太刀川 純孝

This study aimed to measure the thermal effusivity distribution on a lunar regolith simulant (FJS-1) using a thermal microscope and to calculate the average thermal effusivity and thermal conductivity using density and specific heat. Moreover, discussions were conducted based on the results of the microstructural analysis of the sample. The FJS-1 particles were embedded in an epoxy resin and polished to a mirror finish. The samples were analyzed using scanning electron microscopy equipped with energy-dispersive X-ray spectroscopy (SEM-EDS). X-ray diffraction (XRD) was performed to identify the mineral phases in FJS-1. The results of SEM-EDS and XRD showed that a single sand particle was composed of several minerals, such as anorthite and olivine. Then, the thermal microscope was used to obtain the distribution of the thermal effusivity of a particle from the mirror-finished sample in a 1 x 1 mm(2) area with intervals of 10 mu m. The measured thermal effusivity correlates with the SEM image of the sample. Anorthite has a small thermal effusivity of 1.99 +/- 0.31 kJ center dot s(-0.5)center dot m(-2)center dot K-1, while olivine has a large thermal effusivity of 2.73 +/- 0.35 kJ center dot s(-0.5)center dot m(-2)center dot K-1. In both cases, the thermal effusivity was found to be of the same order of magnitude as the reported values. The average thermal effusivity and conductivity of a single particle were determined to be 2.4 +/- 0.6 kJ center dot s(-0.5)center dot m(-2)center dot K-1 and 2.6 +/- 1.3 W m(-1)center dot K-1, respectively, based on the proportion of existing phases.

Abstract In recent planetary exploration space missions, spacecraft are exposed to severe thermal environments that are sometimes more extreme than those experienced in earth orbits. The development of advanced thermal control materials and devices together with reliable and accurate measurements of their thermophysical properties are needed for the development of systems designed to meet the engineering challenges associated with these space missions. We provide a comprehensive review of the state-of-the-art advanced passive thermal control materials and devices that are available for space applications, specifically, variable emissivity thermal control materials and microelectromechanical systems (MEMS), radiofrequency (RF)-transparent and/or tunable solar absorptivity and total hemispherical emissivity thermal control materials, and a passive re-deployable radiator with advanced materials and insulation. Prior to our in-depth review of these thermal control materials, we briefly summarize the thermal environments surrounding spacecraft, the characteristics of thermophysical properties for spacecraft materials that differ from those of materials for ground use, and the significance of solar absorptivity and total hemispherical emissivity for passive thermal control in space. In all four topics of materials and devices, the following subjects are overviewed: the basic principle of passive thermal control techniques in space, the measurement of thermophysical properties of those novel materials, simulation and/or on-orbit verification thermal performance tests, degradation tests in space environments, and some aspects of the implementation of the above-described materials and devices in actual space missions.

© 2017 Elsevier B.V. (La1 − xSrx)MnO3 − δ (LSMO) thin films were prepared on Si(100) substrates using a metal-organic decomposition (MOD) technique for smart radiation devices used for thermal control in spacecraft. The 6 wt% concentration MOD precursor solutions were spin-coated on Si substrates. The coated films were dried at 90 °C and pyrolytically decomposed at 530 °C. Finally, they were annealed in air at 750 °C to obtain dense, polycrystalline, and single-phase perovskite LSMO films. A ferromagnetic-paramagnetic phase transition was observed close to room temperature in a LSMO film. The thermal emittance of the film was changed as a result of this phase transition. The phase transition temperature increased with an increasing Sr ratio at the A-site of the LSMO and decreased after annealing under a reduced atmosphere. Based on this result, we propose that the temperature range where the thermal emittance is widely variable can be finely tuned on demand with post-annealing.

