This is a discussion of the various types of flow which may occur on thin swept wings when the air stream is not attached to the wing surface everywhere. In particular, the behaviour of vortex sheets which arise from such mixed flows is examined.

Air Traffic Control is an all-embracing term and on occasion there is some doubt about its scope. For the present purpose it will be as defined within the internationally agreed principles set out in Annex 11 to the International Civil Aviation Convention. Here it is set out that the Air Traffic Control service is for:–

1. (i) Expediting and maintaining an orderly flow of air traffic.

2. (ii) Preventing collision between aircraft.

3. (iii) Providing advice and information useful for the sake of efficient conduct of flights.

4. (iv) Preventing collision on the manoeuvring area between aircraft and obstructions.

5. (v) Notifying the appropriate organisations regarding aircraft in need of search and rescue and to assist such organisations as required.

There has been considerable activity of recent years, particularly in America, in the field of integral structures for aircraft.

The fundamental problem in aircraft structures is to make a relatively thin skin covering take applied compression, shear and bending loads. This has been achieved so far by the attachment of stiffeners, but this requires a large amount of riveting which in itself causes structural inefficiency and also spoils the smoothness of the skin, particularly when laminar flow is desired. Metal gluing shows great advantages over riveting in this respect but it is doubtful whether the bond strength will be great enough to attach very heavy stringers to very thick sheet. Where structural loadings become higher the required thickness of sheet and stringers may make the continued use of riveting difficult, if not impossible.

In opening this discussion on fatigue it seems that my first duty should be to state what fatigue is – I wish I knew. Just as Alice knew that if you drank much from a bottle marked “Poison” it was pretty sure to disagree with you sooner or later, so do we know that if a component be subjected too often to too high a load it is pretty certain to develop fatigue cracks sooner or later.

It is important to remember that Lewis Carroll was by profession a mathematician, because in a mathematical sense even these heavily conditioned statements should in fact convey a great deal. “Too often to too high a load” infers range of loading, and furthermore it implies a connection between range of load S and cycles N, by which in general N may be expected to increase as S decreases. “Sooner or later” indicates that N is not a unique function of S, and “pretty certain” provides for those pieces which, like certain batteries, obstinately go on long after the others have stopped. Then, lastly “to develop fatigue cracks” is a careful avoidance of a direct reference to failure.

In recent years, aerodynamics and thermodynamics have found common ground in the specialised field of gas dynamics. Developments in this subject have led to a much more complete and widespread knowledge of subsonic, sonic and supersonic flow of gases in the conventional type of convergent-divergent nozzle. When the back pressure is raised above the value against which the nozzle is designed to discharge, oblique and then normal shock waves are set up in the divergent cone at a position along the nozzle axis determined by the magnitude of that back pressure. The gas which has crossed the shock wave is subjected to a process of subsonic compression.

In this paper a theoretical investigation is made of the changes in pressure, temperature, density and Mach number which occur across a normal shock wave, when the position of the wave varies along the nozzle axis. The investigation illustrates the effect of change of medium, for which the relevant property is the ratio of specific heats. This ratio for certain polyatomic gases may approach unity (e.g. for Dichlorodifluoromethane CCl2F2, in gaseous form, γ=1·06), and for the inert monatomic gases γ= 1·667. The analysis is made non-dimensional by expressing such quantities as gas pressure, temperature, and density at any given position along the nozzle axis relative to the values of the particular parameter at entry to the convergent section.

The term “fatigue” is given to the progressive failure of a metal under the repeated application of a cycle of stress. Although investigators have been working for more than a hundred years to determine the mechanism by which fatigue failure occurs, the problem is still not fully understood. The simple slip theories of Ewing and Rosenhain in 1899, Ewing and Humphrey in 1903 and of Gough and his collaborators from 1923 to 1933 appear now to be only partially true, according to some of the modern theories. These suggest that other factors such as polygonisation, deformation bands and crystallite formations, all play an important part.

After two years research work in the Universities of this country and elsewhere, there is need for a comprehensive review of the present position on the suppression of aerodynamic noise from jet engines. This report outlines the fundamental understanding of the problem that has been achieved, the basic experimental work done, and the noise suppression techniques which have been suggested.

A study is made of measured noise levels on advanced types of engines and an analysis is made of reductions required and so far achieved with the proposed noise suppression devices. While the basic principles underlying noise formation, both in subsonic and supersonic jet streams are beginning to be understood, the noise reductions so far achieved are still insufficient and warrant an extended programme of research.

