Trajectory Coupled Aerothermodynamics Modeling for Atmospheric Entry Probes at Hypersonic Velocities

he purpose of this research was to investigate the aerothermodynamics of atmospheric vehicles and to provide a computational database framework for trajectory information for such vehicle. The kinetic energy of a space vehicle as it enters the atmosphere is converted into heat. How much of the heat goes into the vehicle depends on the trajectory flown, size and shape of the vehicle. The choice of the trajectory is actually limited by the realities of vehicle dimensions lift, drag and the mass. A CFD tool can be employed to demonstrate the automatic capability of predicting the flow field around the vehicle. A platform independent GUI based tool ProbeMesh has been developed and integrated to a planetary probe database to study aerodynamic heating and to do trajectory analysis. A comprehensive Relational Database Management System (RDBMS) for atmospheric entry vehicles has been developed from the Planetary Mission Entry Vehicles Manual and also from other web sources. The database comprises vehicle dimensions, trajectory data, and aero-thermal data for many different ballistic entry vehicles. Comparative data analysis capability is achieved through this RDBMS. ProbeMesh GUI provides capabilities to choose from a list of flight vehicles or enter trajectory and geometry information of a vehicle in design so that the trajectory for the vehicle in study can be analyzed. An empirical model of the standard atmosphere GAME: General Atmospheric Model for Earth to calculate pressure, density, temperature, Reynolds number and speed of sound as a function of altitude has been developed. GRAM models have been used for MARS, Venus, Titan and Neptune atmospheres. The vehicular Flight path angle as a function of velocity and altitude to compute the trajectory of the vehicle as it traverses through the atmosphere has been modeled. A fourth order Runge-Kutta integration is employed for trajectory calculations. Fay-Riddell and Sutton-Grave empirical correlations have been used for the stagnation point convective heat transfer and Tauber-Sutton for the stagnation point Radiative heat transfer calculations. An approach is presented for TPS sizing. Nomenclature Greek Cp Specific heat at constant pressure (kJ/kg.K) α angle of attack Cv Specific heat at constant volume ρ freestream density (kg/m) h Enthalpy (J/kg) β Ballistic Coefficient (kg/m) Kn Knudsen number Υ ratio of specific heats Le Lewis number e surface emissivity of the material m Entry mass (Kg) σ Stefan-Boltzmann’s constant M Mach number γ Flight path angle (°) P Pressure (N/m) θ Surface area inclination angle Pr Prandtl’s number κ Thermal conductivity (W/m-°k) &conv q Convective heat transfer rate W/cm 2 Subscripts &rad q Radiative heat transfer rate W/cm 2 &tot q Total heat transfer rate W/cm 2 b forebody Rb Fore body radius (m) conv convective Re Reynolds number e entry Reff Effective radius n nose Rn Nose Radius rad radiative S Base area of the capsule (m) tot total t Time (sec) ∞ freestream T Temperature (K) ∞ V Freestream flight velocity (m/s) § Associate Professor, Department of Mechanical and Aerospace Engineering, SJSU, Senior Member AIAA † Graduate Researcher and Aerospace Engineer, MAE Department, Associate Member AIAA T 44 AIAA Aerospace Sciences Meeting and Exhibit 9 12 Jan 2006, Reno Hilton, Reno, Nevada Page 3 of 17