Multi-layer Insulation with polyimide foam (PF-MLI) has been studied as a new insulator for next generation spacecraft. PF-MLI has higher insulation performance than that of conventional MLI as we showed in our previous work. In the case of the mission which needs high thermal insulation performance such as deep space or lunar exploration mission, it is required to cover an antenna with MLI. However, MLI does not transmit radiowave because it is composed of multiple layers of metalized films with low emittance. In this work, we have been studying a radiowave transmissive MLI based on PF-MLI by using a new thermal control material instead of metalized film. This new thermal control material is called as COSF (Controlled Optical Surface Film). COSF consists of a polyimide film for substrate and dielectric multi-layer coating on it. The solar absorptance and infrared emittance can be freely controlled by the interference of its dielectric multi-layer, and the COSF has radiowave transmissivity. The COSF is designed by genetic algorithm method. The information of multi-layers, such as materials and thickness, is encoded as a gene and evaluated by an evaluation function. The solar absorptance of COSF4 used on the outer cover of PF-MLI is 0.06 and the normal emittance and total hemispherical emittance is 0.76 and 0.70 at 300 K respectively. On the other hand, the normal emittance and total hemispherical emittance of COSF-IR2 in the middle layer of PFMLI is 0.14 and 0.17 at 300 K respectively. These thermo-optical properties has been measured by monochrometer, FTIR at room temperatures and calorimetric method in the temperature range from 173 to 373 K. The Radiowave Transmissive Multi-layer Insulation called as RT-MLI has been made by combining polyimide foam with COSF and its effective thermal conductivity has been measured by guarded hot plate method in the temperature range from 162.5 to 350 K. The density of polyimide foam is 6.67 kg/m3. The prototype of RT-MLI is composed of 2 layers of polyimide foam and 1 layer of COSF in the middle of them. As a result, the effective thermal conductivity of RTMLI is reduced by 40 % compared to that of polyimide foam and similar to that of PF-MLI. This RT-MLI will be used for a lot of mission in near future.

A radiator is a thermal control device which radiates waste heat from the spacecraft to space and cools equipment. Since an infrared space telescope observes weak infrared signals, it is important to shift thermal radiation wavelength to longer values than the observation signals and to reduce the radiation intensity by cooling the equipment. Especially, the next generation infrared space telescope will be cooled at cryogenic temperatures so as to enhance the observation performance. However, a conventional radiator material, such as black paint, shows lower emissivity at cryogenic temperatures. That is why a new radiator with high emissivity at cryogenic temperatures is required. Therefore, we focused on a structure consisting of periodic array of metallic elements separated from a ground plane by a dielectric spacer layer in order to realize a high emittance radiator at cryogenic temperatures. In this work, we reported the performance of the IR emitter with periodic array. It was designed by the FDTD method. The periodic structure is a circle. The metal layer was made by aluminum and the dielectric layer was made by SiO2. The prototype IR emitters were fabricated. Spectral reflectance of the prototype emitter was measured by FT-IR. Moreover, spectral emittance was measured by the blackbody comparison method, and it was compared with spectral absorptance calculated by spectral reflectance.

© 2016 Nakamura et al. AKATSUKI is the Japanese Venus Climate Orbiter that was designed to investigate the climate system of Venus. The orbiter was launched on May 21, 2010, and it reached Venus on December 7, 2010. Thrust was applied by the orbital maneuver engine in an attempt to put AKATSUKI into a westward equatorial orbit around Venus with a 30-h orbital period. However, this operation failed because of a malfunction in the propulsion system. After this failure, the spacecraft orbited the Sun for 5 years. On December 7, 2015, AKATSUKI once again approached Venus and the Venus orbit insertion was successful, whereby a westward equatorial orbit with apoapsis of ∼440,000 km and orbital period of 14 days was initiated. Now that AKATSUKI's long journey to Venus has ended, it will provide scientific data on the Venusian climate system for two or more years. For the purpose of both decreasing the apoapsis altitude and avoiding a long eclipse during the orbit, a trim maneuver was performed at the first periapsis. The apoapsis altitude is now ~360,000 km with a periapsis altitude of 1000-8000 km, and the period is 10 days and 12 h. In this paper, we describe the details of the Venus orbit insertion-revenge 1 (VOI-R1) and the new orbit, the expected scientific information to be obtained at this orbit, and the Venus images captured by the onboard 1-μm infrared camera, ultraviolet imager, and long-wave infrared camera 2 h after the successful initiation of the VOI-R1.

Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved. Japan's Venus Climate Orbiter Akatsuki was proposed to ISAS (Institute of Space and Astronautical Science) in 2001 as an interplanetary mission. We made 5 cameras with narrow-band filters to image Venus at different wavelengths to track the cloud and minor components distribution at different heights to study the Venusian atmospheric dynamics in 3 dimension. It was launched on May 21st, 2010 and reached Venus on December 7th, 2010. With the thrust by the orbital maneuver engine, Akatsuki tried to go into the westward equatorial orbit around Venus with the 30 hours' orbital period, however it failed by the malfunction of the propulsion system. Later the spacecraft has been orbiting the sun for 5 years. On December 7th, 2015 Akatsuki met Venus again after the orbit control and Akatsuki was put into the westward equatorial orbit whose apoapsis is about 0.44 million km and orbital period of 14 days. Its main target is to shed light on the mechanism of the fast atmospheric circulation of Venus. The systematic imaging sequence by Akatsuki is advantageous for detecting meteorological phenomena with various temporal and spatial scales. We have five photometric sensors as mission instruments for imaging, which are 1 m-infrared camera (IR1), 2 m-infrared camera (IR2), ultra-violet imager (UVI), long-wave infrared camera (LIR), and lightning and airglow camera (LAC). These photometers except LIR have changeable filters in the optics to image in certain wavelengths. Akatsuki's long elliptical orbit around Venus is suitable for obtaining cloud-tracked wind vectors over a wide area continuously from high altitudes. With the observation, the characterizations of the meridional circulation, mid-latitude jets, and various wave activities are anticipated. The technical issues of Venus orbit insertion in 2015 and the scientific new results will be given in this paper.

Copyright © 2016 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. This paper describes a new flexible thermal control mirror: controlled optical surface films. Controlled optical surface film exhibits high reflectivity to solar radiation, high infrared emissivity, and high transmittance to radio waves. The target of radiowave transmittance is Ge-coated Kapton, which is used on antenna covers; the target of thermooptical properties is the optical solar reflector. Controlled optical surface film controls spectral reflectance and absorptance in wavelengths ranging from 0.26 to 100 μm with a dielectric multilayer. This multilayer is designed using a genetic algorithm method. The solar absorptance of the latest controlled optical surface film is 0.09, and its total hemispherical emittance is 0.77 at 300 K. These properties are similar to those of the optical solar reflector. In addition, there is little radiowave attenuation in this controlled optical surface film, and its radiowave transmittance is similar to that of Ge-coated Kapton.

© 2015 Elsevier B.V. For application as a variable thermal emittance material in a recently-developed thermal control system for spacecraft, (La1 - xSrx)MnO3 - δ (LSMO) thin films with thicknesses of 1.2 μm, 2.5 μm, and 4.3 μm were fabricated on yttria-stabilized zirconia (100) substrates by a pulsed laser deposition and ex-situ annealing at 1123 K in air. All the films were dense and their surface roughness was much smaller than the thermal infrared (IR) wavelength. The films had (100) and (110)-preferred orientations, and the thicker films showed more preferable growth along the (100) orientation. Temperature-magnetization curves revealed that the LSMO films exhibited a metal-insulator transition near room temperature. The thermal emittance of the films estimated from IR reflectance spectra and black body radiation spectra exhibited large non-linearity near room temperature owing to the phase transition. The change in thermal emittance of the LSMO films that were more than 2.5 μm thick was comparable with that of the Ca-doped LSMO ceramic tiles already used as variable thermal emittance materials. Thus, this result clearly demonstrates that LSMO thin films with thickness of 2.5 μm can work as variable thermal emittance materials in the thermal control system for spacecraft.

© 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. 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°C, the fin efficiency difference between the deployed and stowed positions was 0.35, and the specific heat rejection was 188 W/kg.