The report includes helicopter noise analyses with pressure and pulse jets, and touches on the problem of structural fatigue in the vicinity of a jet stream.

In Rocketry, the definition of Specific Impulse is given variously as Thrust per unit mass of propellants consumed per second and Thrust per unit weight of propellants consumed per second. The second definition is convenient in giving units in seconds, but is unsatisfactory in that the weight of a given mass of propellants depends upon gravitational attraction, and decreases with altitude.

The practice of removing aircraft items for overhaul at specified periods has gradually extended in recent years, so that most types of equipment and instruments have come to be dealt with in this way. The period between overhauls is often called–somewhat inappropriately–the life of the item and, for the sake of brevity, this is the word used in this paper.

The primary object of this procedure is to remove the items for overhaul before defects occur. If this is achieved, two benefits accrue:–

1. (a) the defect rate is minimised, with consequent improvement to safety.

2. (b) the operator does the work at convenient times at the overhaul base instead of being driven to do work when and where defects occur.

The purpose of this paper is to illustrate the research and development work necessary to improve existing products and to anticipate demands for new products facilitating the advance in efficiency and performance of aircraft. The experience of Thornton Aero Engine Research Laboratory, in which the author has worked since it came into operation in 1941, has been drawn on almost entirely.

The helicopter, because of its complicated nature, provides the mathematician with a delightful set of problems on which to exercise his ingenuity, but the purpose of this paper is to interpret the functioning of the helicopter in terms of its physical aspects.

The paper is divided into three main parts: —

1. 1. General Performance

2. 2. Control and Stability

3. 3. Rotor Vibration Problems.

This note considers two cases of bending beyond the limit of proportionality which have not received a great measure of attention in the past. These are Case (i). A beam subjected to a pure bending moment acting in a plane other than one of symmetry. Case (ii). A beam subjected to a bending moment and an axial load.

A mathematical form is used for the stress–strain curve in order that the results shall be applicable to a large number of materials and to avoid the tedious arithmetical summations which would be necessary if actual stress-strain curves were used. All the results are presented in terms of form factors since this notation is convenient for design purposes and is consistent with current British practice. Although the method adopted in the analysis is applicable to a number of different types of beam cross section, only rectangular cross sections are considered in detail at the present time.

Curves are presented showing the variation of the form factor with respect to the angle of the plane of loading, for Case (i), and with respect to the ratio of bending moment to axial load for Case (ii). These curves are based upon a standardised form of the stress–strain curve which may be taken as representative of a range of aircraft materials. Consideration is given to the automatic satisfaction of present military proof loading requirements when these form factors are used for the estimation of maximum permissible bending moments and axial loads under ultimate loading conditions.

The recommendations of F. P. Cozzone, for dealing with these two problems are examined and comparison is made between the theoretical work of this note and a limited number of test results available for Case (i).

Exactly what constitutes a design criterion is a matter on which there may well be differences of opinion. For practical purposes design criteria may be taken to be empirical numerical rules to aid the designer in forming his initial conception of a design. The formulation of such rules for aircraft fatigue is still a matter of some difficulty. Not only is the subject of aircraft structural fatigue still in its infancy, but its study has to be linked with systematic experiment in a somewhat unusual way.

In the criteria described in this paper, the structure has first to be subjected to a standardised laboratory test. This test is itself a gross simplification since it has to represent in a rough way a diversity of conditions occurring in flight.

The pioneer rarely receives that meed of praise which is his due, until those days arrive when his pioneering has become a page or so of history. What meed of praise he may then receive is too often made savourless by critics unable to disassociate themselves from the knowledge of their own technical era. The halting speed of pioneering aerodynamic research cannot be judged by the supersonic speed of secret splittings of the atom.

It was R. L. Stevenson who said that to travel hopefully is a better thing than to arrive. In aviation we have arrived at that streamline flow of technical jargon which deludes so many they have at last reached the boundary layer of aeronautical knowledge, so that they find experiments with plates always a little flat, although their curiosity may still be aroused by flying saucers.

The rapid development of the jet engine has been responsible for a complete change of outlook in air transport. The purpose of this paper is to illustrate some of the problems and to indicate ways in which they have been solved. The subjects discussed are those of temperature control of

1. (i) Pressure cabins of aircraft

2. (ii) Engine intakes (for de-icing) and

3. (iii) Aircraft wing surfaces (for de-icing).

