Measurements made at a Mach number of 0.18 and a chord-based Reynolds number of 4·2 x 106 on a constant-chord model having a NACA 4412 aerofoil section are described and compared with the results of flow field calculations.

Both the experimental arrangement and the difficulties initially experienced in achieving an adequate approximation to two-dimensional flow above the wing are briefly outlined.

The measurements include static pressure distributions on the wing surface and on the wind-tunnel walls above and below the mid-span section of the wing. The main emphasis in the experiment was, however, on defining the development of the upper surface boundary layer through separation (at about 20% chord ahead of the trailing-edge) and on into the wake, making extensive use of laser anemometry to measure mean velocities. In addition, Reynolds stresses were measured in certain parts of the flow field by hot-wire anemometry.

The flow field calculations are of the semi-inverse kind in which an inverse momentum-integral treatment of the shear flow, used to avoid difficulties at separation, is coupled to a direct solution of the inviscid flow problem. The main features of the method are outlined.

A numerical model is developed for predicting low-velocity impact damage in laminated composites. Stacked shell elements are employed to model laminate plies with discrete interface elements in pre-determined zones to model the onset and propagation of matrix cracks and delamination. These interface elements are governed by a bi-linear cohesive failure law. Cohesive element zone size is determined by a separate finite element analysis using solid elements to identify the stress concentration sites. In order to save the computational effort, low-velocity impact load is modelled by quasi-static loading. Influence of contact force induced friction on shear driven mode II delamination is modelled by a friction model. For a clustered cross-ply laminate, calculated impact force and damage area are in good agreement with the test results. It is shown that matrix cracks should be included in the model in order to simulate delamination in adjacent interface. The practical outcome of this research is a validated modelling approach that can be further improved for predicting low-velocity impact damage in other stacking sequences.

It is shown that the flow in a region of recirculation may be divided into a core, in which the vorticity may be constant, and a boundary layer surrounding the core. An approximate solution is given for the flow inside a circular cylinder with partly fixed and partly moving walls.

Flow patterns which include regions of recirculation occur frequently, for example, behind bluff bodies, in sharp bends, and in sudden expansions in ducts. Little is known about this type of flow, which will be referred to, for brevity, as ‘cavity’ flow. It used to be thought that the velocities in the ‘ cavity ’ were small compared with the stream velocity, but it is now known that the maximum velocity in the cavity may be about 30 per cent. of the stream velocity. This implies that the motion within the cavity cannot be neglected.

Key issues relating to the Supersonic Business Jet (SBJ) concept are reviewed with the intent to assess the readiness of enabling technologies and hence the concept itself. The multidisciplinary nature of aircraft design precludes an in-depth analysis of each specific aspect, which could individually be the subject of a separate discipline review, hence an overview is presented.

The review looks at the market, environmental issues, with particular reference to the sonic boom phenomenon & solutions, technological issues, including prediction methods, flight testing, systems, certification and interested aerospace companies and design organisations.

It is apparent that the need to reduce the sonic boom signature is vital if the vehicle is to be permitted to operate over land and hence be economically viable. It is clear that sonic boom acceptability requirements must be set if resources are to be effectively focused and designs are to converge. Despite this challenge, considerable investment is aimed at de-risking many of the enabling technologies and raising readiness levels. Many technologies are moving beyond theoretical and numerical analysis into the experimental and flight test domains. Collaboration between the civil and military sectors is increasing.

Clearly, supersonic air travel is not an efficient means of personal conveyance; however, concerns for the environment are difficult to balance against the ‘value of time’ benefits offered by the SBJ concept. Air travel, of which this is a specialised form, is important to the global economy. Continued effort in the areas of human factors, customer demand and certification & requirements would be beneficial.

Measurements were made of the pressure in a blunt-nosed pitot tube, in an air stream at Reynolds numbers from about 15 to 1,000. The results are expressed in terms of a pressure coefficient , where P is the pressure in the pitot tube, ρ is the density of the fluid, and p and V are the static pressure and velocity in the undisturbed stream. As found in previous investigations, Cp becomes greater than 1 at low Reynolds numbers, the increase being about per cent, at a Reynolds number of 50 (based on external tube radius). In disagreement with the work of Hurd, Chesky, and Shapiro, no decrease of Cp below 1 was found at any Reynolds number.

The main factor contributing to hydroplaning on slick runways is the existence of a squeeze film in the contact area between tyre and surface. The situation may be improved by optimising surface texture so that the asperities break through the water film in a minimum time. Two techniques are described for characterising surface texture: a superposition solution based on viscous channel flow and a dimensionless number approach involving size, spacing and shape factors. The “feel” of a surface is quantitatively obtained, and an outflow meter is described which, by directly measuring the mean hydraulic radius of runways, provides a comparison of their drainage abilities.