Polyimide foam (PF) is a low-thermal conductivity and lightweight material with high resistances against heat, protons, and UV irradiation. A new thermal insulation composed of PFs and multiple aluminized films (PF-MLI) has potential to be used in outer space as an alternative to conventional multilayer insulation (MLI). As fundamental numerical data, the effective thermal conductivity of PF in wide ranges of density and temperature need to be determined. In the present study, thermal-conductivity measurements were performed by both the periodic heating method and the guarded hot-plate method in the temperature range from 160 K to 370 K and the density range from 6.67 kg · m- 3 to 242.63 kg · m - 3. The experiments were carried out in a vacuum and under atmospheric pressure. For confirmation of the validity of the present guarded hot-plate apparatus under atmospheric pressure, the effective thermal conductivity of the lowest-density PF was measured with the aid of the heat flow meter apparatus calibrated by the standard reference material (NIST SRM 1450c) in the temperature range from 303 K to 323 K. In order to cross-check the present experimental results, the temperature and density dependences of the effective thermal conductivity of PF were estimated by means of the lattice Boltzmann method based on a dodecahedron inner microscopic complex structure model which reflects a real 3D X-ray CT image of PF. © 2014 Springer Science+Business Media New York.

Conventionally, multilayer insulation (MLI) has been used for spacecraft, and known for high performance thermal insulation. MLI blanket is composed of multiple layers of low cmittance films with spacers. The layers are stacked and sewn together at the edge, and attached to the spacecraft using Velcro on one side of MLI blanket. Though, heat losses through Seam and Velcro make the performance of MLI degraded. Saving heater power is the key factor for deep space missions or lunar exploratory missions. Therefore, local thermal insulation performance degradation is unacceptable. In order to respond to this situation, we focus on polyimide foam (PF) and propose polyimide foam multilayer insulation (PF-MLI), which shows superior performance compared with conventional MLI blanket. Thermal conductivity of PF measured and estimated in this study change with density, and local minimum value, are ascertained from both measurement and estimation. Based on these results, the optimum configuration (layers, density, etc.) of PF-MLI is determined. PF-MLI is composed of aluminized films and PF between films. Two different types of PF are used for PF-MLI and compared with Standard MLI Blanket in terms of effective emissivity and weight. Type BF301 (6~8 kg/m3) with aluminized films and Type BP 101 (29-33 kg/m3) with aluminized films are the measurement samples as PF-MLI, and two different layer numbers are fabricated in each sample. Guarded hot plate method, which is useful heat flux measurement method with reducing heat loss, is applied to measure effective emissivity in a vacuum steady-state condition. Effective emissivity is calculated from the higher temperature, colder temperature of the sample, and heat flux thorough the sample from Stephan Boltzmann law in a steady-state. As a result, effective emissivity of PF-MLI is lower than that of Standard MLI in higher temperature range from 260K to 340K, and it can be concluded that especially PF-MLI using BF301 (7.2 kg/m3) is the optimum thermal insulation from the standpoint of lighter weight and higher performance. In addition, seam effect on degrading the performance is not serious and high performance is expected compared with conventional MLI, whose heat loss is critical to degrading the performance. Copyright © (2012) by the International Astronautical Federation.

The basic construction of thermal control mirrors, also known as Optical Solar Reflectors (OSRs) or Second Surface Mirrors (SSMs), consists of a thin glass substrate with a rear surface silver coating. It has a high total hemispherical emittance due to the glass, but a low solar absorptance from the highly transmitting glass and highly reflecting metallic surface. This paper will describe a new type of thermal control mirror, Controlled Optical Surface Film (COSF). Although the function of the COSF is the same as the conventional OSRs, the construction is different. The COSF consists of a polyimide film (UPILEX-S) for substrate and multi-layer coating on it. While the solar radiation is reflected on the rear surface silver coating in case of the OSRs, it will be reflected on the front surface multi-layer coating in case of the COSF. The preliminary design will be described and the preliminary test results of a comprehensive space qualification are also reported showing that the COSF is stable under high radiation fluence in space. © 2010 The American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

