Wartime experience with aircraft steels emphasised the desirability of a complete revision of specifications for aircraft steels from both the technical and inspectional points of view. This was undertaken in association with the steel makers and the Society of British Aircraft Constructors under the aegis of the British Standards Institution. The aims were

1. (i) To reduce the number of basic types of steels to a minimum, consistent with satisfying the demands of aircraft engineers.

2. (ii) To transfer as many materials from the D.T.D. series of specifications to the B.S. Aircraft series.

3. (iii) To prepare an Inspection and Testing Schedule, which would exist separately from the Materials Specifications, but applicable to all steels in whatever form they may be required.

4. (iv) To achieve a degree of economy in alloying elements, again without detriment to the satisfying of designer's requirements.

The Inspection and testing schedule was issued as British Standard S100 in January 1949, but significant difficulties in its full application appeared in practice and it was not until the second half of 1952 that agreement was reached to put into practice the current standard 2 S100.

Knowledge of the motions of aircraft control surfaces due to oscillating hinge moments is of great importance for flutter investigations. While the impedance of the power control unit may play a large part in determining these motions, it is not their only cause. The elasticity of the power unit mounting and the effect on the input of distortion of the aircraft structure, an effect which may depend on air speed, may also be important. Nevertheless the influence of the power unit itself is of major importance. No very satisfactory methods appear to have been devised yet, however, for making such tests. The difficulty lies chiefly in applying the large oscillatory loads to the output of the power control unit, which are necessary if the test is to be realistic. The facts that the oscillating loads must be applied to the jack while it is moving, and that a non-oscillatory opposing load is also desirable, add to the difficulties. For convenience in assessing the characteristics of the unit, it is desirable that the oscillating load should be superimposed on a constant opposing load and that the basic jack motion should be of constant velocity.

It has been adequately demonstrated that the reduction of structural and aeroelastic problems to a form suitable for computation is best done using a matrix formulation. The present paper outlines such a formulation of the aeroelastic problems on wings of arbitrary plan form and gives a method of solution which has been found satisfactory over a wide range of practical calculation. In the symmetric case the problem of manoeuvre margin due to aeroelastic effect is considered and in the anti-symmetric case the rolling effectiveness and the reversal speed for aileron controls are dealt with.

The slender–body solution for the flow about low aspect–ratio wings of non–zero thickness yields a pressure distribution which is valid in the vicinity of the leading edge and therefore suitable for comparison with measured distributions. In this note an expression is derived for the theoretical pressure distribution, due to small incidence, on a slender elliptic half–cone in uniform motion. It is shown that this distribution is additive to the zero incidence pressure distribution. A comparison with some low speed measurements shows qualitative agreement.

The method can be simply extended to other bodies of elliptic cross section.

The aerodynamic characteristics of airframes are expressed as aerodynamic transfer functions, giving the relationships between input and output for each of the three separate planes of motion, roll, pitch, and yaw. By assuming no cross-coupling between planes and linear aerodynamics, and by making certain other assumptions, which apply particularly to conventional airframes with fixed wings and rear controls, relatively simple approximate algebraic transfer functions giving the relationships between the control surface deflection (the input) and any airframe motion (the output), are obtained.

The open loop aerodynamic transfer functions thus obtained are used as part of the auto-pilot block diagram, in which the performance of other components, such as actuators, instruments and electrical networks, are also expressed in transfer function form. The aerodynamic transfer functions are useful in auto-pilot evolution and synthesis in that they aid selection of the airframe motions to be measured, modified, and fed back to close the auto-pilot loop.

For mathematical assessment of closed loop performance and stability, open loop transient and frequency responses are used, and curves of airframe responses are plotted in linear, logarithmic and polar form by standard methods from the aerodynamic transfer functions. Some methods of using these curves, which follow the general lines adopted in servo-mechanism and electronic amplifier design, are explained briefly.

Analogue computers are frequently used when the computations to be made are so complicated as to need the use of a computing machine. The aerodynamic transfer functions then form one block of the simulator set-up, and on larger computers the more exact form, including any non-linearities and cross-coupling effects, can be used.

A considerable amount of attention has been devoted to the study of the position of the centre of shear for straight beams but there appears to have been Uttle or no consideration of this position for beams of which the centre line is either shghtly or heavily curved. The reasons are probably twofold. Firstly, no demand for such an investigation has been made by designers, and secondly, and probably not least, any very general inquiry into this problem will lead to complex expressions which would not only be difhult and tedious to evaluate but will also be of uncertain accuracy.

