Graphite is a material with unique high-temperature properties. Its strength and modulus increase with temperature up to about 2000-2500°C., in which range it exhibits a higher strength than any known material. The development of new types of graphite (such as pyrolytic graphite, deposited from the vapour) and impregnated graphites, are both significant to future technology. The highly-oriented pyrolytic graphites in particular, with their marked thermal and mechanical anisotropy, and their superior mechanical properties (60 x 103 lb./in.2 strength at 2750°C, with 60 per cent ductility) offer distinct possibilities, particularly in composite structures. Graphite sublimes directly from the solid to vapour phase at 3700°C. at atmospheric pressure. As the sublimation energy is high, it is, at least in theory, a very efficient ablating material. It also has a high creep resistance. Surface reactions limit its value, particularly in oxidising environments, and coatings, or other methods of surface protection, are of high importance. The modulus of graphite is low, being about one-half of that for lead at room temperature, increasing with temperature to a maximum in the region of 2200°C. The properties vary in detail from one type of graphite to another, and with the direction in which measurements are made. Graphite has a low density (about 2·0) and is comparatively cheap, so that its low temperature limitations (zero ductility at room temperature) may be offset, at least in part, by a re-appraisal of the design factors. It can be joined and fabricated fairly readily, and appears to be a material of great value to advanced high-temperature projects in the future, where composite assemblies of various kinds will be essential.

The main purpose of the paper is to show that standard techniques of heat and mass transfer theory, as practised in other branches of engineering, can be helpfully applied to the corresponding aeronautical problems, particularly those arising in connection with high-speed flight and rocket propulsion.

Stratford's analysis of the laminar boundary layer near separation uses two physical ideas. In the outer part of the boundary layer, where viscous effects are small, the development is given by the condition that the total head is constant along streamlines, apart from a second-order correction for viscosity. Near the wall, however, viscous forces must balance the pressure forces, and the profile adjusts itself accordingly. Quantitatively these ideas yield a simple formula for predicting separation, which has been found to be particularly accurate.

In this paper it is indicated how the same approach may be used to yield the full distribution of skin friction along the wall. Further, the effects of suction may be incorporated into the method. Physically, suction affects the outer part of the boundary layer in that the streamlines are drawn towards the wall when suction is applied. At the wall, the balance between viscous and pressure forces is influenced by the momentum of the fluid which is sucked away. When these effects are accounted for quantitatively, the resulting formula for the skin friction is still very simple.

Several examples are considered, and comparison is made with exact theory and with approximate results by other methods. It is indicated that the method has a useful range of validity.

As part of a general investigation into Coanda effect, a study has been made of the reattachment of a two-dimensional, incompressible, turbulent jet to an adjacent, inclined, flat plate. The jet separates from the boundaries at the slot lips and reattaches to the plate downstream, a phenomenon which is associated with the lowering of the pressure between the jet and the plate accompanying the entrainment of fluid there. It is found that the flow becomes independent of both the length of the plate and the Reynolds number when these parameters are sufficiently large: the flow, scaled with respect to the width of the slot, is then uniquely determined by the plate inclination. Two approximate theories are developed for the mean pressure within the separation bubble, the position of reattachment and the increase in volume flow from the slot: the agreement with experiment is fairly satisfactory. These theories are a development of Dodds's analysis for the reattachment of a jet to a plate offset from, and parallel to, the axis of the slot and, for the purpose of comparison, a limited study is also made of this flow.

A method of calculating turbulent boundary layers on infinite yawed wings is given, making use of a method of calculating turbulent boundary layers due to Spence and of an analogy between three-dimensional and axi-symmetric boundary layers. It is also shown that the displacement thickness is equal to that computed using chordwise components and that the streamwise momentum thickness is approximately equal to the chordwise momentum thickness. Shock-free flow and small boundary layer cross-flow are assumed.

A theory of incompressible flow with axial symmetry through the impeller of a centrifugal compressor is given more fully and more accurately that by previous writers. The theory is applied to compute the flow through two different impellers. For each design the experimental axial velocity distribution at the impeller eye is in good agreement with that computed numerically from the theory. This means that for the impeller assembly—not untypical of modern designs—considered here, theoretical calculation of intake axial velocity should provide a useful guide to designers.

Equations are derived which enable the stresses around a compactly reinforced hole in an infinite sheet to be computed. The only restrictions placed on the shape of the hole are that it has at least one axis of symmetry and that the region outside it can be conformally mapped on to the region outside the unit circle by a transformation function in the form of a polynomial of any order. This admits a very wide range of holes of practical importance, including circles, ellipses, and squares, triangles and rectangles with rounded corners. Two loading cases are treated; the first corresponds to uniform tensions at infinity in directions parallel and perpendicular to the axis of symmetry of the hole, and the second to uniform shear at infinity in these directions. Superposition of these basic loading cases enables the stresses to be determined for any uniform state of stress at infinity. The equations are in an ideal form for use with a digital computer.

This paper considers the problem of a reinforced circular hole in a bent or twisted flat elastic plate having small deflections. The reinforcement is chosen so that the ratio of its radial width to the diameter of the hole is small; this implies that there is no radial stress variation across sections of the reinforcement. Using the classical theory for bending, expressions are obtained for the boundary conditions at the junction of the reinforcement and plate. The expressions are general in so far as they are also applicable to plates under the action of lateral loading. The following particular problems are considered: —(a) uniform bending about one or two axes, (b) uniform twisting about two axes. For the particular case of uniform bending about one axis, data has been obtained for the stress concentration around holes for plates and reinforcements having a reasonably wide range of practical dimensions.

