To substantiate an aerodynamic equivalence between blade flapping and feathering motions which is used widely in trim and performance calculations of a helicopter, numerical study are carried out using the CFD code based on unsteady 3 D compressible Euler equations with a moving grid system. Numerical calculations are conducted for a lifting rotor in hover and forward flight. By comparing unsteady aerodynamic loads on blade which execute pure flapping and feathering motions at various flight conditions, influences of unsteadiness due to blade motion and compressibility on the aerodynamic equivalence are clarified. In addition, effect of wake structure on the equivalence in hover and forward flight are also examined. It is ascertained that the blade flapping and feathering motions are equivalent aerodynamically both in hover and forward flight under the condition if the blade motion can be approximated by simple harmonic motions.

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An experimental study on the improvement of aerodynamic characteristics of a modified arrow wing by applying lateral blowing in subsonic flow has been conducted. The modified arrow wing, which is one of the baseline configurations of the proto-type of next-generation SST, is selected for the experiments. The testing model is the combination of a body of a circular cylinder and ogive and a modified arrow wing with aspect ratio of 1.91. The lateral blowing is realized by injecting a pair of sonic jets parallel to the trailing edge of the wing. The experiments have been conducted in the transonic wind tunnel of ISAS under the testing conditions of free-stream Mach number M_∞=0.3, 0.4 and 0.5, Reynolds number Re=4.77×10⁶∼7.34×10⁶, angle of attack α=−15°∼30° and jet momentum coefficient C_μ=0.0124∼0.0312. The results show that the C_L and L/D is increased by lateral blowing while C_D is slightly increased for positive α. The results suggest that the lateral blowing can be useful for the improvement of aerodynamic characteristics of the arrow wing in subsonic flow.

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This paper describes two-dimensional active flutter suppression to cope with the transonic dip using the sliding mode control. The airfoil model has plunge and pitch degrees of freedom with leading and trailing edge control surfaces. The aerodynamic forces acting on the airfoil, lift and pitching moment, are calculated by solving Euler's equations using computational fluid dynamics. At a specific altitude, flutter occurs between Mach number of 0.7 and 0.88, which corresponds to the transonic dip. The sliding mode control makes the airfoil to be stable all through the Mach number including the transonic dip. The sliding mode controller gives wider flutter margin than a linear quadratic regulator. These characteristics indicate that the sliding mode control is useful for active flutter suppression in the transonic flight.

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The ISAS balloon group has been developing and manufacturing plastic balloons made of very thin polyethylene films which can easily reach an altitude more than 40 km with a payload less than 10 kg. To launch the high altitude balloon safely, it is important to develop a launching method. The balloons with thin polyethylene films are normally launched using the dynamic-launching or static-launching method in which the bottom parts of balloons are folded on the ground. These methods depend greatly on the speed and direction of the surface wind at the time of launching, and the most desirable wind speed is less than 2 m/s. Therefore, there are limited opportunities for launching. To launch the thin film balloon easily with very few persons and with certainty, we have developed a new launching method called the packing launch-method. In this method, we pack the uninflated portion of the balloon in a bag. We can ignore the length of balloon in launching since the packed part of the balloon is extended in the sky after launching, and it is possible to launch a large balloon from a launch field of limited size. The packing launch-method is a revolutionary way of launching of high altitude balloons made of thin polyethylene films.

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An airflow sensor based on the measurement of pressure was proposed. This technique employed the one-dimensional flow equations in finite differential forms. The equations were applied to a cell located at the inlet of a circular pipe. The pressure was atmospheric at the pipe inlet and the pressure at the cell boundary Δx apart from the pipe inlet was measured. The velocity was therefore calculated instantaneously by the flow equations from the measured pressure. This technique has the following features: non-intrusiveness, automatic reverse flow detection, independence of initial conditions and minimum pressure loss. The technique also can be extended to a compressible flow.

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The essential benefit of HardWare-In-the-Loop (HWIL) simulation can be summarized as that the performance of autopilot system is evaluated realistically without the modeling error by using actual hardware such as seeker systems, autopilot systems and servo equipments. The most important requirement at the HWIL simulation test is to set the homing seeker at the 3-axis gimbals center of the flight motion table. But, because of the various reasons such as the length of the homing seeker, the structure of the flight motion table and the shape of attachments, this requirement on setting is not able to be satisfied. In this paper, the effect of this position error on the guidance and control system performance is analyzed and evaluated.

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