The normal control of engine jet pipe temperature has not been included because, until recently, it has been fully covered by the engine companies.

Jet pipe temperature control is a specialised application involving the correction of fuel flow for variations of air flow which may be due to changes in either compressor or turbine characteristics at altitude, or even the incorrect sampling of jet pipe temperature at altitude. The methods of temperature control discussed in this paper are often applicable also to engine jet pipe temperature control.

Experiments were made on an aerofoil of 25 per cent, thickness with a suction slot at the rear and a flap that could be withdrawn into the slot. The main object of the work was to investigate the effect of withdrawing the flap when all the air in the boundary layer was sucked into the slot. In particular, as a test of the type of anti-gust device proposed by Thwaites, it was required to observe the change of lift that occurred during withdrawal of the flap and a subsequent change of incidence (with the flap still withdrawn).

The flow was investigated by the use of smoke and tufts, by pressure plotting on the surface of the aerofoil, and by measurements of total and static pressures in the wake. The effects of varying incidence, flap position, and suction velocity were investigated.

It was found that the provision of a flap downstream of the enclosing streamline, so that the rear stagnation point was not in the free fluid, was all important in relation to the behaviour of an aerofoil which had a suction slot at its rounded rear.

With the flap extended beyond the enclosing streamline and with the ratio of suction-slot velocity to free stream velocity higher than the theoretical value, a stable flow was obtained within the theoretical incidence range. A favourable pressure gradient was observed in the slot entry and no measureable wake existed downstream of the flap. Smoke filaments, emitted from a tube projecting from the slot slightly beyond the enclosing streamline, travelled far downstream with apparently little diffusion.

When the flap was placed so that its trailing edge was inside the enclosing streamline, i.e. so that there was in theory a stagnation point in the free fluid, an instability of a three-dimensional character occurred, with a wide wake and a large drag coefficient. The lift curves were unusual and it was found that movement of the flap caused a large change of lift coefficient. There was no evidence that the lift could be maintained constant, independent of incidence, when the flap was withdrawn into the slot.

In 1949 F. H. Pollicutt, in the course of a discussion on a paper presented by P. B. Walker to the Royal Aeronautical Society, reported fatigue test results which indicated that when the aluminium alloy conforming to specification D.T.D.364 was tested, under , axial loads, in the form of Z-sections, its fatigue properties were extremely low. Z-sections with unmachined surfaces and even Z-sections with completely machined surfaces failed after short endurances at stress ranges under which small polished round test pieces would not have failed even after hundreds of millions of cycles. Pollicutt, in discussing these low values, advised that great caution should be exercised in their interpretation, warning equally against an unqualified acceptance of the testing technique and a too facile acceptance of small polished round bar test results.

Pressure Distribution and liquid-film tests on a 10 per cent, thick aerofoil revealed the presence of separation “bubbles” close to the leading edge. These bubbles are formed beneath the boundary layer which separates near the leading edge and re-attaches farther aft; their existence is usually indicated by localised constant-pressure regions in the pressure distributions. It is also believed that if a thin film of liquid (such as a suspension of lamp-black in paraffin) is spread on the surface, the scrubbing action of the air rotating in the bubble will tend to draw liquid forward into the bubble, and hence the location and extent of the bubble may be indicated approximately by the accumulation of the fluid.

Many boundary layer traverses of bubbles on N.A.C.A. aerofoils have been made, but it was felt that similar measurements of the bubbles on this particular aerofoil would provide useful data, since the separation characteristics of this section appeared to differ from those in the N.A.C.A. tests.

The noise problem associated with an aircraft flying at supersonic speeds is shown to depend primarily on the shock wave pattern formed by the aircraft. The noise intensity received by a ground observer from a supersonic aircraft flying at high as well as low altitudes, is shown to be high, although it is of a transient nature.

A study of the shock wave patterns around an aircraft in accelerated and retarded flight is shown to lead to an explanation of the one or more booms, of short duration, heard by ground observers after an aircraft has dived at supersonic speeds.

The shock wave patterns associated with an aircraft flying in accelerated or retarded flight at transonic speeds are shown in certain cases to be very different from the corresponding patterns observed in steady flight. The significance of these results, with reference to problems of flight at supersonic speeds is briefly discussed.