronautical Science has launched the asteroid explorer, HAYABUSA on May 2003. HAYABUSA includes the world-first asteroid rover MINERVA which will explore the asteroid surface in 2005. MINERVA needs a battery because instantaneous power generation by solar cells is insufficient for the surface exploration and data transmission tasks such as image acquisition by the onboard cameras, movement by the onboard actuator, and communication with the mother spacecraft. The night semisphere of the asteroid will be a very cold environment with the temperature lower than minus 100 C. The rover is too small to keep the inside temperature high enough for the typical batteries to be saved. Also the temperature in the day semisphere will be hotter than 100 C. Thus the onboard battery needs the availability of wide temperature range. The rover adopts the electric double-layer capacitor as the onboard electric power storage device, which can be saved under the very cold environment. The electrolytic fluid of the capacitor was tuned considering the environment of the asteroid, which enables the degradation under the very hot temperature as small as possible. So the rover is provided the ability to pass the night and the life with a couple asteroid days. This paper reports the developed electric double-layer capacitor of MINERVA asteroid rover. The energy density of the electric double-layer capacitor is smaller than the typical primary or secondary batteries under the current technical status. But it will be improved in the near future and installation to another space mission with severe temperature demand will be possible., 資料番号: AA0047899141

In general, a space system is charged negatively due to an action of space plasma. A number of works for charging and discharging phenomena of the solar cell array surface in the space system under LEO (Low Earth Orbit) have been carried out. In the future, it is expected that various structures composed of metal and insulator are installed in the surface of the space system which has various functions. It is important to work out what phenomena occur in the system. In this work, experiments are carried out to clarify charging and discharging phenomena under LEO plasma environment using some samples that composed of metal and insulator. When negative voltage is applied to the metal of the sample, micro discharges are observed in oscillating wave-form pulse current at all samples. When the applied voltage is increased, single polarity pulse discharges are observed. The single polarity pulse discharging occurs only on the metal surface and the charges of discharging from the metal are much less than the total charges of charging to the surface of the insulator. This discharging is taken place between the charges on the insulator surface and ion sheath forming near the metal. If minute projection parts exist on the metal surface, the discharging voltage lowers. In these cases, the charges of the single polarity discharging do not depend on the area of the metal and the existence of the insulator., 資料番号: AA0047275011

In order to protect space structure against space debris impacts, it is indispensable to develop a new bumper shield made of strength materials. Metal mesh material is the one of potential materials because of both lightweight and flexible. The purpose of this study is to investigate protection capability of metal mesh materials. In this study selected were two kinds of metal mesh materials; Stainless and Copper. Hypervelocity impact tests were carried out by a two-stage light gas gun using a polycarbonate projectile with 10 mm in diameter, 10 mm in thickness and 1 g in mass. The impact velocity of the projectile was 3.2 km/s. The images of fragments produced by the impact of projectile were taken by two X-ray cameras. Two cameras had delay time. The images were analyzed and the kinetic energy of fragments was calculated. Evaluated was protecting ability of metal mesh material quantitatively compared the kinetic energy before with after impact. As the result, it was found that a stainless mesh with 1 mm in thickness reduced kinetic energy by up to 38 percent but a copper mesh reduced by up to 97 percent. The stainless mesh generated fragments with large degree in spray angle in the rear of the mesh and reduced velocity by up to 48 percent. The copper mesh generated fragments with small degree and reduced velocity by up to 97 percent. The experimental result showed that the metal mesh material is available as a first bumper material. In addition, mesh materials may reduce the total weight of debris shields drastically., 資料番号: AA0047275022