This paper reports on the measurement of aerodynamic forces that act on a square cylinder which was downstream to an identical cylinder. When the two cylinders were in tandem formation, a critical spacing equal to about four times the side length D of the square cylinder, was found to exist. For L/D < (L/D)cri, where L is the streamwise centre to centre spacing between the two cylinders, the boundary layers which separate from the upstream cylinder reattach onto the downstream cylinder. Only the latter sheds vortices and it is subjected to a thrust rather than a drag force. For LID > (L/D)cri, both cylinders shed vortices and both are subjected to drag forces. When the two cylinders were in staggered formation, due to the presence of various lift generating mechanisms, the lift force that acts on the downstream cylinder can be either positive which points away from the wake centre line of the upstream cylinder or negative which points towards the wake centre line of the upstream cylinder. It was found that in the range 4 ⩽ LID ⩽ 12, an unstable regime exists in that the lift force generated pushes the downstream cylinder away from the upstream cylinder’s wake centre line at small dimensionless transverse spacing (Y/D) where Y is the transverse centre to centre spacing between the two cylinders and reverses its direction only at Y/D > 2. At L/D = 29, the largest streamwise spacing that can be investigated in the present study, the influence that the upstream cylinder has on the aerodynamic forces acting on the downstream cylinder was found to be still fairly appreciable.

Circumstances can arise in which a rectangular panel, forming part of a structure, is subject to non-uniform in-plane compression, and in which it is important that instability of the panel should not occur. Problems of this nature occur, for example, in the structure of modern highspeed aircraft subject to the transient thermal stresses that can arise from rapid changes of speed or altitude.

A double Fourier series solution is presented for the buckling of simply supported orthotropic and isotropic skew plates, subjected to in-plane compressive and shear edge loads. All the boundary conditions, including the vanishing of the normal bending moment along the edges, are satisfied. Numerical results, suitable for design purposes, are presented graphically for a number of independent parameters such as skew angle, ratio of flexural rigidities and aspect ratio. The results show that the buckling strength is substantially effected by the skew angle, and to a lesser extent, by the ratio of the flexural rigidities and aspect ratio. In contrast with positive shear edge loads, which tend to decrease the skew angle, negative shear edge loads appear to reduce the buckling strength of orthotropic and isotropic skew plates.

The development of new technologies – such as rapid prototyping – and the use of materials with improved properties – such as highly resistant extruded polystyrene foam which can be easily and precisely shaped, while conserving its mechanical properties – allow researchers to improve design concepts. This article details the development of a new set of morphing wings for a 15-kg maximum take-off weight Unmanned Aerial Vehicle (UAV) from concept design, to flight tests, including modelling, design optimisation, construction and wind-tunnel tests. A set of comparator-equivalent conventional wings have been used throughout in order to be able to judge any benefits stemming from the adoption of morphing technology. This article shows that the morphing wings provide a controllable aircraft while reducing drag by a factor of 40% compared to the comparator wings with conventional ailerons in a deflected position.

The “frozen stress” techniques of photoelasticity can give a complete knowledge of the stress, system in a solid body, but the examination of the stresses requires more time and care than in corresponding flat plate tests. In tests on tubes with transverse circular holes, sponsored by The Royal Aeronautical Society, all practicable geometrical shapes are examined and the maximum stress is measured in tension, bending and torsion. The results are comprehensive and show the inadequacy of previous results. In all cases the maximum stress occurs inside the bore of the hole. The accuracy of all the graphs of stress concentration factors is better than five per cent.

New wind tunnels appear fairly infrequently and are usually of individual design to meet specific requirements and constraints. Detailed information on existing tunnels can be of great benefit to the designer, enabling him to avoid expensive pitfalls and to capitalise on the successful features that have evolved over the years. For this reason, and to advance wind tunnel design generally, new designs should be reported. The following account is presented in this light, even though the low speed tunnel described is unremarkable, being of conventional design and modest proportions. Indeed, its design follows closely the guide-lines laid down for such tunnels by Bradshaw and Pankhurst, guide-lines which themselves draw on the cumulative experience gained from previous low speed wind tunnels. The satisfactory performance of the tunnel is in no small measure the result of this approach.

An improved version (HLLC) of the Harten, Lax, van Leer Riemann solver (HLL) for the steady supersonic Euler equations is presented. Unlike the HLL, the HLLC version admits the presence of the slip line in the structure of the solution. This leads to enhanced resolution of computed slip lines by Godunov type methods. We assess the HLLC solver in the context of the first order Godunov method and the second order weighted average flux method (WAF). It is shown that the improvement embodied in the HLLC solver over the HLL solver is virtually equivalent to incorporating the exact Riemann solver.

The Dynamic Relaxation method is used to analyse the post-buckling of flat rectangular plates. Extensive results are obtained for two types of problem in which the transverse edges are unloaded; in one case the transverse edges remain straight, in the other they are free to wave. Comparisons are made with alternative experimental and theoretical results. The potentiality of this approach to post-buckling problems is demonstrated by considering a variable thickness plate with mixed boundary conditions.

From Lanchester to the A380; from the era of the great pioneers to the integrated engineering teams and sophisticated analyses required to develop todays airliner. The ‘art’ of Aerodynamics has progressed a long, long way in the intervening century, but even with todays computational power it is still relevant to understand the basics and the early theories as developed by Lanchester and his contemporaries. In this lecture, the relevance of some of that early work and the compromises that must be made in optimising a modern wing design will be presented. These will include optimisation of spanwise loading, including trimming effects and winglets, and high and low speed aerodynamic and structural interactions, with examples drawn from the author's career in helping to develop the wings for the HS146 (now the RJX) and the Airbus family.

