Post ight analysis of the Mars Path nder hypersonic, continuum aerodynamic data base is presented. Measured data include accelerations along the body axis and axis normal directions. Comparisons of pre ight simulation and measurements show good agreement. The prediction of two static instabilities associated with movement o f the sonic line from the shoulder to the nose and back was con rmed by measured normal accelerations. Reconstruction of atmospheric density during entry has an uncertainty directly proportional to the uncertainty in the predicted axial coe cient. The sensitivity of the moment coe cient to freestream density, kinetic models and center-of-gravity location are examined to provide additional consistency checks of the simulation with ight data. The atmospheric density as derived from axial coe cient and measured axial accelerations falls within the range required for sonic line shift and static stability transition as independently determined from normal accelerations.

A method to obtain coupled CFD-DSMC, 3-D flow field solutions for highly blunt bodies at low incidence is presented and applied to one concept of the Mars Sample Return Orbiter vehicle as a demonstration of the technique. CFD is used to solve the high-density blunt forebody flow defining an inflow boundary condition for a DSMC solution of the afterbody wake flow. By combining the two techniques in flow regions where most applicable, the entire mixed flow field is modeled in an appropriate manner.

An aerothermodynamic analysis of the Commercial Experiment Transporter (COMET) reentry capsule has been performed using the laminar thin-layer Navier-Stokes solver Langley Aerothermodynamic Upwind Relaxation Algorithm. Flowfield solutions were obtained at Mach numbers 1.5, 2, 5, 10, 15, 20, 25, and 27.5. Axisymmetric and 5, 10, and 20 degree angles of attack were considered across the Mach-number range, with the Mach 25 conditions taken to 90 degrees angle of attack and the Mach 27.5 cases taken to 60 degrees angle of attack. Detailed surface heat-transfer rates were computed at Mach 20 and 25, revealing that heating rates on the heat-shield shoulder ,can exceed t h e stagnation-point heating by 230 percent. Finiterate chemistry solutions were periormed above Mach 10, otherwise perfect gas computations were made. Drag, lift, and pitching moment coefficients are computed and details of a wake flow are presented. The effect of including the wake in the solution domain was investigated and base pressure corrections to forebody drag coefficients were numerically determined for t h e lower Mach numbers. Pitching moment comparisons are made with direct simulation Monte Carlo results in the more rarefied flow at t h e highest Mach numbers, showing agreement within two-percent. Thin-layer Navier-Stokes computations of the axial force are found to be 15 percent higher across the speed range than the empirical/Newtonian based results used during the initial trajectory analyses.

Abstract

Floweld solutions over the Mars Pathnder Probe spanning the trajectory through the Martian atmosphere at angles of attack from 0 to 11 degrees are obtained. Aerodynamic coecients derived from these solutions reveal two regions where the derivative of pitching moment with respect to angle of attack is positive at small angles of attack. The behavior is associated with the transition of the sonic line location between the blunted nose and the windside shoulder of the 70 degree half-angle cone in a gas with a low eective ratio of specic heats. The transition rst occurs as the shock layer gas chemistry evolves from highly nonequilibrium to near equilibrium, above approximately 6.5 km/s and 40 km altitude, causing the eective specic heat ratio to decrease. The transition next occurs in an equilibrium ow regime as velocities decrease through 3.5 km/s and the specic heat ratio increases again with decreasing enthalpy. The eects of the expansion over the shoulder into the wake are more strongly felt on the fustrum when the sonic line sits on the shoulder. The transition also produces a counter-intuitive trend in which windside heating levels decrease with increasing angle of attack resulting from an increase in the effective radius of curvature. Six-degree-of-freedom trajectory analyses utilizing the computed aerodynamic coecients predict a moderate, 3 to 4 degree increase in total angle of attack as the probe, spinning at approximately 2 revolutions per minute, passes through these regions.