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.

Windtunnel tests on a two-dimensional model of a three element high lift system in a takeoff mode indicate that airjet vortex generators can produce a significant increase in lift at a given angle of incidence and a substantial increase in CLmax. An associated reduction in profile drag is inferred from measurements of momentum defect on the upper surface.

The improvements are much greater than those typically achieved with vane vortex generators and cannot simply be associated with the suppression of boundary layer separation. Instead, the airjets and the vortices which they generate promote enhanced mixing and momentum transfer across the complex shear layers above the main wing, from the external flow right through to the surface. Detailed surveys reveal, for example, that the slat wake is dispersed — or absorbed into a region of greater shear — and that the growth of the main wing boundary layer is significantly reduced.

It is suggested that neither the relatively low Reynolds numbers . and Mach numbers of the tests, nor the effects of windtunnel interference, detract from the general relevance of the novel principle involved. Attention is drawn, however, to the need for further work to elucidate and quantify the flow mechanisms and to establish the balance between the benefits from the improved aerodynamics and the performance costs of installing and supplying the jets.

The Royal Aeronautical Society Aerodynamics Group organised a CEAS European Forum on High Lift and Separation Control which was held at the University of Bath on 29-31 March 1995. This paper gives an idea of the scope of the Forum and of the main conclusions that emerged and it is followed in this special issue of The Aeronautical Journalby four papers that were invited for the Forum.

Measurements of wing buffeting, using root strain gauges, were made in the Nasa Langley 0-3 m cryogenic windtunnel to refine techniques which will be used in larger cryogenic facilities such as the United States National Transonic Facility (NTF) and the European Transonic Windtunnel (ETW). The questions addressed included the relative importance variations in frequency parameter and Reynolds number, the choice of model material (considering both stiffness and damping) and the effects of static aeroelastic distortion.

The main series of tests was made on three half models of slender 65° delta wings with a sharp leading edge. The three delta wings had the same planform but widely differing bending stiffnesses and frequencies (obtained by varying both the material and the thickness of the wings). It was known that the steady flow on this configuration would be insensitive to variations in Reynolds number. On this wing at vortex breakdown the spectrum of the unsteady excitation is unusual, having a sharp peak at particular frequency parameter.

Additional tests were made on one unswept half-wing of aspect ratio 1·5 with an NPL 9510 aerofoil section, known to be sensitive to variations in Reynolds number at transonic speeds. The test Mach numbers were M = 0·21 and 0·35 for the delta wings and to M = 0·30 for the unswept wing. On this wing the unsteady excitation spectrum is fairly flat (as on most wings). Hence correct representation of the frequency parameter is not particularly important.

The program VICONOPT is used to find the optimum (least mass) dimensions of a range of stiffened wing panels which are subject to buckling and material strength constraints and are loaded in axial compression with a sinusoidal manufacturing imperfection. Design plots are presented to show the effects that various rib spacings and stiffener types have on optimum design mass. A simplified model of a complete wing box is used to illustrate the design of a full wing panel and plots of optimum values of design variables at various stations along the wing have been obtained. The results were chosen to illustrate the practicality of optimisation with reference to manufacture of a full wing panel and to show the effect of changing the sophistication of modelling and theory used for the range of panels considered. The important aspects of the choice of design variables and design concepts are highlighted and percentage savings in mass, compared with an optimum metal panel design, are given for the various (global) optima found along with some examples of (rejected) local optima.

An experimental investigation into the wall jets created by a single inclined, twin inclined and a four inclined jet impingement arrangement is described. Wall jet velocity profiles were recorded at various azimuthal and radial locations for a range of nozzle heights at jet pressures representative of current V/Stol technology. Use was made of wall jet properties established for a normal jet impingement to provide a means of comparing the attributes of the various configurations. Non-splayed twin inclined impinging jets generate symmetry plane reinforcement and weakening in regions remote from the interaction. Splayed jet geometries show significant height effects on ground plane flow development. The four-jet configuration also demonstrates velocity reinforcement at the lines of wall jet interaction, together with a weakening of the flows remote from the interaction planes. Reinforcement velocities for the four-jet case are significantly lower than those expected from the constituent twin-jet flows.

The performance of a helicopter rotor has always been dogged by the dissymmetry of air velocity over the blades caused by its translation in a direction substantially parallel to the plane of the rotor disc. Furthermore, in order to make the rotor even usable, this dissymmetry of lift has caused the introduction of major mechanical complications to the rotor hub. In forward flight the rotor is aerodynamically split into two halves either side of the flight direction. The advancing side, where the rotor rotation is in the same sense as the forward velocity, will see a greater dynamic head than the retreating side, where the rotor rotation and forward speed are in opposition. A rigid rotor would consequently suffer a major rolling moment and if a viable proposition is to be achieved the rolling moment must be avoided.