A prototype model of high efficient Active Integrated Antenna (AIA) has been developed, as a candidate of a transmission unit in SSPS (Space Solar Power System). The developed AIA consists of three layers, Receiving part, Phaseshifting part, and Transmitting part. Size is 360 mm x 360 mm x 70 mm. Receiving part has the function to receive an external standard signal with 9 receiving antennas, and adjusts the received signal power level and phase. Phaseshifting part has the function to control the phase of the microwave signal outputted from Receiving part by phase shifter according to the phase control command sent from an external phase control equipment. Transmitting part amplifies the microwave signal outputted from Phaseshifting part with the pre-amplifier and high power amplifier of a semiconductor. And Transmitting part has the phased array antenna composed of 9 microstrip antennas of 3 by 3. Moreover, amplified microwave is transmitted from 9 antennas and a power transmission microwave beam is formed. Fundamental characteristic verification test of the prototype AIA has been also performed to acquire characteristic data. The result shows that the AIA will be an applicable candidate for the microwave power transmission unit of SSPS. This paper describes an outline of the prototype model of AIA., 資料番号: AA0047275027

Laboratory tests on charging and arcing of the solar arrays for satellites are carried out in a simulated geosynchronous Earth orbit (GEO) and low Earth orbit (LEO) environment irradiating a solar array test coupon with an electron beam and gas plasma respectively in a vacuum chamber. In this study, the trigger arc inception on the solar array is examined at the inverted potential gradient. And, also examined is whether the interstring sustained arc occurs between adjacent cells of the array. In order to investigate sustained arc occurrence, applied were various potential differences with 0.55 A of line current in GEO environmental test, and applied were various potential differences, line current in LEO environmental test. The location of arc was observed by video images, and acquired were voltage and current waveform by several probes and data acquisition system. Secondary arc can be determined by those two methods. In GEO environmental test, secondary arc occurred frequently, especially under the severe condition. However, sustained arc did not occur, and other following conclusions were obtained. The possibility of being secondary arc and secondary arc duration also increases as the interstring voltage increases. In addition, it can be found that secondary arc duration increases proportional to the pressure increase. In LEO environmental test, the possibility of being secondary arc increases with interstring voltage. This result is well corresponding to the result of GEO environmental test. And, it can be confirmed that increasing line current may lead to an increase of secondary arc duration. In the case of changing line current, permanent sustained arc occurred at 3.0 A line current, in which case the interstring gap was 1 mm, and interstring voltage was 100 V. It was thought that this value may be used as the reference value of solar array design to avoid sustained arc., 資料番号: AA0047275010

The fuel cell for space applications has special system requirements, operational conditions, and designs due to its isolated low gravitational and closed environment, which is much different from that for ground applications. JAXA is now developing polymer electrolyte fuel cell (PEFC) system for the future space applications. It was found that the counter flow of pure hydrogen and oxygen enabled realization of a fuel-cell system without external humidification. Also demonstrated was closed gas operation to determine the applicability of the fuel-cell system in a closed environment, where hydrogen gas was dead-ended and oxygen gas was circulated with produced water removed. Considering gas/water separation in the micro-gravitational environment, a condenser was installed which condenses the humid gas exhausted from cell stack to water, and also installed was a cylindrical separator which contains wicking material to separate gas/water, in the gas recycling line together with a gas circulation pump. The wicking material which absorbed water was compressed by the piston in the separator and discharged the water into the receptacle. Water discharge had an operational cycle to be controlled in accordance with a water capacity of the quantity of electricity to be generated. The 18 cell stack was manufactured and a fuel cell system was assembled including the components mentioned above. It showed stable performance for 1,100 hours under various operational conditions. These results could realize the simple and practical PEFC system for space application., 資料番号: AA0047275014

As growth in electric power of spacecraft, development of less-expensive and lighter space power system is required. Especially, for future high power spacecraft such as the hybrid solar sail that has the potential to explore the Jupiter and asteroids, or Solar Power Satellite, very light weight photovoltaic solar array system with specific power exceeding 1 kW/kg will be required. It is expected to reduce the mass of the solar array system dramatically utilizing the thin-film solar cells. Recently, as low-cost solar cell, amorphous silicon (a-Si) solar cells and CIGS solar cells are developed for the terrestrial commercial product conducted by New Energy and Industrial Technology Development Organization. In addition, high radiation tolerant of CIGS and a-Si solar cells was reported. In this paper, possibility of the terrestrial thin-film solar cell 's technologies for space applications is described., 資料番号: AA0047275005

