The hot-wire anemometer is one of the few instruments which can be used to make velocity measurements in turbulent and unsteady flows. However, the probe supporting the wire inevitably interferes with the local flow and it has been found that the effect of this interference on the reading of the anemometer varies considerably as the orientation of the probe to the flow direction is changed (the wire itself being maintained in the same direction). This leads to errors in any measurements taken where the instantaneous local flow direction differs significantly at any time from the direction for which the anemometer was calibrated. Such errors are quite separate from, and in addition to, errors due to finite wire length, incidence of the wire to the local stream direction, etc.

Experiments on the influence of added narrow band sounds on the subjective assessment of noise were conducted at the NGTE 1966 Open Week and at the 1966 SBAC Show, with over two thousand people participating. The results clarify the apparently conflicting observations of earlier workers and show definitely that the presence of a narrow band content can significantly influence the degree of disturbance caused by noise of the type generated by aircraft near an airport.

Film cooling performance expressed in the terms of film effectiveness, is divided into initial and final regions containing respectively, the potential core and transition lengths and, the main decay region. The theory of Kutateladze and Leont'ev, which is representative of present approaches, is presented and shown not to apply to the film initial region. To resolve the questions, what is the minimum acceptable value of effectiveness in a given situation, and, what proportion of the resulting maximum permissible cooled distance is initial region, the film cooling of a hypothetical turbojet engine to power a Mach 2-2 cruise supersonic air liner is examined. Experimental data from cooling devices representative of current European practice are used in the analysis to give realistic results.

The solution of the general heat transfer problem for typical aircraft-type structures usually involves a finite element approach involving the solution of a large number of simultaneous equations. This may be performed by calculation using a digital computer or with an analogue simulator. For calculations involving both spatial and time intervals considerable care must be exercised to obtain solutions which are not divergent or oscillatory and this usually involves very small time increments and a large total computation time.

A simplified analysis indicates that the increase in profile drag of an aerofoil due to an isolated roughness element is, in general, different from the drag of the element measured on a flat plate with the same free-stream velocity. This “magnification” effect is caused chiefly by the effect of the pressure gradients on the boundary layer downstream of the roughness element.

The degree of magnification is not closely approximated by the ratio of local to free-stream dynamic pressure and, in many typical cases, the contribution to the drag due to roughness elements may be seriously under-estimated in this way.

Measurements of the effect of the initial boundary-layer thickness on the subsequent development of a turbulent boundary-layer in an adverse pressure gradient support the theoretical conclusions.

The pressure reading of a pitot tube in a supersonic flow corresponds to the position of the nose shock wave. Results of measurements of the shock stand-off distance are presented for three different streamwise Mach number gradients and for a uniform flow. These results are compared with existing theoretical values.

Rates of straining that are met in various engineering parts vary considerably from the very small values appropriate to creep to those amounts many orders of magnitude higher that are necessary during metal working processes. If laboratory testing is to follow the ideas of Taira on creep or McLellan on the effects of strain rate on the Ramberg-Osgood constituitive relationships it is necessary to enquire whether conventional testing procedures are adequate. Bodner and Rosen have shown that the phenomenon of discontinuing yielding is solely a function of a testing procedure which allows strain rate changes because of the interaction of specimen stiffness changes and the machine stiffness.

In ref. 1 an analytical solution was given for the one-dimensional distribution of temperature in a web of an idealised, symmetric, multi-cell wing structure when symmetrically heated. It was assumed that the convective heat transfer coefficient, h, was constant, but the adiabatic wall temperature variation was assumed to be either an instantaneous or a linear function of time.

The purpose of the present paper is to indicate the application of an alternative, approximate method of heat flow analysis due to Biot and to show thereby how the usefulness of the results in ref. 1 can be extended by eliminating the assumption therein that the skin temperature was uniform at any given time. The notation of ref. 1 is used with additions as defined in the paper.

Seide and Weingarten have, recently, shown conclusively that in isotropic cylinders, the critical stress for bending stability is essentially the same as that in the corresponding compressive stability problem although the buckle pattern is quite different. The deflection pattern obtained in ref. 1 as a result of elaborate computations presents certain interesting features. The dominant feature is that the buckle pattern is an asymmetric one, because any axisymmetric mode is incompatible with the bending problem.

