Nonadiabatic Model Wall Effects on Transonic Airfoil Performance in a Cryogenic Wind Tunnel,
DOUGLAS AIRCRAFT CO LONG BEACH CA
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This paper addresses the need to match aircraft surface thermal conditions that exist at in-flight conditions when testing models in a cryogenic wind tunnel. It reviews the effects of non-representative heat transfer for such basic viscous characteristics as the effects on boundary-layer transition location, on turbulent boundary layer integral parameters and skin friction, on the transonic turbulent boundary-layershock-wave interaction, and on separation onset and the extent of separated flow regions. A complementary experimental and computational investigation was conducted to help quantify the impact that nonadiabatic model wall conditions would have on aircraft configurations tested in a cryogenic wind tunnel, and to help establish the allowable deviation from adiabatic wall conditions that can be tolerated if reliable results are to be obtained. It is concluded that the temperature of the model in a transonic cryogenic tunnel should be within 1 of the adiabatic wall value as a maximum in fact, model temperatures should be maintained to even less than a 1 deviation from the adiabatic condition, particularly where flow separation effects are important.