There are three practical aspects to the problem of man powered flight: aerodynamics, physiology and engineering. There are many more impracticable aspects to man powered flight, including the philosophical, the fanatical and the sceptical, which unfortunately have not yet been included in the lectures of the Man Powered Aircraft Group but which will no doubt find their places.

The engineering aspects are fortunately fairly clearly defined, although there is likely to be a slight threat of interference with the aerodynamic problems.

The honour that has been accorded me in being invited to deliver the Second Halford Lecture is very much appreciated. My personal acquaintance with Major Halford began about fifteen years ago, and every step since has been one of increasing respect and esteem. It is a privilege and pleasure to pay tribute to the memory of a man whom I recall as an effective crusader for simplicity and reliability in engine design. In the First Lecture, John Brodie discussed very fully the contribution of Major Halford to aircraft power plants, and little is left for subsequent lectures on this subject. I have therefore chosen to speak on the propulsion problems of high-speed aircraft—a field of primary interest to Halford, as demonstrated by the outstanding success of the Goblin and Ghost engines, and more recently by the Gyron series designed for supersonic propulsion.

As a native South Australian, I feel particularly honoured to have been asked by the Adelaide Branch of the Royal Aeronautical Society to deliver this second lecture in the Memorial Series commemorating the historic flight of the Smith Brothers from the United Kingdom to Australia. Just a few months ago—on 10th December, 1959—we celebrated the fortieth anniversary of the epic arrival of the Vickers Vimy at Darwin. One thing emerges clearly from our aviation record in the succeeding forty years, and it is simply that no single event has had such a profound effect on that record as the remarkable, trail-blazing flight of the Smith Brothers. If we came of age as a nation on the cold dawn beaches of Gallipoli on 25th April, 1915, then just as surely we won our wings as a nation on that hot summer afternoon of 10th December, 1919, when the Vickers Vimy touched down at Fanny Bay aerodrome, Darwin. We have, I think, continued to wear those wings with considerable distinction ever since.

The successful development of a helicopter and its introduction into service may be influenced by several factors but a well co-ordinated development programme is by far the most important. A practical and realistic balance between representative ground tests and the subsequent flight proving is essential if the project is to be completed within a reasonable budget and time scale. The two parts of this paper cover these phases and aim to show how they are related in a particular development programme. The use of major fully representative ground rigs is discussed in relation to pre-flight functioning and fatigue life assessment; this is followed by the initial flight trials and finally the structural and performance clearance of the helicopter over its full design envelope. Although the tests and techniques described are generally applicable, the examples quoted relate to the Bristol Type 192, tandem rotor helicopter. The opinions expressed are personal and do not necessarily represent those of the Bristol Aircraft Limited.

The Short SC.l is a small delta wing V.T.O. aircraft of about 7,800 lb. all-up weight. Its unusual feature is that it is provided with vertical thrusting engines to enable it to take-off and land vertically and to hover. It can also operate normally as a conventional aircraft.

The nature of the aircraft introduces a number of problems in the design of the landing gear. The first of these is that all wheels must be fully castoring since a vertical landing may take place with a certain amount of drift in any direction. Another complication arises because of the possibility of landing with slight backward drift, it being necessary to move the main wheels aft when the pilot intends to make a vertical landing; if the wheels were left in the required rearward position for conventional take-off the elevator power would be inadequate to lift the nose of the aircraft. The landing gear is therefore provided with simple means for moving the wheels fore and aft by a small amount, to move the position of the main wheels in relation to the centre of gravity.

The subject matter of this paper is primarily an account of operational experience over a comparatively short period resulting from helicopter activities at the Westland London Heliport. At the outset it is appropriate to recall publication by the Helicopter Association of Great Britain, in January, 1958, of a widely circulated report entitled “Examination of the Case for a London Heliport”. There in was expressed the opinion that there existed an immediate need for an operating base in the inner metropolitan area for use by single-engined helicopters. Further, that it was mistaken policy for the then Ministry of Transport and Civil Aviation to relate the need for a London heliport to the future advent of multi-engined helicopters. In the meantime the initiative had been taken by Westland Aircraft as a private venture to provide a suitable site, and to accept financial responsibility for all procedural and constructional work associated with its development and operation.

