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Supersonic wind tunnels with much lower stream disturbance levels than in conventional tunnels are required to advance transition research. The ultimate objectives of this research are to provide reliable predictions of transition from laminar to turbulent flow on supersonic flight vehicles and to develop techniques for the control and reduction of viscous drag and heat transfer. The experimental and theoretical methods used at NASA Langley to develop a low-disturbance pilot tunnel are described. Typical transition data obtained in this tunnel are compared with flight and previous wind-tunnel data and with predictions from linear stability theory,

Correlations for transition from laminar to turbulent flow on 45° and 60° swept cylinders based on data obtained in the NASA Langley Mach 3.5 Pilot Quiet Tunnel are presented. Variations of free-stream noise from high levels comparable to those in conventional wind tunnels to more than an order of magnitude lower had no effect on transition. However, when boundary-layer trips were attached to the leading edges, transition occurred at lower Reynolds numbers depending on both the trip height and the wind tunnel noise level. Compressible linear stability calculations have been performed for the boundary layer on an infinite swept cylinder. The boundary layer on the attachment line has a generalized inflection point similar to that present in a flat-plate boundary layer. The results show that Tollmien-Schlichting waves are amplified in the attachment line boundary layer and that oblique waves have the highest growth rates. Wall cooling tends to be stabilizing.