Experiments are described in which the lift on a rectangular element of a two-dimensional wing and on a finite aspect ratio wing has been measured in grid turbulence. By measuring the spectrum of the lift and the spectrum of the turbulence upwash component, an experimental value for the turbulent admittance may be found. This is compared with a calculated value based upon linearised theory.

The effect of swirl on the high speed flow past blunt bodies is analysed by assuming constant density in the region between the shock wave and the body. For small swirl the stand-off distance is only slightly affected, but it is shown that there is a critical value of the swirl parameter which, if exceeded, will cause a jump in the position of the shock. This is demonstrated by solving the full constant-density equations for the flow past a sphere and by a perturbation expansion in powers of the density ratio across the shock for a more general body shape. The perturbation solution shows that the pressure coefficient on the body is constant at the critical swirl number.

The Dynamic Relaxation (DR) method is applied to the solution of geometrically non-linear, elastic, laminated plate, flexural and stability problems. Two categories of plate are considered, namely, those that are symmetrically and unsymmetrically laminated with respect to the plate middle surface. Whereas the former category exhibits a bending – twisting coupling phenomenon, the latter exhibits an extensional – flexural type of coupling. The effects of these coupling phenomena are evaluated by comparing the plate responses with those of corresponding homogeneous, specially orthotropic plates. With the exception of uniaxially compressed plates of the latter category in the post-buckling regime, it is found that for both flexural and stability problems the coupling phenomena cause a reduction in stiffness and the extent of this reduction is dependent on the lay-up of the laminate.

Three computer programmes have been written for the stability analysis of both isotropic and orthotropic, multi-plate structures in compression. The bases of these programmes are the exact and approximate Finite Strip methods. A considerable number of results for a range of multi-plate structures fabricated from both carbon fibre reinforced (orthotropic) and aluminium (isotropic) materials are presented in graphical form. These graphs serve to illustrate not only the influence of material orthotropy on the buckling response, but also the possible savings in weight that may be achieved by employing orthotropic materials in place of conventional isotropic materials.

This paper contains results for the response of a helicopter rotor blade under stationary random excitation when the helicopter is in forward flight. The blade model takes into account the bending and torsional modes as well as the root-rigidity conditions. The resulting linear dynamical system equations with periodic coefficients are treated within the framework of the filtering theory to yield matrix ordinary differential equations for the required response-statistics themselves.

A theoretical and experimental investigation has been made into the effect of swirl on a supersonic jet. In certain circumstances a core of reversed flow can occur along the axis of symmetry and the jet mixes much more quickly with the ambient air than does the corresponding non-rotating jet. A characteristic calculation has been performed for the first cell in the case where the axial flow is not reversed. This indicates a shortening of the cell length and the formation of shocks nearer the nozzle.

If surface fences are to be applied for measuring skin friction in three-dimensional boundary layers they must be calibrated for both magnitude and direction of the shear stress. Results of the calibration for fences of different height are given. Furthermore, a manufacturing process and a mounting procedure are described to obtain surface fences with identical calibration curves.

A flow-restriction gas-heater concept is presented. Basically the gas is heated in multiple compressions, with the help of a reciprocating free piston, and with high entropy production within the gas due to the presence of a flow-restricting orifice. The characteristics of such continuously-operating heaters are studied by means of a simple theoretical model, including a first-order estimate of the significant heat losses. The independent gasdynamic and geometrical parameters are varied over practicable ranges, and steady-state heating cycles are digitally computed. The numerical results indicate interesting applications for the heater when the heat losses can be controlled.

This paper presents a contribution to Watson’s formulae and gives analogous formulae for the higher derivatives of the functions ψm and ϵm of classical aerofoil theory. In Appendix A these formulae are related to Thwaites’s method of solving Theodorsen’s non-linear integral equation. The resultant equations are also given in matrix form, which enables the calculations to be mechanised by using automatic computers.

Existing effective viscosity models which have been very valuable in the mean field closure method for turbulent boundary layer computation have shown certain undesirable limitations for certain realistic but general boundary layer flows. The more general flows usually involve non-negligible considerations of pressure gradients and such wall conditions as roughness, curvature and aspiration or transpiration in varying degrees of importance. The effects of these external and wall influences have, unfortunately, been underplayed by most existing effective viscosity models. The present model of the effective viscosity is developed for a general flow and has shown remarkable agreement with experimentation, without being any more complex than existing models.

When the incompressible potential flow past a three-dimensional body is represented by source distributions on the body surface, these source distributions have singularities near an edge or corner, for example á trailing edge of a wing or the (unfaired) intersection of a body and a wing. The nature of these singularities is discussed. When assuming slow variations of the geometry in the main flow direction we can consider a two-dimensional problem in the cross-flow plane. Here the tangential velocities and source distributions are proportional to certain powers of the distance from the corner. For example at a convex right-angled corner these powers are − ⅓ in the asymmetric case (the bisector is a potential line) and ⅓ in the symmetric case (the bisector is a streamline) for both sources and tangential velocities. At a concave right-angled corner the corresponding values for the source distributions are ⅓ (asymmetric case) and − ⅓ (symmetric case) whereas they are 1 and 3 respectively for the tangential velocities.

