The objective of this study is to ascertain the aerodynamic characteristics of a low aspect ratio wing at Reynolds number Re ≦ 1 × 104 by conducting the wind tunnel test. The effects of the wing planform with the small aspect ratio at very low Reynolds numbers which correspond to an insect flight are largely unknown. The thin flat plate having rectangular planform with aspect ratio (AR) of 1 was selected as the wing model. Although the small lift slope and the large maximum lift coefficient provided by the additional vortex lift were obtained at Re = 1 × 104, the maximum lift coefficient decreased as the Reynolds number decreased and the additional vortex lift component was hardly seen at Re = 1 × 103.

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For planetary exploration missions using a lander, autonomous pinpoint landing capability whose precision is less than 100-meter must be needed in order to land on limited investigation areas. This capability cannot be realized with an inertial navigation system in terms of accuracy, so that the inertial error should be reduced in some way. One of solutions is image-based velocity measurement in a navigation path of the planetary lander. However, images taken by the lander have probrems such as frame rate limit and motion blur. Moreover, the available onboard resources of computation are limited. This paper proposes a velocity estimation method using a single blurred image for the pinpoint planatary landing. We present the method based on cepstral analysis and discuss feasibility through resource evaluation with a space-grade FPGA.

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The longitudinal combustion instability characteristics of a pintle-type injector for a bipropellant rocket engine combustor are investigated experimentally. An optically accessible combustion chamber are used to observe unsteady combustion behaviors under oscillating combustion conditions. CH* chemiluminescence and backlit spray images are observed simultaneously with two high-speed cameras. Two experiments with the propellant total momentum ratio (TMR) of 0.76 and 2.48 are conducted, whereas the combustion pressure and the propellant mixture ratio are 0.45 and 1.6, respectively. The combustion oscillations at the natural frequencies of the chamber longitudinal acoustic mode are observed when the TMR is 0.76. The combustion oscillation is caused by the coupling between the heat release and the combustion chamber acoustics. When the TMR is 2.48, the combustion oscillations at the frequencies of 400 and 800Hz, which are lower than the first longitudinal mode frequency, are observed. Since the 400Hz corresponds to the convective time scale in the combustion chamber, the oscillations could be caused by one of the convective modes such as entropy wave.

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In the research and development of electric propulsion, it is important to evaluate the propellant flow within ground test facility (vacuum chamber), however it is difficult to measure due to the rarefied pressure and the low particle velocity. In this paper, a rarefied dynamic pressure detector with a straight optic fiber as a detection string was proposed. Through the experimental evaluation, it was confirmed that its dynamic pressure resolution was approximately 1.0 × 10-5 Pa, that the propellant flow within a vacuum chamber was appropriately measured, and that the compensated propellant flow was obtained with the preliminary evaluation of the pumping flow.

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This paper presents the recent efforts to characterize the maximum load experienced by a Disk-Gap-Band type parachute to support Mars landing mission currently considered at JAXA. We simulated the structure-flow interaction with CFD solver coupled with a mass-spring-damper structural model. The interaction was modeled using a 3D elastic canopy model and the immersed boundary method. Flow parameters were computed for a Mach 1.6 free stream condition. The flow field solution was post processed to yield the drag coefficient. In addition, experimental campaigns were conducted at JAXA supersonic wind tunnel facilities. The time histories of the drag coefficient were recorded; the shock shapes and the canopy opening process were visualized by Schlieren diagnostics to provide data and validate the numerical models. The computed averaged drag coefficient was found to be within 5% with the measurement. The breathing frequency computed from Fast Fourier Transform analysis was found to be a similar tendency with the experimental data.

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A micro-thruster using stacked solid propellant pellets for nano-satellites has the high versatility. This thruster can generate arbitrary total impulse by changing the number of installed pellets. In order to install the thruster in a nano-satellite, it is necessary to clarify conditions of stable operation. We changed materials and manufacturing methods of the thruster, and measured the combustion chamber pressure and the burning rate. A certain condition of the thruster achieved stable operation. A difference between the burning rates of the propellant pins caused unexpected combustion chamber pressure rise. A silicone rubber contributed stable operation because it made the burning rates of the propellant pins even.

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