This study is devoted to investigate slew maneuver of a flexible space structure. The flexible space structure treated in the present study is a rigid body equipped with a flexible aluminum bean, and the control scheme is defined as a method to minimize the bending-moment at the root of the flexible appendage during slew maneuver. The hierarchical gradient algorithm is employed to obtain the optimal control profile. Original of the algorithm can only treat inequality constraints on the state variables, and not on the maximum value of control input. It is necessary in the present problem formulation to take into consideration constraints on the control input since maximum value of control input is naturally limited. This paper proposes a method to improve the hierarchical algorithm so that constraints on the control input can be treated. The method is analyzed numerically to verify the present modified algorithm, and to show that constraints both on the control input and bending-moment are satisfied. The resulting control profile is implemented in an experiment to verify its feasibility. Results of experiments show that bending-moment is effectively reduced in the case that constraint on the maximum value of bending moment is considered.

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Aerodynamic interaction between a side-jet from a blunt body and its external flowfield was experimentally investigated. Experiments were conducted at the shock tunnel of Nagoya University with a Mach number of 8.1. The model used in this study is a 10° half-angle blunted cone, where a circular sonic air jet is injected at right angles to the body surface on the leeward side. To investigate the flowfield and the effects of attack angle on interaction, detailed surface pressure measurements were performed as well as visualizations. As attack angle increases from 0 to 40deg, the region with aerodynamic interaction extends upstream from the nozzle. The magnitude of pressure changes due to the interaction decreases with attack angle, whereas the interaction amplification factor increases. Furthermore, a relatively large flow oscillation with about 5kHz, which was caused by the aerodynamic interaction, was observed, where a linear relation exists between Mach number and Strouhal number. Finally, a simple empirical model to show a characteristic of shock-wave-induced aerodynamic interaction is proposed here, which shows good agreement with experimental data.

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Effects of solid wall on instability of a supersonic shear layer were investigated with the viewpoint of linear stability analysis. It is known that Mach waves become to be radiated from shear layers, as flow Mach number increases. In this case, because closed lower pressure regions of pressure perturbation, which grows to be core of vortex, cannot be formed in the shear layers, growth rates of disturbances are decreased extremely. In the present study, solid wall was set parallel to a supersonic shear layer. It is shown that, with the solid wall, outgoing Mach waves are reflected and turn back to the supersonic shear layer. The outgoing and incoming Mach waves form standing waves, and closed lower pressure regions are formed in the supersonic shear layer. It is shown that, with the solid wall, growth rates of disturbances are increased compared with that for without the solid wall.

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Dynamics of pilot was studied quantitatively in relation to the change of dynamic stability characteristics of aircraft. Three pilot conducted bank control on flight simulator which simulate aircraft dynamics with large variety of dutch roll characteristics, and data was acquired. Dynamics of pilot was expressed by the transfer function which include dead time, gain, lead time constant, lag time constant. To determine these parameter, dynamics of pilot was identified by the acquired data using output error model. The dead time of the pilot was thought to be 0.3sec, because of that the error between the actual data of the aileron deflection angle in bank control and the output of the transfer function of pilot is smallest in case of assuming it. It has been understood that flying quality preferred by the pilot corresponds to that the airplane is controlled with the gain of the pilot more than the certain value changing with the flight condition, with the lead time constant of the pilot less than 0.8sec and with the lag time constant of the pilot less than 0.3sec.

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NAL and NASDA develop High Speed Flight Demonstrators (HSFDs) to examine an automatic take-off and landing technology (Phase 1) and to measure the transonic aerodynamic characteristics of a reusable space plane for getting the reference data of a CFD (Computational Fluid Dynamics) technology (Phase 2). As a HSFD navigation sensor, we develop GPS Aided Inertial Navigation Avionics (GAIA) whose distinctive feature is the usage of the carrier-phase DGPS (CDGPS)/INS hybrid navigation technology. This paper describes the design of GAIA. We show its accuracy, integrity, continuity and then availability performances that are analyzed by numerical simulations, which we first developed. Ground and flight tests are also carried out to confirm the performance of the GAIA flight model. Both results prove that GAIA meets the requirements of the navigation performance even under the severe flight conditions of HSFD.

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NAL has been promoting diagnostic study of the 750kW arc-heated wind tunnel flow to characterize various important flow parameters, such as heat flux and pressure, necessary for determining materials test conditions. In those measurement techniques, static pressure data is one of the most difficult experimental items to obtain. In this report, design of a static pressure probe and measurement system for the pressure, calibration system and the results, and evaluation of response time of the pressure are presented. The results show that static pressure distributions of the high enthalpy air flow in the NAL/NASDA 750kW arc-heated wind tunnel were obtained with good accuracy using these techniques.

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