The results obtained from exact numerical solutions to the laminar boundary layer equations are compared with those given by the approximate method of Luxton and Young and the comparison shows that the latter method may be used for cases involving non-constant wall temperatures. This satisfactory result arises essentially from the fact demonstrated by the calculations that the skin friction is relatively insensitive to wall temperature gradients. It is also shown from these results that there is a lag in the propagation of wall temperature changes through the boundary layer. In consequence the Stanton number is a heat transfer parameter of doubtful value in cases of non-uniform wall temperature.

Because of the lack of reliable data on heat transfer to long bodies, measurements have been made to cone-cylinder-flares at a Mach number of 5·2. Axial heat transfer distributions were closely predicted by a modified reference enthalpy method for both laminar and turbulent flows at zero incidence and the heat transfer to the flare windward generator at incidence agreed with laminar theory.

The stability of a smooth circular cylinder, free to translate horizontally and vertically against linear springs in the wake from an identical neighbouring cylinder, is studied using quasi-static aerodynamic derivatives and simple flutter theory. It is found that at spacing values between ten and twenty cylinder diameters (typical of the spacings employed on “bundled” overhead transmission lines) classical flutter of the leeward cylinder can occur in a certain critical range of wind speeds at certain orientations of this cylinder in the wake. However, the occurrence of flutter appears to be conditional on a positive difference of natural frequency between vertical and horizontal motions of the leeward cylinder in still air.

Classical static instability (divergence) of the leeward cylinder is also shown to be possible over the entire “incidence” range in the wake, but this occurs in a much higher wind speed range than that associated with flutter.

The paper presents a theoretical investigation of the flow when the wake from a single vibrating aerofoil interacts with a fixed downstream cascade set normal to the mean flow direction. It is concluded that the effects of the cascade extend upstream by a distance of about one quarter of the wavelength of the fluctuations in the wake. It is also found that the wake passes through the cascade unchanged in strength, but that the cascade sheds vorticity of equal and opposite magnitude, so that, if an average is taken in the direction parallel to the cascade, the wake is annulled.

This paper describes an experimental study of the interaction between a hypersonic flow and a transverse jet which issues radially from a slender conical model. Measurements in the region far upstream of the jet show qualitative agreement with free interaction analyses, while downstream a film cooling effect appears to reduce the heat transfer rate substantially. The effect of changes in Reynolds number, jet pressure ratio and mass flux is examined and compared with similar phenomena caused by fixed disturbances such as compression corners. The shape of the interaction shock wave is found to agree with the predictions of the second-order blast wave theory. Finally a correlation is found between the boundary layer separation length and a local mass flux parameter.

A numerical method for estimating the pressure distribution on arbitrary finite symmetrical wings with rounded leading edges at zero incidence in subsonic flow is presented. The subsonic flow equations are linearised so that the value of ∂ϕ/∂x at a point on the wing mean plane can be expressed as the generalised principal part of a singular double integral. This integral is modified so that it can be evaluated numerically. Finally Cn is obtained from ∂ϕ/∂x by using the approximation Cp = −2∂ϕ/∂x. Two numerical methods are described; one method to be used when the field point is close to the leading edge and the other for field points in other regions. Numerical results for two wings (Warren 12 and NPL 55° swept) are given and, wherever possible, the computed results are compared with results from other methods and with experiment.

A method of obtaining the load distribution on thin rectangular wings in non-stationary incompressible flow is presented, with particular reference to the problem of gust-induced loading. The method utilises solutions to a Fourier transform of the downwash integral equation, enabling the problem to be expressed in the form of a set of dual integral equations having a series solution. Some values of the overall lift coefficient are computed and compared with values of lift coefficient obtained by collocation methods. The method is easily extendable to subsonic compressible flows.