I am very conscious of the privilege and honour of being asked to deliver the first R. J. Mitchell Memorial Lecture. The presence of such a distinguished audience, and the fact that I feel the subject is one to which it is difficult to do full justice, are factors which do not tend to fill me with confidence, and I hope you will forgive any inadequacies. From a mass of recollections arising as the result of fourteen years of contact with Mitchell, I hope to give you my impressions of the life and career of that great designer.

A theory of primary failure of compression panels of integral construction with unflanged stiffeners is presented, involving rotation, or translation, or rotation and translation of the stiffeners with the corresponding distortion of the sheet without deformation of the cross sections in their own planes. The investigation shows that the failure may occur generally either by pure torsion of the stiffener with the associated lateral distortion of the sheet or by pure flexure in accordance with whichever mode yields a smaller stress. The theory is compared with the test results covering a fairly wide range of the dimensions involved. The theory is in complete agreement with the experiments. The accuracy of the result thus obtained indicates that the fillets at the junction of the skin and the stiffener are not important and their effects may safely be disregarded.

A simple method of determining local turbulent skin friction on a smooth surface has been developed which utilises a round pitot tube resting on the surface. Assuming the existence of a region near the surface in which conditions are functions only of the skin friction, the relevant physical constants of the fluid and a suitable length, a universal non-dimensional relation is obtained for the difference between the total pressure recorded by the tube and the static pressure at the wall, in terms of the skin friction. This relation, on this assumption, is independent of the pressure gradient. The truth and form of the relation were first established, to a considerable degree of accuracy, in a pipe using four geometrically similar round pitot tubes—the diameter being taken as representative length. These four pitot tubes were then used to determine the local skin friction coefficient at three stations on a wind tunnel wall, under varying conditions of pressure gradient. At each station, within the limits of experimental accuracy, the deduced skin friction coefficient was found to be the same for each pitot tube, thus confirming the basic assumption and leaving little doubt as to the correctness of the skin friction so found. Pitot traverses were then made in the pipe and in the boundary layer on the wind tunnel wall. The results were plotted in two non-dimensional forms on the basis already suggested and they fell close together in a region whose outer limit represented the breakdown of the basic assumption, but close to the wall the results spread out, due to the unknown displacement of the effective centre of a pitot tube near a wall. This again provides further evidence of the existence of a region of local dynamical similarity and of the correctness of the skin friction deduced from measurements with round pitot tubes on the wind tunnel wall. The extent of the region in which the local dynamical similarity may be expected to hold appears to vary from about 1/5 to 1/20 of the boundary-layer thickness for conditions remote from, and close to, separation respectively.

The supply of meteorological information for Comet operations began long before the first prototype was in the air, since information on aerodrome temperature and pressure and ambient air temperature up to 40,000 ft. was required to assess engine performance. Thus it was found that extremes of temperature at 40,000 ft. range from about – 80 °C. over the United Kingdom to – 35°C. over Northwest India and performance curves were constructed to cover this range.

Statistics on low temperatures were required in connection with the performance of the de-icing system on descent; statistics on high temperatures were required for engine performance throughout flight.

In recent years interest has been revived in an apparatus which was first developed more than fifty years ago. This apparatus, which is known today as a shock tube, is a very useful research tool, the versatility of which is not generally appreciated in this country. In this paper the author attempts to give a clear physical description of the flow in a shock tube and a review of the problems which may be investigated by means of this apparatus.

Basically, a shock tube consists of a simple fabricated duct which may be closed or open at one end, and closed at the other end. A diaphragm divides this duct into two compartments which initially contain gases at different pressures. In a conventional shock tube the cross-section is constant, the ends are closed, and air is used on both sides of the diaphragm (Fig. 1).

An experimental determination of the relation between the rate of sink of a sailplane and its air speed is of interest both to the designer and to the pilot. This paper describes tests made to establish this relation for the Slingsby Sky. During the tests, the aircraft was flown in a series of “ partial glides ” at constant speed, while a camera looking over the pilot's shoulder photographed the altimeter and A.S.I., together with a slip indicator and a clock. From an analysis of the photographic recordings, the mean air speed and the rate of sink during each partial glide can be determined; a large number of such results enables the relation between air speed and rate of sink to be established. The final results show that a considerable improvement in high-speed performance can be achieved if the aircraft is fitted with a long skid instead of the short skid and large fixed landing wheel; the superiority of the Sky over the Weihe for competition work is demonstrated. Sources of error in the results are discussed; it is shown that the accuracy with which the rate of sink may be determined is limited by the presence of vertical air movements. Possible future trends in performance testing technique are discussed.

