A velocimeter is described based upon the same principle as the five-hole Conrad tube yawmeter. However it uses a single tube, cut obliquely at 45°, that rotates about its own axis together with an outer closely-fitting concentric tube, cut normal to its axis, that can be slid forward over the mouth of the chamfered tube. The small measuring volume, about 1 mm3, and well-defined measuring point enable the probe to be used with some confidence in regions of high shear. Some typical results obtained with the probe in a highly-sheared, three-dimensional flow are given.

An investigation of the four-bladed Dowty Rotol R212 propeller (NACA 16 sections) has been made at l/5th scale (0·7 metre diameter) in the RAE 1·5 metre acoustic tunnel. Propeller power absorption and thrust have been measured over a range of rotational speeds up to 8000 rev/min at mainstream speeds from 15 m/s to 60 m/s for a range of blade settings. Slipstream wake surveys show outward movement of the position of the peak pressure as propeller loading is increased.

Noise analysis demonstrates the predominance of multiple tones whose number and intensity increase with helical-tip Mach number. An empirical formula shows that the fundamental tone SPL varies with tip speed and power loading in an identical manner to that observed on a modern ARA-D section propeller.

Propeller power absorption and thrust have been measured over a range of rotational speeds up to 1400 rev/min at mainstream airspeeds up to 48 m/s for a wide range of blade setting. In addition, cruise conditions have been simulated by maintaining the appropriate ratio of propeller tip speed to the airspeed. Generally good agreement is shown between these full scale experiments with results obtained at 1/5 scale though the effect of Mach number is observable when the propeller is highly loaded.

Noise analysis demonstrates the predominance of multiple tones whose number and intensity increase with helical-tip Mach number. An empirical formula shows that the fundamental tone varies with tip speed and power loading in a similar manner to that observed on a 1/5 scale model.

For a 15° boat-tailed heat shield or payload shroud, the unsteady, compressible Navier-Stokes equations are numerically solved. For a Reynolds number of the order of one million and for four different freestream Mach numbers, the time integration is performed for 64 x 30 grid computational domain. The entire flow-field is analysed to delineate the gross dominant characteristics. It is observed that the flow in the boat-tail region is completely attached for Mach number 2·47 and beyond whereas it is separated for lower Mach numbers. A comparison between the numerical solutions and the available experimental results is provided and, for incipient flow separation conditions, an evaluation of unsteady surface pressure data is presented.

In the Queen Mary College gust-tunnels, unsteady flows are generated by oscillating flaps attached to the downstream upper and lower surfaces of the contraction nozzle of a semi-open test section, open-return, low-speed wind-tunnel. The flow perturbations produced on the mainstream of velocity U∞, are of the travelling-wave type, with wave-velocity Q. Attention is drawn to the contradictory early measurements of Q/U∞. New data are presented which show Q/U∞ apparently diminishing along the tunnel axis from a high value near the nozzle exit to an asymptotic value of about 0·6 far downstream. Using a digital phase meter especially developed for the purpose it is shown that the explanation for this behaviour lies in the two-dimensional nature of the wave process in the region of the flaps.

A numerical panel method has been developed for the calculation of the two-dimensional potential flow about an aerofoil-spoiler configuration undergoing periodic motions. The cases investigated include a spoiler oscillating in simple harmonic motion on a steady aerofoil, and an aerofoil-spoiler configuration oscillating in overall pitch and normal translation.

The method is a perturbation method about the steady in viscid flow model of an aerofoil-spoiler configuration.

Elements of linear piecewise vorticity are placed on the aerofoil and spoiler surfaces, while the separation streamlines are replaced by elements of uniform vorticity. Time varying vorticities are shed from the spoiler tip and aerofoil trailing edge, along the upper and lower separation streamlines respectively, through the wake closure region, combining in a planar wake downstream of the wake region.

Predicted forces and moments are broadly in agreement with trends of experimental results available in the literature.

