New type of hybrid rocket is proposed. Keywords are safety, environmentally tender and low cost. Staged combustion hybrid rocket was designed and basic characteristics were measured and its applicability was confirmed to the near future use, for example to fully reusable sounding rockets, space tugs etc.

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A fully reusable rocket vehicle is proposed to demonstrate good operability characteristics both on the ground and in flight. For achieving technical readiness for future space transportation systems, design considerations not only for higher performance-related issues but also those for good operability are needed The proposed vehicle is to be used as a sounding rocket and has the capabilities of ballistic flight returning to the launch site, and landing vertically making use of clustered liquid hydrogen rocket engines. Before initiating the development of this type of reusable rocket, a small test vehicle with a liquid hydrogen rocket engine was built and flight tested. A demonstration of vertical landing and exercise of turnaround operation for repeated flights are the major objectives of the test vehicle. Two series of flight tests were performed in 1999 and 2001 and the flight test operation provided repeated flight environment and many lessons valuable for designing the fully reusable rocket vehicle.

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The authors have studied the mechanism, which achieves the low restitution coefficient under the microgravity environment, and realized that the mechanism which is constructed out of a rigid shell with balls stored internally makes the restitution coefficient small. On impact, the balls will dissipate energy relative to each other and hence will dissipate the total energy. High efficiency in energy dissipation means the low restitution coefficient. In this paper, the relations among the parameters in this mechanism and the restitution coefficient is analyzed numerically, and confirmed through the microgravity tests.

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This paper discusses on the minimum fuel trajectories of interplanetary spacecraft for sample-return mission. The spacecraft is propelled by a solar electric propulsion system. Both going and returning trajectories with four boundary conditions are optimized simultaneously. Because available thrust of solar electric propulsion is very low, the propulsion system is firing for a long duration. The available thrust depends on the distance between the spacecraft and the sun. The problem is described as an optimal control problem to find the amplitude and direction of thrust. The optimal amplitude of thrust is maximum or minimum value according to optimal control theory. In order to obtain highly accurate solution, the equations of motion are transformed to a new description in which the amplitude of thrust is treated as a constant value. Two problems are discussed in this paper. The first problem is that the staying period at the target planet is not specified. In this problem, it is assumed that the spacecraft departs to and returns from the target planet at the optimal time. The optimal trajectories and the optimal staying period are shown in the results. The results show how the performance of thruster affects to the fuel consumption and the whole flight time. Higher thrust shortens the whole flight time and higher specific impulse improves the fuel consumption. In the returning problem, the staying period is specified. The staying period is required to be longer or shorter than the optimal period shown in the previous problem. The results show that the going trajectory changes in the case of longer staying period and the returning trajectory changes in the case of shorter staying period. The reason is discussed.

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