This problem of taking beam curvature into account starts from the point where it normally ends for most mechanical engineers, who usually require to estimate the bending stresses in members of large curvature such as hooks and chain-links. Now the classical theory of elasticity applied to these latter members has yet to be seen to conform with experimental results to such a high degree as makes the added labour attending such methods worth while.

When I accepted the invitation to give a lecture on “The Uses of the ACE Computor” my first thought was to ask for the title to be corrected. The “ACE” as such has yet to be built and the machine that has been in use in the Mathematics Division of the National Physical Laboratory is a pilot model. However, to an audience whose chief interest is in aeronautics, the ambiguity of a title including the phrase “Pilot ACE” probably outweighs the advantage of its accuracy.

As its name implies, the Pilot Ace was built,at the N.P.L., with the intention of testing out the practicability of various ideas in the design of a full scale automatic computing engine, or ACE as it was called. The over-riding consideration was economy of equipment but, despite the rudimentary nature of the facilities provided on the machine, it was found to be a fast and powerful computor and has been fully and successfully employed for the past three years on a 13 hour basis.

How much is known about high altitude clear air turbulence? Sufficient to show that aircraft may, with little or no warning, experience bumpiness which the pilot would describe as “ moderate ” or, on occasions, “ severe,” at altitudes up to at least 40,000 feet, thus including altitude bands where commercial jet aircraft will be flying during the next few years. The phenomenon, therefore, is of significance to airline operators and the requirement is for the aircraft commander to be supplied with sufficient information to enable him to avoid such turbulence, or to provide him with adequate warning so that crew and passengers can be properly prepared for it. This does not imply that there is evidence to suppose that clear air turbulence may be of such severity as to affect the safety of a flight. It is important, however, to take good care of the passengers and their comfort is an essential consideration.

The problems discussed here are those of a uniformly inclined and skew loadings on a finite rectangular plate such as would be obtained by the action of wind or rain on a large door of an aircraft hangar or other such building. Such a force may be considered constant in magnitude over the area of the plate but non-normal in its direction with the plate.

It is always difficult and sometimes dangerous to estimate future trends and events. Nevertheless, it is a most satisfactory outlet for one's imagination and it is, to some extent, necessary unless we are to abandon all attempts at shaping our own future.

Before trying to forecast the future of the helicopter, I feel that I must define what I mean by a helicopter. Resort to the English Dictionary does not help greatly; as is so often the case, it has been left behind by progress.

A helicopter is described as “a flying machine that should rise vertically by airscrews revolving horizontally.” Derivation from the Greek, helix—a spiral, and pteron—a wing. This definition indicates that every aeroplane having a thrust loading of less than one is a helicopter, unless it is propelled by rockets, ramjets or pulse jets!

The dictionary definition of an aeroplane is “ a flying machine heavier than air using planes.”

The work reported in this Note originally formed part of a comprehensive investigation into the fatigue strength of magnesium alloy ZW2 (2 per cent. zinc, 0·65 per cent. zirconium) in sheet form. The major portion of the programme was eventually cancelled but the tests herein described are sufficiently self-contained to be repotted by themselves. Direct stress fatigue determinations were made on small round test pieces machined from ¾ in. thick plate, at mean stresses of 0, 4,000, 8,000 and 13,000 lb./in.2. Static and rotating bar fatigue tests were also made.

In this paper a theory of wrinkling in Sandwich Construction is presented in two parts. In the first part, the thickness of the core is regarded as finite. The wrinkling stress is given by a simple square root formula consisting of the Young's moduli of the materials and the ratio of the thickness of the face and core. In the second part of the theory the same procedure is followed, with the main difference that shearing stresses in the core are also considered, and the analysis is extended to the case where the face is supported by a sufficiently thick or a semi-infinite medium. The result for the wrinkling stress is a cubic root formula which consists of the moduli of the materials but indirectly depends upon the geometry of the structure; this formula is valid only if the ratio of the thickness of the core and the face is greater than or at least equal to the ratio of the width of the marginal zone of distortions in the core, and the thickness of the face. The width of the zone is also useful as the criterion which discriminates between the two cases of buckling as envisaged in this paper. The theory is compared with the previous works, and also with the test results already available. The selection of the experimental data comprises steel, aluminium alloy and Papreg (laminated paper plastic) as the face; and onazote, cellular cellulose acetate, granulated cork, sponge rubber and expanded formvar as the core materials over a wide range of core-face thickness ratio. The agreement between the theory and the tests is satisfactory.

