It is my intention to be severely critical of the report insofar as it concerns licensing, but I hope that what I have to say will not set the tone for this symposium! It is extremely easy to be critical of the work of others, particularly in a field in which one feels one has some experience, but I shall do my best only to offer constructive criticism.

The density of air traffic, particularly in the USA, is now X increasing at such a rate that the overloading of the Air Traffic Control system seems inevitable. Fig. 1 shows the expected increase in the size of the US civil air fleet and its operations over an eleven-year span, from 1966 to 1977.

Since January 1st 1968, there have been 58 mid-air collisions in the USA, including 19 this year. The majority of these were under positive Air Traffic Control in the vicinity of airports. Some form of airborne system to prevent collisions seems essential for the safe operation of aircraft in the 1970s.

A considerable amount of discussion and development has been carried out in the USA during the past 14 years, and some prototype equipments are now flying. The purpose of this paper is to describe the technique employed in these systems, and to suggest an original principle as a design basis for a collision avoidance system (CAS).

The use of an expansive area change immediately following the diaphragm has afforded considerable convenience in shock tube applications and has been studied by a number of authors. It allows the use of a smaller diaphragm, thereby reducing the disturbance to the shock tube flow caused by the diaphragm opening, and it also permits greater flexibility in the choice of driver section configurations for a given shock tube size.

It is natural to consider the extension of this idea to shock tunnel operation. This has been encouraged by the possibility of applying area change to operation of an arc-driven shock-tunnel. Some early success in employing arc-heated drivers for shock tunnels has been reported by Warren et al. However, the configuration they studied employed a long driver section to achieve correspondingly long test times. In addition, to achieve a constant shock velocity it was necesssary that electrical energy was added so that the gas was heated uniformly before the diaphragm ruptured. This necessitated a fast discharge condenser bank, with a high discharge voltage, the latter to enable an arc to be sustained over the length of the driver section. In the present application, a constriction between the driver section and the shock tube is employed, allowing the driver section to be treated as a constant pressure reservoir of gas. It is anticipated that this will allow the use of shorter drivers, thereby permitting the adoption of lower discharge voltages, and will also allow energy to be added relatively slowly while the shock tunnel is operating. This will produce substantial economies in provision of the large energy storage facilities required for shock-tunnel operation.

In a recent paper, Portnoy discussed the properties of a slender-wing with a slender half-body of revolution mounted beneath it. Another configuration of the same type has a fuselage of rectangular cross-section underneath the wing. The present note gives results for the lift on this class of configurations in attached flow and compares it with that for other wing-body combinations.

Ward's slender-body theory describes the flows about appropriately slender configurations where viscous effects are negligible (i.e. the boundary-layer thickness is small in comparison with the transverse dimensions of the body and no separation occurs upstream of the station under consideration). A restriction of the simplified version used in this note is that no element of a load-bearing surface may be upstream of that station. This implies that the wing span must not decrease in a spanwise direction. Another assumption we make is that no leading-edge separation occurs.

Control of flight is essentially control of lift. When one proposes—as we shall do here—to tamper with one of the most fundamental principles of aircraft control, it is evident that this may have far reaching repercussions and that such a venture cannot therefore be undertaken lightly. The elevator has been the primary longitudinal control since the earliest pioneering days and has served the aircraft well from its modest beginnings to this age of supersonic flight. Why do we want to change and what do we want to take its place; how will the pilot react to an altogether novel way of controlling the aeroplane? These are the questions one is bound to ask when direct lift control is suggested as an alternative and these are the questions which we shall attempt to answer here.

I think that Sir Ronald Edwards and his small team of very busy people deserve to be congratulated upon producing in a remarkably short time a Report that is well documented, well written, generally realistic in its proposals and free of prejudice on such issues as public versus private enterprise. As against this, it may be objected that the Report is inclined to accept rather too readily the orthodoxies of modern “consensus” politics and economics. By that I mean such views as that a mixed economy necessarily achieves the best of both worlds; that nationalised industries should operate as commercial undertakings; and that where nationalised industries are required for some special reason not to operate commercially, they should be compensated in the form of direct subsidy chargeable to the appropriate government department.

The purpose of my contribution is to explain the need for dynamic testing in terms of defining the design problems where testing is required to assist or substantiate decisions made in the light of mainly theoretical evidence in the initial stages of design before hardware becomes available for testing.

I have interpreted the title in the broadest sense of the word vibration so as not to restrict the presentation only to testing associated with assessing or minimising the vibration of helicopters, but to include problem areas which result in dynamic testing requirements.

There are many approximate methods for solving the laminar boundary-layer equations. They vary greatly in complexity and ingenuity but are similar in that most of them use the momentum integral equation together with data culled from Cohen and Reshotko's valuable catalogue of similarity solutions. Further details may be found, for example, in the book by Curle.

Perhaps one of the simplest approximate methods to use is that due to Luxton and Young. More recently Luxton has extended the method so that heat transfer rate distributions also can be obtained. He is careful to point out that, since his method depends on correlations obtained from similar solutions, it cannot be claimed a priori that the method gives an accurate result. Until recently there have been few “exact” calculations against which approximate methods may be tested. Indeed when Morduchow made his review of boundary-layer calculations in 1965 he quoted only two exact results for non-zero Mach number (by NPL) and cast doubt on one of those. Fitzhugh has now computed a number of solutions of the laminar boundary-layer equations using the finite-difference program of Sells. Further solutions have recently been obtained by F. Smith.

