Experimental Heat Transfer at Hypersonic Mach Number

Citation

DeLauer, Richard Daniel (1953) Experimental Heat Transfer at Hypersonic Mach Number. Dissertation (Ph.D.), California Institute of Technology. doi:10.7907/59SJ-TR03. https://resolver.caltech.edu/CaltechETD:etd-04222003-111626

Abstract

An experimental investigation was conducted in Leg 1 of the GALCIT 5 x 5 inch Hypersonic Wind Tunnel to determine the heat transfer coefficients of the laminar boundary layer on a cooled flat plate at a nominal Mach number of 5.8. As a consequence of the investigation, flat plate recovery factors were determined and the effect of condensation on heat transfer was noted. In addition qualitative results as to the laminar boundary layer transition and separation are also presented.

The tests were conducted with a ratio of wall temperature to free stream temperature (T_w/T_δ) of approximately 6.2; but under stagnation temperature conditions ranging from 200°F to 285°F. The stagnation pressure range of 60 psia to 115.5 psia provided a maximum Reynolds number of 2.1 x 10⁶.

A flat plate temperature recovery factor of .858 ± .004 was determined, and it was concluded that the temperature recovery factor range of Mach number independence could be extended to a Mach number of 5.8. The independence of the recovery factor on Reynolds number up to the beginning of the laminar boundary layer transition was also substantiated.

The heat transfer coefficients were obtained for a negative temperature gradient over a considerable portion of the plate. The effect of these gradients produced values considerably higher than would be expected for an isothermal surface. These results, when related the constant temperature case by a theoretical calculation, were in good agreement, with the theoretical results and the results of a friction investigation carried out at the same Mach number. The accuracy of the results was estimated to be ±10% from a value of Nu/Re^1/2Pr^1/3 - .285.

There was no apparent effect on the heat transfer coefficient by condensation, but the adiabatic wall temperature appeared to be 2% lower than for the condensation free flow. Due to a step increase in thickness of the model at the ten inch station, the shock wave-boundary layer interaction appears to produce laminar boundary layer transition at a Reynolds number of 1.3 x 10⁶, and upon reducing the Reynolds number further, the transition point is subjected to an adverse pressure gradient which results in a boundary layer separation.

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