研究者業績
基本情報
- 所属
- 国立研究開発法人宇宙航空研究開発機構 宇宙科学研究所 教授
- 学位
- 博士(工学)(1996年3月 名古屋大学)
- 連絡先
- ogawa.hiroyuki
jaxa.jp - J-GLOBAL ID
- 200901051344540154
- researchmap会員ID
- 1000253790
- 外部リンク
将来の科学衛星に向けた先進的熱制御システムの研究
科学衛星プロジェクトの経験を基に，現状の課題と将来計画を分析し，将来の科学衛星に必要な先進的熱制御システムの研究開発をおこなっています．研究成果はX線天文衛星ひとみに搭載された熱制御システムにフィードバックされている他，次期科学衛星計画への適用が検討されている等，科学衛星の可能性を広げ，世界一流の成果を創出する活動に貢献しています．
科学衛星プロジェクトの熱制御
日欧水星探査計画BepiColombo等のこれまで経験のない極限環境に晒される探査機や，X線大型望遠鏡衛星ひとみ等の熱流体デバイスを積極的に採用した挑戦的プロジェクトにおいては，従来の衛星開発手法やその延長線上では対応できず，これまで経験のない新しい衛星開発手法が求められます．極限環境に耐える新規材料開発や熱設計・解析手法の構築，試験設備整備や検証手法の開発など，新しい研究開発を熱流体力学の学術的知見をもって先導し，熱の観点でプロジェクトの成功に貢献しています．
熱流体力学の応用
熱流体とその周辺の学術的知見を基に，さまざまな宇宙科学プロジェクト活動に貢献しています．再使用ロケットの研究では，エンジン流れや極低温タンク，外部流等熱流体にかかわる課題解決に貢献しています．衛星推進系ではヒドラジンスラスタ内部化学反応流の研究によりスラスタ解析技術の向上に貢献し，ロケット推進系では固体ロケット内部流解析手法を開発し，M-VロケットやSRB-Aの不具合原因究明に貢献しました．その他，ロケットの飛行安全やロケット排気プルームの電波干渉問題等に関わり，ロケット研究に貢献しています．また高速電磁流体中の衝撃波干渉の理論研究や電磁流体を利用した推進システムの研究をおこないました．
経歴
6-
2017年1月 - 現在
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2003年10月 - 2016年12月
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2002年1月 - 2003年9月
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1998年4月 - 2001年12月
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1996年4月 - 1998年3月
学歴
1-
- 1996年3月
委員歴
1-
2013年3月 - 2015年2月
受賞
1-
2015年
論文
80-
International Journal of Heat and Mass Transfer 221 125037-125037 2024年4月
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Applied Thermal Engineering 121109-121109 2023年7月
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Nuclear Instruments and Methods in Physics Research Section A: Accelerators, Spectrometers, Detectors and Associated Equipment 1049 168102-168102 2023年4月
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Applied Thermal Engineering 219 119573-119573 2023年1月A capillary pumped loop (CPL) is a capillary force-driven heat transport device that uses the phase change of a working fluid. A CPL has excellent properties in terms of heat transport capability, heat transport distance, and flexibility in heat transport path arrangement. However, its startup reliability is low, and its startup characteristics have yet to be determined. This study aims to investigate the startup characteristics of a CPL and improve its startup reliability. The startup procedures are newly proposed and experimentally verified in this paper. In the proposed method, the evaporator was preheated in addition to the reservoir preheating before startup. Verification was done by switching the order of the reservoir preheating and evaporator preheating. The experimental results showed that the proposed methods enable the CPL to start up even in conditions where the CPL was unable to start up by a conventional method (reservoir temperatures ranging from 40 °C to 80 °C). It was also confirmed that the order of the evaporator preheating and reservoir preheating affects the degree of overshoot of the evaporator temperature. The maximum temperature of the evaporator at the startup was reduced by up to 91.5 °C in case the evaporator was preheated before the reservoir preheating.
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JOURNAL OF SPACECRAFT AND ROCKETS 2022年2月Typically, spacecraft development is costly and time-consuming because of the many iterations usually needed to reach optimal design solutions. This paper presents an innovative approach that eliminates the need to iterate the thermal design process using a network of variable conductance oscillating heat pipes (VC-OHPs) on every structural panel. The temperatures of the panels where components are mounted would thus be maintained at constant levels by VC-OHPs, even if the instruments' locations or heat dissipation changes. A structural thermal model was built to verify the proposed thermal and structural design in a simulated deep space environment and in a launch environment. It consisted of two VC-OHPs and six aluminum honeycomb panels. A thermal vacuum test was conducted to demonstrate the temperature control by the VC-OHPs. The test results showed that temperature control by VC-OHPs could maintain the panels operating as evaporators at stable temperatures and follow the reservoir temperature. A vibration test was conducted under the launch environment of a Japanese H2A rocket. The results confirmed that the structural thermal model met requirements for resistance to mechanical launch environment. The VC-OHPs functioned after the vibration test. The structural thermal model showed that the proposed thermal control architecture is feasible in an actual spacecraft in terms of thermal and structure design.
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Journal of Low Temperature Physics 211(5-6) 384-397 2022年LiteBIRD is a future satellite mission designed to observe the polarization of the cosmic microwave background radiation in order to probe the inflationary universe. LiteBIRD is set to observe the sky using three telescopes with transition-edge sensor bolometers. In this work we estimated the LiteBIRD instrumental sensitivity using its current design. We estimated the detector noise due to the optical loadings using physical optics and ray-tracing simulations. The noise terms associated with thermal carrier and readout noise were modeled in the detector noise calculation. We calculated the observational sensitivities over fifteen bands designed for the LiteBIRD telescopes using assumed observation time efficiency.
