The Working Party considered the objects of professional training and of technical education and the parts played by both academic studies and by practical experience; these deliberations led them to a consideration of how the Diploma in Technology helped to achieve these objects. The content of both the academic studies and industrial training parts of the Diploma course was studied. Conclusions reached included the following:—

1. (a) Close liaison between lecturing staff and professional engineers is vital.

2. (b) Standards of supervision and staff engaged in industrial training should be similar to those found in academic studies.

3. (c) The Training Boards should initiate bodies in which the industry and the professional institutions co-operate with colleges in the maintenance of standards of industrial training.

4. (d) Attempts to assess and grade students on their industrial training should be approached with caution.

5. (e) Suggestions are made as to the form of projects to be undertaken by Diploma Students.

6. (f) It is neither necessary nor practical to integrate each part of the studies and practical training in detail.

7. (g) The Working Party is whole-heartedly in favour of the sandwich arrangement of courses and makes suggestions for the best lay-out.

8. (h) Recommendations are made concerning the qualification to be given for successful completion of the Diploma in Technology course; it is concluded that this should differ from that given for the normal full time degree course.

The Working Party concludes by making recommendations for further action.

The current interest in supersonic transport aircraft makes it appropriate to consider the problems of propulsion nozzle systems for high speed flight. The discussion will be confined mainly to this field of activity.

In the absence of external drag and weight factors, the ideal configuration to maintain optimum thrust at all flight speeds would employ a mechanically variable area ratio nozzle. By such means the jet flow could be fully and properly expanded at all times.

A practical design to accommodate such variations is inherently complex, and so an alternative solution using aerodynamic techniques has been sought. It is necessary to remember that the success of whatever method is eventually chosen must rest on its ability to combine a high performance at cruise with very small loss during other critical phases of the flight plan.

This paper describes some of the results of a nozzle research programme aimed primarily at solving the above problem, with an attempt to explain the philosophy behind the aerodynamic techniques tried. Some discussion is given of the experimental difficulties in establishing a sufficiently accurate standard of measurement for design point performance. In a few chosen configurations experimental results have been extensively backed by theoretical studies.

The need to demonstrate off-design behaviour in the presence of an external flow field is stressed, with special reference to the misleading results often obtained from static tests.

Finally, some attention is given to the question of base bleed, with its repercussions on the installation of the propulsion nozzle in an aircraft.

The development of a special model aeroplane technique is reviewed which renders possible easy and illustrative experimenting In the Reynolds number regime from 1 to 5000. The models are built in extremely light-weight construction; with wing span from 0·1 to 1 metre and all-up weights from 0·01 to 2 grammes, flight velocities between 0·1 and 1 m/ sec are reached. The requirement of extremely low wing loading, together with the special air flow characteristics in this region, leads to rather unconventional construction principles with the materials balsa wood, nichrome wire and cellulose film. These principles are reviewed briefly. The problem of optimum propulsion for such model aeroplanes is treated analytically. Some results are given for flight performance optimisation. By combining theory and experiment there has been evolved, e.g. the Schwebeleistungs method (power required for horizontal flight) for optimising the duration of indoor model aeroplanes without exact knowledge of their aerodynamic characteristics. Flight characteristics with some elasticity of the structure present formidable problems and are reviewed in some examples. Finally, some problems of hangar meteorology are mentioned.

We are living at a time of transition in civil aviation, between subsonic and supersonic flight. Subsonic flight has no connection with ground transport and supersonic flight, as now envisaged has no connection either. It is thus rather necessary to explain the title of this paper which would otherwise be irrelevant.

The recent analytical studies of rocket engine combustion based on propellant vaporisation as the rate controlling parameter, especially those by Spalding, and Priem and Heidmann, have contributed greatly to our appreciation of the mechanism of stable combustion. Critical combustion experiments on a small scale engine have shown fair agreement with vaporisation predictions, but have also revealed the need to understand and improve the mixing process.

In parallel with performance measurements in a combustion system, detailed measurements of the atomisation process including the break-up mechanism and drop-size distribution have been made. However, there are many gaps in our knowledge of the subject and further work remains to be done.

It is thought that the high-speed photographic and gas sampling techniques developed for these investigations provide useful research tools for more detailed studies of the problems.

The paper concerns rocket engines of the type currently employed in ballistic missiles and satellite booster vehicles, outlining some advantages to be gained from a reduction in thrust towards the end of the burning period. Several methods of throttling a rocket engine are then examined.

Low frequency combustion instability which can arise when an engine is operated below its normal thrust level is considered in some detail, including the effect of such instability on engine performance. Means of ensuring stable operation both in the main thrust chamber and the turbine gas generator are described and compared.

The mechanism of heat transfer to the main injector has been studied and the modifications to the injection pattern necessary to maintain this at a safe level are described. Also, changes in the overall heat transfer rate to the thrust chamber walls are discussed with special reference to the formation of an insulating layer which occurs with carbon-bearing propellants.

