An investigation has been made into the effects of a surrounding stream on an “under-expanded” sonic jet. Experiments were performed for different Mach numbers of the surrounding stream and for different ratios of the static pressure in the jet at the exit to that in the external stream. It was found that the base-flow region produced at the exit of the nozzle, due to the finite thickness of the nozzle walls, had a significant effect on the flow field. Schlieren photographs suggest that the surrounding stream has little effect on the first cell of the jet but that the second cell is, in general, lengthened when a surrounding stream is present. In most cases it is longer than the first cell. Conditions at the beginning of the second cell are probably different from those at the beginning of the first cell and it is suggested that the lengthening of the second cell is mainly a secondary effect due to the alteration in the velocity distribution of the first cell caused by the external stream. A theoretical investigation was made, using an inviscid model of the flow. The flows in the jet and the stream were calculated independently and the jet boundary determined using an iterative procedure. For the free jet, a study was made of the various approximations to the sonic exit conditions and into the errors involved in the numerical solution of the supersonic region. A comparison of the calculated length of the first cell of the jet with the experimental results showed reasonable agreement for the free jet. For a jet with a surrounding stream the inviscid model predicted that the length would be significantly increased as the Mach number of the stream was increased from zero; this effect was not found in the experiments. A modification to the model is suggested and it has been found that, using this, the results are in better agreement with the experiments.

An approximate analysis of conditions at separation produced by turbulent boundary-layer/shock-wave interaction is presented for swept, cylindrically symmetric flows. An integral boundary-layer prediction method is used, incorporating Johnston crossflow profiles. The results indicate a marked reduction in pressure rise required to produce separation as sweep is increased. At low Reynolds numbers the skin friction at separation is inferred to be small, whereas at higher Reynolds numbers the presence of a vigorous streamwise flow may be detected. In the limiting case of zero sweep, or two-dimensional flow, predictions using the approximate analysis are shown to compare well with experimental results of pressure rise to separation.

The powerplant has always played a vital role in the development of aviation. In the beginning, the crucial need was for an engine of high enough power-to-weight ratio to enable the aircraft to fly at all. The Wright brothers had to build their own, and they took to the air with a four cylinder petrol engine having a specific weight of around 15 lb/hp. Much progress was made over the next two decades, and by the mid-nineteen twenties this figure was down to about 2 lb/hp. The aeroplane was now emerging from its infancy; the Atlantic had been flown, and the Otto cycle engine had established itself as successfully in the air as on the road.

It is considered that the blended wing body may offer the possibility of improvements in performance efficiency over the conventional civil transport. Such configurations will require transonic wind tunnel testing. Hence, a feasibility study has been carried out at ARA to investigate possible model support systems that could be used for this purpose. The study addressed issues arising from the mechanical design of two support systems; namely, a blade or Z-sting and a standard rear sting. In particular, the relatively restricted space available within the model fuselage has presented challenges in the design of the layout of the balance and instrumentation required for the model. CFD analysis has been carried out to assess the relative interference effects of the two support arrangements. This initial work has shown unexpectedly significant interference effects associated with the Z-sting support and attention is drawn to the need for further design work to minimise these effects.

A number of experimental studies have shown that postbuckling stiffened composite panels, loaded in uniaxial compression, may undergo secondary instabilities, characterised by an abrupt change in the buckled mode-shape of the skin between the supporting stiffeners. In this study high-speed digital speckle photogrammetry is used to gain further insight into an I-stiffened panel’s response during this transient phase. This energy-dissipating phenomenon will be shown to be able to cause catastrophic structural failure in vulnerable structures. It is therefore imperative that an accurate and reliable methodology is available to predict this phenomenon. The shortcomings of current non-linear implicit solution schemes, found in most commercially-available finite element codes, are discussed. A robust and efficient strategy, which utilises an automated quasi-static/pseudo-transient hybrid scheme, is presented in this paper and validated using a number of experimental tests. This approach is shown to be able to predict mode-jumping with good accuracy.

It is a curious reflection on the progress of science in our time, that a term which has only come into common use in aeronautics in the last decade commemorates the name of a man who carried out his researches some fifteen years before the Wright Brothers first flew in 1903. The obscurity surrounding Ernst Mach, whose name is given to the ratio of the speed of flow of a gas to the speed of sound under the same conditions, is ill-deserved, for there can be little doubt that he was one of the outstanding scientific personalities of the nineteeth century; this distinction arises not so much from his experimental work, which was extremely valuable, but more from his original conceptions on the philosophy of science which helped to lay the foundation for modern developments in physics.

The ideal of high speed centre-to-centre inter-city air transport remains unrealised with the problems of poor economics and excessive noise dominant and still largely unsolved.

Economics hit particularly hard at VTOL aircraft. Thrust/weight ratios in the order 1·15:1 are necessary with the result that total installed thrust is much larger than is required for the cruise condition and payload and range suffer. It is therefore tempting to compromise and go for a STOL solution and current trends are in this direction.

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.