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APPLIED THERMAL ENGINEERING 198 2021年11月We compared the thermal performance of two configuration types of Oscillating Heat Pipe (OHP) experimentally to investigate making heat transfer routing more adaptable. One type of OHP has the conventional configuration of closed-loop multiple-serpentine routing that uses U-shaped turns of 3.75 mm radius. The other type has the proposed circular ring configuration with a much larger radius of 1289 mm. Both types have the same pipe length so that the circular OHP can be considered an "unfolded" layout of the serpentine OHP. The circular OHP has alternating heated and cooled sections at locations corresponding to the sections of the serpentine model. Each type has ten sets of these sections with check valves; the OHPs were oriented horizontally. The working fluid was HFC-134a. Both types of samples functioned normally, and their thermal resistances were almost identical. Each OHP showed traveling pressure waves during operation, which indicated a phase difference at each measuring point. The experiment showed that OHPs, with check valves and alternating heated and cooled sections, can transfer heat without the need for a serpentine configuration and that heat transfer routing can be designed more adaptably.
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Applied Thermal Engineering 197 2021年10月Achieving a high level of pointing accuracy in spacecraft requires that vibration and thermal distortions not be relayed to the structure. Therefore, a flexible thermal strap is essential to dissipate heat from heat-and-vibration-generating components, such as cryocoolers. This study developed a flexible and highly conductive thermal strap for space application with a micro-Oscillating Heat Pipes (OHP). So far, few studies have focused on the flexibility of a micro-OHP, particularly that of the tube type. This paper presents the results of testing strap's thermal performance and dynamic stiffness, which is a key parameter of flexibility but has been rarely evaluated for heat pipes, including OHP. To achieve the triaxial flexibility and resistance against space environment, the OHP consists of a 10-turn micro-OHP composed of metal tubes with an inner diameter of 0.4 mm or less. It is filled with HFC-134a as the working fluid. Ten electroformed check valves are mounted in the OHP. The liquid slug is induced to flow from the condenser to the evaporator, and the OHP operates stably even in a horizontal position. Thermal performance tests have shown the maximum thermal conductance of the strap to be 0.8 W/K and the heat transfer rate to be at least 7 W. The thermal vacuum test (TVT) showed that, below 20 ℃, the thermal conductance decreases with the condenser temperature and that the OHP does not operate below −15 °C. Dynamic stiffness tests were performed on the OHP to determine the dynamic stiffness in the two directions perpendicular to the flow (along the Y and Z axes). The stiffness parallel to the flow was obtained by a finite element methods analysis using a structural model of a bent OHP based on earlier test results. The dynamic stiffness is less than 0.2 N/mm in Y and Z axis, which is less than that of the graphite thermal strap used for space application. During and after the stiffness test, the OHP continued to operate stably without any change in the temperatures of the evaporator and condenser.
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SPACE SCIENCE REVIEWS 216(7) 2020年10月Mercury has a unique and complex space environment with its weak global magnetic field, intense solar wind, tenuous exosphere, and magnetospheric plasma particles. This complex system makes Mercury an excellent science target to understand effects of the solar wind to planetary environments. In addition, investigating Mercury's dynamic magnetosphere also plays a key role to understand extreme exoplanetary environment and its habitability conditions against strong stellar winds. BepiColombo, a joint mission to Mercury by the European Space Agency and Japan Aerospace Exploration Agency, will address remaining open questions using two spacecraft, Mio and the Mercury Planetary Orbiter. Mio is a spin-stabilized spacecraft designed to investigate Mercury's space environment, with a powerful suite of plasma instruments, a spectral imager for the exosphere, and a dust monitor. Because of strong constraints on operations during its orbiting phase around Mercury, sophisticated observation and downlink plans are required in order to maximize science outputs. This paper gives an overview of the Mio spacecraft and its mission, operations plan, and data handling and archiving.
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INTERNATIONAL JOURNAL OF HEAT AND MASS TRANSFER 158 2020年9月This paper presents an effect of reservoir location on heat transfer characteristics of a capillary pumped loop (CPL). The designed CPL has a cylindrical evaporator with a diameter of 12 mm and a length of 65 mm. A stainless steel porous medium that has a pore radius of 4.4 mu m is used as a wick. Heat transport length is designed as 1000 mm. The experiments were conducted with four types of reservoir location such as (1) Distance from the evaporator, L = 150 mm, (2) L = 250 mm, (3) L = 750 mm, (4) L = 1450 mm. As results, it was found that the CPL of case (1) showed high operating temperature and much smaller heat transfer capability such as 10 W. For cases of (2) - (4), the maximum heat transfer capability reached 140 - 160 W and showed the good controllability of the operating temperatures ranging from 60 to 100 degrees C. The reason for the low heat transfer performance of case (1) and the difference in operational characteristics are discussed. It was found that there is a design limitation of the reservoir location of CPLs. In addition, the operation of the CPL that reservoir was located much far from the evaporator was successfully demonstrated. (C) 2020 Elsevier Ltd. All rights reserved.
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APPLIED THERMAL ENGINEERING 165 2020年1月This paper reports the design, fabrication, and testing of a reversible thermal panel breadboard model (RTP-BEM). 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 degrees 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 degrees C.
