A great deal of interest has been shown in recent years in the problem of flexibility in spacecraft structures. Much published work has been devoted to analysing the dynamic interaction between deformation of flexible appendages and attitude control systems. A limited number of publications have sought to find ways of providing some form of active control of flexible motion. Where this has been attempted idealised systems consisting of flexible coupled pendulums or rotors have been considered.

The problem of flexibility may be tackled in three ways. Where the flexure can be shown to be secondary in nature, careful design of the attitude control system can ensure that interaction is kept to a minimum. To this end surveys of structural flexibility effect in spacecraft are invaluable to the design engineers. Where the flexural motion has a more dominant effect, as will follow from the inevitable increase in power requirements for communications, then a more direct approach will become necessary.

Increasingly stringent operational demands on military aircraft, together with the widening flight envelopes made possible by advances in aerodynamic, structural and propulsion technology, have forced the complication of stability and control augmentation systems (S/CAS) upon the aircraft designer. The early SAS configurations were conceived as devices to enhance the natural response characteristics of the aircraft; later, CAS evolved as a way of making the desired response more independent of the basic aircraft characteristics. Many recent examples of S/CAS are very complex, with equations of twentieth or higher order commonly being used for linear handling qualities analysis. In view of the fact that aircraft of this complexity have been part of the handling qualities picture for ten years or more, it is surprising that no widely accepted criteria for these aircraft have evolved.

The study of a jet impinging on a finite disc has applications in separating the stages of a multi-stage rocket during its flight. For example, the jet plume of the upper stage impinges and therefore induces a pressure field on the upper surface of the lower stage. A knowledge of this pressure field is vital in determining the trajectory of the lower stage after separation.

For many years, a first order understanding of aircraft performance has been possible from lift drag curves or forward speed/sinking speed curves for gliders. These presentations of performance are not sufficiently accurate for detailed aircraft performance analysis, but their understanding has enabled generations of aircraft engineers to grasp the importance of various aerodynamic parameters.

A new method of presenting the sailing vector triangles is described. When the locus of the possible vector triangles is drawn in the new presentation, a large portion approximates to a straight line parallel to or making a small angle with the Xaxis in the neighbourhood of the best speed to windward. A geometrical construction is given, which defines the best speed to windward point. It becomes self-evident that, as is well known, most yachts sail at nearly 45 degrees to the true wind when sailing at their best windward condition.

At the Symposium in April last year the Society met to consider how design and operation of quiet aircraft related to the often conflicting economic and environmental pressures. The prospects for quieter aircraft and engines were discussed from many technical points of view. The message emerging from the discussion was that, compared to the improvements made in the past 10 years, and which would be available now at the cost of replacing existing noisy, older generation installations, further reductions in noise at source would be very much harder to win. Moreover, they would be made against a steeply-rising trend of overall detriment to performance and direct operating cost.

So far as the possibilities are concerned for further reductions in the noise of subsonic propulsion engines, if we are to have more than modest, evolutionary improvements in levels, there will need to be radical changes affecting engine design, such as a definite move towards lower jet velocities.

The Jaguar programme resulted from a successful international collaboration in all the relevant disciplines between companies having very different backgrounds in the aviation field, namely British Aircraft Corporation and Breguet Aviation (since merged with Dassault) for the airframe and Rolls-Royce and Turbomeca for the engines. It also involved collaboration in depth between government officials having different civil service structures and national practices in the two countries. Before entering into a description of the various phases of the programme and some of the problems which were met and solved it is worth briefly reflecting on the background leading to the formal start of the project in May 1965.

At present it is still very difficult to predict for a given incidence, sweep-angle, leading-edge shape and Reynolds Number, whether the flow around a wing will remain attached or separate either from the upper surface as a bubble or from the leading edge as a vortex sheet which rolls up above the wing. Various attempts have been made to remove this uncertainty and it was suggested by Küchemann that it should be possible to design wings which combine coiled vortex sheets, originating from the inner part of the leading edge, with attached flow over the leading edge of the outer part of the wing. In this way some of the properties of slender wings could be combined with some properties of swept wings. His original model consisted of a swept-wing planform modified by inboard leading-edge wing strakes, the purpose of the latter being to provide the inboard vortex sheets. A similar idea has now been used in the wing design of the YF-16 lightweight fighter.

The conventional landing technique for most aircraft involves a constant speed approach along a well denned glidepath. The speed is selected to maintain an adequate margin above the stall. Allowance is made for wind conditions. At the beginning of the landing approach a pilot is quoted a nominal mean surface windspeed and direction by radio contact with the airfield control tower. The local controller is provided with a conventional meteorological office dial-reading anemometer but, because of other duties, does not monitor it continuously.

Aerodynamic properties of oblique wings have been studied intensively at NASA Ames Research Center and also by the Boeing Commercial Airplane Co. Figure 1 shows the design proposed by Boeing for a 200 passenger transport intended to fly in the transonic speed range; the design employs an oblique wing of high aspect ratio.

One feature of the oblique wing design that has not been fully documented is its aeroelastic stability. Since the oblique wing has one panel swept forward and the other swept back, it is perhaps natural to try to relate its behaviour to well known phenomena associated with conventional swept-forward or swept-back wings. This approach would lead one to expect aeroelastic divergence of the forward wing panel.

