The mathematical model of aerodynamic forces and moments is described, based on results from various wind-tunnel experiments with the RAE High Incidence Research Model (HIRM). Simulations of the responses due to longitudinal and lateral control inputs at high angles of attack are compared with the responses measured on free-flight models of the configuration. It is shown that the main features of the flight behaviour are reproduced, in particular such phenomena as roll-off, wing rock and nose slice.

On the afternoon of 14th April 1987, the Society held a half-day symposium on wing-tip flows and devices. Papers were presented by J. J. Spillman of Cranfield Institute of Technology on wing-tip sails, by A. C. Willmer of British Aerospace (in association with R. V. Barrett and J. D. Coleman of Bristol University) on the tip flow of part-span flaps, and by H. P. Horton of Queen Mary College on measurements in the viscous flow regions of streamlined tips. Written versions of two of these talks and a synopsis of the third* have been prepared and are being printed in the Journal following these introductory remarks. The contribution by J. S. Smith of RAE to the discussion session is appearing elsewhere.

The aerodynamics of a circular jet emerging normally from a wall into a cross flow is critically assessed. Some general trends are obtained when viscous effects are neglected. But viscous effects dominate the real flow behaviour. Attempts to predict these real effects are reviewed and their anomalies highlighted. The importance of understanding and specifying the jet exit conditions is stressed, which is relevant to complete numerical solutions.

A method is presented for determining the potential flow around bodies of revolution, at incidence, in uniform, incompressible, flow. This method utilises polynomial distributions of both sources and doublets, in the manner of Fuhrmann, and gives considerable advantage over existing techniques. In particular, when compared with the method of Nielsen, in a typical case, the computer storage requirement is reduced by a factor of eleven. Compared with panel methods the saving is very much greater.

Unlike the method of Landweber, both body surface and flow field conditions are obtainable and it would appear that the new method is ideally suited to the ‘store trajectory’ application in place of the existing discrete source method.

High speeds on water are being attained in racing sport and in attempts on world speed records in various classes. Success, safety and stability of these craft depends upon the favourable interaction of their aerodynamic and hydrodynamic characteristics under the influence of two media, one about 800 times denser than the other. Speed on the straight course and in turns is important.

As its velocity increases, a craft experiences increasing dynamic pressure in water and to maintain the balance, the ‘wetted’ area of the craft reduces as it rises up (‘planing’). Modern fast craft have ‘tunnel-hulls’ and lifting areas to generate aerodynamic lift and to assist the craft to attain planing attitudes rapidly.

The ‘lifting’ areas may not necessarily be in the correct locations however. An example often seen is that of a power boat riding virtually on its propeller and ‘wallowing’ in an unstable manner. In this case the variation in riding height alters the relative positions of centres of aero- and hydro-lift such that the ‘stability-margins’ are near critical in both the longitudinal and the lateral sense.

I must start by thanking the Royal Aeronautical Society for the invitation to present this 27th Lanchester Memorial Lecture. It is an honour and a privilege to follow in the footsteps of the distinguished scientists and engineers who have given the first 26 lectures in this series. These lectures have included many outstanding reviews of a wide range of different topics and I am very conscious that they have set a standard that I, for my part, will find difficult to match. I hope, however, that by choosing a topic that has only been mentioned in passing in just a few of the previous lectures, I will be able to make my own distinct, individual contribution to this tribute to the memory of a man who was not only a great scientist and engineer and talented musician but who, by his writings as long ago as 1907, still carries a message for us today in 1987.

Sadly, recent events continue to keep this subject to the fore — and memories of Dawsons Field, Entebbe, Mogadishu, Stansted, Beirut and Malta, to mention but a few of the most dramatic hijack incidents, are recalled, as are other events such as those of the early morning of 23rd June 1985 when one Boeing 747 of Air India was blown up in mid-air off the West Coast of Ireland, killing all 329 on board, with chance preventing the simultaneous destruction of another at Tokyo.

The seriousness of such incidents has, in recent months, grown from the mere appalling and tragic but occasional incident to one of major world importance having substantial effect on world peace and economic survival. The fact that this is no exaggeration is shown by the bombing raid carried out by the USA on Libya in April 1986, in retaliation for that country’s alleged involvement in or support of international terrorism. The subsequent reported mass cancellation of travel to Europe by Americans with consequential serious commercial implications not only for the airlines operating on the North Atlantic, where, for instance, British Airways reported a 25% drop in passenger traffic, is but one example of the effect of terrorism on the stability of world trade.