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Journal of Low Temperature Physics 194(5-6) 443-452 2019年3月15日LiteBIRD is a candidate satellite for a strategic large mission of JAXA. With its expected launch in the middle of the 2020s with a H3 rocket, LiteBIRD plans to map the polarization of the cosmic microwave background radiation over the full sky with unprecedented precision. The full success of LiteBIRD is to achieve δr< 0.001 , where δr is the total error on the tensor-to-scalar ratio r. The required angular coverage corresponds to 2 ≤ ℓ≤ 200 , where ℓ is the multipole moment. This allows us to test well-motivated cosmic inflation models. Full-sky surveys for 3 years at a Lagrangian point L2 will be carried out for 15 frequency bands between 34 and 448 GHz with two telescopes to achieve the total sensitivity of 2.5 μ K arcmin with a typical angular resolution of 0.5 ∘ at 150 GHz. Each telescope is equipped with a half-wave plate system for polarization signal modulation and a focal plane filled with polarization-sensitive TES bolometers. A cryogenic system provides a 100 mK base temperature for the focal planes and 2 K and 5 K stages for optical components.
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日本航空宇宙学会誌 67(9) 310-314 2019年<p>「Advanced Radiation Material（ARM）」の実験では，宇宙機の最外層に搭載する放熱面材料の軌道上の劣化データを取得する．地上試験では加速試験かつ単一の劣化要因の照射試験となるために地上では再現できない宇宙環境に実時間で曝露することにより，現在使用している熱制御材や新規に採用を検討している熱制御材，新規に開発した熱制御材の宇宙環境下での劣化データを取得する．熱制御材は宇宙機放熱面への実際の搭載方法を模擬して接着された状態で曝露して劣化状況の把握・評価を行い，材料選定や熱設計の参考となるデータを取得する．加えて，地上の設備を用いた評価試験結果と比較し，劣化メカニズムを解明する．新規材料の軌道上実証も目的の1つである．本実験では試料の曝露期間を1年間，2年間，3年間とし，それぞれの経過年での熱光学特性，導電性等の劣化データを取得し，地上試験と比較する予定である．これまでに1年間の曝露を終えた試料が帰還しており，これらの試料を分析している．</p>
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Journal of Low Temperature Physics 193(5-6) 1048-1056 2018年12月1日Inflation is the leading theory of the first instant of the universe. Inflation, which postulates that the universe underwent a period of rapid expansion an instant after its birth, provides convincing explanation for cosmological observations. Recent advancements in detector technology have opened opportunities to explore primordial gravitational waves generated by the inflation through “B-mode” (divergent-free) polarization pattern embedded in the cosmic microwave background anisotropies. If detected, these signals would provide strong evidence for inflation, point to the correct model for inflation, and open a window to physics at ultra-high energies. LiteBIRD is a satellite mission with a goal of detecting degree-and-larger-angular-scale B-mode polarization. LiteBIRD will observe at the second Lagrange point with a 400 mm diameter telescope and 2622 detectors. It will survey the entire sky with 15 frequency bands from 40 to 400 GHz to measure and subtract foregrounds. The US LiteBIRD team is proposing to deliver sub-Kelvin instruments that include detectors and readout electronics. A lenslet-coupled sinuous antenna array will cover low-frequency bands (40–235 GHz) with four frequency arrangements of trichroic pixels. An orthomode-transducer-coupled corrugated horn array will cover high-frequency bands (280–402 GHz) with three types of single frequency detectors. The detectors will be made with transition edge sensor (TES) bolometers cooled to a 100 milli-Kelvin base temperature by an adiabatic demagnetization refrigerator. The TES bolometers will be read out using digital frequency multiplexing with Superconducting QUantum Interference Device (SQUID) amplifiers. Up to 78 bolometers will be multiplexed with a single SQUID amplifier. We report on the sub-Kelvin instrument design and ongoing developments for the LiteBIRD mission.
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Journal of Low Temperature Physics 193(5-6) 841-850 2018年12月1日The high-frequency telescope for LiteBIRD is designed with refractive and reflective optics. In order to improve sensitivity, this paper suggests the new optical configurations of the HFT which have approximately 7 times larger focal planes than that of the original design. The sensitivities of both the designs are compared, and the requirement of anti-reflection (AR) coating on the lens for the refractive option is derived. We also present the simulation result of a sub-wavelength AR structure on both surfaces of silicon, which shows a band-averaged reflection of 1.1–3.2% at 101–448 GHz.
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Earth, Planets and Space 70(1) 2018年12月1日© 2018, The Author(s). The exploration of energization and radiation in geospace (ERG) satellite, nicknamed “Arase,” is the second satellite in a series of small scientific satellites created by the Institute of Space and Astronautical Science of the Japan Aerospace Exploration Agency. It was launched on December 20, 2016, by the Epsilon launch vehicle. The purpose of the ERG project is to investigate how high-energy (over MeV) electrons in the radiation belts surrounding Earth are generated and lost by monitoring the interactions between plasma waves and electrically charged particles. To measure these physical processes in situ, the ERG satellite traverses the heart of the radiation belts. The orbit of the ERG is highly elliptical and varies due to the perturbation force: the apogee altitude is approximately 32,200–32,300 km, and the perigee altitude is 340–440 km. In this study, we introduce the scientific background for this project and four major challenges that need to be addressed to effectively carry out this scientific mission with a small satellite: (1) dealing with harsh environmental conditions in orbit and electromagnetic compatibility issues, (2) spin attitude stabilization and avoiding excitation of the libration by flexible structures, (3) attaining an appropriate balance between the mission requirements and the limited resources of the small satellite, and (4) the adaptation and use of a flexible standardized bus. In this context, we describe the development process and the flight operations for the satellite, which is currently working as designed and obtaining excellent data in its mission.[Figure not available: see fulltext.]
