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A fluid-dynamic investigation was carried out to determine the cause of intense heating observed on the lee meridian of hypersonic delta wings and also to derive means for its suppression. Several experimental techniques were combined with analysis of extensive heat-transfer measurements at M f -6 in a range of Reynolds number to acquire a general description of the lee-flow structure. With attached leading-edge flow on the delta wings, the dominant feature is a pair of embedded vortices on the lee meridian whose interaction with the boundary-layer is responsible for the observed local heating. On the basis of flow visualization results and heat-transfer correlations, a qualitative vortex flow model is proposed which differs essentially from the conventional inboard separation vortex model. This approach leads to successful methods for alleviation of vortex heating. The proposed concept of boundary-layer embedded vortices also appears to explain the hypersonic lee-side heating characteristics observed on a shuttle oribiter configuration. Nomenclature M^ = freestream Mach number P^ = freestream impact pressure P' = local impact pressure R^ = freestream unit Reynolds number R x =*'#oo R x . =(x-x*)RR L = model length Reynolds number St^ = Stanton number based on freestream conditions x = coordinate along center line from apex (or nose), positive downstream x* = distance along x to peak heating (or other flow characteristic) x' = coordinate along center line, originating at x = x* y = spanwise coordinate from center line z = coordinate normal to planar upper wing surface, positive upward a = incidence angle of upper planar surface, positive leeward a 0 = incidence relative to model reference axis A = width of vortex trace in oil pattern (Fig. 3) S -boundary-layer thickness in plane of symmetry 0 = angle between vortex trail and wing center line (Fig. 3) A = leading-edge sweep-back angle 4> = mayimum outflow angle in vortex trace (Fig. 3)

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