The heading aerodynamic noise includes all cases where sound is generated by fluid flows. It includes aeolian harps, edge tones and related phenomena as well as vortex, boundary layer, wake and jet noise. In the strict sense it does not include propeller noise and the direct noise generated by; bodies in motion. However, in view of their importance as sources of aircraft noise some of their more important aspects will be discussed briefly and compared with the similar properties of pure aerodynamic noise.

In general it is audible aerodynamic noise which is of major interest at present, but it would be unwise to neglect entirely the ultrasonic aspects. These have received little attention as yet.

The determination of shear distributions for multi-cell sections can conveniently be considered as two separate problems.

1. (i) the shear flow arising from a pure torsional couple applied in the plane of the section, and

2. (ii) the shear flow arising from a resultant shear force acting on the section so as to produce no resultant twist per unit length of that section, i.e. through the shear centre of that section.

Both problems involve finding the value of one unknown shear flow for each cell, so that in multi-cell construction a set of simultaneous equations is obtained, which must be solved to obtain the shear pattern.

The solution of problem (i) by successive numerical approximations has been dealt with completely by Benscoter. The present paper gives a method for problem (ii).

The usefulness of the method becomes more apparent with increasing numbers of spar webs as may be envisaged for future wing sections of small thickness chord ratios. It is analogous in the procedure used to the Hardy Cross moment distribution method for continuous beams, and has similar advantages. An approximate solution of any required degree of accuracy is quickly obtainable at the design stage and the method is readily adaptable for use by a computer with no knowledge of the structural or mathematical theory involved.

Looking back to the early days of aircraft hydraulic systems, a steady upward trend in the operating pressures used can be observed.

One of the reasons for this increase is weight saving, which has become of particular importance in recent years on account of the increased complexity and power demand of the modern hydraulic systems.

The trend in the direction of higher pressures has been necessarily gradual because of the lack of knowledge and experience in this field, and prudence has precluded the adoption of large steps.

Another step has been achieved in Great Britain by successfully raising the working pressure to 4,000 lb./in.2 on several modern aircraft.

Complicated oscillatory systems may be broken down into component “ sub-systems ” for the purpose of vibration analysis. These will generally submit more readily to analytical treatment. After an introduction to the concept of receptance, the principles underlying this form of analysis are reviewed.

The dynamical properties of simple systems (in the form of their receptances) may be tabulated. By this means the properties of a complicated system may be found by first analysing it into convenient sub-systems and then extracting the properties of the latter from a suitable table. A catalogue of this sort is given for the particular case of conservative torsional systems with finite freedom.

The properties of the composite system which may be readily found in this way are (i) its receptances and (ii) its frequency equation. Tables are given of expressions for these in terms of the receptances of the component sub-systems. All of the tables may easily be extended. The tabulated receptances may also be used for determining relative displacements during free vibration in any principal mode.

A method of presenting information on the vibration characteristics of machinery, which is effectively due to Carter, is illustrated by means of an example. More general adoption by manufacturers of this method (which requires no more computational effort than must normally be made) would lead to enormous savings of labour in calculating natural frequencies of composite systems.

The presence of foam or unduly aerated oil in the lubricating-oil systems of aircraft power plant can be very troublesome—more so in reciprocating engines than in gas turbines, since the rates of circulation are greater and the viscosity grade of the oil is higher. Consequently in this paper, which surveys present knowledge about oil foaming, reciprocating engines are the main consideration.

In the conventional aircraft “dry-sump” oil system, a simplified diagram of which is shown in Fig. 1, oil is pumped from a supply tank to the bearings, and so on, by a so-called “pressure” pump. This oil then drains down into the sump and is picked up by “scavenge” pumps which return it to the supply tank. There are usually at least two scavenge pumps so that the oil is cleared away irrespective of the attitude of the aircraft.