The analysis of a cylindrical shell with cut-outs, which are reinforced to reduce the stress concentrations, is of importance in many engineering fields, particularly in aircraft structures. The most common assumptions made are that the curvature effect is negligible and that the reinforcement is a compact flexible bead. In the analysis of holes in fuselage shells the neglect of the effect of curvature may lead to serious errors as shown theoretically by Lurie and Vandyke and experimentally by Houghton and Rothwell. In some cases the bending stresses are of the same order as the membrane stresses which themselves may be much higher than the flat plate solution. Lurie obtained a perturbation solution for an unreinforced hole by expanding in powers of curvature parameter β, the results being valid for small values of β. Vandyke extended this solution up to β=4 by using a collocation procedure. In fact, if the collocation method is used it seems possible to obtain a convergent series solution for any β, the only limitations being the shallowness of the shell and the restriction that the deflections are not large.

Sandwich panels are utilised quite often in aircraft and missile structures mainly because of their high strength to weight ratio. The core of the panel, made of light material such as balsa wood or synthetic material, stabilises the face skin sheet which is usually metallic, and contributes to the shear strength of the section.

Rectangular sandwich panels have been analysed by various investigators. In particular, Reissner deduced the range of linearity of deflections as a function of the core rigidity relative to that of the face skin. Results for parallelogrammic sandwich panels, however, are not available.

One of the best known results in finite wing theory, which is quoted in aerodynamic text books, states that the distance between two rolled up vortices behind a wing with an elliptic spanwise load distribution is sπ/2, where s is the wing semi-span. Within the framework of linear wing theory this result is correct. But it is of interest to enquire how this result is affected when the standard linear model becomes less valid, for example, at higher wing lift coefficients.

In this note, it is assumed that a continuous trailing vortex sheet rolls up into two discrete vortices. These two discrete vortices are assumed to have cores of finite radius; inside the cores the flow is taken to be solid body rotation, while the flow outside the cores is the standard irrotational vortex field. It is assumed that this rolling up process takes place far downstream of the finite wing which generates the trailing vorticity.

Over the subsonic speed range, the pressure field attached to a ducted rotor decays rapidly with distance away from the rotor. This exponential decay, coupled with the low radiation efficiency of a subsonically rotating pressure pattern, is generally sufficient to bring the radiated sound power well below that produced by interaction of the rotor with disturbances in the flow. This is particularly true for multibladed rotors as used in aircraft-engine fans and compressors, and accounts for the importance of rotor-stator interaction as a source of noise at subsonic tip speeds.

The pursuit of better economics for subsonic transport aircraft has led in recent years to the development of high by-pass ratio turbofan aero engines. The improved efficiency of these engines results mainly from an improvement in the propulsive, or Froude efficiency brought about by the use of an increased airflow with a smaller jet velocity for a given thrust level. The specific thrust (thrust per unit airflow) for a high by-pass ratio engine is thus appreciably lower than for a conventional turbojet so that the ratio of engine gross thrust to installed net thrust is higher. Typically, for an engine of unit by-pass ratio, the ratio of gross to net thrust is about 1·6 under cruise conditions, while for an engine of by-pass ratio five it is between 2·6 and 3. Losses in the exhaust system effectively act on the gross thrust level so that any inefficiency is magnified by the gross to net thrust ratio. Thus the high values of this ratio in the new high by-pass engines render losses in the exhaust system of especial significance.

Many structures are made up of skin material which is stiffened by reinforcing members. Configurations of this type are predominant in the aircraft industry where they are referred to as semi-monocoque structures. When analysing such structures, using finite element techniques, the reinforcing members (stringers) are lumped together to form pure axial elements and the skin is then represented by isotropic plane stress elements.

In several recent papers the bursting pressures of shock tube diaphragms of various sizes and materials have been presented either as discrete values or curves of bursting pressure against thickness. The lack of a unifying dimensionless presentation, however, restricts the general applicability of such results and it is the object of this work to present a suitable dimensionless strength parameter, and to outline a procedure which enables predictions to be made of the bursting pressure of ductile metal diaphragms in square-section shock tubes of any cross-sectional size, without performing bursting tests.

The elementary beam theory for shell structures yields the exact solution for normal and shear stress, if the three following conditions are fulfilled:

• (a) the shell must be prismatic;

• (b) the cross section must not deform in its plane;

• (c) the centres of rotation of the different cross sections must form an axis of rotation parallel to the longitudinal axis of the beam.

The analysis of axisymmetrical thin shells has always been an important part of the theory of structures, and recently the problem of the cooling tower has aroused considerable interest for practical engineering reasons. Various solutions have been obtained, using either the finite difference or the finite element methods, with varying degrees of success. Amongst the latter, elements such as the TRIB 3c from the ASKA library and the axisymmetrical elements from the SABOR programme, with either a straight or a curved meridian line, have given reasonable displacements. But, by the nature of these elements, the stress distributions are discontinuous across element boundaries. Other existing elements, e.g. the arbitrarily curved triangular shell element SHEBA developed by Professor Argyris's team at the ISD in Stuttgart, are also suitable for the purpose and yield an excellent estimation of the stresses.