Large-structured, high-performance spacecraft for scientific observation in space have recently been operated, so that large amount of primary power with solar cells is required to achieve their mission objectives. To resolve such a requirement, the asteroid sample return mission spacecraft, MUSES-C, which carried multijunction solar cells with a high efficiency of 26 percent and was launched at Kagoshima Space Center (KSC) in May, 2003, has paved the way for the application of the multijunction solar cells. Their widespread spectral response to the extraterrestrial sunlight, the so-called air mass zero (AM0) sunlight, can range from ultraviolet to infrared rays. In general, solar cells under development are required to withstand the AM0 sunlight during a long duration in order to be qualified as space solar cells. That's what is called aging. At their practical stage, power generation capability of solar cell panels must be evaluated by means of the AM0 sunlight after the installation of the solar cells in the panels. However, there is not a solar simulator with three kinds of lamp and the AM0 spectral characteristic for the multijunction solar cells in the world yet. It is thought that such a simulator is essential to the evaluation of future multijunction solar cells. Thus, first of all, it is described how to design a new multi-solar simulator with the three kinds of lamp: Hg lamps with high brightness in the UV region, Xe lamps in the visible region, and halogen lamps in the IR region. In its performance, it is possible to achieve an illumination intensity of 1.4 kW/sq m a projection area of 2.5 x 2.5 sq m, and a uniformity of +/- 15 percent. In addition, a spectrometer of 250 to 1,850 nm in wavelength is required to maintain a precise simulator performance specified for more than hundreds of hours., 資料番号: AA0047275007

In Japan, plans for a demonstration SPS experiments were proposed from the Ministry of Education, Culture, Sports, Science and Technology (MEXT), and the Ministry of Economy, Trade and Industry (METI). In the METI's plan assumes that existing launch technology will be used for an early actual proof. In this demonstration satellite, due to the limit of transmitting antenna size and generation of power, power flux density on the ground (PFD) becomes about 0.3 W/sq m. This satellite assumes a subrecurrent orbit and its altitude is 370 km. For this reason, it is considered that the rectenna site is constructed in the place where the satellite passes the zenith. The lowest operation power density of conventional rectenna was about 100 W/sq m, so, present power density is not the sufficient strength so that rectenna may operate. So, receiving power can be increased by using parabola antenna, however, the power reception time from the satellite drastically decreases, since the beam profile becomes pencil beam. In order to lengthen the power reception time, an elliptical beam with wide-angle directivity in the satellite's direction of movement, and narrow directivity perpendicular to it is desired. A rectangular parabola rectenna with an elliptical beam was proposed, which can be formed by changing the reflector's rectangular shape and the directivity of the primary feed. A theoretical analysis was performed by the PO method and confirmed its validity using experiment. The target antenna gain was set to 32 dBi, and it is a condition for applying sufficient input power to the rectifying circuits. The - 3 dB beam width was designed +/- 4 degrees for the major-axis direction. The calculated antenna gain and the aperture efficiency were 32.0 dBi, 48 percent, respectively. Since, this value changed to 31.8 dBi and 46 percent in the experiment, respectively. For low input power flux density, good low input power characteristic rectifying circuits were adopted. The maximum value of the present rectifying circuit efficiency was 75 percent for a 350-mW input. From these data, it was proven that the efficiency of rectenna could be expected to about 33 percent. It is considered that this value is an enough practicable value, when the power density is low with 0.3 W/sq m is considered., 資料番号: AA0047275028

In the future, LEO spacecraft will be larger and higher powered. Because of the balance of leakage currents through ambient space plasma, their main conductive body will have a higher negative potential without plasma contactor operation. When spacecraft operate with a higher voltage, more intensive arcing is suspected to occur on the surface. In this study, ground-based experiment and Direct-Simulation-Monte-Carlo Particle-In-Cell plasma simulation were carried out to understand the arcing phenomenon and to examine influences of ambient space plasma on the arcing process. Simulating plasmas were generated by electron cyclotron resonance discharge. When arcing occurred on negatively-biased anodized aluminum sample (AAS) plates in the plasma environment, the time variations in arc current and bias voltage were measured. Arc spot diameter was also measured. The experimental results showed that both the peak arc current and the total charge emitted by arcing increased with initial charging voltage and neutral particle number density. The diameter of arc spots increased with initial charging voltage although it was almost constant regardless of neutral particle density. The calculated results showed that neutral particles in addition to charged particles around spacecraft played an important role in expansion of arc plasma causing the arcing characteristics. Accordingly, high voltage operation of LEO spacecraft might bring drastic degradation of AAS by arcing, depending on ambient plasma conditions., 資料番号: AA0047275009

