In a paper delivered before the Society in March 1960, Morgan concluded that a Mach number of about 2.0 may prove a natural “fit” for the first generation of long-range supersonic transports. Sub-sequently, at the second Congress of I.C.A.S., Kuchemann developed the theme of fitting an aircraft to a particular duty and concluded that the slender wing provides a natural basis for the layout of the supersonic transport. It is convenient to accept these two papers as a starting point for a more detailed examination of the aircraft, since, taken together, they can be taken as fixing some of the major variables—range, speed and general layout.

This short paper is confined to a discussion on one aspect of the Structural Design of Supersonic Transports. It is chosen because it covers the area in which the impact of supersonics is probably most critical from the structural point of view, i.e. kinetic heating.

It is of utmost importance that the significance of kinetic heating is not exaggerated. Adequate attention must be paid to assessing the degree to which past experience is relevant if excess weight is to be avoided.

It is in this direction that the discussion is aimed.

In view of the need to keep the paper short, attention is concentrated on the Mach 2 class of aeroplane constructed largely of aluminium alloy. Much of what is said will also have relevance to the Mach 3 class of aeroplane.

With increasing speeds of aircraft, problems arise over materials of construction, partly through kinetic heating effects which cause temperature rises and partly through rain erosion effects. Additionally, account has to be taken of a required life of 30,000 hours which is much longer than that over which most materials are normally evaluated.

For speeds of Mach 2.2, the maximum temperature attained will be of the order of 130°C, while for Mach 2.7 the corresponding temperature will be 200°C or higher. For the lower speed, it is probable that aluminium alloys will be satisfactory, although it must be emphasised that no information is yet available on the long-term behaviour of materials at the elevated temperatures likely to be experienced. Test programmes are in hand, but these will necessarily take considerable time for their completion. For speeds of Mach 2.7 or above, steel or titanium alloys will be required, but here again the necessary evaluation for long periods at high temperature has yet to be completed.

An important point to realise in the question of propulsion for aircraft of Mach 2 and upwards is that efficiency of the power plant is no longer a problem. Due to the high pressure developed from the intake the overall pressure ratio of the engine is high, and the aircraft forward speed being high the Froude propulsive efficiency is also favourable. Fig. 1 shows, in fact, that at Mach 2 overall thermal efficiency as high as 35 per cent is achieved which, of course, compares favourably with the best electrical power station generation. Indeed, it is one of the main reasons why economic transport at high Mach numbers is possible. The falling aerodynamic efficiency, as defined by lift/ drag ratio, is offset by this rising propulsive efficiency, so that a somewhat similar overall performance is obtained as compared with subsonic aircraft.

The problem of determining the type of engine which best suits a particular application is not new to the engine designer, but the supersonic transport is perhaps unique in that a number of design criteria have been introduced which cannot easily be taken account of in the usual terms of engine weight, fuel consumption or even aircraft direct operating cost.

However, an examination of the effect of these standard criteria on different types of engines is a starting point in determining what engine will best match the supersonic transport requirement. A survey of a series of thermodynamic cycles both at Mach numbers in the region of 2.0 and at subsonic speeds typical of standoff or diversion suggests that two types of engine merit further study.

The flight plan of the supersonic transport demands efficient aerodynamics and propulsion over a wide range of speed and altitude. As an example, the typical trans-Atlantic flight shown in Fig. 1 demonstrates that apart from cruising efficiently at, say, M = 2·2 for 2.500 n.m. at altitudes from 50,000 ft. to 65,000 ft., the aircraft must also be efficient at subsonic speeds. This is particularly true on the portion of the flight after the normal stage length, i.e. the diversion and hold, since the fuel required here must be carried the whole way and the “ growth factor” on aircraft size is large.

I would like to preface my remarks by gratefully acknowledging the inspiration I have received from my colleagues and others with whom the problems I shall talk about have been discussed, and by making it clear that, although my views are naturally coloured by the work on which I am engaged, they must not necessarily be taken as the official views of the Air Registration Board.

