Abstract
This report presents the latest test results on the aeroheating of an Apollo capsule model under high-enthalpy flow conditions in the free-piston shock tunnel HIEST at JAXA Kakuda. The objective was to obtain benchmark data to validate JAXA in-house numerical simulation codes developed to predict hypervelocity flow fields such as the external flow around a re-entry vehicle. The SUS-304 stainless steel model had a maximum diameter of 250 mm and was equipped with 84 miniature co-axial thermocouples on its windward surface. Twelve thermocouples were also mounted on the leeward side of the model and were used to determine the establishment of flow around the model. Heat flux distribution around the model was measured in a high-enthalpy, high-pressure (i.e. high Reynolds number) flow at 0o and 30o angles of attack. Aeroheating characteristics were observed with a fully laminar boundary layer and with a transition boundary layer. To change the free-stream Reynolds number, stagnation pressure was varied from 14 MPa to 54 MPa and stagnation enthalpy was varied from 3.6 MJ/kg to 21 MJ/kg. The flow characteristics of the graphite nozzle throat, which was designed for high-enthalpy, high-pressure shots, are also discussed.
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