In parallel with the work undertaken at the University of Southampton on the study of the noise from subsonic jets, a further investigation has been made into the noise from supersonic or choked jets. Whereas in the former case the noise is due to the turbulent velocity fluctuations and thus contains a wide spectrum of random frequencies, the latter case contains, in addition to this wide spectrum, noise caused by turbulence-shock wave inter-action.

Once the critical pressure ratio of a jet has been attained, i.e. that pressure ratio which produces sonic or supersonic flow in the jet pipe just before the orifice, the structure of the jet depends crucially upon the nozzle shape.

A problem which occurs frequently is that of choosing a suitable shape for a duct, such as a wind tunnel contraction or an air intake. Basically similar problems, involving potential flow fields, occur in other branches of engineering, particularly in electrical engineering, and the electrolytic tank is now established as a tool which may usefully be employed in their investigation. The use of the simple shallow tank is limited to those fields which can be treated as two-dimensional or axisymmetric, but many problems fall within these categories.

In forming a duct model for the electrolytic tank the walls of the duct are represented by insulating surfaces, and electrodes are positioned to represent two suitable velocity potential surfaces up- and down-stream of the duct. To represent a sector of an axi-symmetric duct the base of the tank must be inclined at 3°-5° to the horizontal and the water line on the tank base then represents the axis of symmetry.

A Similarity in properties will be demonstrated between the moments and product of inertia of a lamina and the components of plane stress. It is now a well established fact that in two-dimensional stress theory it is advantageous to use certain combinationsof the stress components.

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First, it will be shown that the use of similar combinations of the moments and product of inertia leads to a useful graphical means of obtaining these quantities for awkwardly placed axes and also to some results not normally to be found in the text-books. Secondly, an application to photoelastic analysis is also made.

This note describes the design and operation of the rig used to test and develop the Canadair CP 107 “Argus” hydraulic systems. The Argus is a maritime reconnaissance version of the Bristol Britannia, redesigned and built in Canada for the R.C.A.F. by Canadair Ltd., Montreal.

It is common practice on both sides of the Atlantic to test as thoroughly as possible before flight the various systems of a new aircraft. This is usually accomplished by the construction of space frames on which the systems are assembled and subjected to simulated flight conditions.

It was decided during the early stages of the design of the Canadair Argus that the deviations from the original Britannia hydraulic systems were great enough to warrant a full scale hydraulic systems test rig, in as compact a form as possible, which would give the desired test conditions and results.

Relatively simple theoretical procedures enable the pressure loss of the simpler forms of baffle used for flame stabilisation in combustion chambers to be computed, while tests made under non-burning conditions are available for ad hoc testing of the geometrically more complicated forms. Both these methods, with the appropriate corrections for the momentum changes due to burning, have in the past proved suitable for the prediction of the drag of aero-engine combustion chambers under operating conditions. Recently, however, it has been found that under certain conditions, associated especially with very high rates of heat release, the drag of combustion systems can become very much greater than would be predicted either from simple theoretical considerations or from conventional “cold” tests, and vary in an unexpected manner with the normal flow parameters.

In a recent publication the problem of denning the thrust of a jet engine, or the internal drag of a ducted body, is considered in some detail. This Report was compiled by a special Panel of the Aeronautical Research Council and, during its discussions, concern was expressed at the position regarding the definitions of certain quantities having the dimensions of pressure; in some cases the most widely-used definitions were not rigorous, in others a variety of names were in use. In addition, some quantities of considerable value in fluid dynamics were either unnamed or, had been given misleading names. This state of affairs was felt to be undesirable. As a result, the A.R.C. set up a small Panel to consider the matter more fully. The Panel contained representatives from both of the two important groups interested in obtaining uniformity in this field, those concerned with external aerodynamics and those primarily interested in internal flow problems.

The title of the paper infers that, in general, aircraft design is not fundamentally suitable for production and, for various reasons, design for quick and easy production must be the result of compromise. In practice optimum design and easy production never seem to coincide exactly. Refinement of design for production covers many broad conceptions.

