It was of great interest to learn, from their most comprehensive paper “Helicopters in the Royal Navy“*, of the authors’ views on the SAR role, and especially of their outlines of the most beneficial features for ideal SAR aircraft (pages 472-476). This note refers to some aspects of the latter that were not mentioned, on which it would be interesting to know the opinions of the authors and of other helicopter pilots.

Some six years ago, I learned from the US Coast Guard of a system they had adopted with considerable success, to minimise the above hazards- At that time, I drew this arrangement (illustrated in the accompanying sketches) to the attention of RN and RAF officers concerned with rescue, but the idea received little support, for which there may be two reasons.

The American aircraft had, I believe, winch lines 200 ft long, whereas that on the Wessex was only 106 ft long. The Sea King (winch line 245 ft) had not then become operational. For their ideal SAR aircraft, the authors propose one 300 ft long. Thus, on the Wessex, the risk of fouling an obstruction would have been approximately doubled. As the longer lines now available and being proposed would go far to minimise this risk, perhaps the US system may in future be considered more favourably.

Stiffened panels usually consist of a skin of uniform thickness with several identical equally spaced longitudinal stiffeners. This paper demonstrates that the weight of a panel which is designed to have a specified initial buckling load, in uniform longitudinal compression, can be reduced by using more than one size of stiffener.

Initial buckling stresses quoted were all obtained using the computer program VIPAL, which is based on a recent “exact” method which allows for coupling between all possible modes. In this method the panel is considered as an assembly of thin uniform flat rectangular plates which are rigidly connected together along their longitudinal edges and are simply supported at their ends. The in-plane destabilising effects of the longitudinal compressive stress on the behaviour of these plates are included.

The design of a vehicle with cruise capability which can be packaged in a volume limited space, such as a cylindrical tube, centres around the problem of folding the lifting and control surfaces in such a manner that their deployment is simple and reliable. Several folding schemes are illustrated in Fig. 1. A wing fabricated of a flexible material and suspended by lines from the fuselage (parawing, Fig. 1a) can be folded to occupy only a small volume. These wings, however, are difficult to deploy at high forward speeds and have characteristically poor lift/ drag ratios unsuitable for high performance cruise capability. Another concept which allows storing the wing in a volume-limited space is the “scissor” wing (Fig. 1b) which, when deployed, swings out from a cavity in the fuselage.

Velocity defect law has been accepted as a universal velocity distribution for a turbulent boundary layer along a flat plate. The two necessary conditions for such a similarity to exist are not fulfilled with flat plate boundary layer having smooth or uniformly rough surfaces. Rotta has proved that a linear increase of surface roughness in the stream direction is necessary to ensure such a similarity. This note is an experimental investigation into turbulent boundary layer along a flat plate with linear increase of roughness height. It is obvious that, under such conditions, complete similarity characteristics are fulfilled.

There is some evidence to indicate that injuries associated with parachute opening shock are much less frequent than injuries associated with the acceleration loads imposed by ejection seats, even though the stresses presumed to be induced in the human body may often be of the same order. Since all such data is based on operational experience, the details are necessarily vague, and it is hard to draw firm conclusions.

The Aeromedical Research Laboratory, USAF, as part of its studies of aircrew injury, decided to see whether there was any difference between the dynamics of parachute opening when used with a rigid anthropometric dummy of the type normally used to obtain opening shock data, and a human parachutist whose body is naturally more resilient.

Users generally believe that the modern aeroplane can fly on its own wings economically. They see ordinary commercial competition in the European charter market: and in spite of the low prices that have resulted, they see that several of the UK charter operators have made very respectable profits. Sir Sefton Brancker told us that “healthy competition is the only sure preventive to inefficiency, slackness and dishonesty”—and some of us believe he might agree now that air transport is no longer an exception. Within California, by a unique regulatory quirk, the establishment airline fares have been nearly halved by competition; and yet specialist airlines have made profits. Users see also that three of the major American airlines which stood high in your Committee's productivity league, were each able to shed over 5000 employees when financial stress forced management to cut costs. In the present climate here, it may be that the Authority will find that its statutory duty to reward only efficient operators with an economic return may cause some painful problems, notably on the domestic trunk routes, where competition has exposed massive differences in cost levels.

The expansion of a boundary layer about a corner is of primary importance to a near wake flow field. Experimental evidence shows that turbulent boundary layers separate precisely at the corner and methods are available for predicting such an expansion. Inviscid calculations using rotational characteristics have been successfully used. However, one of the simplest techniques of treating the expansion of a boundary layer is the method of isentropic streamtubes. Calculations with such a method may be easily programmed for digital computers or quickly done on an electronic calculator. In this note, results obtained from the simpler streamtube method are compared with results from the more complex rotational characteristic calculations.

The streamtube method assumes that the flow field is isoenergetic, the stagnation pressure is constant along a streamline, and the flow is uni-directional before and after expansion. Obviously, the boundary layer is assumed to expand without interacting with the adjacent supersonic flow. The coordinate system used is shown in Fig. 1. Y is the coordinate normal to the attached flow direction, and W is the coordinate normal to the inviscid jet boundary streamline. The individual streamlines are located in the expanded profile by applying conservation of mass to the streamtubes.