In order to protect space structure against space debris impacts, it is indispensable to develop a bumper shield with strength materials. The high strength fiber is one of potential materials from a viewpoint of strength, lightweight, and flexibility. The purpose of this study was to develop a single bumper shield using high strength fibers and new bumper materials against medium size debris impacts. Developed was a multilayer bumper shield using Vectran fibers, a fiber net, a polyester sheet, and a shock absorber. Hypervelocity impact tests were carried out by a railgun accelerator of the Institute of Space and Astronautical Science. The experimental results showed that the developed shield (the single bumper) could stop the polycarbonate projectile with 13.9 mm in diameter, 0.9 g in weight, and 6.9 km/s in velocity. The thickness of the new shield is about 21 mm, and the total areal density is 15 kg/sq m. The projectile impacted the single bumper shield and produced a hole of 80 mm in diameter. But the projectile was destroyed into the shield, and did not reach on surface of an aluminum block that is located behind the shield., 資料番号: AA0047275021

Electron-beam irradiation experiments concerning the arcing discharges on solar arrays were carried out. It is very important to make this discharge mechanism clear and offer a guideline for designing the solar arrays of high reliability. We used a pair of real GaAs solar cells on a substrate as a sample. The sample was biased to -9 kV and the electron beam of the energy of 10 keV was irradiated to the sample. Under that condition, the voltage from DC battery was applied to the gap between the cells. However, the detrimental arcing did not occur even under several conditions of the cell gap distance and the cell gap voltage below 120 V, although a few hundreds of ESD took place. Therefore, the high voltage use of solar arrays is thought not to cause sustained arcing discharge to destroy the satellite system if the arrays are designed to lower the maximum voltage between cell strings with a reasonable distance., 資料番号: AA0049120017

The space solar power system (SSPS) is required for clean power source in the near future. The SSPS consists of not only conventional spacecraft technologies but also newly developed technologies. In development of the SSPS the new technologies should be inspected both on orbit experiment and on the terrestrial experiment. However the experiment cost on orbit is high, attention should be paid to developing simulation method as alternative to experiment on orbit to cut the cost down. In this paper described is development of a novel simulation method for analyzing microwave beam propagation characteristics from an SSPS orbit to ground beam detection point. Several studies for microwave beam propagation in space plasma important for SSPS had been done. These studies were focused on the nonlinear effects appeared in the extremely high power microwave propagation in the space plasma. In contrast with these studies, overall propagation analysis was focused on from the orbit to the ground. The simulation method developed is based on the beam propagation method (BPM) on a conventional PC platform. To overcome the increasing calculation error in long distance propagation with beam size expansion, alternative direction implicit method (ADI) is introduced. The databases IRI, COSPER, and IGRF are used to take account of space environment (ionosphere) and the atmosphere on the microwave propagation path. Simulated are microwave beam propagation characteristics of 5 m of beam radius from altitude 570 km through ionosphere and atmosphere. Good agreement was obtained with the theoretical expectation but the simulation took 6 days CPU time on PC platform. As the code with ADI method suited for parallel computing, the calculation time could be shortened to 2 days with 4 CPU through 570 km propagation., 資料番号: AA0047275025