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APPLIED THERMAL ENGINEERING 141 20-28 2018年8月In this study, a newly proposed heat pipe system was investigated to transfer heat from a vertical heated plate to a vertical cooled plate arranged in parallel. The heat pipe system comprises 32 loops connected in series and a reservoir. Each square-shaped loop (with a side length of 2 m) comprises a capillary tube with an inner diameter of 1.0 mm without any internal wick. The system's overall thermal performance was investigated at room temperature using R410A as the working fluid. Temperatures, pressures, and reservoir weight were monitored, and thereby confirming that the system transfers heat up to several hundred watts by a passive two-phase flow. Numerical simulations with a simple model were consistent with the data and verified that the saturated pressure of the system is controlled by the reservoir temperature independent of the amount of heat load.
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JOURNAL OF ASTRONOMICAL TELESCOPES INSTRUMENTS AND SYSTEMS 4(2) 2018年4月The Hitomi (ASTRO-H) mission is the sixth Japanese x-ray astronomy satellite developed by a large international collaboration, including Japan, USA, Canada, and Europe. The mission aimed to provide the highest energy resolution ever achieved at E > 2 keV, using a microcalorimeter instrument, and to cover a wide energy range spanning four decades in energy from soft x-rays to gamma rays. After a successful launch on February 17, 2016, the spacecraft lost its function on March 26, 2016, but the commissioning phase for about a month provided valuable information on the onboard instruments and the spacecraft system, including astrophysical results obtained from first light observations. The paper describes the Hitomi (ASTRO-H) mission, its capabilities, the initial operation, and the instruments/spacecraft performances confirmed during the commissioning operations for about a month. (C) The Authors. Published by SPIE under a Creative Commons Attribution 3.0 Unported License.
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Publications of the Astronomical Society of Australia 35 2018年<title>Abstract</title>Measurements in the infrared wavelength domain allow direct assessment of the physical state and energy balance of cool matter in space, enabling the detailed study of the processes that govern the formation and evolution of stars and planetary systems in galaxies over cosmic time. Previous infrared missions revealed a great deal about the obscured Universe, but were hampered by limited sensitivity. SPICA takes the next step in infrared observational capability by combining a large 2.5-meter diameter telescope, cooled to below 8 K, with instruments employing ultra-sensitive detectors. A combination of passive cooling and mechanical coolers will be used to cool both the telescope and the instruments. With mechanical coolers the mission lifetime is not limited by the supply of cryogen. With the combination of low telescope background and instruments with state-of-the-art detectors SPICA provides a huge advance on the capabilities of previous missions. SPICA instruments offer spectral resolving power ranging from <italic>R</italic> ~50 through 11 000 in the 17–230 μm domain and <italic>R</italic> ~28.000 spectroscopy between 12 and 18 μm. SPICA will provide efficient 30–37 μm broad band mapping, and small field spectroscopic and polarimetric imaging at 100, 200 and 350 μm. SPICA will provide infrared spectroscopy with an unprecedented sensitivity of ~5 × 10−20 W m−2 (5σ/1 h)—over two orders of magnitude improvement over what earlier missions. This exceptional performance leap, will open entirely new domains in infrared astronomy; galaxy evolution and metal production over cosmic time, dust formation and evolution from very early epochs onwards, the formation history of planetary systems.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 16(6) 566-571 2018年<p>In the development of the recent scientific satellites, requirements of increases in size and improvements in shape accuracy of observation instruments have become more stringent. In order to satisfy those requirements, our research group has examined a pointing control mechanism utilizing artificial thermal expansion as a linear actuator. Our previous research indicated that the pointing control mechanism could satisfy a design requirement of the next generation scientific satellites under a certain orbital environment. It is desirable to conduct performance evaluation under various orbital environments to design a flight model. In order to do that, a detailed thermal mathematical model needs to be built. In this paper, estimated parameters in a thermal mathematical model were estimated by correlating numerical simulation with experiments obtained by thermal vacuum experiments. As a result, the temperature errors between the numerical simulation and the experiments were minimized when the thermal conductance of elastic hinges of the pointing control mechanism was lower than expected. This implies that the estimation of the thermal conductance of the elastic hinge needs to be carefully done.</p>
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JOURNAL OF SPACECRAFT AND ROCKETS 55(1) 77-84 2018年1月This study presents the evaluation results of the thermal performance for Japan Aerospace Exploration Agency's X-ray astronomy satellite (Hitomi) ASTRO-H mission. Hitomi was successfully launched on 17 February 2016. The thermal design of the spacecraft and scientific instruments was challenging because the high-energy resolution and the sensitivity needed to achieve the desired scientific objectives required small thermal distortion and a strict temperature range, in spite of large heat dissipation. Hitomi housed four telescopes and six detectors. Each detector had its own radiator and heat pipes for heat dissipation. The most essential mission instrument, the soft X-ray spectrometer, had four loop heat pipes to transfer heat from two sets of compressors and the cold heads of two-stage Stirling coolers. Six fans were mounted close to six cryocoolers for ground cooling in the fairing until just before launch. The heat pipes were operated properly in orbit. Two loop heat pipes for the cryocooler compressors continued operating from the time of launch, thereby maintaining cryocooler temperatures within the allowable temperature range. Comparing the thermal analysis results and flight data, the difference between the predicted and measured temperatures was less than 4 degrees C for more than 95% of all measured temperatures. The thermal control system functioned properly in both launch operation and in orbit, and the thermal mathematical model simulated the actual thermal design of Hitomi.
