An expression is derived which relates the pressure on a wing in a supersonic free stream to the pressure on a thin wing with the same surface shape. The expression is used to find the pressure distribution for caret wings and flat delta wings with attached flow at their leading edges. The compression surface pressure distributions found are in good agreement with existing experimental and theoretical results, except when large pressure changes occur in the flow behind the attached shock wave. Some expansion surface results are also obtained for wings with an isentropic expansion at the leading edge. The effects of flow and geometry changes on the pressure distribution are investigated. It is found that a small improvement in the lift/drag ratio of a caret wing can be obtained by halving the anhedral required for the plane shock wave condition.

A unified theory is given of hypersonic and supersonic flow over the lower surface of a caret wing at certain off-design conditions when the bow shock is attached to the leading edges of the wing and when there exists no internal shock. The flow field on the lower surface of a caret wing consists of uniform flow regions near the leading edges, where the cross-flow is supersonic, and a non-uniform flow in the central region, where the cross-flow is subsonic. The basic assumption is that the flow in the central region differs slightly from the two-dimensional supersonic flow over a flat plate at the same angle of incidence as that of the lower ridge of the wing. Based on this assumption, a first-order perturbation flow is first calculated and then strained and corrected so that it matches the uniform flow which is obtained exactly. Slope discontinuities of the pressure curve are found at the cross-flow sonic line. Numerical examples and comparisons with previous theories and experiments are included.

The response of a flexible helicopter rotor blade to random loading is investigated, the random input being the vertical velocity component. The model takes into account blade flexibility in bending as well as torsion, and also general root rigidity. The spectral density and the mean square value of the transverse displacement are computed for both hingeless and hinged rotor blades and the results are evaluated.

Many investigators have studied the aerodynamics of axial flow turbomachinery but none has produced a complete map of the three-dimensional flow behind a rotor row. This is of considerable interest to the aero-acoustician. A system is described which uses a constant temperature hot-wire anemometer to analyse the flow behind such a rotor. Although much information may be extracted by using the technique, its interpretation depends to a large extent on its form of presentation. An analysis of the flow behind a research fan is used as a means of discussing various forms of visual presentation.

Two prediction methods for compressible turbulent boundary layers with air injection have been developed. The results from the two methods are compared with experiment for constant pressure flows at Mach numbers from 1.8 to 3.6. In general the results from the two methods and from experiment are in close agreement. However, there are some differences at the higher Mach numbers and it is suggested that both methods need further development for use at Mach numbers above 3.

Numerical methods are used to study the growth of waves of finite amplitude on a pair of parallel infinite vortices. The vortices are treated as lines except in so far as the detailed structure of the core is needed to remove consistently the singularity in the line integrals for the velocities of the vortices. It is shown that the vortices eventually touch and the shape of the wave at this instant is calculated. The wave is quite distorted at this instant, but it is shown that its gross properties are given roughly by linear theory.

Surface pressure measurements and schlieren photographs at M∞ = 2.50 and M∞ = 4.00 are presented for a family of axisymmetric elliptically-nosed models with cylindrical afterbodies. The results are an extension of those obtained by Holder and Chinneck on the same series of models at lower Mach numbers.

Some comparison is made with theoretical results, and the forebody surface pressures obtained experimentally are found to be in good agreement with results obtained using either the method of integral relations or the time-dependent method.