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One potentially attractive propulsion concept offering significant payload gains for orbit transfer from LEO to higher orbits, station keeping and attitude control of spacecraft is thermal propulsion using light gas (typically hydrogen) as propellant and various kinds of heat energy. Solar Thermal Propulsion (STP) is a typical thermal propulsion with high Isp (500 – 1,000 s) in an appropriate thrust magnitude range and provides possibly much less space pollution than conventional chemical propulsion. This paper presents the test results of a 30 mm dia. (medium-sized) windowless type of single crystal Mo thruster for orbit transfer of 50 kg class microsatellites. The cavity dia. is 20 mm, double the size of the previous model, and can apply to a primary solar reflector of up to 3.5 m dia., which is the maximum size containable in the H-II rocket fairing without segmentation. The performed mission analyses indicate that this size of STP is suitable to orbit transfer of 50 kg class microsatellites, such as LEO to GEO, or only multiple apogee kicks from GTO to GEO or deep space missions.

A couple of 5 cm mercury electronbombardment ion engine systems will be tested on board the engineering test satellite (ETS-III) of Japan, scheduled to be launched in 1982. To ascertain their reliability, the ion thrusters were to be subjected to a series of development tests in each phase of EM, PM and FM (engineering, prototype and flight models). Measurements on ion beam divergence and thrust were particularly planned because the measured quantities are important for ion thruster application to the satellite. Thermal-vacuum and electromagnetic compatibility tests are necessary for an on-board device. A special method must be considered and established for these tests because ion thruster is a peculiar device. In this paper, these tests conducted as research by using preliminary engineering model ion thrusters are described, with a brief summary of the data obtained.

In the present paper the development of electric propulsion systems in Japan has been reviewed. The status of hardware development, ground test and the future applications are presented with reference to activities in the government institutions and industries.

The R & D works on MPD thruster system and ion engine system in Japan are reviewed. A 1 kW class MPD thruster system, for which the flight test will be made onboard a small space platform in early 1990's, has experienced preliminary endurance tests of 1 million shots twice. From these tests, thermal control system, automated operation system and electrode mass loss assessment were established. A 20 mN class xenon ion engine system was determined to be developed and be tested onboard the ETS-VI to be launched by a H-II rocket in 1992. 500 hour operation tests of the thruster were conducted and showed the reduction of erosion rates of accelerator grid, etc. whose surfaces were coated with low sputter materials. A system interface tests were also made to see the compatibilities between subsystems. Performance improvement studies of xenon ion thruster and its test facility are going on.