The 1,142nd lecture to be given before the Society and the 37th Main lecture to be held at a Branch Centre "Research at the College of Aeronautics" by Professor A. J. Murphy, M.Sc., F.R.Ae.S., F.I.M., Principal, College of Aeronautics, was given under the auspices of the Cambridge Branch on 21st January 1960. Professor W. A. Mair, M.A., F.R.Ae.S., Francis Mond Professor of Aeronautical Engineering, University of Cambridge, and President of the Cambridge Branch, opened the proceedings and then handed over to the President of the Society, Mr. Peter G. Masefield, M.A., F.R.Ae.S., Hon.F.I.A.S., M.Inst.T.

Mr. Masefield: As an old Cambridge engineering man it was a great pleasure to be President of the Society in the year in which the first Main lecture was held at Cambridge.

One of the most important and significant developments in aeronautical education since the war had been the founding, and the progress in recent years under Professor Murphy, of the College of Aeronautics at Cranfield. It had been founded just after 1944 as a result of the Fedden Committee of that year, set up by Sir Stafford Cripps, then Minister of Aircraft Production. Cranfield was primarily concerned with two things: post-graduate training in all branches of aeronautics and associated sponsored research by both staff and students of the College. Professor Murphy was to speak on the latter part.

British aeronautics today was passing through one of those periodic upheavals which seemed to happen every few years in this great business in which so many of them were engaged. Some of the changes now in progress he thought could mean a great difference in British aviation in the 50 years to come, compared with the 50 years that had gone by, and many of them he hoped would be concerned in these changes and in the new era of British aviation which was just beginning. He hoped that 1960 would mark the beginnings of a new chapter in aviation history as 1910 had done fifty years ago.

Over the past 50 years about 85 per cent of work in aeronautics in Great Britain had been in the military field; certainly 85 per cent of industrial effort. In future, manned military aircraft would be only a small proportion of the effort. The rest would be in missiles, civil aeronautics and space activities. In all of these they would be working increasingly competitively internationally and particularly with America and Russia. To make real impact their technology must be the highest possible. There was no doubt that the two foundation stones on which everything else had to rest were firstly, sound basic technical education and secondly, first-class research effort. At Cambridge, and 40 miles away at Cranfield, both those aims were being pursued. He could not over-emphasise the importance of these two subjects. The Society was closely concerned in both education and research and, therefore, it was an especial pleasure to introduce Professor Murphy.

Since October 1955 Professor Murphy had been Principal of the College of Aeronautics. Graduating from Manchester University in 1920 with First Class Honours in Chemistry, he had first done some research at University College, Swansea, before going to the National Physical Laboratory Metallurgy Department. In 1931, he had joined J. Stone and Company Ltd. as Chief Metallurgist and spent some years in Industry as a Director of that company and its associates. In 1950 he became Professor of Industrial Metallurgy at Birmingham University and, three years later, Director of the Departments of Physical and Industrial Metallurgy, a post he held until his appointment in 1955 as Principal of the College of Aeronautics. Professor Murphy had been President of the Institute of Metals and of the Institution of Metallurgists.

This paper is a brief summary of an investigation made to check the effect of lift on the shock noise of aircraft flying at supersonic speeds. The method of Hayes has been combined with the theory of Whitham to predict the asymptotic shock strength of wings carrying lift and of combinations of bodies and lifting wings. (A similar, but not quite as general, method was derived by Walkden in Ref. 6.) Whitham's formula, including only the volume term, has been used extensively to predict the boom intensity of aeroplane type bodies and the agreement with experiment has, so far, been quite reasonable. The test data obtained to date extends only up to about 40,000 ft. altitude and the calculations of this paper show that under those conditions the shock noise of the aircraft tested so far will, in most cases, be dominated by the volume term. It is shown that at higher altitudes lifting effects will dominate for even the small fighter and they will dominate over most of the altitude range for large bomber and supersonic transport aircraft. The boom intensity due to lift decreases with altitude as h–¾ which compares to pa½ h–¾ in the volume case (pa = pressure at altitude h). It is insensitive to Mach number (for M > 1·4), wing loading, wing plan shape and lift distribution. A simple rule for calculating the shock noise due to combined volume and lifting effects is proposed which is applicable to configurations with wings located towards the rear of the fuselage. The rule states that the shock noise of an aircraft carrying lift is equal to the shock noise due to volume (neglecting lift) or the shock noise due to lift (neglecting volume), whichever is the greater. A chart is presented from which rapid estimates can be made of the shock noise of lifting wing-body combinations.

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.

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.

The concept of small stage efficiency is introduced when studying one-dimensional gas flow in nozzles in order to permit a closer approximation of real flow conditions than is possible from an isentropic analysis. It is more or less conventional to assume the flow conditions are adiabatic whenever the small stage efficiency is used. That is to say, small stage efficiency is generally considered in relation to flows contained within adiabatic boundaries, in which case it becomes a measure of the heat generated by internal frictional effects alone.

Expressions for the secondary circulation behind a cascade are derived in Ref. 1. A simple derivation of the same results is given in Ref. 2. The secondary flows within the blade passages and their extensions downstream are determined by the “secondary passage circulation”, Г, given by

for a diffusing cascade, and

for an accelerating cascade,

where Ul inlet velocity to cascade

z distance in spanwise direction

s blade pitch

αl, inlet flow angle

α2 outlet flow angle

Ll, distance along lower side of aerofoil

L2 distance along upper side of aerofoil

q1 velocity along lower side of aerofoil

qu velocity along upper side of aerofoil.