The characteristics of the flow field around a circular cylinder in cross-flow placed at various distances from a plane, parallel both to the flow and to the cylinder axis, were analysed using a hot wire anemometer. Experiments were performed in a wind tunnel with Reynolds numbers ranging from 0.85×105 to 3×105. The spectra of the hot wire signals were obtained using a Fast Fourier Transform technique programmed on a PDP 11/40 computer. As regards a smooth cylinder, the main features of the vortex shedding mechanism in the subcritical regime remained unaltered for distances from the plane greater than approximately 0.4 diameters; in particular the Strouhal frequency did not show any significant variation relative to the typical value for an isolated cylinder. As for lower values of the distance from the plane, the regular vortex shedding disappeared and the hot wire spectra showed typical turbulent features. The possibility of obtaining supercritical conditions by roughening the cylinder surface was confirmed together with the importance of the Reynolds number based on the typical roughness size, Rk, in the evaluation of the flow regime around the cylinder. In the case of roughened cylinders, and with values of Rk below-350, the regular vortex shedding disappeared at a distance from the plane smaller than 0.3 diameters. This fact suggests that, at least in part of the supercritical regime, the influence of the plane can be smaller than in the subcritical regime.

Applications of “integral transforms of in-plane coordinate variables” in order to formulate unsteady planar lifting surface theories are demonstrated for both sub- and supersonic inviscid flows. It is concise and pithy. Fourier transforms are exclusively used, except for only Laplace transform in the supersonic streamwise direction. It is found that the streamwise Fourier inversion in the subsonic case requires some caution. Concepts based on the theory of distributions seem to be essential, in order to solve the convergence difficulties of integrals. Apart from this caution, the method of integral transforms of in-plane coordinate variables makes it be pure-mathematical to formulate the lifting surface problems, and makes aerodynamicist’s experiences and physical models such as vortices or doublets be useless.

The aerodynamics of a thin airfoil of arbitrary camber having a porous trailing edge in steady, subsonic, compressible potential flow is investigated. In the special case of a flat plate airfoil with a porous trailing edge, an exact, closed form solution is obtained using complex variable theory. The pressure loading on the airfoil, the lift and pitching moments are exhibited explicitly along with typical numerical results. The corresponding situation in supersonic flow is also considered.

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.

The ‘double integral’ method, employed by Squire and Trouncer to calculate the flow of a round turbulent jet in a moving stream, is generalised. Representation of the velocity profile in simple form and use of an arbitrary upper limit of integration in the second application of the momentum integral equation permits the recovery of the well-known similarity solutions for plane and axi-symmetric laminar jets issuing into a quiescent medium. Additionally, an approximation for the decay of centre-line velocity is obtained for the non-similar situation when there is an ambient co-flowing stream. This agrees well with Wygnanski’s perturbation solution for the plane laminar jet. The results of Squire and Trouncer are re-examined in the light of this generalised approach. The work has application to the operation of fluidic sensors.

A linear stability analysis was made for a radial wall jet. A perturbation method against the two-dimensional wall jet was used for the formulation, from which a non-homogeneous Orr-Sommerfeld equation was derived. The computation showed that disturbances are more unstable in the radial wall jet than in the two-dimensional case, which agrees qualitatively with an experimental evidence.

Base cavities (hollow extensions mounted behind the rear end of a bluff body) are known to reduce substantially the aerodynamic drag of two-dimensional bodies. For axisymmetric bodies their effect is not well established, and the objective of this work was to assess their drag-reducing potential. Experiments were performed using three different types of base cavities, one solid walled and two ventilated, each with six different depths ranging from 0.1 to 0.9 body diameters. All three types of cavities reduced the body drag for small cavity depths, with a maximum drag coefficient reduction of 0.027 being obtained with a ventilated cavity 0.2 diameters deep. The reductions are an order of magnitude lower than those reported for two-dimensional bodies, but are achieved with a much shorter cavity depth. Although the ventilated cavities had a beneficial effect for small depths, at large depths (greater than 0.5 diameters) they had an opposite effect – a very large drag increase. The cause of this increase is as yet not understood. Hot-wire investigation revealed that base cavities suppress wake periodicity. For cavities of small depth this correlated with a reduction in drag, but for large cavity depths the trends of the intensity of the periodic motion and of drag were not always the same.

