In this paper some experimental results related to a calibration model (the Standard Dynamics Model, SDM) are presented.

The wind tunnel test activity is based on static force measurements and flow visualisations. The behaviour of the fuselage on its own is analysed and a correlation with some available prevision theories for the onset of asymmetries on ogival forebodies is proposed.

A complete set of static data for the SDM is introduced, evaluating the main aerodynamic interactions between the various elements of the configuration.

A series of flow visualisations was performed to verify with qualitative results the validity of the conclusions found from the analysis of the static coefficients.

In several recent reports reference is made to spanwise variations in the boundary layers growing in the working sections of various wind tunnels. Systematic investigations by Patel, Bradshaw and de Bray are in general agreement that these variations are caused by the screens which are fitted into the settling chambers of these tunnels to dampen out non-uniformities in the upstream flow. They conclude from their experiments that these variations will be much reduced if the downstream screen (or preferably the two downstream screens) has an open area ratio greater than about 0·57 and that these should be followed by a honeycomb to obtain the most steady and uniform flow, although a honeycomb in this position may increase the turbulence level in the working section.

A theory is presented for the airflow and power requirements of a compressible hovercraft jet operating under equilibrium conditions. It is shown that, even for a partially sonic jet, these requirements may not differ markedly from those of an incompressible peripheral jet.

The problem of unsteady hypersonic and supersonic flow with attached shock wave past wedge-like bodies is studied, using as a basis the assumption that the unsteady flow is a small perturbation from a steady uniform wedge flow. It is formulated in the most general case and applicable for any motion or deformation of the body. A method of solution to the perturbation equations is given by expanding the flow quantities in power series in M−2, M being the Mach number of the steady wedge flow. It is shown how solutions of successive orders in the series may be calculated. In particular, the second-order solution is given and shown to give improvements uniformly over the first-order solution.

The serious mathematical complications of the kinetics of the aeroplane spin have given birth to methods of explanation based on assumptions and simplifications, the application of which, not unfrequently, has led to confusion and misunderstanding on the part of the student, let alone error on the part of the initiator of the method.

One of the problems, is the selection of the geometric parameters that have to be introduced as opposed to those which may be left out of the picture. Obviously, in such complex movement as the spin, the most obscure parameters are the angles of the motion. Unfortunately, there is a minimum number of such angles which is necessary and if one, hopefully, begins with a lesser number, one soon finds that the rest have to be put in place, even tardily, if the puzzle is to be completed.

Downward deflection of a propulsive jet may be employed to augment aerodynamic lift in an advantageous manner. In each situation there occurs a simple optimisation when the losses due to reduced propulsive efficiency are contrasted with the gains which accrue from jet lifting. Typical cases will be considered here for wing dominant configurations, i.e. cases where the lifting surfaces contribute most of the drag, with the body contributing but a minor effect. Variable geometry of both airframe and engine is implicitly assumed so that results may be applied to preliminary design studies which will determine the relative propulsion system size and wing area.

Vortices shed from cooler tubes can excite standing pressure waves in coolers if the vortex shedding frequency coincides with the standing wave frequency(1, 2, 3, 4). If resonant conditions are encountered in a particular cooler the single vortex shedding frequency may be suppressed by fitting in-clined splitter plates into the coolers. These splitter plates also “detune” the coolers by raising the resonant frequencies.

Splitter plates of this type were used recently to eliminate severe noise and structural vibration in a new cooler for the 3 ft wind tunnel at the Royal Aircraft Establishment, Bedford.

Shock control bumps can be used to control and weaken the shock waves that form on engine intakes at high angles of attack. In this paper, it is demonstrated how shock control bumps applied to an engine intake can reduce or eliminate shock-induced separation at high incidence, and also increase the incidence at which critical separation occurs. Three-dimensional Reynolds-average Navier–Stokes (RANS) simulations are used to model the flow through a large civil aircraft engine intake at high incidence. The variation in shock strength and separation with incidence is first studied, along with the flow distribution around the nacelle. An optimisation process is then employed to design shock control bumps that reduce shock strength and separation at a fixed high incidence condition. The bump geometry is allowed to vary in shape, size, streamwise position and circumferential direction around the nacelle. This is shown to be key to the success of the shock control geometry. A further step is then taken, using the optimisation methodology to design bumps that can increase the incidence at which critical separation occurs. It is shown that, by using this approach, the operating range of the engine intake can be increased by at least three degrees.

When I was asked to take part in this Symposium it was suggested to me that I should talk about the Accident Investigator's view on Flight Safety. Now I do not suppose that the accident investigator's view is very different from the view of anyone else working in the safety field. The whole object of flight safety is to prevent accidents and the whole object of the accident investigator is to ascertain the causes of accidents so that remedial measures may be taken.

The usual method of load measurement with strain gauge load cells in a bridge circuit involves the measurement of the out of balance voltage either by the height of a trace displayed on a cathode ray oscilloscope, or the width of a band produced by a tuned galvanometer. Both involve careful regulation of the input voltage or calibration during use. The method described provides for a null balance of the bridge at any instant in the cycle of a uniformly repeated cyclic load up to 10,000 cycles per minute. An accuracy of ±½ per cent, of the capacity of the load cell is obtainable.

The rapid progress made in mechanical engineering during the present century has brought many troublesome problems in its wake. One of the most serious of these troubles arises from resonant and such like vibration. It has, therefore, become increasingly necessary for a designer to take account of vibration. A powerful instrument for this purpose was provided by the late Lord Rayleigh, whose method for finding an approximation to the fundamental frequency of vibration of an elastic system is fairly well known. In this paper an attempt is made inter alia to explain Rayleigh's Principle and to clarify, so far as the author is able to do so, the basis of the principle. Other devices for ascertaining approximate frequencies are also given and an appendix has been added on forced vibration experiments.