In its broadest sense it should be the uppermost consideration from the very beginning of the design of the prototypes. The primary object of designing any new aircraft, civil or military, is to be able to produce a series of the most efficient fully approved and tested production aircraft at the highest possible rate in the shortest possible time. The first part of the paper, therefore, deals with the design and production planning phases of a new aircraft when, by joint consultation with the production departments, the design is viewed from the production angle from the outset.

The effects of end loading on an arbitrarily distorted flexing beam are investigated, with particular reference to bending moment distributions along the span of the beam, and the effects of lengthwise strains. Cross-spring and “ spider ” type flexural systems are treated in detail. Curves are presented which enable the stiffness and stressing of any flexural system to be calculated for any loading conditions, subject only to the restriction that deflections are small, so that the linearised beam equation may be applied.

The solution of a set of m ordinary linear differential equations of the nth order having variable coefficients can be provided by the use of a set of m2 indicial admittances and the solution is derived from that given by the use of impulsive admittances. The indicial admittances are easily derived from the impulsive admittances and vice versa.

An “indicial admittance” is a function representing the response of a system to an applied “input” in the form of a “unit step function” and the concept is due to Heaviside. When the indicial admittance is known, the response to an input which varies with time in an arbitrary manner can easily be found. Similarly, the response can easily be found when the “impulsive admittance” is known, where this is the response to a unit instantaneous impulsive input.

This Note describes the design and construction of a “ Tilting Test Bed ” for helicopter engines for Alvis Ltd., for which some onerous requirements were laid down (Fig. 1). The most important of these were that:—

1. (i) Readings of engine torque and b.h.p. must be given by a dynamometer, so as to provide accurate data.

2. (ii) The engine, and therefore the complete test plant, must be capable of being tilted, while running, through 120°—from the horizontal to the vertical and 30° beyond.

3. (iii) During the tilting process the engine throttle setting must remain unaffected and the readings of torque, b.h.p., fuel consumption, and so on, continue with unaltered accuracy.

4. (iv) A powerful fan must be incorporated, for cooling certain types of engine not provided with selfcontained fan, and the air-blast must continue unchanged during tilting.

While congratulating F. H. Robertson on his technical note in the January 1957 JOURNAL (page 52) on “short cut” methods of estimating aircraft direct operating costs, which certainly give a very rapid approximate answer, I am somewhat uncertain as to its practical value. In the analysis of possible projects, having established the general range of operating costs, which can usually be done by a combination of experience and inspection, interest lies primarily in the effect on cost of operation of such factors as change of type of power plant, fuel used, what complexity of structural designs is permissible from a cost aspect, how much is desirable from a weight saving aspect and so on. Unfortunately, Mr. Robertson's formulae do not allow such comparisons to be made primarily because of the assumptions made.

There have recently appeared a number of papers and notes bearing on the merits of turbo-prop vis-a-vis turbo-jet transport aeroplanes, mostly exhibiting specific favour towards one type or the other, but disagreeing among themselves and all based upon investigations which, it is contended here, are incomplete.

The basis used has invariably been a comparison of the relative direct operating costs per unit of weight (i.e. per passenger or per ton) per mile, and the relative earning capacities have not been considered. This, it is believed, is a serious omission which can only lead to a distorted assessment and erroneous conclusions.

The author has been associated with a thorough-going investigation into the relative merits of twin-engined turbo-prop, turbo-jet, and reciproeating-engined aeroplanes, all designed on a strictly comparable basis to do, as far as the performance of suitable available engines would allow, the same job in terms of capacity payload and associated operational stage length, take-off and landing performance and one-engine-inoperative climb performance in given atmospheric and altitude conditions.

Flight tests were made to verify a method of predicting pressure cabin leakage rates at altitude from ground test results proposed in a Technical Note published in The Journal in August 1954 The results showed very good agreement with the theoretical treatment of the problem.

It is therefore recommended that the method of the Technical Note be used to estimate the change of leakage rates with altitude and also to estimate the time for the excess pressure to fall during a descent with no pressure supply.

