The idea of using the reaction of a jet of water pumped sternwards from a ship is endowed with a respectable antiquity, but tests on the efficiency of such propellers, whether model or full-scale, have not been extensive. Perhaps the most thorough tests were those of S. W. Barnaby (Thornycroft and Co.) in 1880, described by him in the Proceedings of the Institute of Civil Engineers, Vol. LXXVII, page 1, 1883. This investigator measured the reaction on a nozzle delivering water under pressure into still water by a force balance connected to the nozzle itself. As a result, the “ Waterwitch,” a vessel of 1,160 tons, was constructed to work on this principle, the jet being supplied by a centrifugal pump which took water from a scoop amidships. The efficiency of the jet itself worked out at .71, but with the pump the combined efficiency was only .38.

The Ninth Lanchester Memorial Lecture was given by Professor W. A. Mair, MA, FRAeS, on “STOL—Some Possibilities and Limitations” in the Society's Lecture Theatre on 12th May 1966. The Chair was taken by the President, Mr. A. D. Baxter, MEng, CEng, FRAeS. Before the lecture he presented the Society's Gold Medal for 1965 to Professor M. J. Lighthill, DSc, FRS, FRAeS, for “his outstanding original work in many fields of Aeronautics”, explaining that Professor Lighthill had been unable to be present at the Wilbur and Orville Wright Memorial Lecture in December 1965 when the Society's main awards for the year were presented.

The President then said that this was the first meeting of the Society since his installation as President and it was a very pleasant way to start his year in office with first, the presentation of the Gold Medal to one distinguished scientist and second, the introduction of another as the Lanchester Memorial Lecturer.

There would be many members of the Society who would remember Dr. Lanchester, his attendance at lectures and his contributions to the discussions. It was true, however, that the real stature of such men was rarely recognised at close quarters and often only in the light of later developments was the importance of their work realised. It was 20 years since Dr. Lanchester's death and their Memorial Lecture was in its ninth year. Each year, each President had added a tribute to this great man. He was a man of many parts—a scientist, musician, poet and engineer and aeronautics owed much to him. It was fitting that the Lecture had established a tradition of surveying some field of research associated with aerodynamics, in which Lanchester was so eminent. He thought that Lanchester would approve of both the subjects discussed and the distinguished men who had honoured his memory by presenting them. Before introducing Professor Mair he wished to welcome Mrs. Lanchester and Mr. George Lanchester and his wife, and Mrs. Mair.

Professor Mair must be well known to most of them. After graduating in Mechanical Sciences at Cambridge in 1939 he had joined the Royal Aircraft Establishment at Farnborough, where he scent some six years in the Aerodynamics Department doing research on high subsonic speed, both in the wind tunnels and in full-scale flight. In 1946 he had gone to the University of Manchester as Director of the Fluid Motion Laboratory and since 1952 he had held the Francis Mond Chair of Aeronautical Engineering at Cambridge. His chief interest there had been mainly in low speed aerodynamics and his authority in that field was widely recognised. In 1963 he had been appointed Chairman of the Powered Lift Committee of the Aeronautical Research Council.

The linear vibrating system of two degrees of freedom, see Fig. 1, whose mathematical solution is well known (see for example ref. 1), is nevertheless on the verge of being too awkward for manipulation in the design office. It is of considerable interest as the basis of a vibration absorber and techniques continue to be developed for investigating the effects of changing the design constants on the properties of the system and the more complicated ones derived from it.

The stability criteria for systems whose characteristic equations are of the eighth degree are expressed in a convenient form and a detailed scheme for the computation of the complete set of test functions in sequence is given. A special method for the direct calculation of the critical test function T7 is also given; the vanishing of this determines the critical flutter speeds of quaternary systems.

The classic problem of stability and transition to turbulence of an incompressible flow over a flat plate subjected to uniform suction through the surface is considered. This is an important “model” problem for laminar flow control and the results have direct application to the design of low drag aerofoils. Detailed mean boundary layer calculations are presented and the flow stability is determined by obtaining numerical solutions of the Orr-Somerfield equation. Transition conditions are estimated by using the semi-empirical en method. Accurate values have been obtained for the levels of suction necessary to delay instability and transition indefinitely. Consideration has also been given to the effect of surface suction upon the plate drag.

