An Experimental Investigation of Mass Transfer Cooling Techniques for Atmospheric Entry Vehicles

Abstract

An experimental campaign was carried out to investigate the feasibility of using mass transfer cooling techniques such as thermal protection system (TPS) for Martian entry spacecraft’s configurations. The tests were performed using hypersonic shock tunnels HST2 and HST3, at Laboratory for Hypersonic and Shock Wave Research (LHSR), IISc, Bangalore, simulating low (1.5 km/s) and high enthalpy (2.5 km/s) flow conditions, respectively. A large angle (60° apex angle) blunt sphere-cone configuration was used as test model, and heat transfer rate to the model surface was measured using platinum thin film sensors. Film and transpiration cooling were the two techniques investigated using nitrogen and helium gas as coolants. The investigations showed that these techniques resulted in reduction in heat transfer, and it was more effective for the high enthalpy test case.

Reprinted by permission of the American Institute of Aeronautics and Astronautics, Inc, from Experimental Investigation of Heat Flux Mitigation During Martian Entry by Coolant Injection, Journal of Spacecraft and Rockets, Vol. 51, No. 4(2014), pp. 1363–1368 & Experimental Investigation on Transpiration Cooling Effectiveness for Spacecraft Entering Martian Atmosphere, AIAA Journal Vol. 54, No. 9(2016), pp. 2922–2926. Copyright © AIAA.