For realization of Space Solar Power System (SSPS), it's essential to develop Microwave Power Transmission (MPT) System with high DC-RF conversion efficiency and light weight. MPT system for SSPS requires very accurate beam control (less than 0.001 degree) with huge phased arrays whose diameter is over 1.0 km. It also requires total efficiency of more than 80 percent (excluding conversion efficiency of solar cells) to work as a space system. To satisfy these requirements, a Phase Controlled Magnetron (PCM) is suitable for microwave transmitter. The PCM takes advantage of high power, high DC-RF conversion efficiency, low cost, and light weight of a magnetron. Typical output power from the PCM is 300 W, while output power from one antenna element is less than 1 W. Therefore insertion of power divider and phase shifters after PCM output is required to satisfy SSPS parameters and it leads to disadvantage. Because the power dividers and the phase shifters are inserted after high power microwave output, insertion loss in these elements is very critical. The loss after the PCMs must be reduced to realize the PCM-based MPT system. Slot-coupled Active Integrated Antenna (AIA) system is proposed as one of the solution of this Final Stage problem. This system uses waveguide slot antenna as a power divider and AIA as low-loss phase shifter and antenna layer. By using an electromagnetic coupling between an AIA and waveguide slot antenna, thin and light antenna structure can be realized. This system also has low insertion loss and is suitable for high power from the PCM because of the use of the waveguide as an input port. One of the important factors of the slot-coupled AIA is coupling factor of the electromagnetic coupling between microstrip lines and waveguide slots. The aim of this study is to improve the coupling factor by reconsideration of the electromagnetic coupling structure. Coupling factor of nearly 90 percent was achieved by changing waveguide dimensions., 資料番号: AA0047275026

資料番号: AA0047120030

大気をもつ惑星・衛星への探査プローブ投入など、将来の惑星突入ミッションに対応して突入カプセルの耐熱材料の地上試験を行うためには、地上にて突入飛行環境を模擬した高エンタルピ気流を発生させる装置が必要となる。本稿は、主に惑星大気突入環境を模擬する装置としての小型誘導加熱型高温気流発生装置の設計の考え方を示し、実際に製作し、始動試験を行った結果を報告する。, 資料番号: AA0047120029

資料番号: AA0047118070

A degradation test for a solar array coupon against ESD was performed under stimulated Low Earth Orbit environment. All tests were performed in a vacuum chamber with a plasma source. A test coupon was biased at - 400 V with the aim of developing the next generation 400 V high voltage solar array. The LCR circuit was used in order to simulate the arc current gathering the charge stored on coverglass. Tests were repeated until the solar array coupon was damaged. All of arc spots and waveforms were detected. The electrical performance of the coupon was measured after every test without opening the vacuum chamber. Many arcs occurred and caused the cell degradation. The cell was damaged by only one arc that occurred at the edge of cell, not at electrode., 資料番号: AA0046978008, レポート番号: JAXA-SP-03-001

宇宙航空研究開発機構(JAXA)が開発中の静止軌道人工衛星のWINDS(超高速インターネット衛星)とETS-8(技術試験衛星8号機)について、NASA Charging Analyzer Program(NASCAP)を用いて表面帯電解析をおこなった。衛星表面帯電の影響として、衛星構造体(衛星アース)と誘電体、特に太陽電池表面のカバーガラスとの間の電位差により起こる放電(Electrostatic Discharge: ESD)の可能性がある帯電時間と、日照時での放電をトリガーとする持続放電による太陽電池の劣化とを見積もった。静止軌道の宇宙プラズマ環境パラメータとして、Los Alamos National Laboratory(LANL)の3つの静止衛星の延べ約12年間分のデータの統計解析結果を用いた。その結果、衛星により帯電特性が異なることが明らかになった。, The total duration is evaluated in the condition that the Electrostatic Discharge (ESD) could occur and the number of expected trigger arcs for the secondary (sustained) arcs on solar array for the geostationary satellites, the Wideband InterNetworking engineering test and Demonstration Satellite (WINDS) and the Engineering Test Satellite 8 (ETS-8). For the evaluation, NASA Charging Analyzer Program (NASCAP) is used with the statistical plasma parameters observed by the Los Alamos National Laboratory geosynchronous satellites., 資料番号: AA0046978015, レポート番号: JAXA-SP-03-001