The sound pressures, associated with a jet efflux, set up stresses in an aircraft structure which are essentially those occurring in the resonant condition of the structure. These sound pressures contain components in a wide range of frequencies and the resultant stresses in the structure have low magnitude by normal structural standards but, because of the high frequency, they are very important from the fatigue standpoint. A rigorous theoretical approach is being developed at Southampton University which, although extremely difficult to apply to actual aircraft structures, emphasises the factors which must be considered. However, for immediate practical use it is suggested that an approximation can be made to render the problem tractable. It is considered that in an actual aircraft the effects are primarily local and on this assumption an approximation is developed and compared with test results. Reference is made specifically to the de Havilland Comet aircraft.

This paper considers the problem of a reinforced elliptical hole in a plate under the action of a principal stress system of the type found in cylindrical and ellipsoidal pressure vessels. That is, stress systems in which the ratio of the principal stresses is not greater than two to one. It is shown that when the ratio of the major and minor axes of the ellipse can be chosen arbitrarily, practical reinforcements can be designed to give a maximum stress around the hole which is only slightly greater than the maximum stress in a similarly loaded plate with no hole. General expressions are obtained for the stress distribution in the plate around the hole, for the stress acting on a normal cross section of the reinforcement, and for the cross-sectional area of a reinforcement which gives a small stress concentration. These are used to find the variation in the stress distribution around the hole due to reinforcements having different cross-sectional areas when the applied principal stresses are in the ratio of two to one and Poisson's ratio for the material of the plate and reinforcement has practical values.

It is suggested that a convenient way of presenting the results of fatigue tests in which two different stress amplitudes are applied alternately is to plot log N against log (n1/n) where N is the total cycles to failure and (n1/n) is the fraction of cycles run at the high stress.With these co-ordinates, a simple geometrical construction gives a safe design method for the two-stress level system using only the conventional S-N curve and the value of (n1/n) expected to be encountered in service. If N1 and N2 are the lives at the high and low stresses as read from the S-N curve, one point may be plotted at (log N1, 0) since this represents the programme when all cycles are at the high stress. On the assumption, shown to be justified, that less than one cycle of high stress per 10,000 total cycles would not significantly affect the life at the low stress, a second point is plotted at (log N2, 4). The straight line joining these two points is always found to predict safe values of N for any value of (n1/n).This conclusion is checked against a wide range of experimental results taken from six different sources in the literature covering rotating-bending and push-pull tests, ferrous and non-ferrous metals, any order of stressing and length of programme cycle from 50 up to 5 million. This last feature means that the length of the programme cycle in service need not be known. All that is required is the proportion in which the two stress amplitudes are mixed. The average value of the ratio (experimental life/predicted life) for the data examined is 1·8, the extreme values being 1 and 56. By plotting in three dimensions an equation is also developed for the three-stress level spectrum and a suggestion is made for an extension of the method to multiple stress levels.

It is well known that when a plane air jet of high velocity blows over a convex surface, such as the upper surface of a lowered trailing edge flap, it flows along the surface, remaining attached for a considerable length; in the case of the flap, it adheres right to the trailing edge. This effect, usually described as the Coanda effect, has been utilised in the attempts to increase lift at constant incidence by “flap-blowing” in which the high velocity air jet is ejected either through slots in the flap shroud, or tangentially out of the flap surface.

The low velocity air stream approaching the flap over the wing upper surface, which in the absence of “blowing” would tend to separate at the flap shroud, is entrained by the high velocity jet, and the incipient flow separation over the flap thereby suppressed. The energy contained in the “blowing” jet can, in addition, also produce “supercirculation.“

It is shown that when heavily curved circular beams of constant cross section are subjected to a compressive load along the “open” diameter, the load for the start of significant deformation as indicated by the reduction in diameter along the line of action of the load, may be estimated by the use of rectangular plastic stress blocks based on the yield stress of the material. Comparison of the experimental values obtained on specimens of duralumin and mild steel with those predicted theoretically are, on the whole, only slightly underestimated.

It is shown that, on the basis of the data provided by two sets of influence coefficients for a wing (or other surface)–an “elastic” set giving deflections (and hence incidence angles) in terms of applied loads, and an aerodynamic set giving aerodynamic loads in terms of incidence angles (or other linear function of the displacements)–all “static” aeroelastic problems can easily and expeditiously be solved by the use of a digital computer. It is also suggested that the same method of approach may well be used for solving oscillatory aeroelastic problems such as flutter.