Experiments were conducted to study shock-induced separated13; flows on the lee surface of delta wings with sharp leading edge at supersonic speeds. Two sets of delta wings of13; different thickness (10xB0; and 25xB0; normal angle). each with13; leading edge sweep angles varying from 45xB0; to 70xB0;, were13; tested. The measurements. carried out in a Mach number13; 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 shockinduced 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.
展开▼
