This paper discusses elliptic formation control of spinning tethered formation flying system. A virtual structure method is applied to control the formation, and a simple operation of an expansion ratio of the virtual structure is proposed to achieve the formation. Two types of elliptic formation are treated in this paper. One is the formation in which three spacecraft are placed on a same elliptic trajectory. The other is the elliptic formation in which each spacecraft draws an each elliptic trajectory. Accuracy of the formation, thruster force and tether tension during elliptic formation are considered with numerical simulations. From the results, these elliptic formations conclude to be useful in some situations.

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As future space vehicles, Reusable Launch Vehicle (RLV) needs to be developed, where there are two kinds of RLV: Single Stage To Orbit (SSTO) and Two Stage To Orbit (TSTO). In the latter case, the shock/shock interaction and shock/boundary layer interaction play a key role. In the present study, we focus on the supersonic flow field with aerodynamic interaction between a delta wing and a hemisphere-cylinder, which imitate a TSTO, where the clearance, h, between the delta wing and hemisphere-cylinder is a key parameter. As a result, complicated flow patterns were made clear, including separation bubbles.

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In this study, a Computational Aeroacoustics (CAA) solver, UPACS-LEE, is developed as an extension of UPACS, a multi-block CFD solver developed in ISTA/JAXA that can run on parallel computers. Using UPACS-LEE code, acoustic propagation from non-compact sound source can be calculated, including non-uniform mean flow effect and reflection at the wall. Also, high order scheme is implemented in order to resolve the waves with a small number of grid cells. Various benchmark problems are solved, and the results are compared with the analytical solutions in order to validate the reliability of the code. At the end of this paper, acoustic scattering around high lift devices of an aircraft is shown. The effect of the mean flow is investigated by comparing the result with that of the no-flow case. Including the effect of the mean flow results in the Doppler effect, and it also significantly influences the directivity of the sound.

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For a main propulsion system of small satellite or precise attitude control of various size satellites, a miniaturized microwave ion thruster based on HAYABUSA μ10 neutralizer was fabricated. This thruster has a 1.6cm beam diameter grid system with high hole number density for the improvement of beam extraction. Through the experiments, it was evaluated that the typical thruster performance was the thrust of 0.34mN, the thrust/power ratio of 16mN/kW, the propellant utilization efficiency of 68% and the specific impulse of 3,200s. Judging from the comparison with the other small thruster performance, it was estimated that this thruster can be a candidate for the small thruster.

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A scramjet engine with a wall-mounted hypermixer injector, which generates streamwise vortices for enhancing supersonic mixing and combustion, is examined at a Mach 8 simulated flight condition in the High Enthalpy Shock Tunnel (HIEST). The engine and the fuel injector are full scale models of the HyShot-IV flight experiment planned for 2005 by JAXA and University of Queensland (UQ). Main purpose of the present study is to clarify the combustion and operation characteristics of the hypermixer scramjet owing to the ability of the streamwise vortices for mixing enhancement and boundary layer control. For comparison, two injectors with normal and parallel injection without streamwise vortex generation are also examined. The results show the superior performance of the hypermixer injector in scramjet mode obtaining higher pressure rise in a shorter distance compared to the other two injectors. In the case of the hypermixer injector, a 1D analysis of an inviscid nozzle flow shows the increment in the specific impulse due to combustion to be 2,649 and 2,224 sec for the equivalence ratio, Φ=0.3 and 0.6, respectively. At Φ=1.0 and 1.5, sudden rapid combustion of the premixed fuel at the end of the combustor generates a strong pressure wave, which propagates upstream up to the injector location and decades there. As a result, a new quasi-steady combusting flow is established throughout the combustor downstream of the injector. The pressure wave is identified as a kind of detonation wave, which is suggested to propagate upstream mainly through the streamwise vortices. As a driving force of the upstream propagation of the detonation wave, mixing and combustion enhanced through the interaction between the detonation wave and the streamwise vortices are considered.

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This paper presents experimental results of H infinity tracking control for slewing flexible space structures. The model of flexible space structure employed in the experiment consists of a rigid body rotating around a shaft and a flexible beam attached to the rigid body. The minimum time optimal control for flexible structure, which is a multi bang-bang control, is used to obtain a reference optimal trajectory. Two full-bridge strain gauges are located at the root of the beam in order to sense the bending-moment of the flexible beam and a tachometer is located at the rotary shaft of the torque motor in order to measure the angular velocity of the rigid main body. The angular velocity and bending-moment are utilized to calculate the control torque provided to the main body. The effectiveness of the H infinity feedback tracking control on suppression of residual vibration and achieving the goal attitude is studied for two types of flexible beams experimentally.

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