A turbulent hot jet is calculated and its entrainment effect is modelled by a line sink with the objective of estimating the downwash effect on rectangular lifting surfaces of various aspect ratios. The line sink density is determined by numerical solution of the free shear layer equations with a turbulent kinetic energy model. The ‘incompressible equivalent jet’ concept is verified within the context of practical engineering calculations.

This is a summary of work detailed in a more comprehensive paper which contains a major re-examination of, and challenge to, the assumptions of aerodynamics theory as applied to infinite flow fields at the speed of sound and their consequent relevance to the reality of free flight.

In steady-state flow fields generally, fore-bodies are embedded in the flow by a process based on the contraction of the streamline cross-sections. In subsonic flow this is achieved by velocity increase whereas in supersonic flow it is by velocity decrease. A primary reason for the challenge to accepted theory lies in the fact that, in a steady-state sonic field streamlines can only increase from their undisturbed area (Fig. 1) and this leads to fundamental difficulties in the construction of the forward flow field which in turn lead to the various theoretical anomalies.

The major achievements of the Royal Air Force during the last war were made possible very largely by three aeroplanes – the Hurricane, the Spitfire, and the Lancaster. They became very well known to the British public at that time, and indeed to the enemy. Less widely known, but nevertheless publicly acknowledged, were their chief designers – S. Camm, R. J. Mitchell and R. Chadwick.

Two other aeroplanes that found a place in the public eye were the Wellington bomber and the Sunderland Flying Boat. A. Gouge, the chief designer of the latter, was hardly noticed, but B. N. Wallis, who contributed most to the design of the Wellington, became a national figure when he moved on to the design of his great bombs.

This study examines the performance of two axisymmetric nozzles which were designed to produce uniform, parallel flow with nominal Mach numbers of 4 and 8. A free-piston-driven shock tube was used to supply the nozzle with high-temperature, high-pressure test gas. The inviscid design procedure treated the nozzle expansion in two stages. Close to the nozzle throat, the nozzle wall was specified as conical and the gas flow was treated as a quasi-one-dimensional chemically-reacting flow. At the end of the conical expansion, the gas was assumed to be calorically perfect and a contoured wall was designed (using Method-of-Characteristics) to convert the source flow into a uniform and parallel flow at the end of the nozzle. Performance was assessed by measuring Pitot pressures across the exit plane of the nozzles and, over the range of operating conditions examined, the nozzles produced satisfactory test flows. However, there were flow disturbances in the Mach 8 nozzle flow that persisted for significant times after flow initiation.

A standard flat facet approximation for finite element analysis of general thin shells has been sought after for many years, but fundamental difficulties hindered the development of a definitive solution. Recent advances in the finite element method, however, have provided more rational resolutions of some of these difficulties than the kind of artifices which were employed formerly. This paper completes the theory of a basic triangular flat shell element which incorporates modern developments and meets practical criteria of mathematical simplicity and general applicability. The flat facet shell element is also, potentially, an appropriate basis for constructing a better curved approximation to a shell surface in order to achieve improved accuracy. Numerical results are given for a variety of shell test problems to provide comparisons with accepted benchmark solutions and to assess the available accuracy relative to other finite element formulations.

A systematic parametric investigation of the flutter behaviour of a binary wing-with-engine nacelle system in inviscid incompressible flow is performed. The equations of motion in coupled bending and torsion are initially set up and the aerodynamic model for the calculation of the motion-induced unsteady airloads is described. The aeroelastic stability equations are then derived in non-dimensional form and the procedure used for the numerical solution of these equations is outlined. For a typical large commercial aircraft configuration, numerical results are presented for a variety of systematic parameter variations with special emphasis on the effects of the location of the elastic axis and the motion-induced unsteady airloads on the engine nacelle.

An AGARD calibration model has been tested in the subsonic cryogenic windtunnel (KKK) of DLR Köln, model support interferences are calculated in the form of corrections to the aerodynamic coefficients. Two different model support systems are investigated: the model mounted on a horizontal sting (rear sting), and the model mounted on a vertical sting (ventral sting). Since for these configurations earlier experiments already exist, the corrections obtained are applied to these measurements for further verification. Furthermore the sting influences on the longitudinal static stability of the model are investigated in more detail.

