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A Mars orbiter/lander mission using a micro-satellite (less than 100 kg at Mars arrival) with a deployable membrane aeroshell for the orbit insertion by the aerocapture and the electric propulsion for the trajectory maneuver was considered. The aerodynamic heating environment during the atmospheric flight was investigated solving the thermo-chemical nonequilibrium axi-symmetric Navier-Stokes equations around the spacecraft with the flare-shaped aeroshell. To obtain the appropriate amount of deceleration at the atmospheric pass, the drag modulation technique, in which the aeroshell is timely jettisoned from the spacecraft, was assumed. The hypersonic wind tunnel experiment successfully demonstrated that the backward jettison works well without significant time delay and attitude disturbance. Finally the corridor width for the entry path angle was estimated considering the ability of the orbit insertion, the peak aerodynamic heating, and the capability of the electric propulsion to raise the periapsis altitude. The combination of a low ballistic coefficient flight with the membrane aeroshell and an electric propulsion works well to acquire a finite width of the entry corridor for the aerocapture.

Aerothermodynamic Studies on Low-Ballistic-Coefficient Mars Aerocapture Vehicle with Drag Modulation and Electric Propulsion