Aerothermodynamic environments for Mars entry, Mars return, and lunar return aerobraking missions

Details Aerothermodynamic environments for Mars entry, Mars return, and lunar return aerobraking missions Aerothermodynamic environments for Mars entry, Mars return, and lunar return aerobraking missions

Jun 1990

adsabs.harvard.edu/cgi-bin/nph-data_query?bibcode=1990jtht.confv....r&link_type=abstract

AIAA and ASME, Joint Thermophysics and Heat Transfer Conference, 5th, Seattle, WA, June 18-20, 1990. 23 p.

Other

Aerobraking, Aerothermodynamics, Convective Heat Transfer, Radiative Heat Transfer, Return To Earth Space Flight, Spacecraft Reentry, Aerocapture, Boundary Layer Equations, Carbon Dioxide, Earth (Planet), Mars (Planet), Moon, Navier-Stokes Equation, Nitrogen

Scientific paper

The aeroheating environments to vehicles undergoing Mars aerocapture, earth aerocapture from Mars, and earth aerocapture from the moon are presented. An engineering approach for the analysis of various types of vehicles and trajectories was taken, rather than performing a benchmark computation for a specific point at a selected time point in a trajectory. The radiation into Mars using the Mars Rover Sample Return (MRSR) 2-ft nose radius bionic remains a small contributor of heating for 6 to 10 km/sec; however, at 12 km/sec it becomes comparable with the convection. For earth aerocapture, returning from Mars, peak radiation for the MRSR SRC is only 25 percent of the peak convection for the 12-km/sec trajectory. However, when large vehicles are considered with this trajectory, peak radiation can become 2 to 4 times higher than the peak convection. For both Mars entry and return, a partially ablative Thermal Protection System (TPS) would be required, but for Lunar Transfer Vehicle return an all-reusable TPS can be used.

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