With cruciform missiles it is usual to express overall aerodynamic forces and moments in terms of aileron, elevator and rudder angles. It is shown that an alternative formulation in terms of individual control angles is both more general and more revealing.

The fundamental assumption is that the overall coefficients can be expressed sufficiently accurately as the sum of independent control effects and of the mutual interference between adjacent controls. This assumption has been found to be valid over quite wide regions of supersonic Mach number, body incidence and control deflection for some typical cruciform rear control missiles. In principle the method applies at all speeds and to all types of layout, but it seems likely to be less widely applicable at subsonic and transonic speeds and to moving wing and canard layouts. The region of applicability is ultimately a matter of experimental investigation for each missile.

This new approach gives a clearer understanding of cross-coupling effects, simplifies and improves the accuracy of a mathematical model of the aerodynamic characteristics, and facilitates the interpolation of wind tunnel test measurements. These measurements are made on configurations with two adjacent controls deflected and the other two controls either always undeflected or always removed. The test programme can be matched more easily to the aerodynamic characteristics, and the testing time is much the same or significantly less than for a comparable programme involving combinations of aileron, elevator and rudder settings. Greater demands on experimental accuracy are imposed, however.

There is considerable scope for research work in investigating the applicability of the new technique to various types of missile layouts and to assessing the relative merits of having the two dummy controls undeflected or removed.

Before examining the Airline Pilot's views on Flight Safety, a brief look at this Airline Pilot will be taken to get a rough idea of the viewpoint from which he speaks.

The Chairman, in his opening remarks, referred to pioneering. In the early days of aviation the pilot was a pioneer, and as such he had to take risks. At the present time, and in some parts of the world many years ago, the pioneering adventure of flying has grown up into a serious transport industry, and the airline pilot has grown up with it. Today, the airline pilot is an executive of that transport industry; a highly skilled employee or a highly specialised manager depending on the way you look at him.

The first basis for the design of any power-driven mechanism is an exact knowledge of the characteristics and capabilities of the power unit available. Without such information no designer is able to start on the initial sketch of the mechanism. The man powered aircraft is no exception and its designers have suffered in the past by insufficient knowledge about man's power. This may seem remarkable because until about two centuries ago almost all the world's work was done by muscle-power, and much of the muscle was human.

The paper mainly discusses recent R.A.E. and N.P.L. research on the aerodynamics of Direct jet lift (turbo-jet and turbo-fan), Propeller lift (tilt-wing and deflected slipstream), Jet flaps, and Boundary layer control (blowing and suction). Some of the work at Firms and Universities is also mentioned where the information has been made available for general circulation. Attention is drawn to the present state of knowledge in relation to V.T.O.L. and S.T.O.L. aircraft applications and to some major aerodynamic problems which need further study.

There are many problems to be solved when it is required to fly an aircraft at speeds approaching the speed of sound. When the speed of the aeroplane exceeds its critical Mach number, regions of supersonic flow appear at points on its surface, where the velocity of the displaced air is greatest. These regions of supersonic flow increase in size with increasing speed until the flow is wholly supersonic. The range of speed in which this mixed flow occurs is called the transonic range. In this range an aircraft will experience a sudden increase in drag, trim changes, loss of control effectiveness and possibly instability, or buffeting.

The advent of the Skylark firings and the more recent trials of ballistic missiles such as the Black Knight firings, have demanded a continuous development of both ground and airborne instrumentation at Woomera Range; this in turn has raised the need for sighter rockets and instrumentation test vehicles with altitude capabilities in the 200,000-500,000 ft. range.

The need is currently being met by the rocket vehicles Long Tom and Aeolus. These rockets are of interest because they represent the first designed and built at W.R.E. that can be classed as high altitude rockets; also, they make possible the provision of relatively cheap and readily available vehicles for upper atmosphere research.

