The possibility of obtaining two-dimensional turbulent jet flow from a series of closely-spaced uniform holes in line has been investigated both theoretically and experimentally. The case studied was that of a jet discharging into still fluid of similar density at incompressible speeds. Such a quasi-two-dimensional jet is a particular example of a multiple-interfering jet group.

This paper describes rig tests made to evaluate the effectiveness of the cockpit insulation and cooling system of a high-speed aircraft.

The tests showed the dependence of cockpit internal temperature distribution and heat pick-up on the cooling air mass flow and inlet temperature. Analysis of the test data showed that there was considerable heat leakage into the cockpit; the heat leakage increased with cooling air flow and constituted two-thirds of the heat entering the cockpit when the air flow was moderately high (20 lb/min). Some of the leakage heat entered the cockpit through equipment mountings but it was evident that other leakage paths existed. One more obvious heat leakage path, at the canopy, is illustrated.

The tests also showed that the internal heat transfer coefficient increased with air flow, reaching a value of 2·5 C.H.U./hr ft2°C when the flow was 20 lb/min.

The flow past an unyawed, flat, delta wing with subsonic edges, carrying a slender half-cone mounted centrally underneath, is found by using the method of supersonic conical flows. The wing plane is taken to be at zero incidence, but it is shown that wing incidence, camber, twist and thickness effects may be incorporated by superimposing on this solution the field of the isolated wing with these properties. The results are shown to agree with previous work on configurations with sonic or supersonic leading edges for the common case of sonic edges. The leading edge loading may be made zero to secure attached flow at any lift coefficient by using camber and incidence, but it is shown that there is a certain positive lift coefficient for which a negative incidence alone will suffice.

The stress distribution in rotating circular discs containing a central hole and a symmetrical array of non-central holes is determined by numerical solution of the equations of generalised plane stress. Particular attention is given to an annulus containing the holes and of width approximately eight hole diameters, in which the full two-dimensional equations are solved. The region outside this annulus is treated as radially symmetric and the stresses there are determined from a simpler one-dimensional model. Stress distributions are reported for uniform discs of fixed geometry containing 10, 20 and 45 holes. Results are also obtained for 20-hole discs of non-uniform thickness comprising a uniformly tapered disc, a disc with a thickened annulus containing the holes, and a uniform disc with each hole surrounded by thickened bosses. As a check on the numerical method, calculations have been carried out on a disc of identical geometry to one examined photoelastically bv Leist and Weber with good agreement. The effect of changing Poisson's ratio for this particular disc is also examined.

The following theoretical investigation is concerned with the stability of the flow through a system composed of a multi-stage axial flow compressor followed by a throttle.

Such an investigation was carried out by Pearson and Bowmer in 1949. In 1962 Pearson’s work on the analysis of axial flow compressor characteristics, and the accumulation of empirical data regarding factors affecting the surge line, re-awakened interest in the possibility of predicting the surge line of a multi-stage axial flow compressor-throttle system.

In this paper the equations governing the stability of flow at any operating point in such a system are obtained by applying Kirchhoff’s laws to the associated electric circuit at that operating point, and the analysis is applied to a wide range of flows of the calculated characteristics of a seven-stage axial flow compressor.

A study of the simplest compressor-throttle system is given, in which the equations of motion of the system are derived mechanically and electrically, and the range of validity of the equations and their stability are discussed in order to bring out the relation between the mathematics and physics of the simple system before applying these methods to multi-stage axial flow compressors.

For the relatively simple electrical representation used in this paper for an axial compressor of n stages, there are shown to be 2n possible values of p, the transient rotational frequency, and these are determined over a sufficiently wide range of flows on the seven-stage compressor studied.

As a result, a region of the compressor characteristic map can be marked out in which all the values of the transient rotational frequency have their real parts less than zero, corresponding to stability of operation, a region where at least one of the values of p is real and positive corresponding to non-oscillatory instability of operation, and an intermediate region where some of the values of the rotational frequency p are complex with positive real part, corresponding to oscillatory instability of operation.

It is suggested that the non-oscillatory instability found here is associated with the surge and the line of inception of non-oscillatory instability with the surge line.

A large deflection theory is established for the bending and stretching of aeolotropic plates with elastic moduli that vary in the x, y, z directions. The theory is formulated in terms of two simultaneous non-linear differential equations, with variable coefficients, for the transverse deflection w and an Airy stress function F. Small deflection heterogeneous plate theory and some other sub-classes of plate problems are formulated by specialising the general non-linear plate equations. Effects of initial irregularities and of a variable elastic foundation are considered.

A combined experimental and theoretical study has been made of radial compressible flow without swirl between parallel discs when the fluid velocity is everywhere subsonic. It constitutes an extension of a previous analysis, wherein the flow was considered to be incompressible.

A similarity solution for the radial pressure distribution is shown to be possible only in special cases where certain terms in the equations of motion can be neglected. Approximate solutions are obtained for the laminar and the turbulent radial pressure distributions, using an integral momentum method. Both theories agree well with experiment. The critical Reynolds number for reverse transition is found to be approximately the same as that for incompressible flow in radial channels and circular pipes. The radial pressure level for compressible radial flow between parallel discs is found to be less than that for incompressible flow at the same mass-flow rate. The non-linear form, in pressure, of the compressible solution can be shown to account for this result.

Jet flap theory for thin aerofoils has been extended to include the effect of jet entrainment on the external flow when the jet is blown over the upper surface of the aerofoil. The effective camber of the aerofoil is increased by the sink effect due to entrainment and the increase of lift at zero incidence is proportional to the square root of the jet momentum coefficient. Formulae and charts are presented to facilitate the determination of the increments of lift and pitching moment due to this effect. The theory is shown to be in first-order agreement with the exact solution for a circular-arc aerofoil of small camber with distributed sinks on the upper surface.