In rocket flights, ionized exhaust plumes from solid rocket motors may interfere with RF transmission under some conditions. In order to clarify the important physical process involved, microwave attenuation and phase delay due to rocket exhaust plumes were measured during sea-level static firing tests conducted on two types of full-scale solid propellant rocket motors. The measured data were analyzed by comparing them with numerical results such as flowfield simulations of exhaust plumes and by employing a detailed analysis of microwave transmission by using a frequency-dependent finite-difference time-domain (FD2TD) method. The results revealed that either the line-of-sight microwave transmission through ionized plumes or the diffracted path around the exhaust plume mainly affects the received RF level, which depends on the magnitude of the plasma RF interaction. For the actual launch vehicle flight, the transmission process is dominated by the diffraction effect so that we applied a two-dimensional diffraction theory to analyze the communication between a vehicle and a ground station. The attenuation levels estimated using diffraction theory agree with the data recorded in-flight. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

The Smart Radiation Device (SRD) is a new thermal control material that decreases the temperature variation by changing the emissivity without using electrical instruments or mechanical parts. The emissivity of the SRD changes physically depending on its temperatures. Bonded only to the external surface of the spacecrafts, the SRD controls the temperature. The drawback of the SRD is the high solar absorptance. The multi-layer film for SRD was designed in order to decrease the solar absorptance from 0.81 to less than 0.2 by putting multi-layer film on it and the optical performance of the Smart Radiation Device with Multi-layer film (SRDM) was evaluated. Copyright © 2008 SAE International.

A new thermal control material named the Smart Radiation Device (SRD) has shown improvement in development. The SRD can be used as a variable emittance radiator that controls the heat radiated into deep space without assistances of any electrical instruments or mechanical parts. Its total hemispherical emittance changes from low to high as the temperature increases. This new device reduces the energy consumption of the on-board heater, and decreases the weight and the cost of the thermal control system (TCS). Space environmental simulation tests on the ground were performed, and the first generation of the SRD has been demonstrating success on the MUSES-C ‘HAYABUSA’ spacecraft that was launched in May 2003. During its cruise on the orbit, the distance from the spacecraft to the sun varied from 0.86AU to 1.70AU. As the spacecraft experienced solar intensity variation by a factor 4, it was effective to use the variable emittance radiator for decreasing the heater power. In-orbit temperature indicated that the SRD had successfully minimized component temperature variation and saved heater power, as expected. With the opportunity to validate the SRD in space, this lightweight and low cost thermal control device offers a possibility for flexible thermal control on future spacecrafts.

The Hinode satellite (formerly Solar-B) of the Japan Aerospace Exploration Agency's Institute of Space and Astronautical Science (ISAS/JAXA) was successfully launched in September 2006. As the successor to the Yohkoh mission, it aims to understand how magnetic energy gets transferred from the photosphere to the upper atmosphere and results in explosive energy releases. Hinode is an observatory style mission, with all the instruments being designed and built to work together to address the science aims. There are three instruments onboard: the Solar Optical Telescope (SOT), the EUV Imaging Spectrometer (EIS), and the X-Ray Telescope (XRT). This paper provides an overview of the mission, detailing the satellite, the scientific payload, and operations. It will conclude with discussions on how the international science community can participate in the analysis of the mission data. © 2007 Springer Science+Business Media B.V.

The Smart Radiation Device (SRD) decreases the temperature variation by changing its emissivity depending on the temperature. The first generation of the SRD has been demonstrated on the MUSES-C 'HAYABUSA' spacecraft launched on May 9th 2003. This new thermal control device reduces the energy consumption of the on-board heater, and decreases the weight and the cost of the thermal control system. With the opportunity to validate the SRD in space, lightweight and low cost thermal control devices offer a possibility for flexible thermal control on interplanetary spacecraft. Copyright © 2007 SAE International.

