A method for the estimation of the moments of inertia of small aircraft such as Remotely-Piloted Research Vehicles (RPRVs) has been developed. When both the moments acting on a vehicle pivoted at or near its C. G. and its angular accelerations are measured simultaneously, the moments of inertia can be estimated by directly fitting the equations of motion to the measured data with multiple regression. This method, which is a variation of the equation error method for parameter identification, has great advantages over the usual output error method, in the sense that the existence of the structure which holds the vehicle has nothing to do with the measurement accuracy, and that the integration of the equations of motion is not necessary. The result of our experiments shows the accuracy and ease of application of the method.

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Flutter experiments on six flexible slender body models like a spaceplane were conducted in a low speed wind tunnel. The model was supported vertically in its wing surface so as to be able to move freely in the horizontal plane. Theoretical analysis has been carried out by using Doublet-Point method for non-planer surfaces in calculation of unsteady aerodynamic force. The followings were revealed. 1) The so-called body flutter occurred in four modes with a rigid wing connected with its fuselage at one point different in its longitudinal position and the wing span and one model with a flexible wing connected at four points. 2) In another model with a flexible wing, a wing-fuselage coupled flutter occurred. 3) In all experiments the body divergence did not occur. 4) For the flutter-speed, -frequency and -mode, a resonable agreement is observed between experimental and theoretical results. 5) The theoretical analysis shows that the structural damping coefficient yields an influence on the flutter speed. 6) The flutter vibration in experiments was mild.

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A new formulation of solving the three dimensional laminar boundary layer equations was made using finite element method. The code was tested on a flat plate with attached cylinder and calculated results gave fairly good agreement with Cebeci's results by Box method. The spanwise flow in boundary layer along a thin rectangular wing at low angles of attack was, also, calculated.

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Projectiles were fired at hypersonic speeds into stoichiometric H₂-air mixtures. The flowfields around the projectiles were visualized using a high-speed framing schlieren technique. We comprehensively studied oblique detonation and oscillating combustion phenomena around the hypersonic projectiles, where the varied parameters were the projectile velocity, projectile-nose shape, projectile diameter and initial gas pressure. We studied, in detail, the dependence of the oscillating-combustion cycles on the projectile velocity. It has been found that the observed four oscillation modes are composed of three fundamental modes. We compared the present experimental results with the numerical results of Matsuo, Fujii and Fujiwara, and discussed a mechanism generating oscillating combustion.

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An aerodynamic design method that treats multiple wings as well as their mutual interaction is proposed. This method provides section shapes of wings which realize specified surface pressure distributions. It can be applied to three dimensional design problems in potential, invisid and viscous, subsonic and transonic flow fields. The basic idea is an iterative residual reduction method using a new inverse problem solver which is the extension of Takanashi's integral equation method. An inverse problem is reformulated to be integral equations which express the relation of pressure differences to geometrical changes with consideration of interacting effects among wings. The inverse problem is solved numerically by introducing piecewise linear/constant function approximation. This method works well on several preliminary design problems. From both viewpoints of quality of solutions and efficiency, the method is found to be promising for complicated design problems.

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Objective of the present study is to clarify the characteristics of the worst disturbance on flexible space structures. “The worst disturbance” is defined as a disturbance which most effectively causes the vibration of a dynamical system and is formulated theoretically by using the idea of H_∞ control theory and the optimal control problem. Flexible structures are modeled by an Euler-Bernoulli beam and the worst disturbance on the flexible space structure model is obtained mathematically by numerical simulation.

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As ignition method in supersonic air streams, direct injection of dc arc discharges has been proposed and empirically examined. Since the electrodes are buried parallel to the streams in the same level of one of the chamber walls and the discharged arc is pushed into the main air flows by carrier gases, chemically active species are effectively produced and exposed in the combustion chamber. The experiments conducted so far showed 1) this direct arc injection method (DAIM) can ignite methane fuel in the flow of M=2.3 and total temperature T₀ as low as 300K and 2) methane air stoichiometric premixed gas is most effective for ignition as the carrier gas used to produce active species and pushed into the main flow.

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