The new theory is compared with four old sets and one new set of experimental data. It greatly improves the accuracy of prediction for cases where the incidence and flap angle are small. The new theory substantiates the usefulness of a small flap in applications of the jet flap principle.

The existing data relating to stresses round pins in holes are reviewed and co-ordinated.

The first part of the paper deals with pins of which the diameter is less than half the plate width and which are no better than a push fit in the hole; the second section deals with larger pins; and the third section deals with interference fit of the smaller pins. The theoretical background is sketched in two appendices, and the results of three series of photoelastic measurements of the effects of clearance of the pin in the hole are described.

An analysis is presented of the buckling of a rectangular plate under combined biaxial compression, bending and shear. The sides of the plate to which bending is applied are simply-supported, the other two sides being supported by arbitrary edge members. The solution is expressed in a form suited to automatic computation and some sample results are given.

Donnell type stability equations for thin circular orthotropic conical shells are presented and solved for external pressure loading. The solution is likewise applied to stiffened conical shells by consideration of equivalent orthotropic shells. Typical cases are computed and compared with corresponding isotropic shells. Correlation with equivalent cylindrical shells then yields a simple approximate stability analysis for orthotropic or ring-stiffened conical shells under hydrostatic pressure.

The concept of an idealised “unique” stage characteristic is employed to analyse and understand the performance of axial compressors. It is shown that there exists a “matching” line across the compressor characteristics at any point of which all the stages are operating at the same point of their “unique” stage characteristic, and that this matching line is readily obtainable, almost from inspection. Simple calculations lead to a derivation of both this “unique” characteristic and the effective area ratio of the compressor. The behaviour of the stages at other points than the matching line is readily understandable and presents a simpler picture of compressor performance than is often obtainable from actually measured stage characteristics.

An analysis is presented which enables the boundary-layer thickness parameters of a re-attaching shear layer to be determined when the free-stream flow upstream of the base is supersonic, the base pressure is known, and die initial boundary layer is turbulent. The application of this analysis to some experimental results, on the flow behind blunt-trailing-edge wings and over a back-step where both the base pressure and the initial boundary layer are known, would appear to indicate that the re-attached profile could be specified by one parameter, namely the transformed shape parameter, the transformation used being a turbulent analogue of the well-known Stewartson-Illingworth transformation of the laminar boundary layer and where the shape parameter is defined as the ratio of boundary-layer displacement to momentum thickness. By adopting a value of the shape parameter in advance, it is possible to use the analysis to determine the base pressure by an iterative process and so, on this basis, it is suggested that this analysis is used to replace the existing recompression criterion of the Chapman-Korst model of separated flow which aims to predict base pressures and is known to be capable of improvement.

As part of this investigation, an improvement had to be made to the existing compressible turbulent shear layer velocity profiles of Korst and others and this was achieved by means of the compressibility transformation.

The flows induced by incompressible jets have been analysed for two-dimensional and axisymmetric jets emerging at right angles from an infinite plane surface. In particular, the pressure distribution on the surface has been predicted by replacing the jet with a line sink of variable strength. The theory takes account of the size of the blowing slot or nozzle and is in good agreement with experiment in both cases.

This paper covers an investigation into a new way of assisting the pilot to hover an unstabilised jet-lift aircraft. In a normal control system the aircraft is controlled by a moment which varies linearly with stick position. In this paper control systems were considered in which the controlling moment was a linear function of the position and rate of change of position of the stick. The investigation was primarily limited to optimising such a system with freedom restricted to rolling motion and the associated lateral displacement. Effects of limited control power and engine failure were considered. With the optimum control the mean errors in position recorded when hovering were under half those achieved with a normal manual control and the results were almost as good as for an ideal rate-stabilised aircraft. Extending the investigation to both pitch and roll planes led to an even greater improvement, even when the effects of gyroscopic coupling between the two planes were taken into account. The precise engineering of a “mixed” control system of the type discussed needs much thought and it will require considerable development, but the results achieved indicate that it may prove an attractive alternative to an autostabiliser

A study has been made of the deformation of skew slabs using a pure resistance electrical analogue with a square mesh. The three commonly occurring boundary conditions of simply-supported, clamped, or free edges have been represented on the analogue, and the results of typical problems with these edge conditions are given.

The same fluid kinetic-chemical parameter is shown to describe the peak velocity conditions at which cylindrical flameholders can sustain a flame in both gaseous and liquid fuel/air mixtures. Experimental data indicate the existence of three possible aerodynamic conditions which may aflect the flame stabilisation process. In both gaseous and liquid fuel flame stabilisation an increase in stabiliser size is shown to increase the peak stable burning velocity.

Existing experimental data on the occurrence of attached flow around the leading edges of swept plane thin wings is summarised. It is shown that the change from separated to attached flow occurs while the Mach number normal to the leading edge is subsonic and that the Mach number required for the change increases with incidence. It is also shown that the Mach number and incidence conditions which hold on the boundaries dividing the two types of flow on wings with straight edges also apply locally at points on the leading edges of “Gothic” wings. The implications of this result on the flow over general curved-edge wings is briefly discussed.

A simple circuit is described which gives a direct meter reading of the intermittency factor in a turbulent flow. The operation and setting-up procedure of the instrument is discussed.

The paper deals with the behaviour of three shock waves meeting at a point in a perfect gas. It is shown that the equations of motion can be reduced to a single polynomial equation of degree 10. The real roots of this equation are studied to determine their physical significance. In addition, the appearance of degenerate shock systems is shown to be associated with the formation of certain multiple roots of the polynomial equation.