Even at small angles of incidence, the flow separates from the sharp leading edges of a slender wing. These flow separations usually lead to the formation of two free vortex layers, joined to the leading edges of the wing and rolling up to form spiral-shaped vortex sheets above the upper surface of the wing. This vortex formation is illustrated schematically in Fig. 1. The streamlines on the vortex sheet follow helical paths. Smoke injected near the wing apex for flow visualisation remains concentrated close to the axis of the core of the vortex sheet.

Extensive use of reinforced plastics has contributed significantly to the success of small VTOL engines, notably the Rolls-Royce RB 162 series in whose compressors it has been possible to demonstrate their versatility and ruggedness. The choice of reinforced plastics for such obvious reasons as lightness, specific strength and corrosion resistance has been fully justified, while apparent deficiencies in stiffness and rain erosion resistance have not proved embarrassing.

Considerable use has been made of epoxy resin pre-impregnated sheet incorporating non-woven parallel aligned glass fibre, in particular the “Scotchply” series of “pre-pregs”. Exceptionally high fibre contents have been achieved, resulting in higher strengths and moduli than are usual with glass-reinforced plastics and furthermore, since Scotchply has proved so uniform in composition and ply thickness it has been possible to precision-mould most components with such accuracy that machining is eliminated—a major economic factor. Further economies have been achieved by minimising press curing times by the use of latent catalytic type hardeners for the resin systems which, in addition, confer long storage life and good heat resistance. The use of very thin sheet material has allowed much ply orientation within laminates, resulting in substantial and controllable variations in anisotropy, allowing properties to be tailored more nearly to meet the conditions imposed.

Every structural member deforms under the forces acting upon it. In this paper the Maximum Principle, recently developed by Pontryagin is applied to the problem of determining that profile of a beam that deforms least under its own weight. The beam considered is a solid horizontal cantilever having constant density, modulus of elasticity and width. The height is taken as an independent variable subject to the restrictions that it lies between an upper and lower bound and is symmetrically distributed about the centre-line. The method is attractive in that it can readily be extended to take account of variable density, width and/or hollow members by extending the control vector u defined in the analysis.

This note describes how a single constant temperature hot wire may be used for measurements of direction, velocity and turbulence in a two-dimensional flow. The wire probe is rotated by a servo motor which automatically sets the wire with its axis either in the stream direction or normal to the flow. The accuracy of setting the wire in the direction of the stream is about , and across the stream is about 1°. If the higher accuracy is demanded the velocity and turbulence measurements require a second setting of the probe, at 90° to the previous one. When less precision is acceptable, the angle, velocity and turbulence measurements may be taken at the single setting, normal to the stream.

The cylindrical tankage which forms part of the primary structure of missiles and space vehicles may be subjected to high compressive axial and bending loads. Since allowable compressive stresses for monocoque construction are generally low because of small bending rigidity and sensitivity to geometric imperfection, it is advantageous to use some type of wall stiffening. One such choice is the integrally milled-out 45° waffle. This consists of a square gridwork of ribs which form 45° helices over the cylindrical surface. The panel is generally milled out in the flat and then bent up to form a circular cylinder. An advantage of the 45° construction is the exclusion of buckling modes between hoop reinforcement; such modes are frequently critical for the 0° to 90° type of stiffening.

Experimental results are presented of the measurement of skin friction in fully developed turbulent flow in concentric annuli using Preston tubes situated on the inner and outer annular surfaces. Both Preston's calibration and Patel's calibration were used to evaluate the results. It was found that the latter gave excellent results. Several radius ratios were investigated with a reasonable range of the annulus Reynolds number. The good agreement was not affected by radius ratio or smallness of core tube within the range of these parameters investigated here.

Some experiments on the effect of boat-tailing on the pressure distribution round blunt-based aerofoils are described. The experiments were carried out at low speeds at a Reynolds number of 1.5 X 105. The wake was investigated with attention being paid to the vortex shedding, and to the distance downstream of the base where vortices form.

It is shown that the theory due to Nash predicts the effect of boat-tailing on base pressure quite well and that a correlation of drag coefficient, Strouhal number and base pressure proposed by Bearman applies to the models tested here.