This paper develops a method to determine the lift-drag ratio of air cushion craft and specifically the maximum value of the lift-drag ratio, its associated design speed, and related performance and economic measures. The method works from well-established equations for drag and power for air cushion craft. Such values are required to determine the performance and economic efficiency of the craft in many modes of operation. The method covers both amphibious and non-amphibious craft of the types used in both military and commercial operation. A basis is developed for an optimisation procedure to design future craft for maximum efficiency.

Effective control of combat aircraft at high angles of attack requires the flight control system to operate to higher frequencies and gains. This increases the possibility of excitation of the structural modes by altering the aeroservoelastic stability of the complete system. This paper details the results of tests carried out on a rigid, 55° leading edge sweep delta wing, half model fitted with a half span elevon operated through a closed loop control system. The effects of steady elevon deflections up to 40° angle of attack are described. The unsteady pressures resulting from elevon oscillations at frequency parameters up to a value equivalent to the first wing bending mode of a typical combat aircraft have been measured. The effect of the elevon oscillation on the elevon itself and on the bending mode have been estimated by calculating the unsteady hinge and root bending moments about the wing root. The lower surface, unsteady elevon effects were found to be independent of incidence and frequency parameter. The amplitude of response of the weak vortex on the upper surface of the wing reduced rapidly with frequency parameter. Above an incidence of 10°, the vortex breakdown was above the wing and led to a further reduction in the vortex response. The vortex response lagged the elevon motion.

This paper is an endeavour to present, to the members of this Society and others, as complete information as possible on the subject of a much discussed anti-knock material and to try and dispel any misconceptions which some may still have regarding the use and behaviour of ethyl.

The development of the modern aviation engine has made rapid strides during the last two or three years. The very rapidity of this development, however, has brought an exceedingly important fact home to the engine manufacturers; this is, the influence of the fuel in governing engine performance and reliability.

The two parts of this paper summarise the turbomachinery research undertaken at Pyestock during the 50 years since the National Gas Turbine Establishment was formed in 1946. The theoretical and experimental activities are described, and their influence on UK military and civil aero engines is assessed. The way in which NGTE supported non-aeronautical gas turbines is also explained. Part 2 covers turbines.

Screen separations as little as 0.02 chamber diameters do not adversely affect suppression of mean-flow non-uniformity provided the overall pressure drop coefficient, K, does not exceed about 3.5. At larger K a velocity overshoot develops near the wall boundary layers, which decreases only slowly with wider separation. A simple approximate extension can be made to the theory of Taylor and Batchelor for any number of closely-spaced screens. This shows good agreement with measurements and with low-drag screens predicts perfect attenuation when the overall pressure-drop coefficient is 3.2 or 7.6. Some inferences regarding the effect of close screens on turbulence suppression are also drawn.

In many practical circumstances in aeronautical engineering and in related industries, flows arise where concentrated streamwise vorticity is embedded within a turbulent shear flow. The juxtaposition of these features makes such flows particularly challenging to compute using CFD methods. In particular, one may be forced to abandon strategies for modelling the turbulent stresses that prove adequate in two dimensional shear flows (i.e. in flows where the vorticity vector is orthogonal to the mean velocity). The paper reviews developments in modelling turbulent stresses from their transport equations (rather than by appeal to an eddy viscosity hypothesis) and demonstrates, by way of a range of examples, the capabilities of this approach for handling flows with concentrated streamwise vorticity.

Real gas effects are among the major complications of hypersonic flow and this is illustrated by examples of the temperatures and pressures attained when a flow is brought to rest. Once dissociation appears, these depend to a marked extent on the type of compression.

In designing for the alleviation of aerodynamic heating there are two distinct cases. The first is the uncontrolled re-entry of a freely falling body, for which it is best to have a shape with a high pressure drag. The second is sustained flight, for which a shape with low pressure drag could be better, radiation giving appreciable control of surface temperature.

Low pressure drag accords with design for aerodynamic efficiency and there is scope for research on three-dimensional lifting shapes. There is some discussion of this and also of the philosophy of securing maximum amounts of separated flow.

Finally, there is continual emphasis on the need for experimental research, since hypersonic flow fields may differ considerably from those that would be expected by extrapolation of conventional supersonic experience.

The first ten years following the end of the First World War saw a steady “run down” in the armed forces and during this period what still existed in the field of aero-engine development was inspired, mainly, by participation in long range pioneering flights, distance records and international speed events—such as the Schneider Trophy Contest.

Research and development were limited by the funds available and it was not until about 1935, when the Government decided on an “all out” policy to re-equip the Royal Air Force with equipment to match the growing aerial strength of Germany, that aero-engine development received the support it needed. It was somewhere in this period that marked the “renaissance” of the development of aero-engines.

This paper describes the design optimisation study used to aerodynamically optimise the fairings that cover the rear wheels of the Land Speed Record vehicle, BLOODHOUND SuperSonic Car (SSC). Initially, using a Design of Experiments approach, a series of Computational Fluid Dynamics simulations were performed on a set of parametric geometries, with the goal of identifying a fairing geometry that was aerodynamically optimised for the target speed of 1,000 mph. Several aerodynamic properties were considered when deciding what design objectives the fairings would be optimised to achieve; chief amongst these was the minimisation of aerodynamic drag. A parallel, finite-volume Navier–Stokes solver was used on unstructured meshes in order to simulate the complex aerodynamic behaviour of the flow around the vehicle’s rear wheel structure, which involved a rotating wheel, and shockwaves generated close to a supersonic rolling ground plane. It was found that the simple response surface fitting approach did not sufficiently capture the complexities of the optimisation objective function across the high-dimensional design space. As a result, a Nelder–Mead optimisation approach was implemented, coupled with Radial Basis Function design space interpolation to find the final optimised fairing design. This paper presents the results of the optimisation study as well as indicating the likely impact this optimisation will have on the ultimate top speed of this unique vehicle.