This note is concerned with the response of a pitot tube in the region of its interaction with a weak shock wave. It will be helpful if we distinguish immediately between our usage of the terms ‘pitot pressure’ and ‘normal shock recovery pressure’. We use the former to mean the experimental reading obtained from a pitot tube and the latter to mean the theoretical stagnation pressure downstream of a normal shock at the Mach number of a point in the flow.

Across an oblique shock, the normal shock recovery pressure changes discontinuously. However, in the course of a recent study of the impingement of supersonic jets on flat plates, the authors noticed that, as a pitot probe was traversed across a weak shock wave, not only was the discontinuity smoothed out by the finite size of the probe but also values of pitot pressure were obtained which were larger than the higher of the two values of normal shock recovery pressure.

As one means of achieving the required movement of a flying control the simple skew-hinge mechanism appears attractive because it has a significant range of approximately linear operation and because all of its mechanical components can be accommodated easily within the sort of shallow depth available in some wings. The mechanism is kinematically similar to that of the well-known Z-crank and Fig. 1 shows one possible arrangement of it inside a wing. For this particular case an axial translation of the control push-pull rod AA causes rotation of the flying control H about its hinge axis BB and the mechanical connection between these driving and driven members consists of the skew-hinge itself, which is attached to the flying control H and which has an axis CC which intersects the axis BB at an angle α, together with the other components of the mechanism. These are the hinge driving-block F, the driving-link G, and the drive transmission-collar I.

In the theoretical and experimental work on the buckling of sandwich panels with laminated faces reported in Refs. 1 and 2, panels which had certain coupling terms present in the constitutive equations for the faces were solved in an approximate manner. The previous theoretical solution for angle-ply lay-ups assumed that there were sufficient layers present to treat the faces as orthotropic, with the Bij coupling terms in the constitutive equations put equal to zero. Further simplifying assumptions were made when comparing theoretical and experimental results in Ref. 2 where tests were carried out on three different types of carbon fibre faces; orthotropic, balanced angle-ply and unbalanced cross-ply. It was only in the first case that an exact theoretical analysis was carried out, and the second case was again treated as an orthotropic panel neglecting coupling between direct and shear deformations and bending and twisting curvatures. The unbalanced cross-ply lay-up, which consisted of two fibre reinforced plastic orthotropic layers, one at 0° to the plate axis and one at 90°, gave rise to large B11 and B22 coupling terms which were taken into account approximately using “reduced stiffness” terms with the orthotropic equations. However, the disparity between theory and experiment was greatest in this case, particularly as far as wrinkling was concerned, where the experimental failure load was 78% higher than the theoretical wrinkling load. This type of cross-ply layup is important since it is used when making carbon fibre sandwich floor panels; it being chosen from the point of view of manufacturing convenience rather than from considerations of strength and stiffness. Furthermore, the bending stresses in the faces, which result from the normal loads, predominate in one direction so that one face is subjected to essentially a uniaxial compressive stress which might cause the faces to wrinkle. In order to study this problem in greater depth, it was decided to carry out further theoretical and experimental work on the simpler problem of the wrinkling of sandwich columns with carbon fibre cross-ply faces.

Advances in computational methods are producing a need for accurate and detailed experimental results for test cases. At the same time economic trends force many of these experiments to be made in small wind tunnels. In some cases the small scale of the model does not provide enough space for the instrumentation and so the model must be distorted from the ideal shape. This paper is concerned with one aspect of this problem, that is, with the need to make wings thicker than required, either for strength or to provide volume for pressure tubes and other measuring devices. In particular the paper deals with the conditions under which the pressures measured on the flat surfaces of wings with bevelled or triangular cross-sections as shown in Figs. 1(b) and 1(c) are a good representation of the pressures which would be measured on the corresponding flat surface of thin wings (Fig. 1(a)). This problem has been touched on in a number of papers reporting experimental results but it does not appear to have been tackled systematically. The discussion of this paper is based on these earlier tests and on some new test results.

During the course of the preceding decade, numerous solutions to laminated plate problems have appeared in the literature. This has been the direct result of the development of new materials, e.g. Carbon Fibre Reinforced Plastics (CFRP), for the aerospace industry, where lightness allied with stiffness are characteristics of paramount importance. However, it appears that few, if any, solutions have been published for the case of plates on elastic foundations, even though such solutions may be useful when analysing the face behaviour of certain types of foam-filled, sandwich panels. In attempting to provide such solutions and to establish the effect of the foundation on the plate response, the subsequent analysis was undertaken. As a first step towards this goal, it appeared sensible to consider the simple case of a plate strip, before considering the more complex case of a finite plate. Consequently the analysis and results presented herein are restricted to plate strips.

A problem, commonly met in scientific and engineering experiments, is the determination of the damping of a system, given a recorded trace (such as that from a uv recorder) of the free damped oscillation following a disturbance. The simplest and most tedious method is to plot the logarithm of the peak amplitude at each half cycle against cycle number. Other methods have been given by Hogley and Gimmestad, Bratt and Wilcox and Crawford. However, the problem is so common, and the experiments and methods of recording the results so diverse, it is felt that a further method could find a place in the range of those available.