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Journal of Astronomical Instrumentation 06(02) 1740006-1740006 2017年6月25日 査読有りThe oscillating heat pipe (OHP) is a novel heat-transfer technique used in thermal engineering. Although the OHP offers many technical advantages, it has not yet been actually used in the sky. Motivated by the need to develop a cooling system for use in the balloon-borne General Anti-Particle Spectrometer (GAPS) project, we are developing OHP technologies. To demonstrate the thermal performances of an OHP in real balloon flight conditions, a scaled-down OHP model was launched by a stratospheric balloon in Japan. In this study, we report the results of the flight demonstration.
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Journal of Heat Transfer 138(12) 2016年12月1日© 2016 by ASME. It is reported that the operating temperature of an oscillating heat pipe (OHP) at an operating limit is not dependent on the ambient temperature but that the maximum heat transfer is dependent on this. In this study, using different ambient temperature conditions, a 15-turn OHP filled with HFC-134a as a working fluid was operated until it dries out. The maximum heat transfer was found to vary with changes in the ambient temperature, but the operating temperature at an operating limit, which depends on the filling ratio (FR) of the working fluid, was found to be constant. At the operating limit, the operating temperature decreased with an increase in the FR when the ratio was greater than 50 wt.%. Visualization experiments and calculations were used to confirm that there is an increase in the liquid volume in the OHP in accordance with an increase in the heat input and that ultimately the OHP fills with the liquid, resulting in the failure of OHP operation. In contrast, at the operating limit, when the FR was less than 50%, the operating temperature increased in line with an increase in the FR. In this case, it is assumed that the volume of liquid slugs decreases as the heat input increases, thus causing the OHP to dry out. This theory is explained using α P-V diagram of the working fluid in the OHP. The OHP thermodynamic cycle reaches a saturated liquid or vapor line before it reaches a critical point if a specified volume is shifted from the specified volume at the critical point. The optimum FR for maximum heat transfer is therefore decided by the void ratio at the critical point of the working fluid.
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JOURNAL OF PROPULSION AND POWER 32(5) 1087-1094 2016年9月The main engine of the reusable sounding rocket has been developed at the Kakuda Space Center of the Japan Aerospace Exploration Agency. A total of 54 engine firing experiments were conducted between June2014 and February2015. Throughout these experiments, the advanced capabilities and functions of this engine, such as wide-range throttling, accurate controllability, and health monitoring, were proved to be feasible. This rocket was also required to be reusable for over 100 flights. To confirm the long-life durability of this engine for over 100 flights in a limited experiment period, sequential multiple-firing tests were planned and conducted. Numerical analysis to evaluate damage of the main chamber due to firings was also conducted. The results of the durability tests and numerical analysis showed reuse of this engine for over 100 flights to be feasible.
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INTERNATIONAL JOURNAL OF MICROGRAVITY SCIENCE AND APPLICATION 33(3) 2016年The Reusable Sounding Rocket is a fully re-usable sub-orbital sounding rocket that takes-off and lands vertically from/on the same launch site. The Reusable Sounding Rocket recovers its payload: one can make repeated use of instruments, recover samples, have a big data storage onboard, and so on. The Reusable Sounding Rocket can be launched daily. The target cost for operation is 1/10 of the existing sounding rocket. S-310. The duration of mu g environment is about 3 minutes. The payload bay is in the nose-cone and its volume is 0.8m in diameter and 1m in height. The maximum mass of the payload is 100kg.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 14(ists30) Pi_17-Pi_26 2016年
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SPACE TELESCOPES AND INSTRUMENTATION 2016: ULTRAVIOLET TO GAMMA RAY 9905 2016年The Hitomi (ASTRO-H) mission is the sixth Japanese X-ray astronomy satellite developed by a large international collaboration, including Japan, USA, Canada, and Europe. The mission aimed to provide the highest energy resolution ever achieved at E > 2 keV, using a microcalorimeter instrument, and to cover a wide energy range spanning four decades in energy from soft X-rays to gamma-rays. After a successful launch on 2016 February 17, the spacecraft lost its function on 2016 March 26, but the commissioning phase for about a month provided valuable information on the on-board instruments and the spacecraft system, including astrophysical results obtained from first light observations. The paper describes the Hitomi (ASTRO-H) mission, its capabilities, the initial operation, and the instruments/spacecraft performances confirmed during the commissioning operations for about a month.
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APPLIED THERMAL ENGINEERING 91 1176-1186 2015年12月This paper presents a study on the optimization of a lunar lander for surviving extremely cold lunar nights by using completely passive thermal control. The effects of parameters such as shape, thermal conductance, and effective emissivity on temperature swings of the lunar lander were numerically analyzed. On the basis of the results, some candidate models were formed and the temperature swings of each model were compared. Then, from the viewpoint of feasibility, a particular model was selected. The model was scaled and experimental comparisons were performed. These comparisons proved the efficiency of the proposed model, both numerically and experimentally. (C) 2015 Elsevier Ltd. All rights reserved.
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JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER 29(2) 403-411 2015年4月This study proposes a passively deployed radiator for use as a new thermal control device for small satellites. This radiator can control the amount of heat dissipation by varying its heat rejection area depending on the temperature. This radiator consists of a fin, a baseplate, and an actuator. First, thermal diffusivity, specific heat capacity, total hemispherical emissivity, and solar absorptance of graphite sheets, which are a part of this radiator fin, were measured. From these measurements, it was determined that the thermal conductivity of the graphite sheet varied from 950 to 1490 W/m K, the total hemispherical emissivity varied from 0.22 to 0.31, and the solar absorptance was 0.66. Second, the performance of the radiator was calculated via thermal analysis as a function of size and thickness, and a half-scaled test model of this radiator was designed. Third, the test model was fabricated and its thermal performance was tested under vacuum conditions. The testing showed that the thermal dissipation of this half-scaled radiator was 54 W at 60 degrees C, the fin efficiency difference between the deployed and stowed positions was 0.35, and the specific heat rejection was 188 W/kg.