A computational study is presented, which examines the performance of variants of second-moment closure and non-linear eddy-viscosity models when used to predict attached and separated flows over a high lift aerofoil for a range of incidence angles. The capabilities of both model types, especially in respect of resolving the onset of suction-side separation at high incidence, are contrasted with those of a low-Re k-ε model based on the linear Boussinesq stress-strain relationship. The second-moment model contains a conventional linear approximation of the pressure straining process; a cubic (realisable) variant has been investigated in an earlier study and found to offer no advantages. The quadratic eddy-viscosity model features coefficients which are sensitised to the strain and vorticity invariants. While both models, in the form originally proposed, are superior to the linear eddy-viscosity variant, neither performs well in respect of resolving separation, unless modified so as to return the requisite low level of shear stress in the boundary layer approaching separation. Once separation is resolved with sufficient realism, the near wake aft of the trailing edge is also well represented. All models return poor representations of the far wake which is characterised by low levels of turbulence production to dissipation ratio.

Recent advances in aeronautical technology are reviewed, with particular reference to work in the United Kingdom, and the process is outlined whereby these have progressed from fundamental research through to exploitation in marketable products. Aerodynamic, materials, propulsion and avionic aspects are discussed and successful applications described in commercial and combat aircraft and helicopters. Current and anticipated advances in technology are considered in the context of possible future applications, including brief consideration of the possibility of a new generation of supersonic transport aircraft. Lessons are drawn from past experience on actions necessary now if these advances are to be exploited successfully in improving the performance, or reducing the acquisition or operating cost, of the next generation of aircraft, without introducing an unacceptable level of risk.

This paper describes an experimental investigation into the effect of miniature vortex generators (VGs) on the longitudinal aerodynamic characteristics of a highly-swept wing with drooped leading edges. The experiments were performed in a low-speed windtunnel on a wing of the same delta planform as that used in a previous study. The latter wing was of fixed camber, while the subject of the present study was a wing with three different leading edge droop angles. The maximum reduction in drag due to the VGs was found to be about half that for the fixed-camber wing. This is reconciled with the different behaviour of the flow on the upper surface for the two types of wing. Without control, the drag of the variable-droop wings is much lower than that of the fixed-camber wing. As a result, with control, the variable droop wings have lower drag than that of the fixed-camber wing. Compared to that of the variable-droop wings without control, the fixed-camber wing, with control, has lower lift-dependent drag for lift coefficients between 0·4 and 0·7. However, at higher lift coefficients, the reverse applies.

The design of control systems for helicopters in hover and at low speed is a basic requirement for the extension of mission profiles and new mission demands. A special task for various applications is the position hold under wind and gust conditions above a ground fixed or moving target, like a shipboard reference, or a small vessel or lifeboat in rescue missions. For the solution of this problem a controller concept was developed and the feasibility was proven and successfully demonstrated in flight tests.

The in-flight helicopter simulator ATTHeS of the DLR has been equipped by the Institute of Flight Mechanics with an innovative measurement system for the hover position above a target. A video camera in combination with a highly parallel computer system for processing the optical information was used as an integra ted sensor system for the measurement of the relative position of the aircraft to a target. Based on the existing well-proven flight control laws of ATTHeS for the forward flight condition, which are implemented for handling qualities investigations, these control laws were modified and adapted to fulfil the special requirements of the position hold task, including altitude hold and heading hold capabilities. The integrated system of optical position sensor and control computer enables the helicopter to hover automatically above a defined target in constant altitude and with constant heading. Flight tests above a moving car under wind and gust conditions underline the future potential of the overall system to be used under operational conditions.

A rigid, 55° sweep, half delta wing has been oscillated in pitch at subsonic speeds, and the unsteady pressures on both the upper and lower surfaces recorded for pre-stalled conditions. The oscillations were of low amplitude and at frequencies equivalent to a typical wing first bending mode for full scale applications.

When compared to a quasi-steady approximation, the unsteady pressures on the upper surface of the wing lag the steady case along the line of the primary attachment. The lag represents a constant convective time from the leading-edge with increasing frequency of oscillation. A further localised area of lagging flow is observed beneath the vortex burst point, the location of which is a function of mean angle of attack.