The total dose data from a Tsubasa (MDS-1) satellite, flying in a highly eccentric orbit, is analyzed. The total dose is measured by the small dosimeter utilizing RADFETs. The RADFETs have been calibrated with the Co-60 gamma ray source and 56 devices (denoted DOS-S) were mounted in several experimental modules in the satellite. The total dose data behind certain Aluminum shield domes have been analyzed initially. The total dose change is affected by electron flux in the thin shield. In thicker shield, electron and proton affect the total dose profile., 資料番号: AA0046978014, レポート番号: JAXA-SP-03-001

Space environment is complex, and spacecraft surfaces can be simultaneously subjected to bombardment by energetic charged particles, solar ultraviolet radiation, atomic oxygen, and hypervelocity debris and micrometeoroids. Differences between the results of ground-based and on-orbit materials tests are often ascribed to synergism between space environmental factors, which are generally difficult to reproduce in ground tests. In 1997, King et al. reported that the yield of volatile products formation (primarily CO and CO2) by the interaction of hyperthermal (4.8 eV average energy) atomic oxygen with polysulfone and polyimide under charging conditions (AIAA 97-3901). When the target was exposed simultaneously to atomic oxygen and high-energy electron beams, the increase in reaction yield was clearly observed compared to atomic oxygen exposure alone. They concluded that this phenomenon was due primary to target charging. The research project, which investigates the relationship between surface charging and atomic oxygen-induced reaction, is under planning at Kobe University using a laser detonation atomic oxygen beam source and a quartz crystal microbalance. The purpose of this project is to clarify the detail of this phenomenon quantitatively. The overview of this project and preliminary results obtained are presented herein., 資料番号: AA0046978011, レポート番号: JAXA-SP-03-001

Spacecraft in space environment are exposed in plasma or are irradiated by high energy cosmic rays. Dielectric materials on the surface of spacecraft are usually charged up in such environment. The accident gives on unrecoverable damage to the spacecraft. Therefore it is necessary to analyze the mechanism of surface charge distribution on insulating materials in space environment. The measurement system has been developed for surface discharge pattern on BSO cell which is used as an image detecting device applying Pockels effect. By using this optical method, the 2-dimensional charge distribution can be measured during surface discharge. Furthermore with high speed CCD camera, the time dependent 2-dimensional images can be detected. Here, typical results of surface charges on Kapton film putting on BSO cell will be shown., 資料番号: AA0046978010, レポート番号: JAXA-SP-03-001

This paper deals with the electron-beam irradiation experiments concerning the arcing discharge on the solar array. It is very important to investigate this discharge phenomenon and offer a guideline to design solar arrays with high reliability. A pair of real GaAs solar cells on a substrate were used as a sample. The sample was biased to - 9 kV and the electron beam of the energy of 10 keV was irradiated to the sample. On that condition, the voltage from DC battery was applied to the gap between the cells. The detrimental arcing did not occur even at the gap voltage of about 80 V, in spite of the occurrence of ESD of a few hundreds times. Therefore, the high voltage use of solar array, for example 100 V, is thought not to cause sustained arcing discharge to destroy the satellite system if arrays are designed to lower the maximum voltage between cell strings with reasonable distance., 資料番号: AA0046978009, レポート番号: JAXA-SP-03-001

An overview is given on the 6 experimental components of the Mission Demonstration Satellite-1 (MDS-1, Tsubasa in Japanese) mission in the space radiation environment. Various aspects related to the radiation shielding and to the radiation-induced effects on instrumentation will be also discussed. For the benefit of future space missions, relevant lessons learned concerning space radiation environment effects, and on use of COTS (Commercially available Off The Shelf) devices and components in harsh radiation environment will finally be summarized., 資料番号: AA0046975009, レポート番号: JAXA-RM-03-022

Tsubasa (MDS-1) was launched by H-2A launch vehicle TF-2 on Feb 4th 2002. Tsubasa exceeded the planned 1-year orbit mission life. To satisfy the short term, low cost and reliable satellite, new technology and approach are applied for Tsubasa system. This paper describes the new technologies of Tsubasa bus development, the evaluation of each approach and comparison with NASA LEWIS satellite., 資料番号: AA0046975002, レポート番号: JAXA-RM-03-022

