In the design of structural frames it is often the practice to assume the joints to be pinned, or hinged, and to calculate the forces in the members on this assumption; actually, the joints are made with considerable rigidity. The following are the results of an analytical and experimental investigation into the stresses in a rigid-jointed space frame and of the magnitude of the errors due to the assumption oi pinned joints.

The list of the Wilbur Wright lecturers is a long and distinguished one, and I feel it a great honour to have been asked to join this eminent company. I was invited just before Christmas, 1939, to address you to-night on the future of British civil aviation. I felt then that though only with great difficulty would it be possible to produce a lecture under war conditions, the invitation must be accepted, firstly because no greater compliment can be paid an aeronautical engineer, and secondly, to support the Council's policy of continuing the Wilbur Wright lectures, which were not interrupted by the last war, through this one.

My experimental work in flight refuelling goes back over a period of ten years, when I became interested in its possibilities as a means of removing the hazard' of a heavily loaded take-off. My idea was that an aircraft should have a normal and safe take-off, and that when it was comfortably in the air, it should receive its load of fuel, thus being given both an extended range and an increased1 pay-load.

In 1932, initial experiments were carried out on two D.H. 9's, and early in 1933, we continued experiments with a D.H.9 and a Handley Page W.iotransport machine.

The field of viscosity forces in fluid currents of low viscosity—i.e., high Reynolds number—is frequently confined to a narrow zone in the vicinity of the surface; consequently, calculation can be simplified by the following assumptions :—

1. In the friction zone, customarily called the “ boundary layer,” only the most effective components of the viscosity force μΔω, are considered—e.g., for the x component of this force, μΔu (calling the direction of main flow parallel to the surface, the x direction, and that perpendicular thereto the y direction) the expressions μ (∂2u/∂x2) and μ (∂2u/∂z2) can be neglected for μ (∂2u/∂y2).

The performance of large aero-engines is limited by the weight of the charge that can be inspired at the high speeds and by the tendency to knock, which is due chiefly to the hot exhaust valve. The performance can be considerably increased when a valve control system is employed, which offers small resistances and large passages for the gas flow. Besides this, the walls of the combustion chamber must be kept at a low and uniform temperature, in order to ensure freedom from knocking even with high compression.

As tests show that the knock resistance and the piston pressure in engines with rotary valves are sensibly greater than in the same engines with poppet valves, measurement results are used in order to ascertain the possible limiting values obtainable for the compression ratio, the mean piston pressure and the specific fuel consumption of large aero-engines.

The main object of this paper is to put forward a routine method for taking acccount of the variation in section in the stressing of rigid jointed frames. The formulæ in general use which mainly apply to members of uniform section are suitably modified so as to be applicable to cases where members may be of variable section. The underlying feature of the method is to base the treatment of a rigid jointed frame on the properties of cantilevers loaded at their free ends. Certain deflection coefficients are involved and as shown may be found separately as required for each member. These coefficients may be computed either analytically or graphically. The importance of taking account of variation in section is brought out by practical examples. In certain cases it is shown that very considerable errors may arise if instead of having regard to the actual variable section a mean value is assumed as being sufficiently close from the point of view of the resulting loads. From experience already gained in the application of the method to an actual structure the labour involved is not much in excess of that for the case of the same structure with the sections averaged.

The method does not take end load deflection effects into account and an extension of the method to do this seems desirable. There are, however, many cases in which end load deflection effects are not serious, and in any case it seems useful to have a solution in which these are neglected in the absence of means for obtaining a more comprehensive analysis.

Total air transport operating costs are divided into three general classifications : (1) Significant operating costs which are defined as those resulting from the flight of an aeroplane or those affected by variation of aeroplane size, power, initial cost, and aerodynamic efficiency. Equations for the development of significant costs are presented. (2) Passenger service costs which are defined as those which vary with the number of passengers aboard an aeroplane are enumerated. (3) Overhead costs in air transportation are also dwelled upon and it is concluded that these costs change their proportion to significant costs when the physical and performance characteristics of the aeroplane being considered change.