Ghosh's large-deflection hypersonic similitude and piston theory for plane flows have been extended to axisymmetric flows: a new piston motion in conico-annular space as an extension to Sedov's axisymmetric piston has been proposed and a set of ordinary differential equation for 1-D unsteady motion obtained. Wedge and cone flow solutions are obtained without the constant-density assumption of the earlier theory; the pressure ratio is expressed as a closed form function of the similarity parameter whereas the previous constant density theory gave an iterative procedure. The present theory unifies small disturbance theory and constant density theory.

We should like to preface our essay on the subject of spinning by mentioning the circumstances under which our investigations were carried out and the sources of our information. The Panel of the Aeronautical Research Committee which has been appointed to deal with all questions connected with the stability and control of aeroplanes was requested in 1924 to consider the urgent problems connected with the alarming accidents due to certain machines failing to retover from a spin. After the issue of a preliminary reporton the situation by the Panel, the writers of this paper were asked to go into the whole question as far as existing information from full-scale and wind tunnel experiments would permit. We have had ready access to all available data, coming chiefly from Farnborough on the full-scale side, and from the N.P.L. on the model side.

The lightweight and compact separate jet lift engine has been with us long enough to have built up an acceptable level of confidence in its abilities, but there have been misgivings over the magnitude of some of the operational problems, in particular those of ground erosion and hot gas reingestion due to the high velocity jet.

The object of this paper is to present a Series of tests which has been made by Rolls-Royce to ascertain the nature of the problems and to put forward some simple solutions. It is not suggested that any of these solutions is unique but merely that they are adequate for our present needs.

Prediction of aeroengine exhaust plume near-field development requires knowledge of velocity and turbulence distributions at nozzle exit. The high Reynolds number nozzle inlet boundary layers of engineering practice are fully turbulent, but acceleration can induce re-laminarisation. Thus, to reproduce nozzle exit conditions accurately, large eddy simulation (LES) for plume prediction must be capable of capturing re-laminarisation and any subsequent boundary layer recovery. Validation is essential to establish a credible LES methodology, but previous studies have suffered from lack of nozzle inlet/exit measurements in the test cases selected. Validation data were here taken from an experiment on a convergent round nozzle with a parallel exit extension to allow boundary layer recovery. LES inlet condition generation applied a rescaling/recycling method (R2M), whose performance was validated against measurements of first and second moment statistics as well as the turbulence integral length scale. Simulations employed two sub-grid-scale (SGS) models; these produced similar predictions up to the end of the nozzle convergent section, but marked differences appeared for the nozzle exit turbulence field. The Smagorinsky model predicted much lower turbulence levels than measured, whereas the Piomelli and Geurts model revealed the presence of a small separation region at the convergence/parallel section corner, which led to higher exit turbulence and much better agreement with measured data.

Before beginning my contribution to this symposium, I must make it plain that I do not claim to speak here for employers of pilots, other than myself. For that matter, I do not even claim to be speaking for my own Company. The opinions I shall express are my own and they result from a modest experience in the management side of the air transport industry. They are not the product of any discussion or agreement with other employers of pilots.

Concerning the title of my contribution, I shall assume the word ‘employer’ to mean an air transport undertaking, operating under an Air Operator's Certificate, and the word ‘pilot’ to mean a qualified aircraft captain engaged in public transport operations.

This paper discusses the various aspects of flow separation on wings and aerofoils at subcritical Mach numbers. It is shown that in general the onset of non-linearities in the aerodynamic characteristics, for plane constant-section wings, can be related to the initiation of regions of flow separation somewhere on the wing and often near the wing tip.

Empirical correlations are produced, tor plane wings of constant symmetrical section, enabling estimates to be made of the lift coefficient for onset of separation at either the leading or trailing edge of the wing.

A correlation of stall types on aerofoils, due to Gault, is extended in application to the estimation of initial stall and separation types for plane wings of constant symmetrical section. Suggestions for extending the application of this correlation to the estimation of section stall type for cambered, twisted and non-constant section wings are made.

In this paper, a robust three-dimensional guidance law against manoeuvering targets is designed using the approach of discrete-time partial stabilisation. In the proposed method, the equations of the guidance problem are divided into two subsystems where the asymptotic stability is desired only for the first one. The control input of the second subsystem is designed such that the collision to be ensured in a short time. Despite recent advances in technology and implementation of digital controllers, the design of guidance laws with the approach of discrete-time partial stabilisation has not been done, till now. One of the advantages of this paper is to design a discrete-time guidance law even with the difficulties of the discrete-time Lyapunov theorem. Moreover, the Lyapunov function is chosen based on the physics of the guidance problem (making the rate of line of sight (LOS) rotation close to zero), and it is shown that it is not possible to asymptotically stabilise the system in the case of manoeuvering targets. Nevertheless, to guarantee the collision with the target, it is enough to limit the rotation rate of LOS to a small value. Finally, simulation results are given to show the appropriate performance of the proposed guidance law.