We are developing a wireless multi-channel measurement system in order to measure temperatures during thermal vacuum tests of spacecraft, accelerations during vibration tests, etc. As a developing measurement system, we introduce the prototype model of the wireless temperature measurement system. The temperature data are transmitted to the data logger by radio instead of thermocouple wires, thus it will be easier to prepare thermal vacuum tests, and the test system will become very simple. Details of the system configuration, its specification and performance are presented. Copyright © 2005 SAE International.

We have developed a variable-emittance radiator device, made of thin, and fight ceramic tiles, for thermal control applications on spacecraft. The ceramic material used is La1-xSrMnO3 with a perovskite structure, which shows a phase transition from a ferromagnetic metal with a low thermal emittance of 0.35 to a paramagnetic insulator with a high thermal emittance of 0.75 at around 300K. This device automatically controls the temperature of a spacecraft without requiring electrical or mechanical instruments. The fabrication process for the ceramic tiles was designed to reduce production costs and weight. The tiles are less than 70-microns thick and weigh 450g/m(2). Both a conventional type of ceramic wafer and a thick film type were developed and tested for their durability under various environmental and radiation conditions. The results showed no degradation. These variable emittance radiators made of ceramic tiles have been used for the MUSES-C spacecraft designed to probe asteroids, and are also scheduled to be used for the INDEX spacecraft designed to observe the earth.

Two types of thermal control materials developed for space use are described. These new materials use a variable emittance radiator called the smart radiation device (SRD). The SRDs are based on La0.825Sr0.175MnO3 and La0.7Ca0.3MnO3 and comprise a thin and light ceramic tile. The materials undergo a metal-insulator transition near room temperature, and this allows the infrared emissivity of the device to change from low to high as the temperature is increased from 173 to 375 K. This is beneficial for thermal control applications on spacecraft. For example, bonded only to the external surface of the spacecraft, the SRD controls the heat radiated to deep space without electrical instruments or mechanical parts. Its function is similar to the thermal louver, but the SRD is lighter. This new device reduces the energy consumption of the electrical heater for thermal control and decreases the weight and the cost of the thermal control system. Optical properties, such as the total hemispherical emittance and the solar absorptance, have been measured. Space environmental simulation tests were performed with independent irradiation by protons, electrons, and UV on the ground. Degradation of the materials' optical properties is discussed.

A variable-emittance radiator device, made of thin and light ceramic tiles, has been developed for thermal control applications on spacecraft. The ceramic material used is La1-xSrxMnO3 with a perovskite structure, and shows a phase transition from ferromagnetic metal to paramagnetic insulator at around 290 K (Tc). This device automatically controls a spacecraft's temperature without electrical or mechanical instruments. Below the Tc, the device is metallic with a low thermal emittance of 0.3, and above the Tc, it becomes insulative with a high thermal emittance of 0.7. For the ceramic tiles, two different fabrication processes were studied to reduce the cost and weight; one is a conventional ceramic wafer process and the other is a thick film process on zirconia substrates. Total thickness of the ceramic tiles obtained is less than 70μm and the weight is 450gr/m2. © 2002 The Japan Society of Applied Physics.

Variable-emittance radiators based on the metal-insulator transition of (La,Sr)MnO3 thin films have been developed. The emittance property of the films was evaluated from infrared reflectance spectra; that is, the (La,Sr)MnO3 thin films show low emittance at low temperature but high emittance at high temperature. Moreover, the emittance property significantly changes at the metal-insulator transition temperature, where the material changes from a highly reflective (i.e., low emissive) metal to a less reflective (i.e., high emissive) insulator. The (La,Sr)MnO3 thin films fitted on a spacecraft surface can, therefore, be used to automatically control the emissive heat transfer from the spacecraft without the need for any electrical power. The developed radiators also greatly reduce the weight and production cost of the thermal control devices. The dependence of the emittance property on film thickness reveals that 1500-nm-thick films can be used for variable-emittance radiators. © 2002 American Institute of Physics.

