In this study, numerical method for a global stability analysis based on a linear perturbation equation and an approximate eigensystem analysis is introduced by an improvement of Chiba’s method, which is a numerical method for a global stability analysis. The advantage of the improved method is that we can reduce the computational procedure approximately half with increasing the accuracy of velocity disturbances of the global stability analysis. In order to examine the validity of the modified method, the onset of the oscillation in the flow around a two-dimensional cylinder is calculated. The mode of the transition from steady flow to oscillatory flow calculated by the improved method is consistent with the mode calculated by Chiba’s method. The influence of the size of the computational domain on the result of the global stability analysis is checked by calculating the stability of the flow in a certain region of the flowfield. At a supercritical condition, two types of the basic flows, symmetric-type and averaged-type, are used to calculate the stability of the flowfield. The result of the symmetric-type basic flow agrees with Zebib’s result whose basic flow is given by the symmetric flow. The averaged-type basic flow is obtained by averaging the CFD result of the oscillating flowfield at a supercritical condition. The frequency of the flow oscillation which is observed in the CFD result agree with not the result of symmetric-type but the result of the averaged-type basic flow.

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An interferometric CT technique has been applied to observe a 3-D unsteady and shock/vortex interacting flow field induced by shock waves. A small duct model with a pair of circular open ends is introduced in the test section of diaphragmless shock tube, which can be rotated around its central axis to change the observation angle. The estimated shock Mach number is about 2.3 at the exits of the model in nitrogen gas. Computational fluid dynamics (CFD) simulation by TVD scheme is also applied to the unsteady 3-D inviscid flow field. We developed a novel presentation method of flow field, which is named as ‘Distribution Combined Schlieren Images (DCSI) method’ to the CT results to demonstrate the 3-D features of complex flow field in our study. The results of CT measurement and CFD simulation are discussed.

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We designed two electrothermal pulsed plasma thrusters (PPT). A preliminary PPT using a poly-tetrafluoroethylene (PTFE) tube as a propellant showed thrust efficiencies of 10–12% with stored energy of 21.4J at cavity lengths between 14mm and 29mm. High specific impulse and small impulse bit were obtained with short cavity, and low specific impulse and large impulse bit with long cavity. A PPT with a propellant feeding mechanism using two PTFE bars showed impulse bit per joule of 43–48μNs/J, specific impulse of 470–500s and thrust efficiency of 10–12% with stored energy of 4.5–14.6J. Both the PPT and capacitors were mounted on a thrust stand with an 1-m-long arm developed for precise measurement of an impulse bit. With a repetitive 10,000-shot operation, a total impulse of approximately 3.6Ns was obtained, and each PTFE bar of approximately 2mm was supplied.

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An air-breathing pulse laser powered launcher has been proposed as an alternative to conventional chemical launch systems. The trajectory from the ground to a geosynchronous transfer orbit by pulse laser propulsion is calculated by modeling the thrust during pulsejet, ramjet and rocket flight modes, and the launch cost is estimated. The results show that the pulse laser powered launcher can transfer 0.085kg payload per 1MW beam power to a geosynchronous orbit, and the cost becomes quarter of existing systems if one can divide a single launch into 22,500 multiple launches.

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This paper considers the design problem of flight controllers that simulate gust responses and pilot input responses simultaneously, and proposes a method to realize those; robust model-matching controllers are designed as feedback controllers using scaled H_∞ problems to simulate the gust responses of target aircraft, and right inverse systems are designed as feedforward controllers using H_∞ problems to simulate the pilot input responses of the aircraft. For the lateral/directional motions of an experimental aircraft, we design such controllers for two aircraft models. To confirm their performance, hardware-in-the-loop simulations and flight tests are conducted, and they show that designed controllers have good performance.

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Recently, as advanced drag prediction method in transonic region, a drag decomposition method is watched with keen interest. This method is based on and extended from the momentum conservation theory on the closed integral surface around the airplane, which is usually called ‘Control Volume Method’ (CVM). In this paper, aiming the next target which is the drag decomposition in supersonic flows, the validation study of the CVM in the supersonic region was conducted. Two dimensional structured mesh computation of NACA0012 airfoil was used for the investigation. At the lift and drag prediction using the CVM in the supersonic region, the discontinuous variation or oscillation of the lift and drag value was observed when the integral surface was set to some particular positions related to the generated shock waves. By the avoidance of the inappropriate positions, however, the good performance of the lift and drag prediction using CVM was achieved.

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