Vandrey’s procedure, based on the method of singularities (ring vortices), for finding the pressure distribution on an axisymmetric body in a uniform stream is extended to the case of flow past a similar body in a uniform stream within a cylindrical duct of infinite length. The final form of the integral equation for the velocity distribution on the body is the same as that given by Vandrey; however, its kernel possesses additional terms representing the influence of the duct. Numerical solutions are worked out for varying radii ratio between a sphere and a duct and also between the more general-shaped axisymmetric body and a duct.

The effects of contraction geometry on non-isotropic free-stream turbulence were investigated experimentally. Eight nozzles of circular cross-section, having different contour shapes and various non-dimensional lengths, were tested. The contraction ratio was 10:1. The results are compared with test data for isotropic turbulence and with a well-known linear theory. It is found that the variation of the lateral component of turbulent energy along the nozzles is similar to that with isotropic free-stream turbulence, whereas the behaviour of the longitudinal component is entirely different. The change of turbulence depends on local contraction ratio only.

The effect of the operation of a 15 per cent trailing edge control surface on the flow around a two-dimensional aerofoil in a high subsonic shock-free condition is investigated using the integral equation method developed by Nixon and Hancock. The effect of retaining the non-linearities in the transonic potential equation is to increase considerably the magnitude of the pressures over the front part of the aerofoil in comparison with the pressure found using a modified linearised theory in which significant second-order terms in the boundary conditions are retained. The magnitude of the lift coefficient and the pitching moment coefficient are increased by 10-15 per cent over the values found using the modified linear theory, and by 20 per cent over the values found using standard linear theory. However, the magnitude of the hinge-moment coefficient is decreased by the order of 20 per cent compared with the modified linear values and by 30 per cent when compared to the standard linear values.

Equations are derived for the total lift about low aspect ratio wings in an inviscid, incompressible, uniform flow. The analysis extends Bollay’s analysis, which followed the line suggested by Prandtl, and it is shown experimentally that Bollay’s simplifying assumptions can still be applied. Agreement is good between the present theory and the author’s experiment for values of total lift with small tip clearance.

A study is made on the aerodynamic interference between a wing and a surface of velocity discontinuity in a non-uniform potential flow field. Actually, the surface of velocity discontinuity is deformed around the wing, which penetrates the surface, but no theoretical predictions of the surface deformation can be found in existing papers. In the present paper, a deformation of the surface is theoretically predicted, leading to the conclusion that a “sectorial region” will be formed on the wing. Formation of the sectorial region is recognised in new experiments and in many existing experimental results. Both theoretical and experimental analyses are made of the physical roles of the sectorial region.

The power spectral technique has been extended to show the effect of aerodynamic non-linearities on the normal acceleration response of a rigid aircraft in the cruise configuration. Non-linearities in the normal force and pitching moment variations with incidence have been considered. The resulting changes from the linear root-mean-square values of normal acceleration were only 3 to 5½% for a root-mean-square vertical gust velocity of 20 m/s.

Solutions to the problem of separated flow past slender delta wings for moderate values of a suitably defined incidence parameter have been calculated by Smith, using a vortex sheet model. By increasing the accuracy of the finite-difference technique, and by replacing Smith’s original nested iteration procedure, to solve the non-linear simultaneous equations that arise, by a Newton’s method, it is possible to extend the range of the incidence parameter over which solutions can be obtained. Furthermore for sufficiently small values of the incidence parameter, new and unexpected results in the form of vortex systems that originate inboard from the leading edge have been discovered. These new solutions are the only solutions, to the author’s knowledge, of a vortex sheet leaving a smooth surface.

Interest has centred upon the shape of the finite vortex sheet, the position of the isolated vortex, and the lift, and variations of these quantities are shown as functions of the incidence parameter. Although no experimental evidence is available, comparisons are made with the simpler Brown and Michael model in which all the vorticity is assumed to be concentrated onto an isolated line vortex. Agreement between these two models becomes very close as the value of the incidence parameter is reduced.

This paper gives two improved tapered elements for vibration analysis derived using quintic polynomial displacement functions. The elements employ different nodal continuity conditions. Results are compared with those given by the basic cubic approximations and analytical solutions for various end conditions. The advantages of the new elements over the basic element are discussed.

The concept of a fully-developed flow based on the hypothesis of selective memory is here applied to general wall-jet type flows. In the presence of a (constant) external stream, the free-stream velocity and the jet momentum flux are taken to be the chief quantities governing the development of the wall jet: two additional non-dimensional parameters, representing a momentum flux Reynolds number and the relative momentum defect in the initial boundary layer, are shown to have only a secondary effect on the fully-developed flow. The standard correlations so determined are also found to predict quite well the flow development in Gartshore and Newman’s experiments on wall jets in adverse pressure gradients; possible reasons for this somewhat surprising result are discussed. Finally it is shown, by application to the still-air case, that the parameters discovered in incompressible flow are, with appropriate but straightforward modification, successful in describing compressible wall jets also.