Advances in computational methods are producing a need for accurate and detailed experimental results for test cases. At the same time economic trends force many of these experiments to be made in small wind tunnels. In some cases the small scale of the model does not provide enough space for the instrumentation and so the model must be distorted from the ideal shape. This paper is concerned with one aspect of this problem, that is, with the need to make wings thicker than required, either for strength or to provide volume for pressure tubes and other measuring devices. In particular the paper deals with the conditions under which the pressures measured on the flat surfaces of wings with bevelled or triangular cross-sections as shown in Figs. 1(b) and 1(c) are a good representation of the pressures which would be measured on the corresponding flat surface of thin wings (Fig. 1(a)). This problem has been touched on in a number of papers reporting experimental results but it does not appear to have been tackled systematically. The discussion of this paper is based on these earlier tests and on some new test results.

It is a pleasure and indeed an honour to be invited to contribute this paper to the 70th Birthday Tribute to Professor A. D. Young, in recognition of his longstanding friendship and as a colleague of the author for many years.

It was Professor A. D. Young, at the College of Aeronautics in 1949, who gave the necessary guidance and encouragement to the author in starting research into aerodynamics noise, which was possibly one of the first attempts made in making a quantitative assessment of the noise from a turbulent air jet.

The paper discusses the early experimental work and its conclusions, and how it led to attempts to reduce jet noise through the use of vortex generators and the corrugated nozzle. The comparison between theory and experiment is commented on with reference to the early work on acoustic sources and the questions this comparison raised. A final section is included which attempts to close the gap in our understanding of jet noise generation.

Pilot-Induced Oscillation (PIO), although an old issue, still poses a significant threat to aviation safety. The introduction of new systems in modern aircraft modifies the human–machine interaction and makes it necessary for research to revisit the subject from time to time. Given the need of aircraft manufacturers to constantly perform PIO tests, this study analysed the feasibility of using three different computational pilot models (Tustin, Crossover and Precision) to simulate PIO conditions. Three aircraft models with different levels of propensity to PIO (original, low propensity and high propensity) were tested, as well as two pilot gain conditions (normal and high). Data were collected for a purely longitudinal synthetic task through simulations conducted in MATLAB®. PIO conditions were detect using a tuned PIO detection algorithm (ROVER). Data were analysed in terms of both whether the pilot models triggered a PIO condition and for how long the condition was sustained. The results indicated that the three pilot models only provoked PIO conditions when high gain inputs were applied. Additionally, Crossover was the only pilot model to trigger a PIO for the three aircraft models. There were also significant differences between the pilot models in the total PIO time, as the Tustin model typically sustained the oscillatory condition for longer.

Head-up display (HUD) is an electronic presentation of flight information projected into the pilot's line of sight, enabling him to see both the outside world and relevant instrument information at the same time (see Fig. 1).

Using HUD the pilot no longer needs to alternate his attention continually between his cockpit instruments 2-3 ft (0·6-0·9 m) away and the outside world effectively at infinity. This continual change which present day pilots must make is first, very tiring and secondly it takes time; time to scan instruments, assimilate the relevant information, move the head and eyes, refocus the eyes on the outside world, adjust perhaps to a different light level and then take in the changed surroundings before returning to check the instruments. This time may add up to several seconds. In normal flight this is acceptable.

The light gyroplane is a class of aircraft popular with amateur constructors and pilots. As a result, there is limited design guidance available since formal technical resources are not available to the community. Rule-of-thumb, intuition and historical experience tend to influence design evolution. Empennage configuration is a prime example of this paradigm, and the objective of this Paper is to explore those factors that influence horizontal stabiliser effectiveness with particular reference to dynamic stability. An individual-blade rotorcraft mathematical model is coupled with a vorticity-based flowfield code, necessary to capture the highly interactional aerodynamics associated with empennage location at the rear of the airframe. A parametric study of horizontal stabiliser location shows that maximum benefit from the energising influence of the propeller slipstream is obtained if the surface is placed near the edge of the propeller wake. Further, traditional design parameters such as tail volume ratio offer an incomplete indicator of empennage effectiveness without consideration of airframe blockage and propeller slipstream. It is concluded that empennage sizing calculations can be straightforward, but require due consideration of the impact of the close-coupled nature of the vehicle on stabilising surface aerodynamic effectiveness.