Recent free flight tests have provided information on the applicability of the sonic area rule to non-lifting configurations. An analysis of results for the sonic wave drag of swept-back, tapered wings has suggested that the sonic area rule is applicable to thin wing and slender body combinations provided that the product of the wing span/length ratio and the cube root of the thickness /chord ratio is less than unity. The wing span/length ratio was found to be a much more useful slenderness parameter than the aspect ratio.

When the wing has a round-nosed section, theoretical considerations predict the existence of a so-called leading edge drag force varying with Mach number. One would expect such a wing to have the same sonic drag-rise as the equivalent body with an identical cross-sectional area distribution. Some test results do not confirm this drag-rise equivalence. It appears that the transonic drag-rise of the leading edge force is much greater than that predicted by theory.

Tests on one basic wing-body combination, with different body cross-sectional shapes of the same area, suggest that the cross-sectional shape, of a smooth slender body, has no influence on the transonic wave drag of a configuration.

The fundamental frequencies of flexural vibration are determined for uniform isotropic rectangular plates that have free edges and pinpoint supports at the four corners. For the particular case of a square plate the lowest five frequencies and mode shapes have been determined. The second frequency is most unusual, since an infinite number of mode shapes exist with identical frequencies. Finite difference expressions, which simplify the treatment of the free boundaries for definite values of Poisson's ratio, are used in conjunction with extrapolation procedures to obtain the approximate solutions.

Experimentally it is shown that variation of axial spacing between neighbouring blade rows can alter the stall characteristics of a compressor stage. The stall cell patterns of this compressor stage are described in detail and changes of slope in stage temperature and pressure rise characteristics are explained in terms of stall propagation. Strain gauge tests on stator blades are used to indicate the possibility of large dynamic stresses occurring when the stall cell frequency is equal to the blade natural bending frequency.

The paper describes the development of a small-bore probe microphone, the frequency response of which is smoothed by arranging that the probe tube is terminated in its characteristic impedance.

In Reference 3, attention is drawn to the difficulties of measuring the streamwise extent of the roughness from insects. The present paper deals with the problem theoretically for an aerofoil in two-dimensional, incompressible flow. A tentative approach to the determination of effective excrescence height downstream of the leading-edge zone is also advanced. The application of these investigations, in conjunction with the analysis from Ref. 3 regarding the critical conditions for premature transition, leads to estimates of the amount of significant roughness which are in good agreement with flight observation.

The slender-body theory is applied to the determination of the unsteady aerodynamic forces on slender wing-body combinations when the body is stationary and the wing is deforming harmonically. The uniform body is circular in cross section, the wings are placed symmetrically and the wing deformation mode is necessarily symmetrical about a vertical plane through the body axis. The frequency parameter is restricted to that range for which the “cross-flow” potential equation is Laplace's equation. The case of an allmoving slender control surface on a body is treated in detail and numerical results are given for the forces and moments on body and control surface when the control plan form is triangular: the presence of a gap is neglected.

The analysis of the stability of second-order non-linear systems by Poincare's method of singular points in the phase plane is briefly presented and used to determine stability criteria for the short-period motion of an airframe having non-linear normal force and pitching moment curves. These results are then compared with those obtained by the application of quasilinear stability theory.

The pressure distribution on the wing of a centrally mounted wingbody combination due to anti-symmetrical body shaping (with respect to the wing plane) can be calculated in the case where the wing leading edge is supersonic by using the fact that the upper and lower halves of the flow field are independent and that each half can be treated by quasi-cylinder theory. When the leading edge is subsonic the anti-symmetric body shaping produces flow round the leading edge. This paper is concerned with wing pressure distributions including the effect of this flow round the leading edge.

Analytical solutions for the deflections and bending moments are derived for elastic struts with linearly varying axial loading applied either (a) along the undeformed straight strut axis or (b) at the strut axis in the deflected position. The instability of struts with various end conditions is discussed and the stability criteria are given as a series of curves relating the maximum compressive axial load with the maximum distributed loading. Furthermore, explicit formulae for the deflections and bending moment distributions have been compiled for several typical cases of lateral loading combined with the axial loading of the type (a), which is of some practical importance in aircraft structures. A numerical example is included to show the practical application of the general analysis to a typical strut problem.

In assessing the merits of multiple versus single engines the possibility of a loss in reliability due to the use of more working parts must be considered. This paper gives a mathematical treatment of this problem and numerical examples illustrate the method.

Transformation functions, in the form of three-term polynomials, are derived for the conformal mapping of the region outside a square or an equilateral triangle, having rounded corners with varying and controllable degrees of curvature, on to the region outside the unit circle.

A set of design charts is presented for the calculation of transient temperature and thermal stress distributions in thermally thick plates subjected to aerodynamic heating.

The method is particularly useful for determining temperatures and thermal stresses in plates with an arbitrary variation of the heat transfer coefficient and the adiabatic wall temperature of the boundary layer. The present method is based on repetitive applications of the exact analytical solution to a unit triangular variation of the adiabatic wall temperature and a constant heat transfer coefficient. The actual variation of the adiabatic wall temperature is represented as a series of straight lines while the heat transfer coefficient is approximated by a step function. The temperature distribution through the plate is separated into linear and “self-equilibrating” temperature distributions to facilitate thermal stress calculations; these distributions can be obtained directly from the design charts presented in this paper.

The general principle of this semi-numerical method is also applied to thermally thin plates subjected to arbitrary heating conditions.