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SummaryExperiments were conducted to study the types of flow that occur on the lee surface of delta wings at supersonic speeds. Two sets of flat topped delta wings of different thickness (wedges with 10° and 25° normal angle respectively), each with leading edge sweep angles of 45°, 50°, 60° and 70°, were tested. The measurements, carried out at Mach numbers of 1·4, 1·6, 1·8, 2·0, 2·5 and 3·0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wing only). In addition, a 60° sweptback delta wing with a normal angle of 40° was also tested. The tests on this wing included both oil flow visualisations and static pressure measurements. From these and other existing measurements, the leeside flows have been classified into nine distinct types, namely (i) leading edge separated flow with secondary separation, (ii) leading edge separated flow with secondary and tertiary separation, (iii) leading edge separated flow with a shock wave beneath the primary vortex, (iv) leading edge separated flow with shock-induced secondary separation, (v) fully attached flow, (vi) flow attached at the leading edge with inboard shock-induced separation, (vii) mixed type of flow, (viii) flow with a leading edge separation bubble and (ix) leading edge separated flow with a shock wave lying on the lee surface in between the leading edge vortices. These types of flow have been displayed in a plane of Mach number and angle of attack normal to the leading edge. The experimental results indicate that increasing wing thickness has no qualitative effect on the types of flow observed but does shift the boundaries between some of the types of flow.

SummaryThe concept of a fully-developed flow based on the hypothesis of selective memory is here applied to general wall-jet type flows. In the presence of a (constant) external stream, the free-stream velocity and the jet momentum flux are taken to be the chief quantities governing the development of the wall jet: two additional non-dimensional parameters, representing a momentum flux Reynolds number and the relative momentum defect in the initial boundary layer, are shown to have only a secondary effect on the fully-developed flow. The standard correlations so determined are also found to predict quite well the flow development in Gartshore and Newman’s experiments on wall jets in adverse pressure gradients; possible reasons for this somewhat surprising result are discussed. Finally it is shown, by application to the still-air case, that the parameters discovered in incompressible flow are, with appropriate but straightforward modification, successful in describing compressible wall jets also.

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Experiments were conducted to study shock-induced separated flows on the lee surface of delta wings with sharp leading edge at supersonic speeds. Two sets of delta wings of different thickness (10° and 25° normal angle), each with leading edge sweep angles varying from 45° to 70°, were tested. The measurements, carried out in a Mach number range from 1.4 to 3.0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wings only). Using the test results, some features of shock-induced separated flows, including in particular the boundary between this type of flow and fully attached flow, have been determined. The experimental results indicate that this boundary does not seem to show any significant dependence on wing thickness within the limit of thicknesses tested. It is shown that this boundary can be predicted for thin delta wings using a well known criterion for incipient separation in a glancing shock wave boundary layer interaction, namely that a pressure rise of 1.5 is required across the shock. Comparison of the predicted boundary with experimental results (from oil flow visualisations) shows good agreement.