The discrepancy between the inlet heads of the propellants at high vehicle accelerations is noted and the relative effect of this on engine operating conditions for the throttled and unthrottled cases described.

Finally, the effect of pressure change on the thrust chamber performance is considered with reference to studies made on the Rolls-Royce RZ.2 engine.

In the past, attempts have been made to evaluate injectors for rocket engine combustion chambers by the use of water analogy rigs and model combustion systems that simulated the injection and combustion phase change occurring in the actual engine. To confirm that conditions in the engine were being correctly simulated, a technique was evolved for determining the mixture ratio distribution achieved by the combustion system of a Spectre variable thrust rocket engine. Gas samples extracted from the rocket-efflux were analysed, and the technique has been applied to evaluate the Spectre's standard central mushroom type injector and also a multi-head injector.

Tests have been conducted over a thrust range of 2000 lb to 8000 lb and at oxidant/fuel ratios from 7·5 to 13·0.

In parallel with this external sampling, a probe has been designed and developed for extracting gas samples from selected points across a diameter of the combustion chamber itself. This probe has been successfully operated for several minutes under combustion conditions of 500 p.s.i.a. and 2600°K, without sustaining any damage.

Analysis of the oxidant/fuel ratio pattern within the combustion chamber and in the efflux, at comparable operating conditions, indicates that little change in distribution occurs between these two points of the system. Also, the distribution found with the standard injector was that for which the combustion system was designed. It is demonstrated that loss of performance depends on the degree of non-uniformity of mixing. A 5 per cent loss in performance at full thrust and optimum mixture ratio occurs with the standard injector.

It is now generally recognised that liquid hydrogen offers a considerable increase in vehicle performance over the more conventional fuels, despite its relatively low density.

A design study on an upper stage, which uses a liquid hydrogen/ liquid oxygen chemical propulsion system, revealed a number of problem areas which extend beyond the current levels of experience and knowledge in rocket technology existing in the U.K. This study was made on the third stage of a communication satellite launching vehicle, and a number of the problems high-lighted during the course of the investigation, covering both the propulsion system and the structure, are described and discussed in terms of their importance and their effects on the launcher system. Solutions are suggested where possible—although, lacking practical confirmation, these must be tentative in the circumstances.

While no fundamental difficulties are anticipated in the development of a rocket engine using hydrogen, it is essential to accumulate some background of data and experience to ensure that early policy and designs are sound; the importance of beginning practical work as soon as possible is emphasised.

The equation for the buckling of a homogeneous elastic plate is well known

For a simply-supported rectangular plate without shear (Nxy=0), the critical loads are usually found by substituting into equation (1) and determining the loads at which nontrivial solutions exist. In the presence of shear, however, the difficulty with the above approach is that the shear term would involve cosines instead of sines.

In the stability analysis of cylinders under external loading, the axial compression and lateral pressure cases are relatively well established: see, for example, ref. 1. However, from a design point of view, a biaxial system of forces due to a combination of axial compression and external pressure is often encountered in launch vehicle structures. While many other combined loading cases have appeared in the literature, the case under present consideration has not; therefore, this note is devoted to a general treatment of this problem. It is to be noted that Radhakrishnan presented some specific results for this loading combination for elastic and plastic buckling.

Using the Donnell equation for small deformations, the present report considers the effect of various compressive loading combinations on the stability problem of an un-stiffened circular cylinder.

The solution of some non-self-adjoint eigensystems describing the elastic instability of structural elements requires essentially the determination of the minimum value of a parameter associated with the eigenvalues. Usually, a substantial programme of computation is implicit in this type of system.

During a preliminary consideration of possible methods of practical computation for a certain problem of this kind, which is currently being investigated by the author, the question arose concerning the suitability of a finite-difference approximation for the purpose. The object of such an approximation was to provide a relatively rapid, if crude, means for a first estimate of the solution by desk-calculation methods.

The following is an attempt to show the gains and losses achieved when a rotating cylinder is used to create additional lift during the take-off and landing of an aircraft.

A hundred-year old theory lies behind the results (the “Magnus” effect), which, up to the present day has not been put to practical use in flight for increasing CL. Many proposals and tentative designs have been studied at major Research Establishments, notably A.V.A. Gottingen and the RAE but experimental studies on aircraft converted for the purpose have yet to be made.

Limited test results show that the CL on a rotating cylinder can be as great as 15 providing that sufficient care is taken to approach infinite aspect ratio by means of end-plates on a high aspect ratio cylinder. The one great disadvantage with the cylinder is the high drag coefficients produced. However, there seems no reason why a fairing should not be placed in the wake to reduce the drag to that corresponding to a blunt-nosed thick aerofoil section.