Aerothermodynamic Analysis on Low-Ballistic-Coefficient Aerocapture Vehicle with Membrane Decelerator
- Paper number
IAC-05-E2.2.03
- Author
Ms. Kumiko Nakamura, University of Tokyo, Japan
- Coauthor
Dr. Kojiro Suzuki, University of Tokyo, Japan
- Year
2005
- Abstract
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 the atmospheric entry corridor is significantly extended and the aerodynamic heating is much reduced thanks to ultra low ballistic coefficient (smaller than 0.1 kg/m2) of such vehicle design. The aerocapture vehicle considered here consists of main body and spin-stabilized 100 meters diameter membrane disk which can possibly be used as solar sail or as an area for solar cells. The trajectory simulation of an ultra low-ballistic-coefficient aerocapture vehicle shows that there is sufficient allowance of aerocapture corridor by using jettison technique of the membrane decelerator even if the uncertainty in atmospheric density is considered. The preliminary analyses using the empirical prediction method for aerodynamic heating and aerodynamic force on the membrane decelerator show that the temperature and tensile stress of the membrane are below the limit of polyimide film material during the aerocapture flight. To predict the aerodynamic performance of such vehicle more accurately, the coupled numerical analyses of the flowfield and the structural dynamics are made. In the structural analysis model, the membrane is described as a network of pseudo-particles. For the flowfield, the Direct Simulation Monte Carlo (DSMC) method is used, since such low-ballistic coefficient vehicle flies in quite low density atmosphere at high altitude and a rarefied flow around it must be analyzed. The shape of the membrane is determined by the balance among the aerodynamic force, the centrifugal force caused by the spin and the elastic force on the membrane. If the disk shape of the membrane decelerator can be sustained during the atmospheric flight just by spinning the membrane around its axis, the vehicle mass is expected to be significantly reduced because any additional support structure is not necessary. Numerical results show that large membrane decelerator can survive under the aerodynamic force with the assist of the centrifugal force of the spin as long as it flies through low density atmosphere of high altitude at an appropriate spin rate. Parametric studies are made for various flight conditions and membrane decelerator design. It is shown that the present vehicle concept of a low-ballistic-coefficient aerocapture vehicle with membrane decelerator has much advantage for a planetary exploration probe on the trajectory design, thermal protection and vehicle mass in comparison with all-propulsive vehicles and aerocapture vehicles with conventional rigid aeroshells.
- Abstract document
- Manuscript document
IAC-05-E2.2.03.pdf (🔒 authorized access only).
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