To describe steady, two-dimensional, constant-density flow in a laminar boundary layer on a curved surface, a single equation may be derived from the complete Navier-Stokes equations, with no approximations being necessary. The conditions under which similar solutions are attainable are discussed, and the validity of some previous calculations is upheld.

Results are presented from low-speed windtunnel tests on an isolated generic chined forebody with pneumatic methods of yaw control at high alpha. Measurements of forces and moments were made on the forebody at angles of attack in the range 0° ≤ ɑ ≤ 60°. Slot blowing through the chine edge at various angles and upper surface tangential blowing have been compared, and optimum configurations have been identified. The effects of forebody cross-sectional shape have also been investigated and have been found to be highly critical in determining blowing effectiveness. Limited flowfield visualisation is also presented in order to gain some understanding of the flow field characteristics.

It is well known that, under certain conditions, duralumin and other wrought aluminium alloys are liable to suffer intercrystalline attack, such corrosion being frequently accompanied by little visible corrosion product on the surface of the material. The necessity for careful temperature control during the solution heat treatment of duralumin and the beneficial influence of a rapid rate of cooling after such heat treatment on the corrosion-resistance of this material are now too widely appreciated to need emphasis here.

This paper investigates the impact of the wall-heating conditions on the heat transfer performance of a rotating channel with one side smooth and one side roughened by 45° inclined ribs. Previous experimental and numerical studies for single-ribbed wall-heated channels showed that rotation has a significant negative impact on heat transfer performance. In order to investigate this uncommon behaviour, RANS simulations were conducted under three different wall-heating conditions in the present study: ribbed wall heated, all walls heated and adiabatic conditions. Numerical results show that the presence of uneven wall-heating conditions has a negligible impact on the stationary case, but it has a large influence on rotational cases, in both, the heat transfer and the flow field. The underlying reason is that in rotating cases, uneven heating results in different buoyancy effects on the trailing and leading walls of the channel that alter the main flow velocity profile. As a consequence, also secondary flows and heat transfer performance are affected.

An investigation of design procedures in high technology engineering organisations is being undertaken by the Divi-sion of Industrial Systems Engineering at University College of Swansea. To ensure comparable levels of en-gineering sophistication the study has been confined mainly to the aircraft component industry.

Initial work in this study has revealed some interesting facts about the prediction of design and development costs. Detailed information about design and development expenditures, estimated and actual, has been collected from one company and less precise data indicating similar trends has been gathered from three other companies.

The determination of shear distributions for multi-cell sections can conveniently be considered as two separate problems.

1. (i) the shear flow arising from a pure torsional couple applied in the plane of the section, and

2. (ii) the shear flow arising from a resultant shear force acting on the section so as to produce no resultant twist per unit length of that section, i.e. through the shear centre of that section.

Both problems involve finding the value of one unknown shear flow for each cell, so that in multi-cell construction a set of simultaneous equations is obtained, which must be solved to obtain the shear pattern.

The solution of problem (i) by successive numerical approximations has been dealt with completely by Benscoter. The present paper gives a method for problem (ii).

The usefulness of the method becomes more apparent with increasing numbers of spar webs as may be envisaged for future wing sections of small thickness chord ratios. It is analogous in the procedure used to the Hardy Cross moment distribution method for continuous beams, and has similar advantages. An approximate solution of any required degree of accuracy is quickly obtainable at the design stage and the method is readily adaptable for use by a computer with no knowledge of the structural or mathematical theory involved.

The aerodynamic characteristics of two-dimensional transient aerofoil motions in low-speed flows in a wind tunnel with either closed wall or open (jet) walls, including the effect of a downstream closed wall diffuser, have been investigated. The mathematical formulation for the aerofoil and its unsteady wake is based on linear theory and is solved by a piecewise linear vorticity method; the wall boundaries are represented by distributions of sources. Numerical calculations have been made for various values of tunnel height to chord ratio.

Interference effects on the rate of build up of lift to a steady state following a step change in incidence can be large, especially for open jet tunnels.

The commercial Computational Fluid Dynamics (CFD) software STAR-CCM+ was used to simulate the flow and breakup characteristics of a Liquid Jet Injected into the gaseous Crossflow (LJIC) under real engine operating conditions. The reasonable calculation domain geometry and flow boundary conditions were obtained based on a civil aviation engine performance model similar to the Leap-1B engine which was developed using the GasTurb software and the preliminary design results of its low-emission combustor. The Volume of Fluid (VOF) model was applied to simulate the breakup feature of the near field of LJIC. The numerical method was validated and calibrated through comparison with the public test data at atmospheric conditions. The results showed that the numerical method can capture most of the jet breakup structure and predict the jet trajectory with an error not exceeding ±5%. The verified numerical method was applied to simulate the breakup of LJIC at the real engine operating condition. The breakup mode of LJIC was shown to be surface shear breakup at elevated condition. The trajectory of the liquid jet showed good agreement with Ragucci’s empirical correlation.