With the exponentially increasing computer power available over the last 30 years, the ‘art of compromise’ has progressively become the ‘science of optimisation’ and the lecture closes with a brief reference to techniques used to optimise current wing designs for the Airbus family.

Recent advances in coaxial rotor design have shown benefits of this configuration. Nevertheless, issues related to rotor-head drag, aerodynamic performance, wake interference, and vibration should also be considered. Simulating the unsteady aerodynamic loads for a coaxial rotor, including the aerodynamic interactions between rotors and rotor blades, is an essential part of analysing their vibration characteristics. In this article, an unsteady aerodynamic analysis based on a vortex particle method is presented. In this method, a reversed-flow model for the retreating side of the coaxial rotor is proposed based on an unsteady panel technique. To account for reversed flow, shedding a vortex from the leading edge is used rather than from the trailing edge. Moreover, vortex-blade aerodynamic interactions are accounted for. The model considers the unsteady pressure term induced on a blade by tip vortices of other blades, and thus accounts for the aerodynamic interaction between the rotors and its contribution to the unsteady airloads. Coupling the reversed-flow model and the vortex-blade aerodynamic interaction model with the viscous vortex-particle method is used to simulate the complex wake of the coaxial rotor. The unsteady aerodynamic loads on the X2 coaxial rotor are simulated in forward flight, and compared with the results of PRASADUM (Parallelized Rotorcraft Analysis for Simulation And Design, developed at the University of Maryland) and CFD/CSD computations with the OVERFLOW and the CREATE-AV Helios tools. The results of the present method agree with the results of the CFD/CSD method, and compare to it better than the PRASADUM solutions. Furthermore, the influence of the aerodynamic interaction between the coaxial rotors on the unsteady airloads, frequency, wake structure, induced flow, and force distributions are analysed. Additionally, the results are also compared against computations for a single-rotor case, simulated at similar conditions as the coaxial rotor. It is shown that the effect of tip vortex interaction plays a significant role in unsteady airloads of coaxial rotors at low speeds, while the rotor blade passing effect is obviously strengthened at high-speed.

Operation of vehicles over dusty terrains and unprepared airstrips has highlighted the erosion problems incurred by ingestion of dusty air. Hibbert summarised the experience from British helicopter trials up to that date and drew attention to some of the factors that influenced the extent of erosion damage and to methods of minimising it. He cited a trial during 1963 in which a twin-engine Wessex Mk 2 helicopter having one intake protected by felt filters was tested in North Africa. Some idea of the severity of the problem may be gained from the fact that the unprotected engine surged after 3h 40 min while the other was found to have lost half its surge margin. However, operation can usually be arranged so that time spent in the dusty environment is minimised and service lives under desert conditions may be typically around 100h. In attempts to improve lives to a more acceptable level, different types of filtration have been investigated and the mechanism of erosion studied in some detail. This note describes some laboratory experiments to determine the influence of particle size on the erosion and attempts to relate this to filtration and engine life.

When a vibrating structure encloses a volume of fluid, the acoustic effects within this volume modify considerably the response characteristics of the structure, provided that the cylinder is vibrating in radial modes only. Measurements made of the displacement caused by a particular sound wave are of the same order as the values predicted. The calculation of the response of the cylinder to an acoustic wave also yields the sound field inside the cylinder and, again, the results are in general agreement with practical experience.

Free vibration characteristics of stiffened plates possessing symmetrical stiffeners have been investigated by the finite element method. The main elegance of the element used lies in that the formulation takes into account the arbitrary orientation of a stiffener inside the plate element. The element being isoparametric in nature caters effectively for the irregular edges of the boundary as well as the transverse shear deformation of the plate and the stiffeners. Rectangular and skew stiffened plates having various boundary conditions and with stiffeners varying in number and in spacing, have been analysed. The correlation of the natural frequencies obtained by the present approach with the published results have been found to be very good.

The prediction of blade deflections and vibratory hub loads concerning helicopter main rotors in forward flight is the objective of this work. They are determined by using an aeroelastic model derived through the coupling between a nonlinear blade structural model and a boundary integral equation solver for three-dimensional, unsteady, potential aerodynamics. The Galerkin method is used for the spatial integration, whereas the periodic blade response is determined by a harmonic balance approach. This aeroelastic model yields a unified approach for aeroelastic response and blade pressure prediction that may be used for aeroacoustic purposes, with the possibility of including effects from both blade-vortex interaction and multiple-body aerodynamic interaction. Quasi-steady aerodynamic models with wake-inflow from the three-dimensional aerodynamic solver are also applied, in order to perform a comparative study. Numerical results show the capability of the aeroelastic tool to evaluate blade response and vibratory hub loads for a helicopter main rotor in level flight conditions, and examine the sensitivity of the predictions on the aerodynamics model used.