A 3D-panel method is used which allows simultaneous calculation of several displacing and lifting bodies of nearly arbitrary shapes in the flow. The aerodynamic coefficients are obtained by numerical integration of the surface pressure distribution on the model.

The imposition of swirl on co-axial jets exhausting into confined space, akin to a combustor model, has far reaching effects on their mixing and flow development. In the present paper, the influence of swirl on both the jets has been determined in terms of velocity and pressure distributions in the confined space. It has been shown that swirl in the central jet leads to faster mixing whereas higher swirl in the annular jet improves both mixing and development.

Experiments were made on a two-dimensional low speed diffuser with passive control. The passive control consisted of porous surfaces made of holes/slots connected by a breather passage. Passive control postpones the stall and reduces buffet levels in the diffuser. This test confirms that passive controls to prevent flow separation are not restricted to transonic speeds.

A new free-wake aerodynamic model of a rotor hovering in ground effect is briefly presented. The numerical characteristics of the model are investigated. The convergence between consecutive iterations and convergence with respect to different parameters that define the model, are presented and discussed. The model is then validated by comparing the numerical results with existing experimental data. Very good agreement of the thrust coefficient is obtained, whereas discrepancies in the torque coefficients are discussed. At the end of the paper the capabilities of the model are shown by considering various aspects of the ground effect.

The flying wing supersonic transport is simply a straight wing of high aspect ratio and of sufficient size to contain passengers comfortably. Such a wing can be steered to different sweep angles and thus adapted to efficient flight over a range of Mach numbers.

The writer demonstrated such a wing by flying models at the First International Congress in the Aeronautical Sciences. The potential advantages of such a flying wing were discussed by G. H. Lee of Handley Page Aircraft at a subsequent Congress. The present paper gives recent studies of flight stability, efficiency and operational problems. It is suggested that by limiting the Mach number to twice current jet speed fuel efficiency approaching that of present day subsonic transports could be achieved. This goal would probably also require the development of variable bypass engines.

An eddy viscosity relation for the inner region is combined with an empirical wake function for the outer region of a turbulent boundary layer. The calculated distributions for the mean velocity and Reynolds shear stress are compared with both direct numerical simulation and experimental data in a zero pressure gradient boundary layer, mainly at low Reynolds numbers. The comparison is favourable in both the inner and outer regions of the flow.

In recent years, interest has grown in the use of small satellites for a variety of applications. In certain missions, where propulsion systems are necessary, electric propulsion technology can offer advantages. Small satellite technology and applications are reviewed and some of the features of various types of electric propulsion are discussed: missions where small satellites might use electric propulsion are described.

The improved thin aerofoil method, which features extended expressions for lift and moment coefficients, is considered for further investigation and validation. The procedure to calculate the singularity coefficients is improved by using all aerofoil coordinates as control points in a least squares scheme. The classical NACA 0012 and NACA 65012 sections, the modern aviation aerofoil LS(1)—0417 and the extremely thick Kennedy-Marsden aerofoil are validated in place of the previously cited Karman-Trefftz aerofoil. This selection covers thickness ratios of up to 27·9%, camber ratios up to 7·69% and incidence up to 16·7°. Comparisons of velocity (or pressure) distributions and aerodynamic coefficients are made with two panel methods and with exact solution or experimental results whichever is available. Results indicated that the accuracy of the extended method is much better than expected and compares well with panel methods except for the extremely thick aerofoil. Additional results in the form of a systematic investigation of a weighted global error in the pressure distribution for the Karman-Trefftz aerofoils used in the previous study, are also included. Such an error shows similar trends and in many cases comparable magnitude to the errors generated by panel methods.

Analysing the action of atmospheric turbulence on aircraft for the determination of limit and fatigue loads is an essential stage in the structural design process. These calculations are governed by regulations, the evolution of which is of great importance to the aeronautical industry. The aim of this paper is to present the various calculation methods used for the preliminary dimensioning, substantiation and certification of aircraft, and to assess, on the basis of aircraft response studies and comparisons of loads calculated, the repercussions of the progress which has been made in recent years. Moreover, the regulatory requirements currently in force internationally, though often disparate and unbalanced, are now aimed at standardisation as a result of the impetus of both the airworthiness authorities and the aeronautical industry. Against this background, we propose a new calculation method developed on the basis of results obtained at Aerospatiale.