This paper describes outline of the piggy-back satellite "INDEX" for demonstration of advanced satellite technologies as well as for observation of fine structure of aurora. Aurora observation will be carried out by three cameras(MAC) with a monochromatic UV filter. Electron and ion spectrum analyzer (ESA/ISA) will measure the particle phenomena together with the aurora hanging. INDEX satellite will be launched in 2002 by Japanese H2-A. The satellite is mainly controlled by the high-speed, fault-tolerant on-board RICS processor (three-voting system of SH-3). The attitude control is a compact system of three-axis stabilization. Although the size of INDEX is small (50Kg class), several newly-developed technologies are applied to the satellite system, including silicon-on-insulator devices, variable emittance radiator, solar-concentrated paddles, lithium-ion battery, and GPS receiver with all-sky antenna-coverage. © 2001 International Astronautical Federation. Published by Elsevier Science Lt d.

New thermal control material named Smart Radiation Device (SRD) has been studied to apply for spacecraft. Infrared radiative properties of the SRD change depending on its own temperature without electrical or mechanical instruments. The SRD however shows too high solar absorptance to apply it as a radiator. To overcome such drawback, it is necessary to design and apply the mutilayer films on the SRD for reflecting solar radiation, keeping infrared radiative properties of the SRD. To perform optimum design of the multilayer films, the genetic algorithm (GA) was employed. In this paper, we propose a design of the thermal radiative properties of the SRD with the multilayer films. Copyright © 2001 Society of Automotive Engineers, Inc.

Radiative and optical properties of polycrystalline La1-xSrxMnO3 (0≤ x ≤ 0.4) in the vicinity of the metal-insulator transition are presented. The temperature dependence of the total hemispherical emittance εH of La1-xSrxMnO3 was measured by the calorimetric method in the temperature range from 173 to 413 K. It was confirmed that εH showed unexpected variation as a result of changes in the hole concentration (x). Especially in the case of La0.825Sr0.175MnO3, εH remains high above the transition temperature TC due to insulator-like behavior; on the other hand, it decreases sharply below TC because of metallic behavior. The spectral reflectance was measured by FT-IR in the wavelength range of 0.25 to 100 μm at room temperature. The optical constants were calculated by Kramers-Kronig analysis of the spectral reflectance data. An insulator-like character of the optical properties appears at lower Sr2+ doping levels while a metallic one exists at higher Sr2+ doping levels. © 2001 Plenum Publishing Corporation.

La1-xSrxMnO3 is a manganese oxide with a perovskite-type structure, LaMnO3, by substitution of La3+ sites with Sr2+, which shows a variety of phenomena with changing hole concentration (x). The total hemispherical emittance, εH, of La1-xSrxMnO3 (x = 0, 0.175, and 0.3) in the vicinity of the metal-insulator transition temperature was measured by the calorimetric method in the temperature range 173-373 K. Remarkable differences in εH resulted with changing hole concentration (x). It was observed that εH of La0.825Sr0.175MnO3 changed notably at about 280 K accompanied by the metal-insulator transition and the maximum amount of variation in εH was 0.42. The effects of surface conditions and thicknesses of samples were also observed. The uncertainty in the measurement of εH was estimated to be at most ±2.2% from 173.15 to 373.15 K.

The Smart Radiation Device (SRD) is a thin and light tile whose infrared emissivity is varied proportionally by the temperature of the radiator. Bonded only to the external surface of the spacecraft's instruments, it controls the heat radiated to deep space without electrical or mechanical instruments used for changing emissivity. Its function is similar to the thermal louver which has been used for a lot of spacecraft, but the SRD is lighter than it. Thus, by using this new device, we can control the temperature of the instruments on the spacecraft more easily. The materials of the SRD are La0.825Sr 0.175MnO3 and La0.7Ca0.3MnO 3. In this paper, design and preliminary test results of the SRD will be presented. The optical properties for the materials of the SRD, such as the total hemispherical emittance and the solar absorptance, have been measured. In addition the degradation by protons has been investigated. © Copyright 1999 Society of Automotive Engineers, Inc.