Every control system employs some computing elements and it is customary to designate systems after the type of computing predominantly used in them. Thus, we have hydromechanical systems, pneumatic systems, fluidic systems, analogue electric systems, and so on. Traditionally, as described by Watson, the aircraft gas turbine has used hydromechanical control (see Fig. 1). The diagram does not depict any particular engine component but illustrates some of the principles used in a variety of present-day systems. A mixture of bellows, springs, cams, orifices, etc, compute according to their particular arrangement; frequently the fuel itself is used as the computing medium. Essentially such systems compute continuously using physical representations of the measured parameters. They are, therefore, of analogue form.

It is unnecessary to enlarge on the importance of the fuels used in aviation in a paper before this Society. Not only is fuel quality the most important individual factor to be considered in the economics of civil aviation, but engine development, aircraft design and performance depend upon it.

An apology is offered, in view of the title of the paper, that as it is barely possible in the confines of a single paper to deal adequately with the volatile fuels (known as petrols and gasolines) used in the normal carburettor or homogeneous charge engines, other types of fuels have necessarily been omitted.

An overview of some of the problems, which have to be overcome in the design of hypersonic missiles or projectiles, is given. In the main, these are connected with the high levels of heat transfer rate suffered by such vehicles, particularly those induced by aerodynamic control surfaces. Some of the aspects of heat transfer, induced by, and suffered by such controls are presented. Data are presented and discussed, regarding the effects of flow separation due to flaps or flares, the effects of sharp and bluff fins (both swept and unswept). Finally, the aerodynamic interactions and heat transfer rates induced by reaction control jets are presented.

Much of the data published on the subject of cost prediction for aircraft has dealt with empirical methods of estimating the total engineering hours of design and development. Less seems to have been written by or for the practising estimator, concerned not only with the total estimate at the initial stages of the project but also with its continuous refinement to full go-ahead, on to certification and later still to variants. In this paper the authors describe some of the methods they use, the historical background against which the methods have evolved and discuss some of the problems which are met.

Turbulent airflow in a supersonic intake model of rectangular cross-section is studied numerically. Instability of shock waves formed in the intake and ahead of the entrance is examined. Solutions of the Reynolds-averaged Navier–Stokes equations are obtained with a finite-volume solver of second-order accuracy. The expulsion and the swallowing of the shocks with a variation of free-stream parameters are studied at both subsonic and supersonic conditions prescribed at the outlet. Hysteresis in the dependence of 2D flow on the free-stream velocity and angle-of-attack is documented. An influence of 3D effects on the flow is examined.

Linear elastic buckling strength assessments are presented for a range of thin plate arrays with optimal strength to weight ratio configurations subject to a range of arbitrary in-plane stress states. The outcomes of the assessment are presented as design curves demonstrating relative buckling strength increases with respect to the classical square plate datum of equal mass.

A stiffness matrix method is adopted for the initial buckling strength predictions. The method is based on exact flat plate theory and assumes that the plate is continuous over supports, whereby deformations in one cell of the plate array influence the deformations in adjacent cells. The supporting webs of each cell are approximated in this study by simple rigid supports that enforce nodal lines or lines of zero out-of-plane displacement in the buckled panel, i.e. the bending and torsional stiffness provided by the supporting webs of the real structure are ignored. Selected results are validated by FEM prediction.

Probably few readers realize the advance which has been made towards a practical solution of the problem of aviation within the last eight years. It is greater than all that was accomplished during the preceding two centuries. Not only has the subject been rescued from the contempt brought upon it by past failures and eccentric proposals, but very able men—Langley, Maxim, Pilcher, Hargrave, and Lilienthal—have shown by their experiments and writings that man may fairly hope eventually to fly through the air.

Hundreds of failures during the last 300 or 400 years have well shown what arrangements will not answer, but the labours of these latter experimenters now enable us to forecast that success is most probably to be won with aerocurves, driven by screws or flapping propellers, actuated by a steam or a petroleum engine.

What is the best course to be now followed to make still further advance? Most of the above experimenters seem to have proceeded upon the assumption that it is practicable to evolve a complete flying machine at one operation, and to apply propellers and motors from the first, before we are quite certain as to the stability in a wind.

A method is proposed by which a direct numerical simulation of the compressible Navier-Stokes equations may be embedded within a more general aeronautical CFD code. The method may be applied to any code which solves the Euler equations or the Favre-averaged Navier-Stokes equations. A formal decomposition of the flowfield is used to derive modified equations for use with direct numerical simulation solvers. Some preliminary applications for model flows with transitional separation bubbles are given.