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Proceedings of the International Astronautical Congress, IAC 11 8980-8988 2015年In order to make the access to space for researches by the sounding rocket much easier and make the opportunities of the rocket launches much frequent, a fully reusable sounding rocket is proposed, and its preliminary system design study is made progress in ISAS/JAXA. Technical demonstrations and system/subsystem designs for a reusable sounding rocket are also underway. The technical key issues to develop an operational system of reusable sounding rocket are 1) a reusable engine, 2) aerodynamics and returning flight mechanics, 3) cryogenic propellant management, 4) health management system, 5) repeated operations. In order to verify these important technologies, 1) repeated engine tests, 2) a model flight demonstration, 3) a reusable insulation development, 4) sloshing tests of cryogenic liquid propellant, 5) a landing gear development, 6) a construction of health management system for hydrogen gas leak, have been completed or are planned now. Progress of these recent activities for development of reusable sounding rocket is summarized in this paper.
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JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER 29(1) 65-73 2015年1月This paper proposes a lunar long-duration method that uses a characteristic of a very low thermal conductivity of lunar regolith. The principle of this method is to put a heater in at the desired depth of the regolith and heat up the regolith layer during lunar daytime. Because of the very low thermal conductivity of regolith, stored heat in regolith propagates gradually and raises the surface temperature during the cold lunar night. By this method, a lunar lander will be kept warm passively during a cold lunar night. A temporospatially small-scale experimental apparatus that simulates the lunar surface environment was fabricated and the feasibility of a passive thermal control method with no electrical power during lunar nighttime was evaluated. The effectiveness of the proposed method was analytically evaluated by considering the relation between experimental result and actual lunar environment.
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SPACE TELESCOPES AND INSTRUMENTATION 2014: ULTRAVIOLET TO GAMMA RAY 9144 2014年12月3日The joint JAXA/NASA ASTRO-H mission is the sixth in a series of highly successful X-ray missions developed by the Institute of Space and Astronautical Science (ISAS), with a planned launch in 2015. The ASTRO-H mission is equipped with a suite of sensitive instruments with the highest energy resolution ever achieved at E > 3 keV and a wide energy range spanning four decades in energy from soft X-rays to gamma-rays. The simultaneous broad band pass, coupled with the high spectral resolution of Delta E < 7 eV of the micro-calorimeter, will enable a wide variety of important science themes to be pursued. ASTRO-H is expected to provide breakthrough results in scientific areas as diverse as the large-scale structure of the Universe and its evolution, the behavior of matter in the gravitational strong field regime, the physical conditions in sites of cosmic-ray acceleration, and the distribution of dark matter in galaxy clusters at different redshifts.
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JOURNAL OF ASTRONOMICAL INSTRUMENTATION 3(2) 2014年11月A meter-scale Oscillating Heat Pipe (OHP) has been developed for the General Anti-Particle Spectrometer (GAPS) project. Two types of OHP routing, U-shaped and O-shaped, have been investigated. For the operation at low temperature, R410A was used as the working fluid. As the result of the investigation, we verified for the first time that both the meter-scale U-shaped and O-shaped OHPs can transfer heat under gravity in a wide temperature range between 20 degrees C and -60 degrees C. Generally, the Oshaped OHP showed better performance than the U-shaped OHP. Both OHP models showed good thermal conductance and a good amount of heat transport under the particular sets of conditions which meet the design requirements. In order to clarify the drive force to operate OHP to further improve the OHP design, the performance difference between the U-shaped and the O-shaped models has been interpreted in terms of the gravity effect and the pressure loss.
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Journal of Astronomical Instrumentation 3(2) 2014年11月1日© 2014 World Scientific Publishing Company. A meter-scale Oscillating Heat Pipe (OHP) has been developed for the General Anti-Particle Spectrometer (GAPS) project. Two types of OHP routing, U-shaped and O-shaped, have been investigated. For the operation at low temperature, R410A was used as the working fluid. As the result of the investigation, we verified for the first time that both the meter-scale U-shaped and O-shaped OHPs can transfer heat under gravity in a wide temperature range between 20°C and -60°C. Generally, the O-shaped OHP showed better performance than the U-shaped OHP. Both OHP models showed good thermal conductance and a good amount of heat transport under the particular sets of conditions which meet the design requirements. In order to clarify the drive force to operate OHP to further improve the OHP design, the performance difference between the U-shaped and the O-shaped models has been interpreted in terms of the gravity effect and the pressure loss.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 12(29) To_4_1-To_4_10 2014年The thermal condition of ASTRO-H under air-cooled environment before launch is investigated with a thermal testing and a computational analysis. The thermal testing shows that the temperatures of devices are confirmed to be within the operating range if the additional fans are used. Moreover, the results of the thermal testing are compared with those of computational results. The computational results of temperature of the devices around the dewar with the additional fans are in good agreement with those of the thermal testing. The good agreement in the condition with the additional fans is because the forced convection, which is a dominant effect, is well captured in the computational analysis. Meanwhile, the computational results of temperature on the side panels are in very good agreement with thermal testing despite the difference in the flow outside satellite by air conditioner: computational analysis models the air flow from the air-conditioner while thermal testing does not. This is because the air-flow is very slow (0.1[m/s] at the side panel locations) and forced convection effects are very small.