The marked effect on significant cost of varying aeroplane utilisation is brought out. Utilisation is also shown to be chiefly a function of trip length, with block speed, seasonal schedule variation and other factors being secondary functions.

The effect on significant cost of varying block speed either by increasing cruising power or improving aerodynamic efficiency is investigated for one type of transport aeroplane. Increased block speed obtained by increasing cruising power is found to increase significant cost very rapidly. Increased block speed obtained by improving aerodynamic efficiency is found to decrease significant cost appreciably and relatively large increases in aeroplane first cost are found to be justified if aerodynamic efficiency is improved. The adverse effect on significant cost of increasing trip length is investigated. The effect of changing specific fuel consumption is discussed.

It is pointed out that economic factors apart from those associated with fundamental aeroplane design and performance, such as interior arrangement, safety records, competition and general economic conditions, play the major role in the economic outcome of air transportation.

When the all-metal cantilever monoplane brought to a close the era of wood, tubes and fabric for aeroplane structures, the new development opened up broad fields for the improvement of aerodynamic efficiency, with resulting great increases in all-around performance, military effectiveness, and economy of operation.

The all-metal fuselage represented a new type of construction in that the skin formed a shell to replace trusses in carrying the bending loads imposed at the tail in flight and on the ground.The scheme of using a shell to give–at once–structural strength and aerodynamic fairing, previously accomplished by two elements, appears to be economical and sound. If the labour now needed for the fabrication and assembly of the multiplicity of internal members is to be eliminated the stresses must be taken in the skin and the principle should, if practical considerations permit, be extended to include the wing and fixed tail units.

Tests were made in the N.A.C.A. gust tunnel to determine the effectiveness of a long-period dynamically overbalanced flap in reducing aeroplane accelerations due to atmospheric gusts. For two gust shapes, one gust velocity, one forward velocity, and one wing loading, a series of flights was made with the flap locked and was then repeated with the flap free to operate. The records obtained were evaluated by routine methods.

The results indicate that the flap reduced the maximum acceleration increment 39 per cent, for a severe gust with a representative gust gradient distance of 8 chord lengths and that, for an extreme gust shape (a sharp-edge gust), the reduction was only 3 per cent. The results also indicate that the flap tended to reduce the longitudinal stability of the aeroplane. Computations made of the effectiveness and the action of the flap were in good agreement with the experimental results.

Explicitly or implicitly the yield strength of a material is often used as a measure of incipient structural damage. With the yield strength determined by conventional methods, however, it cannot be said in general for two structural elements geometrically alike but of different materials that similar loads, producing maximum stresses equal to the yield strengths in the two cases, are simply related to the yield strengths. A definition of yield strength is proposed in this paper which often has the advantage that, for geometrically similar structures of different materials, loads producing maximum stresses equal to the yield strength are proportional to the yield strength.

1. Piercy, Piper and Preston (1) and Piercy, Preston and Whitehead (2) have developed a family of aerofoils which are the inverses of hyperbolas. These calculations give a family with a finite tail angle, equal to the angle between the asymptotes. It may sometimes be preferable to have a rounded trailing edge. It is the purpose of this note to describe a family whose members are inverses of ellipses.

2. It will be seen that, if the centre of inversion is on the major axis produced, symmetrical sections will result; whilst, if the centre is just off the major axis, a cambered section will be produced; and the inverse of the major axis (which is a circle) can be taken as the camber line.

A new graphical method of predicting aeroplane performance has been evolved by utilising the unique properties of Eiffel's logarithmic propeller chart. Its substitution for rigorous numerical methods results in the saving of labour without sacrifice of accuracy. Particularly convenient when constant speed propellers are involved, this method yields performance characteristics for altitudes below, as well as above, the critical ones for supercharged engines.

Although coupling of the word “ rigorous ” with any method of aeroplane performance prediction may be somewhat of a misnomer, the title of this paper is intended to identify a method which falls short of complete rigour only to the extent dictated by acceptance of the following simplifying assumptions :—

1. (1) That the propeller thrust acts along the direction of the flight path, and

2. (2) That lift is equal to weight in steady rectilinear flight.