The magnitude of the unsteady pressure variations was seen to increase linearly with the amplitude of the pitching oscillation while the phase lag was seen to increase linearly with frequency parameter.

An analysis technique capable of simulating the effect of engine failure during takeoff from offshore platforms is presented. Use is made of inverse simulation whereby the control actions required for a subject helicopter to follow a particular trajectory can be established. A mathematical representation of the towering takeoff procedure, and details of the modified inverse simulation technique needed to cope with the modelling of an engine failure are described. Detailed piloting accounts of the strategy used to fly the towering takeoff (with and without engine failures) are also given and are used to give qualitative validation of the analytical approach. Simulation results for a single engine failure of a transport helicopter during critical phases of a towering takeoff are presented. Finally, some directions for future work and potential applications of the technique are discussed.

The process of designing the Boeing 777 high lift system is described. Particular attention is given to design constraints and examples are given of the trade studies completed. The theoretical design tools and experimental test program are described. A survey of the high lift elements is presented and flight test results are compared with predictions.

Two numerical methods are presented for the computation of steady and unsteady Euler flows. These are applied to steady and unsteady flows about the NACA 0012 aerofoil, using structured grids generated by the transfinite interpolation technique. An explicit central-difference scheme is produced based on the cell-vertex method of Ni modified by Hall. The method is second-order accurate in time and space, and with flow quantities stored at boundaries the boundary conditions are simple to apply. This is a definite advantage over the cell-centred approach of Jameson, where extrapolation of the flow quantities is required at the boundaries, making unsteady boundary conditions difficult to apply. An explicit upwind-biased scheme is also produced, based on the flux-vector splitting of van Leer. The method adopts a three stage Runge-Kutta time-stepping scheme and a high-order spatial discretisation which is formally third-order accurate for one-dimensional calculations. The upwind scheme is shown to be slightly more accurate than the central-difference scheme for steady aerofoil flows, but it is not clear which is the more accurate for unsteady aerofoil flows. However, the central-difference scheme requires less than half the CPU time of the upwind-difference scheme, and hence is attractive, especially when considering three-dimensional flows. The transfinite interpolation technique is ideal for generating moving structured grids due to its simplicity, and grid speeds are available algebraically by the same interpolation as grid points. The method is also ideal for use in a multi-block approach.

Although mini-tufts are considered generally to show the surface flow direction on windswept surfaces, there is some uncertainty about their correct interpretation, particularly as three-dimensional flows approach separation. This note provides some guidance on this controversial question, based on experience with mini-tufts on wings, fins and canards in the DRA 13 ft x 9 ft low speed wind-tunnel at Bedford.

An illustration of the value of the technique is given. Mini-tuft photographs are used to identify the trailing edge condition which relates to the classification of the steady and fluctuation pressure distributions observed a swept wing with a conical, separated flow.

The optimal takeoff performance of a vectored thrust V/Stol aircraft is analysed by a nonlinear programming method based on a sequential unconstrained minimisation technique. The study is focused on ship operations which involve VTO and ski-jump STO launches. Once the optimal ramp geometry is determined, minimum-fuel and minimum-time problems are solved. The optimal control sequences are evaluated and the effects of ship motion and atmospheric winds are discussed. A microburst encounter following an STO launch is also dealt with in terms of aircraft survival.

A model of the viscous flow past a flexible sail section operating near the ideal incidence is described. Viscous effects are calculated via weak viscous-inviscid interaction of a panel method and an integral boundary layer method, and a new model for the leading edge separation bubble is introduced. The flexible section is allowed to deform in response to the pressure and shear stresses acting on it. Results are presented for the effect of incidence, excess length and Reynolds number on the development of the boundary layers on each side of the section and the consequences for the lift and drag of the section are described. The numerical model is compared with experimental results, giving in general good agreement and shedding light on the physics of the viscous flow past flexible membranes.

The aerodynamic interaction between a wing and a couple of propellers in tractor configuration is investigated by means of a model based on the lifting line concept and under the assumption of quasi-steady incompressible motion of inviscid fluid. The ways the propellers influence the wing performances, particularly the induced drag, are analysed. It turns out that the lift increase and the possible drag reduction depend strongly on the direction of the blade rotation and on the propeller distances from midspan, and result from two main effects: the direct propeller induction on the wing and the modification of the lift distribution along the span. An additional nonlinear (mixed) contribution affects the drag but has only negligible influence on the lift. The described model allows one to evaluate separately all these effects and helps understanding their influence on the global modification of the wing performances.