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Advances in the Astronautical Sciences 146 697-702 2013年A fully reusable sounding rocket is proposed and conceptually designed in ISAS/JAXA. In phase A in the proposed project, technical demonstrations for key technologies to develop the reusable sounding rocket are planed as follows; 1) reusable engine development and repeated operations, 2) reusable insulation development for cryogenic tank, 3) aerodynamic design and model flight demonstration for returning flight, 4) cryogenic liquid propellant management demonstration, 5) landing gear development and 6) health management system construction. In this paper, the present system design of the reusable sounding rocket and technical demonstrations are summarized.
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SPACE FOR OUR FUTURE 146 697-702 2013年A fully reusable sounding rocket is proposed and conceptually designed in ISAS/JAXA. In phase A in the proposed project, technical demonstrations for key technologies to develop the reusable sounding rocket are planed as follows; 1) reusable engine development and repeated operations, 2) reusable insulation development for cryogenic tank, 3) aerodynamic design and model flight demonstration for returning flight, 4) cryogenic liquid propellant management demonstration, 5) landing gear development and 6) health management system construction. In this paper, the present system design of the reusable sounding rocket and technical demonstrations are summarized.
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Science and Technology of Energetic Materials : journal of the Japan Explosive Society 73(5) 169-174 2012年12月31日
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42nd International Conference on Environmental Systems 2012, ICES 2012 2012年 査読有りA thermal control system (TCS) of a microsatellite is proposed with loop heat pipes (LHPs) including bypass valves. "Free from restrictions in thermal design," all instruments can be mounted anywhere on the internal side of the six structure panels making up the satellite without concern for the thermal design of the entire satellite and other instruments. The temperatures of all instruments are maintained under any attitude (i.e., external thermal environment) by concentrating dissipated heat in a "center heat source" (CHS) using six LHPs mounted between the CHS and structure panels and other heat transport devices. An experimental study and numerical simulation are conducted to validate the microsatellite TCS. In the experimental study, two LHPs are connected to a heat source and a heat load is input to a condenser to simulate the heat input to a radiator in orbit. The heat from the heat source is successfully transported via one LHP if heat is input to the radiator connected to the second LHP. Orbital thermal analyses of the microsatellite are also conducted. Typical low-Earth, geostationary, and polar orbits are investigated for spinning and three-axis stabilized satellites. Heat from the CHS is transported via the LHPs, and the CHS temperature is maintained within the required temperature range in all analysis cases as a result of bypass valve operation. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 10 1-4 2012年The aerodynamic characteristics of a vertical landing rocket are affected by its engine plume in the landing phase. The influences of interaction of the engine plume with the freestream around the vehicle on the aerodynamic characteristics are studied experimentally aiming to realize safe landing of the vertical landing rocket. The aerodynamic forces and surface pressure distributions are measured using a scaled model of a reusable rocket vehicle in low-speed wind tunnels. The flow field around the vehicle model is visualized using the particle image velocimetry (PIV) method. Results show that the aerodynamic characteristics, such as the drag force and pitching moment, are strongly affected by the change in the base pressure distributions and reattachment of a separation flow around the vehicle.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 10(28) Pc_27-Pc_33 2012年A loop heat pipe (LHP) is a two-phase heat transfer device that utilizes the evaporation and condensation of a working fluid to transfer heat, and the capillary forces developed in fine porous wicks to circulate the fluid. LHPs have been gaining increased acceptance for spacecraft missions, and recently, small LHPs on the order of a few hundred watts have been investigated for this purpose. In this study, a 100W class small LHP with a polytetrafluoroethylene wick as the primary wick was designed and fabricated for thermal vacuum testing. The LHP has a thermoelectric converter (TEC) to control the loop operating temperature. The thermal vacuum test was conducted to evaluate the LHP's thermal performance under a space-simulated environment such as ultra-high-vacuum, and black body radiation, except for a gravitational effect. The loop showed large thermal hysteresis before and after the large and small head loads. The TEC was able to control the loop operating temperature with a small amount of electrical power.
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JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER 25(4) 547-552 2011年10月This paper reports the development of an experimental small loop heat pipe with polytetrafluoroethylene (PTFE) wicks. Three kinds of PTFE porous materials were fabricated, and their wick properties were evaluated. It was clarified that the peak pore radii of the PTFE porous materials ranged from 0.8 to 2.2 mu m, and their porosities ranged from 27 to 50%. A small loop heat pipe with PTFE wicks was designed using these properties, and an experimental small loop heat pipe, for which the wick in an evaporator is exchangeable, was fabricated. The experimental loop was tested with 1.2 and 2.2 mu m PTFE. wicks under atmospheric conditions. The test result with a 2.2 mu m PTFE wick showed a high operating temperature, possibly due to vapor leak from the evaporator grooves to the compensation chamber. The test results with the 1.2 mu m PTFE wick showed a lower operating temperature, lower thermal resistance, and stable operation of the loop heat pipe.
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JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER 25(3) 386-392 2011年7月Experiments are conducted to investigate whether an oscillating heat pipe with a liquid reservoir can serve as a thermal control device in space. The oscillating heat pipe consists of a stainless steel capillary tube (inner and outer diameters of 0.8 mm and 1 mm, respectively), which meanders between a heating section and a cooling section 15 times in each direction (30 times in total). A 50 mL reservoir is connected to the oscillating heat pipe via another capillary tube. The 1,1,1,2-tetrafluoroethane (HFC-134a) is used as the working fluid. The heat input to the heating section is increased from 0 and 70 W, in 10 W increments. When the oscillating heat pipe is set horizontally, the temperature of the heating section remains at about the reservoir temperature of 40 degrees C for three orientations of the reservoir with respect to gravity: vertical, horizontal, and vertically inverted. In the top-heating mode, the temperature of the heating section also remains at about the reservoir temperature. The oscillating heat pipe with a liquid reservoir is confirmed to operate as a variable conductance heat pipe, and its operating temperature can be controlled to almost be the liquid reservoir temperature for each investigated orientation of the reservoir. The oscillating heat pipe with a liquid reservoir is confirmed not to lose its temperature control function in gravity; thus, the operating temperature can be controlled by regulating the liquid reservoir temperature, not only on the ground but also in space.