To the layman the parachute is often no more than a large pocket handkerchief which brings men safely to earth without a bump, but to the aeronautical engineer it is a complex aerodynamic decelerator, which has a large number of degrees of freedom and dynamic characteristics which are often difficult to account for by classical aerodynamic theory. Whichever view one may care to take of the parachute what is certain is that advances in parachute technology are nearly always the result of a balanced programme of both theoretical and experimental work, and for this reason equal attention will be given in this paper to both the experimental and theoretical aspects of parachute technology.

The short take-off capability is of paramount importance for a fighter airplane to enable its operation from short and damaged runways. This paper analyses the airplane take-off process from the viewpoint of reducing the ground roll/take-off distance with the use of thrust vectoring. The airplane take-off is modelled incorporating the ground reactions on the landing gear and the thrust vector forces and moments. The take-off problem is formulated as an optimal control problem with appropriate constraints. Though many researchers have applied optimal control techniques for designing airplane manoeuvres, its application to the airplane take-off problem is rarely available in the open literature. It is expedient to use such methodology to understand the use of thrust vectoring features of an aircraft to maximise the benefits in shortening the ground roll/take-off distance. An optimal control methodology has been applied in this paper with the objectives stated above to a twin-engine fighter nonlinear aircraft model popularly known as F-18/HARV. Computation of flight path and control schedules using optimal control has been carried out with and without the use of vector nozzles. A reduction of about 6% in take-off distance and about 29% in ground roll distance is obtained with the use of thrust vector for the configuration studied.

An experimental investigation was conducted to determine the flow characteristics of shallow swept cavities at a free-stream Mach number of 2. The investigation described herein focused on obtaining information on both time-averaged and time-dependent flow features. The data consisted of steady and unsteady pressure measurements as well as some surface oil and schlieren flow visualisation. The effective length-to-depth ratios l/d of cavities investigated ranged between 5·65 and 8 for shallow ‘open’ cavities and between 11·31 and 16 for ‘closed’ cavities. The cavity sweep angles were 0°, 15° and 45°. The results of the swept cavities, when compared to the datum cases of the straight (0°) cavities, showed some distinct differences. With regard to time-averaged flow characteristics, in an open swept cavity (5·65 < l/d < 8), the flow displayed quasi-open flow behaviour, distorted by the spanwise cross flow within the cavity. In the case of the closed swept cavity (11·31 < l/d < 16), flow features resembling the ‘closed’ to ‘transitional closed’ cavity flow types were seen to exist simultaneously across the span. Unsteady pressure data indicated that for an open cavity at 15° sweep angle, the discrete frequencies observed were similar to those of a straight open cavity. In contrast to this, at 45° sweep angle, the frequencies were broadband with no discrete frequency across the cavity length. For the closed cavity, the frequencies were all broadband irrespective of the sweep angle. The investigation also showed that the influence of the sweep angle on the pressure drag of the cavity strongly depends on the type of cavity flow: a sweep angle increase from 0° to 45° approximately doubled the pressure drag for an open cavity, while it led to a drag reduction of about 37% for the closed cavity.

As aviation is one of the fastest growing industrial sectors world wide, air-traffic emissions are projected to increase their stake in the contribution to global warming. According to studies, both carbon dioxide and contrails are the principal air-traffic pollutants, whereas the impact from contrails in terms of radiative forcing is possibly larger than that of all other air-traffic pollutants combined. New regulations with the objective of mitigating contrail occurrences might cause a change in the design requirements of aircraft. In light of this, a method considering contrail formation during the aircraft design process is presented in this paper. Aircraft performance and optimisation is carried out with NASA’s flight optimisation system. Combining historical meteorological data with air-traffic data enables an assessment regarding contrail formation. As an example, a particular aircraft type in terms of range, speed and payload is optimised for minimum block fuel consumption considering different altitudes. The change in contrail formation in terms of contrail-km formed is calculated. The results suggest that if aircraft of the considered class were designed for higher altitudes, contrail occurrences would diminish slightly at a non-negligible fuel burn penalty.

Dynamic stall is a complex process encountered by an aerofoil in the unsteady flow environment such as a helicopter rotor in forward flight as well as a fixed wing aircraft in manoeuvres and in other unsteady situations. The onset of dynamic stall effectively determines the flight envelope of the helicopter. Significant effort are being made to develop CFD to capture the dynamic stall behaviour, however traditional engineering models based on lifting line theory still offer fast turn-round and broad understanding required for the rotor design process. This paper describes a new engineering model for dynamic stall, developed originally for wind turbine application at a typical Mach number of 0·12. The new dynamic stall model, with a better definition of stall onset, is based on improvements made to Beddoes’ original trailing edge stall model. This paper will describe and demonstrate the improvements in identifying both the stall-onset and the pitching moment break at high pitch rates, when being applied to a generic rotor aerofoil RAE 9651 at M = 0·3. Further validation against oscillatory tests and other Mach numbers are still required. However the study has provided sufficient confidence for it to be employed in a rotor analysis code.