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The MUSES-B spacecraft will be launched in 1996 by the Institute of Space and Astronautical Science (ISAS). Its primary mission is experiments on space Very Long Baseline Interferometry (VLBI) for radio astronomy using a large deployable antenna. A challenging thermal design must be compatible with a wide range of sun angle and an 86 minute eclipse. The thermal design and verification has been performed separately for the major modules of the spacecraft; a main structure, a deployable antenna and Reaction Control System (RCS). Special attention is paid to the exposed RCS whose solar input varies significantly depending on the sun angle. This paper describes the thermal design concept for MUSES-B and verification results of its thermal model test focusing on the main structure and the RCS.

The satellite ‘HALCA’ was launched in February 1997 by the Institute of Space and Astronautical Science (ISAS), Japan. HALCA is a scientific satellite with a large deployable antenna to make experiments on space VLBI. The deployable antenna, developed under the requirements of large aperture area and accurate reflector surface, is formed of complicated structures and lots of mechanisms. Though heat exchange in the antenna was complicated and difficult to evaluate, antenna thermal performance in orbit was within expectations and all mechanisms were adequately controlled until the antenna deployment. This paper describes thermal control design of the large deployable antenna, thermal design verification in the thermal balance test, and evaluation of performance in orbit.

Japanese X-ray astronomy satellite ASTRO-E2 named “SUZAKU“ was successfully launched on July 10, 2005. SUZAKU is the fifth Japanese X-ray astronomy satellite to observe X-ray coming from hot and active regions in the universe in collaboration with NASA GSFC, MIT and University of Wisconsin. “SUZAKU” has achieved the high sensitivity wide energy band X-ray spectroscopy than ever before. It is equipped with X -ray telescopes (XRT) and three kinds of focal plane instruments, X-Ray Imaging Spectrometer (XIS), X-Ray Spectrometer (XRS) and Hard X-Ray Detector (HXD). A radiation-cooling system, connected to XIS and HXD with heat pipes, is provided to cool them below −30 C and −20 C respectively. Furthermore, a side panel has a large cut out to expose XRS cryogenic Dewar for direct cooling. Flight temperatures indicate that the three sensors are kept below their cooling-requirement temperature.

Variable emittance radiator, called SRD, is a thin and light ceramic tile whose infrared emissivity is varied proportionally by its own temperature. Bonded only to the external surface of spacecrafts, it controls the heat radiated to deep space without electrical or mechanical parts such as the thermal louver. By applying this new device for thermal control of spacecrafts, considerable weight and cost reductions can be achieved easily. In this paper, the new design and the new manufacturing process of the SRD and its optical properties, such as the total hemispherical emittance and the solar absorptance, are described. By introducing this new design and manufacturing process, the weight of the SRD is easily decreased, keeping its strength and the optical properties.

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.

The ISAS's space VLBI satellite HALCA was successfully launched in February 1997. The spacecraft HALCA consists of a box shaped main structure and a large deployable mesh antenna with 8 m effective diameter. The integrated spacecraft with the mesh structure antenna is so large and complex that the thermal design and tests had been performed separately for the main structure and the large antenna. No thermal vacuum test had been conducted in the fully integrated spacecraft configuration. The complex heat exchange between the antenna and the main structure had been taken into account in the numerical thermal analysis. Good correlation between in-orbit temperature and flight prediction has proved validity of the design and the verification method where no integrated spacecraft thermal vacuum test was performed.

This paper describes a new passive thermal control device-a Reversible Thermal Panel (RTP)-which changes its function reversibly from a radiator to a solar absorber by deploying/stowing the radiator/absorber reversible fin. The RTP consists of Highly Oriented Graphite Sheets (HOGSs), which have characteristics of high thermal conductivity, flexibility and light weight, as thermal transport units, which can transport the heat from equipment to reversible fin, and of a Shape - Memory Alloy (SMA) as a passively rotary actuator to deploy/stow the reversible fin. The RTP prototype model was designed and fabricated using HOGSs, a honeycomb base palate, and a prototype reversible rotary actuator. The heat rejection performance of the RTP as a radiator and the heat absorption performance as an absorber were evaluated by thermal vacuum tests and thermal analyses. The autonomous thermal controllability achieved using the prototype rotary actuator was also evaluated.

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.

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.

The Smart Radiation Device (SRD) which is made from a ceramic material is a thin and light tile. The material undergoes a metal-insulator transition at around 290K and this allows the infrared emissivity of the device to change from low to high as the temperature is increased from 175K to 375K. This is beneficial for thermal control applications on spacecraft. For example, bonded only to the external surface of the spacecraft's instruments, SRD controls the heat radiated to deep space without electrical instruments or mechanical parts used for changing emissivity. It reduces the energy consumption of the electrical heater for thermal control, and decreases the weight and the cost of the thermal control system. In this paper, the design of the new material for SRD and the ground test results such as the radiation tests of electrons and UV will be described.

A lightweight 100 W-class deployable radiator with environment-adaptive functions has been investigated. This radiator - Reversible Thermal Panel (RTP) - is composed of flexible high thermal conductive materials and a passive reversible actuator, and it changes its function from a radiator to a solar absorber by deploying/stowing the reversible fin upon changes in the heat dissipation and thermal environment. The RTP is considered one of the candidates of thermal control methodology for the Japanese Venus mission “Planet-C”, which will be launched in 2010 to save its survival heater power. In this paper, design and fabrication of the RTP proto-model (PM) and the test results of deployment/stowing characteristics in an atmospheric condition are reported. Thermal performance estimation with thermal analytical model of the RTP PM is also presented.