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Heat Transfer Asian Research 39 39(6) 355-364 2010年9月
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40th International Conference on Environmental Systems, ICES 2010 2010年 査読有りThe pressure oscillation in oscillating heat pipes is caused by mutual excitation between pressure and void fraction. The pressure oscillation forms a progressive wave traveling in the heat pipe. The wave velocity is determined by the mechanism of self excited oscillation. The amplitude and the frequency are determined by the heat transport performance, that is, the heat load and the heat transport length. Based on the theory, pressure oscillation characteristics are obtained as a function of design parameters and operating conditions of the heat pipe. © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
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航空宇宙技術(Web) 9(9) 49-55 2010年This paper reports test results of active control of heat transfer performance of a small loop heat pipe (LHP) with a thermoelectric converter (TEC) in an atmospheric condition. The TEC is attached to a reservoir of the LHP to control the operating temperature by heating and cooling using a bipolar power supply. Tests were conducted by varying the evaporator power, the reservoir set point temperature, and the condenser sink temperature. The test results showed that the TEC is able to control the LHP operating with small electric power. The test results also demonstrated enhancement of the start-up success by maintaining a constant reservoir temperature during the start-up process. Forced shut-off of the loop operation by rapid heating of the reservoir was also demonstrated.
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Thermal science and engineering 18(1) 9-15 2010年1月1日 査読有りLoop Heat Pipes (LHPs) are two-phase thermal control system, which works only by heat from its cooling target. In order to utilize the LHPs in various fields, it is requested to be smaller, more reliable, and higher performance. In the present study, a miniature LHP has been fabricated, and the effect of amount of working fluid charged on thermal performance of the LHPs has been investigated. Tests were conducted including start-up, power step up, as function of amount of working fluid. The test results showed that under-charging of working fluid caused start-up failure, while over-charging of working fluid made the LHP less stable.
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Thermal Science and Engineering 18(4) 127-132 2010年Experiments were conducted to verify whether an Oscillating Heat Pipe (OHP) with reservoir works as a temperature controllable thermal control device in space. We set the reservoir in 3 ways (vertical, horizontal, and vertically inverted directions) and confirmed that the OHP operates as a variable conductance heat pipe in each case. We also put the OHP vertically and confirmed that the OHP works even in the “top heat mode”. It is concluded that the OHP does not lose its temperature control function of reservoir and heat transfer function of OHP by gravity.
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Science and Technology of Energetic Materials : journal of the Japan Explosive Society 70(5) 117-121 2009年12月31日
MISC
380-
大気球シンポジウム: 2022年度 = Balloon Symposium: 2022 2022年11月大気球シンポジウム 2022年度（2022年11月7-8日. ハイブリッド開催（JAXA相模原キャンパス& オンライン)) Balloon Symposium 2022 (November 7-8, 2022. Hybrid(in-person & online) Conference (Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan 著者人数: 26名 資料番号: SA6000177012 レポート番号: isas22-sbs-012
書籍等出版物
1講演・口頭発表等
33-
46th International Conference on Environmental Systems 2016年7月
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第16回宇宙科学シンポジウム 講演集 = Proceedings of the 16th Space Science Symposium 2016年1月 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)第16回宇宙科学シンポジウム (2016年1月6日-7日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)相模原キャンパス), 相模原市, 神奈川県資料番号: SA6000046247レポート番号: S4-010
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45th International Conference on Environmental Systems 2015年7月
所属学協会
5-
2020年9月
共同研究・競争的資金等の研究課題
9-
日本学術振興会 科学研究費助成事業 2023年4月 - 2027年3月
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日本学術振興会 科学研究費助成事業 2023年4月 - 2026年3月
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日本学術振興会 科学研究費助成事業 基盤研究(B) 2018年4月 - 2021年3月
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日本学術振興会 科学研究費助成事業 挑戦的萌芽研究 2016年4月 - 2018年3月
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日本学術振興会 科学研究費助成事業 基盤研究(C) 2008年 - 2010年
産業財産権
6学術貢献活動
1-
パネル司会・セッションチェア等, 査読2003年7月 - 現在
● 指導学生等の数
6-
年度2018年度(FY2018)博士課程学生数1
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年度2019年度(FY2019)博士課程学生数2修士課程学生数1学術特別研究員数1
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年度2020年度(FY2020)博士課程学生数1修士課程学生数1学術特別研究員数1
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年度2018年度(FY2018)博士課程学生数1
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年度2019年度(FY2019)博士課程学生数2修士課程学生数1学術特別研究員数1
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年度2020年度(FY2020)博士課程学生数1修士課程学生数1学術特別研究員数1
● 専任大学名
2-
専任大学名東京大学（University of Tokyo)
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専任大学名東京大学（University of Tokyo)
● 所属する所内委員会
6-
所内委員会名研究所会議
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所内委員会名プログラム会議
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所内委員会名信頼性品質会議
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所内委員会名環境・安全管理統括委員会
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所内委員会名ISASニュース編集小委員会
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所内委員会名宇宙科学プログラム技術委員会