Two interesting flow phenomena have been observed during the course of some transition measurements carried out on a swept-back wing with a curved tip. Firstly, the influence of leading-edge sweep angle on the type of instability responsible for transition has resulted in two distinct types of transition occurring simultaneously over different portions of the wing under certain critical conditions. Secondly, unlimited spanwise turbulent contamination of the flow was observed with an excrescence located close to the leading edge. The circumstances in which this occurred are believed to be somewhat exceptional, although much further work is desirable on this point.

In ref. 1 Bennett derived the well-known expression for the power required to overcome the profile drag of the rotor blades for a helicopter in forward flight:

The power given by (1) is composed of two parts: the contribution of the torque of the profile drag with respect to the rotor axis and that of the drag of the rotor blades due to the forward motion (H-force).

Several methods for performance calculation of helicopters require the two parts of the profile-drag power separately, as they make use of the equation for the equilibrium of forces in the direction of flight, to which the rotor drag gives a contribution. In this note expressions for the two parts of the total profile-drag power are derived using the same method as indicated in the original work of Bennett. Numerical results are given.

The difficulties of computing a solution to the Prandtl lifting line equation for large values of A/a0 are well-known, at least to any investigators who have tried. In attempting to deduce the loading over straight-tapered wings, like those used on sailplanes, the author has used an iterative method by which the nth. approximation to the downwash is obtained by a numerical integration based on the (n–l)th. approximation to the distribution of circulation. The main difficulty here is essentially one of making a good initial guess of circulation for the first approximation to the downwash.

Using a Stewartson transformation defined by

it is straightforward to show that in the two cases for which it is possible to find similar solutions to the laminar boundary layer equations with Pr≠1, Le=1, that is stagnation flows Ue∝X and flat plate flows Ue=constant, it is not necessary to assume anything about the specific heats to derive the equations given and solved by Beckwith.

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.

At its fifth meeting, the Commonwealth Advisory Aeronautical Research Council discussed the need for a better understanding of the operational requirements for aircraft in different parts of the Commonwealth. It therefore asked for papers from the respective countries, bearing in mind particularly the possibilities of V.T.O.L. and S.T.O.L. aircraft.

A survey has been made in Australia. The general impression gained is that within the Australian continent there are no outstanding future operational needs, if the term outstanding is interpreted in the sense of V.T.O.L. or S.T.O.L. or other extraordinary condition of operation. This is not surprising since Australia has, on the whole, excellent conditions for flying; there are no high mountains, very little icing or fog, and the country tends generally to be flat and low. In external operations conditions are not so favourable and certain definite requirements emerge. Theses together with some general observations, have been recorded as of interest to designers. It is clear that simpler forms of V.T.O.L. aircraft than rotary winged types are needed, and this may well indicate a need for more research.

The first International Agricultural Aviation Conference, held at the College of Aeronautics at Cranfield between the 15th and 18th of September 1959, was well timed to mark one stage in the development of the art of airborne farming—it was the stage when the art ceased to be mainly experimental and became essentially a commercial business.

Intermittently for the past thirty years, in various parts of the world, attempts have been made with varying degrees of economic and practical success to do certain operations connected with farming, forestry or other allied activities. Two main factors within the past decade have served to intensify the interest and activity in the art. One important factor is the general improvement in aircraft, including helicopters, coupled with the very large number of relatively suitable ones which became redundant (and therefore cheap) at the end of the Second World War. The other factor, equal in importance, concerns the development of the science of agricultural chemistry which has given the farmer a new and wide range of fertilisers, selective weed killers and other chemical forms of pest control which are effective in reasonably small bulk.

During a recent open week at the National Gas Turbine Establishment at Pyestock, an experiment involving just over 1,000 subjects was made to establish some measure of the disturbance value of such rhythmic or amplitude modulated noises as are generated by helicopter rotors. The results were (i) that, as may be implied from Niese's results with modulated tones, the subjective assessment of such a noise depends upon its peak sound pressure level (i.e. 82·5 db. in Fig. 3) rather than upon its r.m.s. value, (ii) that the effect of the modulation frequency is negligible in the range from 4 to 12 cycles/sec. and (iii) that these results are invariant with peak sound pressure level in the range from 85 to 95 db. re 2 x 10–4 dynes/cm2.