The Classical Kirchhofif–Love Theory for the deflection of thin plates leads to fourth order Lagrange's differential equation,D△4w — q = 0 for which a general solution is not always possible. Exact solutions are known so far only for a few special cases and, therefore, numerical solutions have often been tried. The advantage of numerical solution is that it can be applied easily to any plate plan form which is in marked contrast to the analytical method where, for mathematical reasons, definite restrictions have to be imposed on the geometrical shape of the plate. Among the various numerical methods, relaxation is the easiest, but when applied to solving a biharmonic equation, the process becomes extremely difficult and laborious as convergence is very slow and the unit relaxation operator cumbersome to deal with.

Some five years ago the author was privileged to deliver a Section Lecture to the Royal Aeronautical Society on the subject of reheat. The present paper attempts to summarise the problems which now arise and to give some idea of the progress which has been made in the intervening years.

In 1949, reheat was in its infancy in Great Britain. A certain amount of progress had been made in the United States but the information from that source was scanty and vague. Tests at the National Gas Turbine Establishment (N.G.T.E.) had given some engine data but this was in the nature of preliminary information only and was by no means complete. In fact the majority of the problems which now beset us were then completely unknown or were considered unimportant. The N.G.T.E. work was valuable, however, in that it demonstrated the practicability of reheat, although at the time the comments of many who saw this and other schemes in operation were somewhat sceptical and definitely unflattering.

In a recent paper, Yates has shown that it is convenient to distinguish between two main types of stimulus which induce torsional vibration in geared-shaft systems. The first, which is dealt with in detail in the literature, is “force excitation” by which periodic torques act upon the system and throw it into oscillation; for the purpose of this note, it will be convenient to refer to a previous article as dealing with this problem.

The second type of excitation to which Yates refers is “displacement excitation.” It may be thought of as arising from errors in the cutting of gears which cause periodic variations in the velocity ratio of mating components. A Hooke's joint which couples two rotating systems which are capable of torsional oscillation would give rise to this form of stimulus.

The effort required to produce very accurate supersonic nozzles has suggested the need for criteria from which the order of accuracy required for quantitative experiments can be deduced. The purpose of this note is to indicate a way of obtaining such criteria for measurements of aerodynamic forces, but also to point out that “ sufficiently accurate ” nozzles for such measurements are indeed very difficult to produce and to suggest that a remedy might first be sought in interference corrections.

The problem of the buckling under bi-axial compression is considered for flat rectangular isotropic plates with simple edge conditions and no lateral restraint. The buckling stress coefficient is plotted against the side ratio for various conditions of edge restraint and a known compression or tension on the sides of the plate. It is found that there is some reduction in the buckling stress if the sides of the plate are subjected to compressive stresses and, conversely, there is an increase for tensile stresses. Furthermore, for plates with large side ratios, there is a rapid decrease of the spanwise buckling stress as the chordwise compressive stress approaches its appropriate Euler instability value.

A structure with cut–outs or other discontinuities is generally considered to be more difficult to analyse than a similar structure without such features, although it is in fact less redundant. This is largely due to the special conditions which have to be introduced to allow for these discontinuities, tending to make the calculations less adaptable to routine computation . Another and very practical reason is that it is sometimes fairly easy to—dare we say it?—make a reasonable guess at the load distribution in a structure without cut–outs, whereas a much more critical approach is necessary for the treatment of structural discontinuities. Be that as it may, it will be assumed here that the structure without discontinuities, which will henceforth be termed the “original” structure, can be analysed exactly, in the sense that all the effective redundancies are taken into account. It will now be shown how these calculations may be modified to deal with the corresponding structure with discontinuities, which for this reason will be referred to as the “modified” structure.

The work described was undertaken under a Ministry of Supply Contract to investigate the fatigue strength of certain materials where the expected endurance in use was less than the range of values usually examined in fatigue testing.

The material was supplied in the form of a 6 ft. x 3 ft. x 1/4 in. sheet of aluminium alloy MG5.

Tensile test pieces having a parallel length one inch wide by 4 1/2 in. long were taken in both directions from two diagonally opposite corners of the sheet. Later, similar longitudinal test pieces were taken from areas near the other two corners.

The plane bending fatigue test pieces were of the form shown in Fig. 1 and were all cut in the longitudinal direction. The waisted test portion was polished on the edges, the original sheet surface being retained on the faces. Steel plates were clamped to the ends of the test pieces by bolts passing through the holes in order to localise the strain in the test portion as much as possible.