The supersonic aeroplane, due to its novelty, presents the three main types of airworthiness problem in a more pressing form. These are: knowledge of the environment; knowledge of what the aeroplane must be able to do in that environment (which includes a clear idea of acceptable risk); and obtaining, before passenger service, the necessary confidence that the intended safety level will be achieved. I can only touch on a few of the individual problems.

The rate of error in the performance of any task increases rapidly after a certain degree of complexity or rate of operation is exceeded. As aircraft speeds increase, therefore, the safety and efficiency of the operation will depend largely on the extent to which the job of the crew can be simplified and on the work-load being kept within tolerable limits. In recent years the problem has been tackled by the addition of specialist crew members and by the use of automatic and semi-automatic equipment. Distribution of the total load over an increased number of people is no longer a satisfactory solution, however, since it imposes an additional burden on the Captain in his capacity as overall monitor and the time available for this vital task is all too little. While the pilot by tradition tends to reserve judgment on automatic devices generally, where a high standard of reliability is proved the system will receive acceptance in due course and there is little doubt that it is in this field where the greatest contribution to operating efficiency will be found.

A great deal has been said and written about the necessity to launch into a completely new standard plateau of systems engineering for supersonic, as distinct from subsonic, aircraft. So far as I have been able to gather, one reason for all this could be the argument that, as the supersonic aircraft will be so costly to operate, i.e. so uneconomical, it must be operated at a very high utilisation with very short turn-rounds and that therefore all the systems must be very much more durable than has previously been the case. Another reason could well be that the flight régimes are such that there will have to be much more dependence on certain systems, e.g. pressurisation at 70,000 ft., cabin cooling at very high speed and so on.

The argument for the supersonic transport seems to run as follows: it offers a new and faster way to go places; it might not cost more to do so; it ought to create new wealth, and its influence on human affairs, although not clearly foreseeable, surely will be beneficial. There is a certain amount of evidence to warrant optimism; there are issues of national prestige and well-being; the technical challenge is exciting; so “ Let us somehow find finance and proceed.”

Operators agree that the design and operational problems of supersonic aircraft, one of the possible great advances in transport, are complex and interesting; they might also speculate on its likely social implications, though this is an unrewarding pastime. The operators accept that the aircraft probably will be developed, for it is as much a challenge of the times as have been all the great innovations of routes and vehicles from Columbus right through to our day; but as has always been the case, before any commitment, they must seek assurance of ultimate economic soundness.

The greatly increased performance of both civil and military aircraft since the Second World War has raised problems undreamt of at an earlier period and, from the simple measures originally introduced to improve pilot’s well-being, have sprung the complex installations which are now essential for human survival and comfort and to ensure the proper functioning of vital equipment carried in present day machines.

The term “Environmental Control System” is used for convenience to cover collectively the various cabin and equipment bay, pressure, temperature and humidity control systems that are fitted.

In a note published in the June 1960 JOURNAL, 1 pointed out certain deficiencies in the initial strain concept introduced by Messrs. J. H. Arygris and S. Kelsey. In the August JOURNAL, those authors have disputed my conclusions with an argument which, I feel, does not pay due regard to current literature and I believe may be based on a misinterpretation of certain algebraic manipulations contained in my note. For they concluded that my note contained misprints, which it, in fact, did not. Due to these unfortunate circumstances Messrs. Argyris and Kelsey failed to illuminate the basic disagreement between our two formulations. For this reason I should like to set out in simple detail the exact nature of our disagreement and to point out, as clearly as possible, where the initial strain concept as developed by Messrs. Argyris and Kelsey becomes, I believe, somewhat inadequate and how the concept of fictitious thermal loads achieves greater validity.

We must thank Dr. Grzedzielski for the clarification of his argument and notation contained in his second note. Unfortunately, this note only shows more clearly the fundamental misconceptions and errors in his arguments. Our own argument as to the generality and validity of the initial strain concept remains unaffected.

In our previous note, we were concerned to explain the initial strain concept, which Dr. Grzedzielski had criticised as “ an oversimplification not generally admissible “ both in the calculation of thermal stresses and as a device to simulate the effects of structural cut-outs and modifications. Here, we prove independently by a direct consideration of the compatibility conditions in the final structure, that the modified stress distribution given by our analysis is the correct one—regardless of the extent of the coupling in the flexibility matrix.