This Paper has been written in the framework of a multipartnership between the French Space Agency (CNES) and the Japanese Space Agency (JAXA). One aim of this partnership is to encourage the production of common standards of benefit to the entire European space industry and Japanese space industry. This proposal documents deal with the general requirements for using commercial parts in space applications. These requirements are defined in terms of what must be accomplished, rather than in terms of how to organize and perform the necessary activities. This allows existing organizational structures and methods to be applied where they are effective, and for the structures and methods to evolve as necessary without rewriting the standard. This proposal documents define requirements for all aspects of the selection, procurement and usage of EEE commercial components for use in space applications. This proposal documents are applicable to all parts which are defined and procured using specifications and qualification systems other than the ones traditionally associated with space quality parts (e.g. MIL level S or equivalent). It is also applicable for all actors involved either directly or indirectly and at any level in the design or production of hardware for use in all space segments and mission types. The objective of this proposal documents are to define part selection, procurement, control and usage requirements which, when applied to a space project, will ensure that any commercial EEE components used do not prevent the project from meeting its defined functionality, environmental conditions, quality, reliability, schedule and cost., 資料番号: AA0046975007, レポート番号: JAXA-RM-03-022

Demonstration results of radiation resistance of six types of solar cells developed for terrestrial use in space by MDS-1 Tsubasa satellite are shown. The satellite has flown in the geostationary transfer orbit where radiation environment is quite severe. Monitored open-circuit voltage and short-circuit current values of the cells for duration of about 600 days after launch are analyzed. Super radiation tolerance of CuInGaSe2 solar cell is firstly demonstrated in actual space. Also, most of the cells are recognized to have capability or applicability for space use. Comparison of degradation trend of the flight data with that of ground irradiation test results using the displacement damage dose method reveals good agreement, indicating sufficient validity of ground test technique and accuracy of end-of-life prediction., 資料番号: AA0046975008, レポート番号: JAXA-RM-03-022

Seven kinds of COTS memories were installed in Commercial Semiconductor Devices (CSD) and their single event effects were observed. Single Event Upset (SEU) and Multi-Bit Upset (MBU) were observed in DRAMs and SRAMs, and no Single Event Latchup (SEL) was observed in all memories. The results showed that most of SEU and MBU was caused by protons in Van Allen belt, and that the number of MBU was considerably large to neglect. They also showed that some of the on-orbit data were not consistent with the results obtained from the ground test. This implies that the structure of recent memories have become too small to predict accurate SEU rate by present model., 資料番号: AA0046975005, レポート番号: JAXA-RM-03-022

The total dose data from Tsubasa (MDS-1) satellite, flying in a highly eccentric orbit, is analyzed. The total dose is measured by the small dosimeter utilizing RADFETs. 56 devices (denoted DOS-S) were mounted in several experimental components in the satellite. The total dose data behind certain Aluminum shield domes and in some experimental components have been analyzed. The total dose change is affected by electron flux in the thin shield. In thicker shield, electron and proton affect the total dose profile. The comparison between the flight data and model data were implemented., 資料番号: AA0046975004, レポート番号: JAXA-RM-03-022

Space radiation environment has become a primary consideration in designing a modern spacecraft and in mitigating the risk factors for astronauts. Modem electronic devices and optical sensors are becoming more sensitive to high energy particle environments. This paper addresses present state and problem of the radiation environment model, and comparison between MDS-1 observation data and predicted value. Finally, a new radiation design standard handbook is reported., 資料番号: AA0046975003, レポート番号: JAXA-RM-03-022

This paper shows the analysis results of the Total Ionizing Dose (TID) Effects on COTS Memories by MDS-1 (Tsubasa) satellite. Seven types of commercial devices are installed and evaluated. The measurement result basically consists with the prediction from the ground test data in all devices. But, in DRAMs, the change of the electrical characteristic caused by TID effect surpasses the prediction. This result indicates that the criteria of the ground TID evaluation test should be carefully determined. These results will be reflected in the establishment of guideline for the ground test technique., 資料番号: AA0046975006, レポート番号: JAXA-RM-03-022

資料番号: AA0045496004

資料番号: AA0045496003

資料番号: AA0045496005

資料番号: AA0045441067