This informal history of the ARC outlines the background to its establishment in 1909, the subsequent growth and widening of its research activities for the advancement of aircraft design and construction, particularly around the two great wars, and finally its gradual decline until brought to a close in 1980. The story is told from a personal viewpoint and touches on the many personalities involved.

Quite by chance after this lecture had been written I saw at first hand part of the fire-fighting operation in the Tanneron in the South of France using flying-boats. This provided an excellent point from which to begin.

On 23rd August 1986 a vast fire was started deliberately, it was alleged, in the valley of the River Siagne, between Grasse and Mandelieu. There was a strong mistral blowing at the time. It was brought under control two days later with the aid of a small force of around nine ‘Pelicans’: Canadair CL-215 flying-boats (Plate 1). Each aeroplane can scoop up over 5000 litres (more than five tons) of water, in 10 seconds while planing at 70 knots (Plate 2). At times three aircraft would land in loose formation to do so. In the end the fire, said to be the worst since 1706, was reported to have destroyed 10 000 hectares (say 25 000 acres) of farm, forest and homes. Four people died, 700 suffered from burns and asphyxia, 8000 were left homeless.

The aim of axiomatic aerodynamic modelling is introduced, namely to understand the validity of aerodynamic modelling in the context of aircraft dynamics with particular emphasis on understanding the relationship between full-scale flight behaviour as compared with that predicted from data based on wind tunnel tests.

An axiomatic aerodynamic model is formulated for the attached flow past a wing and tailplane, based on a vortex lattice representation, with the wing wake fully relaxed at each instant of time. Conventional derivatives are estimated based on this model.

Predicted responses, using the above derivatives, are compared with ‘exact’ responses, using the axiomatic model for various elevator inputs and for both non-swept and swept aircraft configurations.

Surprisingly, even in these simpler cases, differences appear between prediction and ‘exact’ responses.

Studies begun over ten years ago have shown that the swirling air flow about the tips of a lifting wing can be redirected using small aerofoil surfaces, called sails, to reduce the overall lift-dependent drag of the wing. Tests on Paris, Jetstream and other windtunnel models, confirmed by flight tests on Paris and Centurion aircraft have shown marked reductions in drag and fuel consumption. Handling tests show improvements in roll control and stalling speeds.

Wing tip sails have the effect of reducing the peak adverse effects on other aircraft flying through the wake of a Paris aircraft but reduce the decay rate so the far field effects are worsened. Sails fitted to a crop-spraying aircraft have significantly reduced the amount of spray lifted above the height of the aircraft by the tip vortices and subsequently blown off-target by a cross-wind.

Windtunnel tests on a model with variable incidence sails show that differential sail incidence can be used effectively to provide roll power, reducing the size of the ailerons required, so allowing a greater flap span. The maximum root bending moments of the wings can be reduced by at least 10% by using negative sail movements in positive manoeuvres or gusts.

Sails can reduce the drag of an aircraft at all speeds leading to fuel savings estimated at US$225 per annum per 1000lb take-off weight per 1000 hours annual utilisation. For an airliner of 300 000lb weight and operating for 3000 hours this means a fuel saving of over a fifth of a million dollars per year!

The longitudinal stability and control of a large receiver aircraft was considered during air-to-air refuelling. A simple horseshoe vortex was used to model the tanker wake and approximate expressions were derived for the additional aerodynamic derivatives due to the position and attitude of the receiver aircraft within the downwash field. These derivatives were shown to depend on the mean variation of downwash with vertical displacement at the receiver wing and tailplane. The mean downwash gradients, in turn, depend mainly on the vertical separation between the tanker and receiver aircraft and the ratio of the tanker-to-receiver aircraft wing spans. Solutions of the linearised equations of motion were obtained for a range of values of the downwash gradients. The large receiver aircraft, considered in the paper, typically exhibits two divergent modes which appear to be controlled in flight by frequent alternate movement of the elevators and engine throttle.

Simple first order surface singularity methods based on:

• (i) Smith and Hess uniform source panels plus a uniform vorticity around aerofoil profile,

• (ii) piecewise linear vorticity around aerofoil profile, with different assumption for Kutta condition,

have been applied to attached flows and separated flows past an aerofoil/spoiler configuration, assuming an inviscid model.

