Forebody and aftbody heat transfer rates of a small-sized Martian aerocapture demonstrator are assessed for the preliminary design study of an aeroshell equipped with lightweight thermal protection system. Flowfields around the vehicle in hypersonic lifting flight are computed in a three-dimensional manner for representative flight conditions along the aerocapture trajectory. Finite rate chemistry and thermal relaxation for the Martian atmosphere are taken into consideration. Radiative heat transfer rates are calculated as well for both forebody and aftbody, taking into account not only electronic transitions producing vacuum ultra violet to visible radiation but also molecular ro-vibrational transitions generating infra-red radiation. The results indicate that the radiative heat transfer can be comparative to the convective heat transfer in a certain part of the aftbody aeroshell, though its absolute magnitude is much less intense than that of the forebody aeroshell.
NomenclatureC A , C N Axial and normal coefficient, respectively C D , C L Drag and lift coefficient, respectively C m Pitching moment coefficient h Altitude, km K n Knudsen number L/D Lift-to-drag ratio n Number density, m −3 p Pressure, Pa q Heat transfer rate, W/m 2 R n Radius of local curvature at stagnation point, m R s Radius of local curvature at shoulder, m T Atmospheric gas temperature, or translational-rotational temperature, K T v Vibrational-electronic temperature, K t Time from atmospheric entry interface point, s V a Velocity relative to the ground, m/s V inf Velocity of approach at infinity, m/s X,Y ,Z Cartesian coordinates, m α Angle of attack, deg. ΔV Velocity increment required for periapsis raise maneuver, m/s Subscript c, r Convective and radiative heat transfer rate, respectively