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.

This paper presents a semi-empirical theoretical model for calculating the effect of nose bluntness on the stability derivatives of oscillating slender cones at hypersonic Mach numbers. It is based on a hybrid blast wave analogy/shock-expansion flow model and is used to obtain closed form analytic expressions for the derivatives for oscillating slender cones. Two models based on zero thickness and finite thickness entropy layers are proposed which are seen to be appropriate to the cases of very small and large nose bluntnesses, respectively. The results are compared with new and existing experimental data and with the predictions of previous theoretical methods.

In Part 1 expressions are derived for the coupled torsional, extensional and flexural stiffnesses of a fibre composite tube, such as a helicopter blade, which is subjected to torsion, longitudinal tension, chordwise and flapping bending moments and shear. The theory is an extension of Batho-Bredt engineering analysis and is valid for a cylindrical tube of arbitrary cross-section with an arbitrary circumferential distribution of fibre composite plies. Particular attention is paid to the coupling effects in which an asymmetric fibre lay-up results in a twisting of the tube under bending and/or tension. Consideration is also given to the influence on the stiffness characteristics of an initial twist in the tube.

In Part 2 the analysis of Part 1 is developed numerically for tubes representative of GFRP blades. Static and dynamic aspects are considered which pave the way for aeroelastic tailoring studies of such blades.

This paper investigates the possibility of increasing the maximum lift of a single element wing section by the introduction of camber near the trailing edge. The optimum upper surface velocity distribution consists of a constant high velocity “rooftop” over the forward part of the wing section followed by transition and an optimised turbulent boundary layer recovery region to the trailing edge. The height of the rooftop, which essentially determines the maximum lift coefficient, is limited by the amount of recovery attainable without separation of the turbulent boundary layer. An increase in the upper surface velocity at the trailing edge would allow the whole top surface velocity distribution to be raised resulting in a considerable increase in maximum lift. The introduction of camber over the rear part of the wing section should produce the required increase in trailing edge velocity on the upper surface. The practical possibilities of using this technique for increased maximum lift were examined by designing a high lift wing section and testing it in the University of Alberta Low Turbulence Wind Tunnel. The results showed that the new wing section behaved qualitatively as expected. A difference of about 0.5U∞ between upper and lower surface velocities at the trailing edge was achieved, even though the thick boundary layers near the trailing edge tend to reduce the effect of camber. A maximum lift coefficient of 2.64 was measured at a Reynolds number of 1.0 × 106.

The effects of finite chordwise flexibility on the development of lift on a rapidly started airfoil are studied. This is a generalization of the so-called Wagner problem of unsteady aerodynamics. It is found that the lift varies with time in a non-monotonous fashion, including more than one extremum. The lift at large times is found to be lower, the more flexible the foil. Changes in the wake vortex sheet are also discussed.

Some velocity profile and skin friction measurements for turbulent boundary layers at subsonic and supersonic speeds are reviewed to derive the magnitude of the wake component and the effects of its variation with Reynolds number on predictions of skin friction. As Reynolds number increases the wake component increases rapidly, reaches a maximum and then decreases. The measurements considered suggest that this maximum may be higher at supersonic speeds than at subsonic speeds. This wake component variation has a significant influence on predictions of skin friction in the Reynolds number range of interest on wind tunnel models.

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.

The behaviour of streams of solid particles when ejected into a fluid crossflow suggests that there are considerable forces caused by aerodynamic interaction between closely spaced particles. The overall effect is a dispersion of the solids normal to the fluid flow direction and a bunching up along the axis. This effect is apparent in aerial work where particulate solids are released from aircraft and also in dilute two phase flow especially at the onset of choking or saltation. The experimental work reported here contrives to measure these forces for spheres at the appropriate Reynolds numbers. The results are considered to be sufficiently accurate and comprehensive to give some understanding of the phenomena occurring and to permit an explanation, at least qualitatively, of the behaviour mentioned.