For separated flows, piecewise linear vorticity methods give reasonable results as long as small panel elements are taken in the region of the separations at the spoiler tip and aerofoil trailing edge. The Smith and Hess method gives results which do not agree too closely with the vorticity methods. There is doubt concerning uniqueness.

Results have been compared using two different wake models; in one, the total head inside the wake is taken to be uniform, in the second, the static pressures along the separation streamlines are taken to be uniform. There appears to be a difference of about 5% in CL. It is not known why.

The paper describes the air-injection method of fixing transition in wind-tunnel tests on a two-dimensional aerofoil and a swept panel at high subsonic free-stream speeds. The method is shown to provide a repeatable level of disturbance which, for a given flow condition, may be generated without interrupting the test and which appears not to cause a significant excess drag. Examples are given illustrating the effectiveness of the technique in demonstrating scale-sensitive features of flows, and a correlation is presented which allows the minimum value of air-injection mass flow needed to fix transition to be estimated from a knowledge of the behaviour of natural laminar flow on the wing.

Measurements are described of the drags of various forms of three-dimensional excrescence mounted on balances installed in the walls of the working section of the RAE 8 ft x 8 ft wind tunnel. The tests cover a range of Mach numbers between 0·2 and 2·8 and a range of Reynolds number.

The excrescences tested include circular cylinders and wings mounted normal to the surface, fairing shapes and mushroom shaped rivet heads. For excrescences which were small compared with the boundary-layer thickness the scale effects were correlated in terms of the wall variables of the turbulent boundary layer. For circular cylinders and wings which protruded well into or beyond the boundary layer the drag may be determined by using a drag defect relationship. The fairings tested comprised half-bodies of revolution with pointed or rounded ends and bodies of rectangular, including square, section with pointed ends.

Three low-order panel methods developed for the analysis of supersonic flows have been used to make calculations for a number of test configurations, with the aim of establishing the range of problems for which this class of computational method yields adequate solutions. The programs investigated were the Woodward USSAERO/C program, the related NLRAERO program, and the British Aerospace Warton supersonic panel program. Results obtained using these programs have been evaluated against theoretical and experimental data for a number of test cases covering wing-alone, body-alone and wing-body geometries. It is concluded that low-order panel methods can provide adequate solutions for supersonic flows about wings, bodies and wing-body combinations, provided the assumptions implicit in the linearised potential flow model are not violated. Examples show that the prediction of lift and pitching moment curve slopes for quite complex configurations may be acceptable, but that the detailed pressure distributions are not always predicted satisfactorily. In particular, serious problems are encountered in calculating the flow about wings with rounded supersonic leading edges due to the linearised flow model which is used.

In Part II*, an inviscid model using a panel method has been formulated to solve separated flow past a steady aerofoil/spoiler configuration in a low speed stream. In this note, an investigation has been carried out to formulate an incremental model to solve the steady incremental problem when the aerofoil/spoiler changes its geometry by a small mount (eg for a small angular change of the spoiler).

It has been found that to satisfy the incremental boundary conditions on the initial aerofoil/spoiler geometry, incremental changes in the influence coefficients have to be included for the small change in aerofoil/spoiler geometry and the change in separation streamline geometry. A first order procedure is satisfactory; a second order procedure is also presented.

Minitufts are increasingly used for flow visualisation on static wind tunnel models because they can be applied in larger numbers for increased detail and with less interference compared with conventional tufts. This Memorandum describes an extension of their use to continuously rotating models where the heavier conventional tufts are inadequate.

Tests on two combat aircraft models in the 4 m x 2.7 m Low Speed Wind Tunnel are described. Measured variations of rolling moment are explained with the aid of the minituft photographs.

Experiments were conducted to study shock-induced separated flows on the lee surface of delta wings with sharp leading edge at supersonic speeds. Two sets of delta wings of different thickness (10° and 25° normal angle), each with leading edge sweep angles varying from 45° to 70°, were tested. The measurements, carried out in a Mach number range from 1.4 to 3.0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wings only). Using the test results, some features of shock-induced separated flows, including in particular the boundary between this type of flow and fully attached flow, have been determined. The experimental results indicate that this boundary does not seem to show any significant dependence on wing thickness within the limit of thicknesses tested. It is shown that this boundary can be predicted for thin delta wings using a well known criterion for incipient separation in a glancing shock wave boundary layer interaction, namely that a pressure rise of 1.5 is required across the shock. Comparison of the predicted boundary with experimental results (from oil flow visualisations) shows good agreement.