In the year 414 before Christ a play was produced in Athens in which Aristophanes described how the birds who wished to build a free city in the air beat off the legislators who came to offer their services. Commenting on this event in an article published 17 years ago, I suggested that, if Aristophanes could look at the situation in the air at that momentand see in what way the legislators had prevented the birds of that dayfrom unfolding their wings, he would have prided himself on his foresight.

This paper concentrates on helicopter operations in the London area. There are of course, other fields of helicopter operations—most topically those to and from oil rigs in the North Sea—but times does not permit consideration of them on this occasion. I shall first then, attempt to set the scene by describing the layout of the controlled airspace which contains the helicopter route system for the London area.

This consists of the London Control Zone (Fig. 1), centered on London (Heathrow) Airport and it measures approximately 26 nm east to west and 16nm north to south. Its vertical extent is from ground level to FL 110 (11 000 ft approximately). This Zone is surrounded by the London Terminal Control Area, the base of which varies between 1500 ft and 2500 ft with a top limit of FL 250 (25 000 ft).

In presenting experimental results on turbulent wall jets, most workers have adopted as basic scales the slot height b and jet exit velocity Uj. Although this is a natural choice, experience with other flows like jets and boundary layers, where the details of the initial conditions are eventually (if not always rapidly) forgotten by the flow, prompts the question: cannot a gross parameter like the jet momentum flux Mj rather than Uj and b separately, suffice to determine the flow, at least in some region where the wall jet may be said to have attained a “fully-developed” state? If the answer to the question is yes, it would give a more compact and practically useful description of the flow than is currently available.

The problem of a delta wing in a high speed stream has been approached using linearised potential flow theory for low supersonic Mach number, and using thin shock layer theory for hypersonic Mach number. Recently the author has given a new theory of supersonic and hypersonic flows with attached shock wave over the compression surface of an unyawed delta wing. It has the advantage of being unified for both supersonic and hypersonic flows, and it gives almost identical results compared with exact numerical solutions. The purpose of this note is to extend the theory of Ref. 4 to include the effects of yaw.

In recent years considerable advances in the field of “optimal control” have been made. The impetus behind many of these advances has been derived from a growing demand that not only is it desirable, but also essential, that the design of a control system be optimised. This is particularly true of the aerospace industry in which the specification laid before the control systems designer usually demands the best possible performance for minimum cost (be that cost financial, on fuel, or on weight). Short of a major technological innovation, the problem becomes one of finding some means of meeting an ever increasing demand for performance, using existing technology. More attention has therefore been given in recent years to the study of optimal control, particularly since the advent of high speed, sophisticated scientific digital computers. Two of the major areas which have been explored have been the optimal design of systems, according to a quadratic performance criterion, and non-linear programming techniques. These are briefly discussed in Section 3 of this paper.

The possibility of failure by buckling of structural members subjected to compressive loadings concerns every structural engineer and designer. In the design of aircraft, these compression members may be stringers, spars, or other stiffeners. Various geometrical sizes and shapes of stiffeners are used in aircraft construction. They are either extruded in the desired shape, machined, or formed out of sheet metal on a hand brake. They may be I-beams, Z-beams, channels, or angles. They may be used singly or may be joined to form a built-up or composite member. The actual choice of the size and shape of the stiffener depends upon the forces and moments to which it will be subjected.

This lecture is devoted to small gas turbines defined as those in the power range up to around 3000 shp and 4000 lb thrust, ie engines for general aviation, corporate and executive aircraft and limited trainer applications and helicopters. I do not propose to deal with APUs as these were the subject of the excellent paper given by C. H. Paul in a recent Halford Lecture.

I intend, as you may have gathered from the title, to treat this subject more in a total business context rather than as a strictly technical lecture. Recognising both the strong technical traditions of the Society and the man, Major Halford, whose memory we honour, I shall, however, take some time to deal with selected technical aspects of the subject.

The task attempted in this paper is to exhibit a contemporary perspective of aeronautical research trends, needs and opportunities and to draw from it such pointers to the future as may contribute in a manner both interesting and helpful to the overall theme of the meeting.

The contribution at which I am aiming is a broad presentation, not going deeply into technical detail but standing sufficiently back from the research scene that, while major details are still discernible to give the picture definition, at the same time the scene as a whole can be viewed in some relationship to its environment of project development, market opportunities, social considerations and other factors.

This paper gives details of a method for measuring local heat transfer rates and shows that the results obtained are in fair agreement with the expected values.

The experiments were performed in the 8 in x 3·5 in supersonic intermittent wind tunnel at Queen Mary College (University of London). A maximum Mach number of 1·6 is obtainable at which the free stream Reynolds number in the working section is 4·92 x 106 per foot under normal conditions. The tunnel has a running time of approximately 20 seconds at a Mach number of 1·6.

The series of heat transfer measurements were restricted to flat plate conditions with non-constant wall temperatures.