The aerodynamic performance of supercritical aerofoils at transonic speeds is strongly influenced by the shock wave-boundary-layer interaction. Passive shock control is one of the techniques used for controlling the undesirable effects of strong shock wave-boundary-layer interaction leading to extensive separation. Using passive shock control, the stall margin is increased and the onset of buffeting is delayed. Passive shock control is modelled by introducing a closed plenum chamber underneath a perforated surface at the foot of the shock wave where a combination of blowing and suction is generated. As a result, a strong normal shock wave is changed into a series of weak shock waves with lambda shape. Since the Baldwin-Lomax turbulence model has been used extensively for passive shock control modelling, which exhibits poor predictability for separated flows, an attempt has been made to modify the Reynolds normal stresses for this model so as to improve the accuracy of numerical results for flows with separation. Further modification to the Baldwin-Lomax model has been employed so that the mass transpiration effects are taken into account for the passive shock control computations. In this paper, a brief description on an implicit finite-volume TVD scheme in general coordinates is given and the details of the Balwin-Lomax turbulence model and its modifications are presented. The validated numerical results for several RAE 2822 aerofoil problems plus corresponding results for the modelled methods are presented and compared with some experimental data and other numerical results.

This paper describes the development of particle image velocimetry (PIV) to measure instantaneous spatial distributions of velocity in flows generated around models in a low speed wind tunnel with a test section 3m wide by 1·5m high. The difficulties associated with using PIV in air at this scale, and how they can be overcome, are discussed. Stereoscopic PIV is used to correct errors due to parallax that are present in velocity components measured in the plane of a light sheet when there is an accompanying flow through the sheet. This technique is also used to measure all three components of velocity in planes across a flow. It is found that apparently good estimates of a mean flow field can be obtained by averaging as few as ten instantaneous measurements of the flow field. The flows studied are generated by a 1/8 scale model of an aircraft for which the wing sweep can be varied. Measurements of velocities associated with the main vortex structures generated over the wing are presented for a sweep angle of 70°. Measurements of velocities recorded in the wake of a large fuel tank mounted below the wing, for a wing sweep of 45°, are also shown.

The Paper describes a metal and composite recurrent cost model of a large civil aircraft wing structure for a multidisciplinary design, analysis and optimisation (MDO) environment. The work was part of a recent European MDO project (BE95-2056) which investigated methods for the integration of structures, aerodynamics, dynamics and manufacturing cost at the preliminary design stage.

The paper discusses the cost modelling approach, which is based on parametric and process cost model methods, and the integration of the cost models into an MDO process. Results for the cost models are shown.

A framework has been successfully developed which allows the incorporation of manufacturing cost models into an MDO environment. It allows a designer to evaluate cost changes with respect to specific design changes such as rib pitch, stringer pitch, wing area and wing sweep.

An aircraft may reasonably be said to be structurally safe if it is known that during the manoeuvres which it is likely to perform

• (i) The stress in no part of the essential structure approaches the failing stress of the material,

• (ii) It does not become unstable or uncontrollable on account of structural distortion.

I intend to discuss in this context the wing structure of the aeroplane.

For many years it was felt both in the design and airworthiness approval of the majority of aeroplane wing structures that the main responsibility for attempting to secure the former desideratum could rest upon the determination of the stresses induced in the wing structural material by load systems appropriate to a rigid aeroplane in certain conditions denned by specified combinations of flight attitude, acceleration, and factor of safety. These combinations — familiarly referred to as the C.P. forward, C.P. back, nose-diving and inverted flight cases—are designed to produce for a particular aeroplane a range of steady load systems which is more severe on the structure than the range of load systems corresponding to the manoeuvres contemplated for the aeroplane.

The optimisation results for composite and metallic versions of a regional aircraft wing are compared using the multidisciplinary optimisation (MDO) program CALFUNOPT. The program has been developed for the conceptual design stage and models the wing using just 11 beam elements. The wing has been optimised for three combinations of the following constraint cases: static strength; aeroelastic roll efficiency (represented by limiting the twist of the wing for an aileron loading) and aeroelastic divergence. As expected, comparison shows that the composite wing designs are significantly lighter than the metallic ones, due to the well-known tailoring of the composite material. However, the simple model reveals some insight that may be useful to the designer, and which could be lost within a more detailed finite element approach.

The upper-skin compression panels produced by the conceptual MDO program, for both versions of the wing, have then been optimised using the more detailed and accurate panel sizing tool VICONOPT, which takes buckling into account. Such optimisation increases the panel mass by 5-10% and also provides a suitable ratio of stiffener to skin area for use in the conceptual MDO model.

The twentieth century has seen man’s mastery of the air, and fundamental to that achievement has been the provision of power to sustain the process of flight. For the first forty years the piston engine reigned supreme, yet for even longer now the power to fly has been provided by the gas turbine. In the conception and development of this prime mover Britain can justly claim to have led the world.

Two servants of the Crown quite independently had the vision and that amalgam of creative and analytical powers to see what might lie beyond the piston engine — one a scientific civil servant working at the Royal Aircraft Establishment (RAE) and the other, remarkably, a young RAF officer already distinguishing himself as a pilot. But the birth of the jet engine proved to be a painful process. While early aircraft designers had embraced the piston engine, neither Government nor the aircraft or aero engine industry showed interest initially in the gas turbine. The change in attitude began when war approached in 1939 and thereafter accelerated in the effort to move from bench demonstrator status to a full-production engine capable of contributing to the outcome of the war. It was in this difficult and complex process of transition that Dr Harold Roxbee Cox — Lord Kings Norton — was to play such a vital role. To appreciate his achievements it is appropriate to remind ourselves of the progress — or lack of it — on the gas turbine prior to his becoming involved.

This paper presents a critical study related to the evaluation and breakdown of total wing drag based on solutions of the Euler equations for subsonic and transonic three-dimensional flows.

This study clearly identifies the false entropy production present in all discretised Euler solutions as the main source of the discrepancies between the drag predictions produced by body-surface pressure integrations and far-field methods. A good understanding of the entropy production mechanisms present in discretised threedimensional Euler solutions has lead to an original procedure for the total drag breakdown and a new method for the evaluation of the induced drag. This new technique is based on body-surface pressure integrations and Oswatitsch’s expression integrated over suitable integration surfaces. In contrast with the various methods for the drag breakdown and the evaluation of the induced drag available in the literature, the proposed technique uses an exact formulation and is simple to implement. Using detailed drag calculations on an elliptic wing (the Xt = 1·0 wing) and the ONERA M6 wing, it has been shown that this new technique is not strongly sensitive to grid refinement and to the level of false entropy production in the calculation domain.