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In this paper, a newly developed Active Noise Control (ANC) system is introduced, that effectively reduces road noise, which becomes a major issue with electrified vehicles, and that enhances vehicle interior sound levels matching seamless acceleration by electric drive. Conventionally, reducing road noise using ANC requires numerous sensors and speakers, as well as a processor with high computing power. Therefore, the increase in system cost and the complexity of the system are obstacles to its spread. To overcome these issues, this system is developed based on four concepts. The first is a modular system configuration with unified interface to apply to various vehicle types and grades. The second is the integration and optimal placement of noise source reference sensors to achieve both reduction in number of parts and noise reduction performance.

The Smart Radiation Device (SRD) is a new type of thermal control material for spacecraft. By changing its emissivity without using electrical instruments or mechanical parts, the SRD decreases the temperature variation of the applied place. The SRD changes its emissivity physically depending on its temperature. The drawback of the SRD is its high solar absorptance. In order to reduce the solar absorptance while keeping its emissivity variation, the wide band filter was designed for the SRD. The function of the wide band filter is to reflect sunlight and to transmit infrared light. The SRD with the wide band filter is called as the Smart Radiation Device with Multi-layer films (SRDM) and the target value of its solar absorptance is 0.13.

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) 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.

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 thermal feasibility of BepiColombo MMO (Mercury Magnetospheric Orbiter) is shown. The temperatures of upper and lower decks on which instruments are located are controlled with in the allowable limit: -30 to 60 degrees C. The battery temperature is also controlled from 0 to 20 degrees C. Because MMO is spin-stabilized spacecraft, MMO requires the sun-shield and the radiator-shield in the cruise phase (Earth to Mercury) in which MMO is 3-axis stabilized, to avoid the exposure to the direct solar flux.

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) 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.

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 La0.825Sr0.175MnO3 and La0.7Ca0.3MnO3. 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.

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.

The thermal design of the LUNAR-A spacecraft, investigating the internal structure of the moon, is a challenge due to the widely varying thermal environments. A passive thermal control philosophy of radiators, multilayer insulation (MLI) and heaters was used for the LUNAR-A thermal design. Verification of the design was performed separately for the major spacecraft modules which consisted of a main structure, a propulsion subsystem and penetrators. As consequence, acceptable thermal performance can be expected for all phases of the mission. This paper describes the thermal design, analysis, thermal model testing and in-orbit temperature predictions.

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.

An effective thermal verification method for spacecrafts is introduced, in which their partial model or structural segment can be used in the test. The essence of this concept is to evaluate thermal design of the spacecraft by integrating a thermal vacuum test into thermal analyses system as a data generating subroutine. The measured temperature data are used for modification of the thermal mathematical model, which in turn requires to generate new test data for further refining of the model. Practically, the verification process is a kind of numerical simulation using the in-situ temperature measurements obtained from the thermal vacuum test run simultaneously with the simulation analysis. This direct coupling or “hybrid simulation” process is terminated when the thermal mathematical model is finalized to determine the spacecraft hardware design.