AEROTHERMODYNAMIC ANALYSIS ON LOW-BALLISTIC-COEFFICIENT AEROCAPTURE VEHICLE WITH MEMBRANE DECELERATOR

摘要

For an exploration of a planet with an atmosphere, aerocapture is known to be a promising technique to save the spacecraft fuel for orbit insertion and to augment its payload capability. However, very tight entry condition required for successful aerocapture under large uncertainty in the atmospheric properties of a planet and severe aerodynamic heating which results in a heavy heat shield prevent us from realizing it. In the present study, we propose an aerocapture vehicle with a large spinning membrane decelerator. Considering a mission to Saturn, various feasibility analyses are made and it is shown that there is sufficient atmospheric entry corridor and that the aerodynamic heating is much reduced thanks to ultra low ballistic coefficient (smaller than 0.1 kg/m~2 ) of such vehicle design. To predict the aerodynamicrnperformance of the vehicle accurately, the coupled numerical analyses of the flowfield and the structural dynamics are made. Numerical results show that large membrane decelerator can survive under the aerodynamic force with the assist of the centrifugal force of the appropriate spin or the light inflatable device as long as it flies through low-density atmosphere of high altitude. Parametric studies are made for various flight conditions and the membrane decelerator design. It is shown that the present concept of a low-ballistic-coefficient aerocapture vehicle with membrane decelerator has much advantage for a planetary exploration probe on entry corridor, thermal protection and vehicle mass in comparison with all-propulsive vehicles and aerocapture vehicles with conventional rigid aeroshells.