Two alternative presentations are given of Patel’s Preston tube calibration. In the first, Δp/τw is tabulated as a function of Δpd2/ρν2; in the second, cf is plotted as a function of Δp/½ρU2 and Ud/ν, so that values of cf, can be read off directly.

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 non-linear mathematical model for the study of the dynamics of an extensible cable subjected to aerodynamic forces generated by a uniform flow field is formulated. Solutions are found considering large displacement caused by suddenly applied loads (i.e., gusts, shock waves, turbulence) for a range of flow speeds and cable lengths. Transition from overdamped to oscillatory motion is observed when flow speed and cable length are increased and decreased respectively. The stability of the system is discussed.

The impingement region produced by directing a uniform, axisymmetric, supersonic jet of air normally onto a large, flat plate has been investigated experimentally and theoretically for four jets in the Mach number range 1·64 to 2·77. A qualitative theoretical description of the flow in the neighbourhood of the sonic line is given. A single-strip version of the method of Polynomial Approximation and Integral Relations (PIR) is applied to the flow, using two alternative methods of determining the centre-line shock height. The PIR predictions are compared to experimental shock shapes and pressure distributions. It is found that a PIR method in which the sonic line is assumed to intersect the shock at the jet edge leads to very good agreement with experiment at the higher jet Mach numbers, but accuracy is much reduced at the lower Mach numbers, the shock height being in error by about 62 per cent at a jet Mach number of 1·64. A change in flow pattern at small nozzle-to-plate distances is reported.

A closed-form expression is derived which gives an approximate solution to the lift generated on a two-dimensional thin aerofoil in incompressible flow with a normal velocity component of the form exp [i (ωt–xx+yy)]. The inaccuracy of the solution when compared with other published work is compensated by the simplicity of the final expression, particularly if the result is required for the calculation of the sound power radiated by an aerofoil in a turbulent flow.

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.

A study is made of the interaction of a combination of free-vortex and source flow with a stationary surface. The laminar boundary layer flow can be expressed in ordinary differential equations by choosing suitable similarity transforms for the Navier-Stokes equations. When simplifying boundary-layer approximations are included, the equations do not yield any unique solution. Solutions to the complete equations are calculated numerically for the special case of equal source and vortex strengths for a limited range of Reynolds number. The results show the presence of “super” velocities and large pressure variations within the viscous layer.