The tri-propellant engine is an attractive propulsion system for a near-term Single-Stage-to-Orbit (SSTO). This paper presents the results of an analytical study to determine the optimum fuel mass ratio of high-density kerosene and low-density liquid hydrogen for SSTO rockets. We performed flight simulation analyses to calculate payload capabilities for tri-propellant combinations with varying kerosene-hydrogen ratios, based on the current, matured Expendable Launch Vehicle's (ELV’s) technology both in propulsion and structural performance. A series burn was assumed for two cases of typical vehicle size; the same amount of total propellant mass and the same total impulse. The present study has shown that a tri-propellant combination with about 85% kerosene and 15% liquid hydrogen provides the highest vehicle performance. The trl-propellant SSTO performance gain at this optimum combination is about 800 m/s in vehicle velocity increase compared to a pure liquid oxygen/liquid hydrogen propulsion system. This performance gain is significant even if the increased complexities are taken into account for the tri-propellant propulsion system.

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To overcome defects of conventional hybrid rocket motors such as low combustion efficiency and low thrust level, the authors propose a new idea of design. The point of this idea is that the motor uses dry towel propellant, in which oxidizer gas flows the gap space in the fuel bed consisting of plastic fibers, tubes, or rods. Combustion experiments were carried out to clarify the basic mechanism of dry towel propellant combustion. As a basic sample of the propellant, the authors employed two types of cylindrical fibrous filters as a fuel in which gas oxygen is flowing. Ambient pressure was varied from atmospheric to more than 10 kgf/cm²G. As the results, the oxygen gas flow rate was found to have little influence on the burning rate. With one type of fuel samples, jump-up of the burning rate was observed in the pressure range of 5 to 6 kgf/cm²G, indicating the change of the combustion mechanism in this pressure range. With the other type of fuel samples, this jump-up was not observed. Up to now, the reason and controlling factor of this jump-up phenomenon are not clear. It is indispensable to clarify the basic mechanism of this jump-up phenomenon to realize the proposed idea.

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The performance of elecrodynamic tether orbit transfer system was analyzed to examine its possibility of the application to the round trip transportation mission between low earth orbits. The mass, the electric power and the mission time of tether orbit transfer system were obtained by the principle of operation of electrodynamic tether and the mission analysis based on the simplified orbit transfer analysis model of ion thruster for representative parameters and were compared with that of ion thruster orbit transfer. The results showed that there is a lower limit of mission time for each altitude of arrival orbit or higher limit of altitude of arrival orbit for each mission time and the most effective mechanical parameter on the improvement of system performance is the tether length. The results also showed that the tether orbit transfer system had mass advantage compared with the ion thruster orbit transfer system for the same missions.

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In order to send a spacecraft to a near-earth small body such as a comet or an asteroid, high-Isp electric propulsion based on the solar electrical power (usually called SEP) is very attractive because SEP can realize faster mission trip time with a smaller spacecraft and a launcher, which cannot be accomplished with conventional chemical propulsions. For the asteroid sample and return program promoted by the Institute of Space and Astronautical Science, a 1kW-class ion engine system is used. Adopting the ion engine that features high Isp around 3,000s and high thrust efficiency 30-40%, such an ambitious mission can be realized by a relatively small satellites weighing 450 kg with a medium-class launcher, Japanese M-5 rocket. The newly designed engine features microwave discharge, which is free from degradation of the thermionic cathode ordinarily used for conventional type ion engines, hence, very long lifetime required for the sample and return mission becomes possible. An engineering model of the thruster head was fabricated, and its thrust performance, thermal design, mechanical design, and lifetime of more than 13,000 hours were successfully demonstrated.

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