Conventional monopropellant thrusters for spacecraft utilize the pelletized catalyst bed to decompose the propellant, typically Hydrazine. The catalyst bed consists of Iridium catalyst supported by porous Alumina ceramics pellets. During the long-term operation of the thruster, however, the catalyst is occasionally damaged, which causes several critical problems like a deterioration of the decomposition performance or choke of the capillary tube. In order to solve or mitigate these problems, a new decomposition device was designed and its preliminary model was fabricated and evaluated in our research group. Preliminary tests were conducted to examine the new reaction mechanism using pulsed or stationary AC discharge plasma instead of pelletized catalysts in a bucket type reaction chamber. In the reaction chamber filled with Hydrazine, sudden increase of the temperature and the pressure was observed immediately after the production of plasma discharge. These results show that the plasma is capable of the decomposition of Hydrazine, and acts as catalysts, and that it is worthwhile developing a new monopropellant thruster system using plasma assisted reaction in order to eliminate the disadvantages associated with the pelletized catalyst bed.

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An orbit transfer system that transports a payload from LEO to GEO using spinning-momentum-tether is introduced in this paper. This system is analyzed in consideration of the change of the angular velocity before and after the catch or the release of the payload for the various parameters of the tether system. The time and the propellant mass required for reboosting the stage are also calculated in detail. As the new thruster for reboosting, the electrodynamic tether (ED tether) system is considered in this study and this system is compared with the conventional ion-thruster system. In addition, the total system mass is optimized for the mission interval that the next transportation of a payload is possible. The results also show that ED tether system can reduce the total system mass compared with ion-thruster system for the same orbit transfer mission.

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A Space Solar Power System (SSPS) has been proposed at the Institute for Unmanned Space Experiment Free Flyer and ISAS/JAXA. The objective of this paper is to evaluate the attitude stability of the SSPS under construction. At first, we investigate "line construction", where the modules are arranged in the shape of line. Through approximate estimation, it is shown that the roll and yaw angular frequencies of attitude motion are twice and once of the orbital angular velocity in the first step of construction, respectively. Furthermore, through dynamical analysis, roll and yaw angles tend to divergence in first step. To avoid such a divergence, we consider another construction way, "construction with similarity", where satellite is constructed keeping similar configuration. From the analysis, it is shown that divergences in the roll and yaw angles can be avoided in the construction with similarity. However, the resorting torques around the pitch and yaw axes are weak in the first step when the SSPS is constructed with similarity. In particular, maximum amplitude around the yaw axis becomes about 29 degrees. Therefore, for both construction ways, active attitude control is desired in first step of the constructions.

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A completely new solar electric propulsion concept, the Air Breathing Ion Engine (ABIE), has been proposed for spacecraft drag makeup at very low altitudes, ranging from 150 to 200 km. ABIE scoops up neutral atoms and molecules traveling at an orbital velocity of approximately 8 km/s, ionizes them by means of an electron cyclotron resonance plasma source that is efficient in a wide range of low gas pressures, and accelerates the ionized air particles electrostatically to exhaust velocities larger than 100 km/s. The key technology of this thruster is the design of a propellant inlet which allows the incoming flow to enter the discharge chamber, yet it prevents the thermalized gas from escaping upstream. In this system, an air-breathing-type neutralizer may also be employed, in which case the need to carry on-board xenon propellant is eliminated and results in gains in payload mass if the mission duration is longer than 2 years. This technology should give researchers access to a part of the atmosphere that is currently very difficult to measure and is thus called the "ignorosphere." Promising applications other than aeronomy include science missions involving accurate gravity and magnetic field mapping, and high-resolution Earth surveillance.

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This paper is devoted to study the optimal control of re-orient the gravity-gradient-stabilized space structure from one equilibrium attitude to another attitude, which is to orient 180 degrees in rotation. The space structure on an orbit is assumed to be able to move a part of the mass in the direction of the axis of the minimum moment of inertia. Two analytical studies are applied to the present problem. One method employs an optimal control to minimize the length of the movement of the particle in the course of the re-orientation. The other one is the orthodox feedback control, which refers to the rotation angle of the system to define the control tension. The numerical result shows that the re-orientation trajectory in the minimum metric control is identical both for the movement.

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For an application of PDE to hypersonic air-breathing engines, thrust performances of air-breathing PDEs are estimated analytically in this study. Thrust/weight ratio of the engine decreases with increasing flight speed because of a low density of atmospheric air. In order to improve this disadvantage, new concept air-breathing PDE is proposed. By closing the exit of the PDE combustion chamber during a filling phase, ram pressure of the intake can be utilized effectively. Furthermore, by cooling the intake exit air by the cryogenic fuel, flight range can be extended for M 0.5. After system analysis of the air-breathing PDE, PDRJE (Pulse Detonation Ram Jet Engine) with air precooler is proposed. Size of PDRJE is 0.3 m in cowl inlet diameter. Engine thrust and weight are estimated to be 4.6x10³ N and 74 kg respectively. To evaluate the basic performances of the PDE, firing test of the scaled model was performed using GH2 and GOX. Heat flux of the combustion chamber was approximately 5 kW/m²/Hz.

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Recently, various kinds of planetary explorations have become more feasible, taking the advantage of low thrust propulsion means such as ion engines that have come into practical use. The field of space activity has now been expanded even to the rim of the outer solar system. In this context, the Japan Aerospace Exploration Agency (JAXA) has started investigating a Deep Space Port built at the L2 Lagrange point in the Sun-Earth system. For the purpose of making the deep space port practically useful, there is a need to establish a method to making spaceship depart and return from/to the port. This paper first discusses the escape maneuvers originating from the L2 point under the restricted three-body problem. Impulsive maneuvers from the L2 point are extensively studied here, and using the results, optimal low-thrust escape strategies are synthesized. The paper presents, as the major results, the optimal escape and acceleration maneuvers schemes using Electric Delta-V Earth Gravity Assist (EDVEGA) technique, in which the comprehensive delta-V evaluation method is presented. The results obtained are highly useful, ushering the future solar system exploration era.

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