Wind-tunnel experiments were conducted to determine the interference characteristics of an inclined turbulent jet exhausting from a flat plate into a turbulent subsonic crosswind. The jet was not yawed with respect to the undisturbed free stream. The suction force (lift loss), the jet penetration into the crosswind, the jet deflection and the rate of total pressure decay along the jet centre line were all reduced by an increase in the inclination of the jet. The centre of pressure moved downstream. These results were attributed to a decrease in the entrainment rate of the jet as the inclination increased in a downstream direction.

The formulation of curved finite elements to represent a two-dimensional circular sandwich ring with honeycomb core and laminated faces is investigated. Assumed stress hybrid and equilibrium methods are found to be easier to employ in this case than the displacement approach. Using these methods, an element stiffness matrix is developed. The approximations of membrane faces and an infinite core normal stiffness are then used to develop simpler elements. Test cases show that these assumptions may become invalid, but that they are adequate for most practical cases where the core thickness to radius ratio and the face thickness to core thickness ratio are both low.

This paper describes four techniques that can be applied to predict the response of an aircraft (including STOL and VTOL vehicles) to wind shear and turbulence during the landing approach. A comparison among the techniques is carried out by predicting the longitudinal response of a light STOL transport to atmospheric conditions within the earth’s planetary boundary layer. The turbulence data were generated by a boundary-layer wind tunnel and by an instrumented helicopter. Of the four techniques described, the flight path correlation technique is the most recent, having been developed to handle some of the problems facing the analysis of STOL and VTOL aircraft response. The other three approaches are the power spectral density technique, the equivalent deterministic input technique, and the in-flight turbulence measurement technique. The practical features of each of these are outlined and typical response predictions are plotted as a function of height during the landing approach. In general terms it was concluded that the techniques requiring the greatest amount of data-collection sophistication and computation (the in-flight data measurement and flight path correlation techniques) would produce the best results.

The method of Beecham and Titchener is extended to systems with n degrees of freedom and is shown to be a combination of the averaging principle and the method of variation of parameters. In this extended form, the method provides a powerful solution algorithm for non-linear problems such as those which arise in aircraft structural dynamics and aeroelasticity. The method is exemplified in application to a two-degree-of-freedom damped non-linear oscillator and to a binary (flexure-aileron) non-linear flutter system. The method is finally extended to non-linear differential equations in first-order form such as those which arise commonly in flight mechanics.

The flow field produced by the intersection of two plane solid surfaces in a supersonic stream is a complex interference flow. These flows can be fully compressive, fully expansive or of mixed compression-expansion nature. This paper presents a comparison of the flow field structure in an axial corner obtained experimentally with that predicted numerically by using a shock-capturing finite-difference method. The effect of sweep and surface deflection are evaluated and the general influence of each is presented for the three classes of corner flows. The results of this study show that the numerical method is a valuable aid in understanding the flow structure for simple configurations. In addition confidence in the numerical method is gained for use in solving the more general three-dimensional configurations where the flow is non-conical and several wave interactions may be present.

A series of advanced aerofoil sections suitable for aircraft propellers has been developed. The sections combine advanced supercritical now characteristics with good low-speed performance and are also directly applicable to other rotating aerodynamic machinery, such as low-solidity ducted fans, windmills, etc. The work has focused initially on propeller sections of around 6 per cent thickness/chord ratio, typical of the fundamental blade station where peak loading is usually carried (0.7 × radius). The complete section family covers a wide range of thickness/chord ratios, from 3 per cent to over 20 per cent. The profiles are defined by mathematical formulae and distort with thickness to achieve a shape having optimum performance characteristics at the flow conditions expected for that thickness. A subset of the family is defined for higher Mach number penetration, up to near sonic values. Practical applications of the new sections have shown notable gains.

The numerical calculation of the subcritical, steady, three-dimensional, potential flow past a semi-infinite swept elliptic cylinder attached to a plane wall is described. The full equations of motion are written in terms of the velocity potential and the exact body-surface conditions are applied. The equations are expressed in coordinates defined by the body shape and further transformations are necessary to reduce the infinite domain to a finite working space and to concentrate the grid points in regions of greatest variation of potential. The resulting equations are solved using a finite-difference scheme. The Mach number distribution for a typical case is presented.