It is shown, by the successive application of coordinate transformations involving complex analytic functions, how a general curvilinear coordinate system can be constructed which is suitable for geometries consisting of a finite surface together with two semi-infinite cylinders. A typical example of such a geometry is considered which comprises a wing-body combination in which two semi-infinite unswept wings are attached to a finite body. The potential problem associated with non-lifting incompressible flow past the wing-body combination is formulated in the general curvilinear coordinate system and is solved numerically. The distribution of the pressure coefficient over the wing-body surface is presented.

Results are presented of velocity and pressure measurements made in the initially laminar boundary layer in a streamwise corner formed by two flat plates at 90° to each other set at various incidences. The leading edges of the plates were sharp in contrast to earlier tests with an aerofoil type leading edge. It was found impossible to obtain a steady enough flow for useful measurements to be made at zero incidence and pressure gradient, a small incidence associated with a favourable pressure gradient was necessary. This is believed to be because of the development of separation bubbles at the sharp leading edge at very small incidences due to small variations of flow direction to be expected in a wind tunnel. The profiled nose used in earlier tests afforded flow conditions much closer to the ideal theoretical model involving zero pressure gradient, but it is argued that any nose however shaped may introduce disturbances in the form of characteristic secondary flows that may well determine the downstream response of the boundary layer. In any case the corner flow is highly unstable at all but very low Reynolds numbers, and in the absence of a region of favourable pressure gradient a Reynolds number in terms of distance downstream of the leading edge greater than about 105 is unlikely to be attained in practice with the flow remaining smooth and laminar.

An account is given of a numerical method for calculating transonic flow past an aerofoil with an allowance for viscous effects, providing that the boundary layer remains fully attached over the aerofoil surface. The method has been developed by combining, in an iterative manner, calculations of the inviscid flow with calculations of the compressible boundary layer and wake. The solution for the inviscid flow is obtained by an iterative scheme, originally established by Garabedian & Korn, which has been modified to give a more realistic representation of shock waves. The boundary-layer development is treated as laminar initially; at a certain transition position a turbulent boundary layer is assumed to develop, and this is determined by the lag-entrainment method of Green et al. Comparisons of the results from the numerical scheme with some experimental measurements are shown for various examples in which shock waves of moderate strength are present. The method predicts, with reasonable accuracy, both the detailed pressure distribution and the variation of drag coefficient with lift coefficient.

An investigation has been carried out to examine one aspect of supersonic jet-engine inlet flow, namely shock-wave and boundary-layer interaction in the presence of an expansion corner. The experiment was undertaken at Mach numbers of 1.8 and 2.5. The local Reynolds number per metre varied from 2.79 to 3.13 × 107 for M = 1.8 and from 4.20 to 6.4-8 × 107 for M = 2.5. Adiabatic turbulent boundary layers of about 6 mm thick on an expansion-corner plate were disturbed at different streamwise positions by oblique shocks with deflection angles of 4°, 6° and 8°. The pressure distributions, wave patterns and the extent of the separation regions were strongly influenced by the presence of the expansion corner, especially when the shock impingement positions were upstream of the corner. The pressure rise induced by the incident shock began to decrease in level when the shock impingement position was about 3-4 boundary-layer thicknesses upstream of the corner. As the incident shock moved towards the corner, the pressure rise and the extent of the separation region decreased. When the shock impingement position was at the corner, the reflected shock was effectively neutralised by the Prandtl-Meyer expansion; especially when the deflection angle of incident shock matched that of the expansion corner. The three dimensionality of the incident shock and separation region is due to the interactions between the glancing oblique incident shock and the tunnel’s sidewall boundary layers.