The amount of labour involved in predicting the performance of a modern aeroplane with even this degree of rigour is so great that methods which incorporate additional simplifying assumptions have come into common use—despite their sacrifice of accuracy to convenience—and the more laborious methods are now rarely employed except in studies of very painstaking character. The desirability of minimising the drudgery of rigorous prediction has long impressed the writer and it is the purpose of this paper to describe a graphical method of attaining that objective.

The principal departure which characterises the new method is the adaptation of Eiffel's logarithmic propeller chart to the construction of available power curves. The use of “ indicated airspeed ” (σ1/2V) and the analogous quantity “ indicated power ” (σ1/2 h.p.) as co-ordinates causes a single curve to represent the power requirements for level flight at all altitudes. Other simplifications arising from the recognition of certain engine characteristics, utilisation of the unique properties of the Eiffel chart and development of a convenient method for evaluating the excess available power (from logarithmic curves) make it possible to eliminate a large portion of the labour previously required for rigorous performance prediction. The method is not only applicable to, but especially convenient in, the case of the constant speed propeller; moreover, it yields complete level flight and climb characteristics for altitudes below, as well as above, the critical altitudes of supercharged engines.

During the past few years an extensive amount of experimental data on split flaps has been made available to the aircraft industry, through the publications of aeronautical research laboratories, both in this country and abroad. In general, each publication deals with one particular aspect of the problem, and when the effect of wing flaps on the performance of an aircraft is being estimated a certain amount of difficulty may be experienced in deciding which of the many reports available gives results most readily applicable to the case being considered ; and what allowances, if any, should be made for wing taper, flap cut-out, fuselage, etc.

In this report the available data has been analysed with a view to answering these questions, and presented in such a form that it may be readily applied to determine the most probable change in the aerodynamic characteristics of a wing that may be expected from the use of this type of flap.

From the appendix an estimate of the accuracy of the method can be obtained, as a comparison with full-scale data is given for lift and drag, and for the other flap characteristics the original curves have been reproduced.

Upon occasion it is required to find the drag of aircraft components which are normally retracted in level flight, such as flaps and undercarriages. The usual method of obtaining this drag from flight tests, is to measure speeds in level flight under known engine conditions with the component retracted and then with the component down. The difference of speed so obtained, together with the known test conditions then yields an answer. The disadvantage of this method is that, as the difference of speed is usually large, incidence changes and airscrew and engine assumptions cause discrepancies in the results. The difference of power between different engines of a type can be ± 5 per cent. A method is proposed which will enable most of the disadvantages of the older method to be overcome.

The structural problems of metal aircraft design largely centre round the difficulty of making efficient compression members. This difficulty is accentuated when loads are small in relation to the size of the structure. For example, the diameter of an aeroplane fuselage cannot usually be less than the height of a man, which results in such small forces at the surface of the shell that the lightest practicable beam is quite disproportionate to its strength.

As a measure of load in relation to size, it is convenient to use a quantity that we suggest may be called the “ structure loading.” This quantity, due to H. Wagner (8), is simply the square root of the applied load divided by a characteristic dimension (such as the length) of the member.

Petrol is a highly inflammable fluid which can be ignited easily and whose vapour forms with air a highly explosive mixture.

There are, however, possibilities of giving a certain amount of protection against such fire risks created by damage to petrol tanks.

The three main problems which have to be dealt with are:—

1. (i) To exclude an ignition which may be caused by bullets hitting a petrol tank;

2. (ii) To seal leaks caused by bullet holes;

3. (iii) To prevent fire after a crash.

These problems are by no means new. A large amount of research and experimental work has been devoted to their solution.

A general method of solution for rectangular plates with clamped edges and any kind of loading has been developed by Professor S. P. Timoshenko. The present paper gives the results of calculations using this method for the maximum deflection, moment, and edge shears for rectangular plates of various proportions with all four edges clamped and loaded by a single concentrated load at the centre. Similar data for a clamped rectangular plate with a uniformly distributed load have been given by I. A. Wojtaszak and also T. H. Evans. A report of an experimental investigation of this problem with some analytical results has been given by R. G. Sturm and R. L. Moore.