The performance of aero gas turbine combustion chambers is discussed under the following headings : Combustion efficiency, combustion stability, ease of ignition, deposits, exhaust temperature variation, and smooth combustion. It is shown that, as assessed by these criteria, combustion chamber performance can be significantly affected by fuel characteristics; also that the effects of fuel type can be greatly modified by equipment design changes. The conclusion is that most of the problems- aggravated by fuel characteristics are better solved by modifications to equipment, if fuel availability and cost are not to be adversely affected.

The 890th Lecture to be given before the Royal Aeronautical Society was held at the Institution of Mechanical Engineers, Storey's Gate, London, S.W.I, on 12th November 1953, and was presided over by Sir William Farren, C.B., M.B.E., M.A., F.R.S., F.R.Ae.S., President of the Society. Introducing the lecturer, Sir William said that Mr. Wimpenny had been with the de Havilland Company throughout his career in aviation and had been responsible for some years for the design and development of stability and control of the de Havilland aircraft, including the Comet, so that clearly he spoke with a wide knowledge of his subject.

A method is developed enabling rapid determination of the optimum cross-sectional dimensions of compression surfaces having unflanged integral stiffeners, and consideration is given to the effects of practical limitations on the design. The theoretical efficiency of the optimum integral design is found to be only 85 per cent, of that of optimum Z-stringer design.

A simple cumulative fatigue damage law is used to estimate the endurance of a number of typical structural components when subjected to the alternating wing loads encountered by an aircraft in flight. A method is developed to study the fatigue damage in terms of infinitesimally small intervals of alternating load and the corresponding fatigue strength. Recent gust data are used in conjunction with the complete fatigue strength diagrams of 13 simple components. The fatigue damage is studied for the ranges of altitude (i) up to 12,000 ft., and (ii) above 30,000 ft. In each case it is found that the damage rate curve, which gives the intensity of fatigue damage at any gust velocity, takes a characteristic form, with a well-defined range of gust velocities giving the greatest fatigue damage. This leads to a simplification of the whole problem, and it is found, finally, that the life of the structure is governed—in so far as the cumulative law is correct—almost entirely by that gust velocity, which, when applied alone to the structure, gives an endurance of about two million cycles. The results of the survey are compared with the design criteria suggested by Walker and the lives estimated by Williams.

This report presents an abbreviated summary of theoretical studies made by a group at the Research Laboratory of Electronics, Massachusetts Institute of Technology, under contract with the Air Navigation Development Board of the Department of Commerce. The details of the analysis and the results have been treated more comprehensively in two technical reports of the Research Laboratory of Electronics.

It is currently believed in some parts of the world that the problem of structural fatigue in the case of aircraft having a modern multiple spar wing structure can be safely dealt with by allowing fatigue cracks to develop in service. The typical multiple spar wing has its bending tension area distributed between the “ caps “ of two or more spars and a large number of independent heavy stringers. It is believed in some quarters that complete failure in flight of any one spar boom or stringer due to fatigue will leave the wing with sufficient strength for the aircraft to land safely. This note does not discuss the validity of this belief, nor the level of safety so achieved, but attempts an objective assessment of a problem that arises as a consequence of this belief.

The 899th Lecture to be given before the Royal Aeronautical Society was held on 11th February 1954 at the Institution of Mechanical Engineers, Storey's Gate, London, S.W.I. Mr. G. R. Edwards, C.B.E., F.R.Ae.S., Vice-President of the Society, presided. Introducing the Lecturer, Dr. D. Williams, M.I.Mech.E., F.R.Ae.S., Mr. Edwards said that Dr. Williams was really too well known to need any introduction; his activities had been devoted in recent years to structural problems as applied to aeroplanes. Without reciting his attainments in detail, the job he had tackled in recent years at the Royal Aircraft Establishment and about which he had written papers, covered such subjects as vibration, wing flutter, sandwich construction, the effect of blast on aeroplanes, energy theorems, fatigue, dynamic loading by gusts and landing impacts, water-borne runways and nose-wheel shimmy, a variety of subjects one or all of which had caused most of them a great deal of anguish at one time or another.

Mathematical and experimental methods for estimating the relative eddy velocity in a radial bladed impeller channel have been determined to replace the empirical velocity so far used in designing and predetermining the performance. The effect of this relative eddy velocity on rotor performance has been experimentally determined on a specially designed compressor and compared with mathematical calculation.