An axiomatic aerodynamic model has been developed for the general motion of a two dimensional aerofoil as it passes in and out of stall, which gives realistic unsteady loads as compared to experimental values. A non-linear set of aerodynamic derivatives with time delays have been derived from the axiomatic aerodynamics. ‘Actual’ and ‘predicted’ dynamic responses of an aerofoil, spring restrained in torsion, following an impulsive input show similar trends, including limit cycle oscillations, although there is a slight difference in frequency and a difference in the magnitude of the initial impulse required to trigger the limit cycle.

A surface singularity method has been formulated to predict two-dimensional spoiler characteristics at low speeds. Vorticity singularities are placed on the aerofoil surface, on the spoiler surface, on the upper separation streamline from the spoiler tip and on the lower separation streamline from the aerofoil trailing edge. The separation region is closed downstream by two discrete vortices. The flow inside the separation region is assumed to have uniform total head.

The downstream extent of the separated wake is an empirical input. The flows both external and internal to the separated regions are solved.

Theoretical results have been obtained for a range of spoiler-aerofoil configurations which compare reasonably with experimental results.

The model is deficient in that it predicts a higher compression ahead of the spoiler than obtained in practice. Furthermore, there is a minimum spoiler angle below which a solution is not possible; it is thought that this feature is related to the physical observation that at small spoiler angles, the separated flow from the spoiler reattaches on the aerofoil upper surface ahead of the trailing edge.

The work to be described forms part of an experimental study into the overall influence that the end region of a flap has on the flow of a high lift wing system. Here, our attention will be restricted mainly, but not entirely, to the tip flow of the flap itself.

The study has been a collaborative venture over the past three years between British Aerospace-Civil Aircraft Division and the Aerospace Engineering Department of Bristol University, the latter funded by an SER.C research grant. BAe's major contribution was the undertaking of the main test programme, with University participation, using a large swept wing, high lift half model in the Filton low speed wind tunnel. The model was loaned by RAE for the study.

The axiomatic aerodynamic model in Part 2 has been extended to an aircraft moving in a general coupled longitudinal-lateral motion but with the flow attached throughout the motion. Results are presented for the responses of a swept wing aircraft due to separate rudder and aileron inputs. For some aileron responses there are inexplicable significant differences between actual response and predicted response, based on derivatives, in roll rate and pitch rate. No such differences appear with rudder inputs.

Wing-tip vortices are of considerable practical importance, primarily because of the associated induced drag. A number of investigations have been made in the past into the development of vortices trailing behind wings, but little attention has been given to the mechanism of generation of such vortices. In the region of a wing-tip the boundary layers, which are usually turbulent at full-scale, separate from the wing surface in a rather complex manner to form a vortex sheet which rolls up into the tip vortex. Although the gross effects of the tip vortex can be predicted by inviscid theory, a detailed understanding of the behaviour of the viscous layers around wing-tips could lead to better optimisation of their design.

A planar vortex sheet model for analysis of thin wings in inviscid incompressible flow is presented. In this model a network of spanwise quadratically varying semi-infinite doublet sheets is introduced which produces a continuous trailing vortex wake. The present method has been applied to wings fitted with partial span trailing edge flaps after appropriate modification to account for the flap juncture. Also, the problem of wings in sideslip is attempted by incorporating the zero load condition at the down stream wing tip. The comparison of results shows that the proposed model retains the simplicity of the vortex lattice model to a large extent while overcoming the limitations of the standard vortex lattice model regarding lattice arrangement and number.

While the bending analysis of cylindrical containers with wall thickness varying linearly has attracted much attention, it seems the general variation in wall thickness has not been considered. This is because of the difficulties which have been encountered due to the complex nature of the differential equations involved. In this paper, the container’s wall thickness is chosen to be of general variation and the differential equation is perturbed to give rise to a sequence of differential equations. It is shown that this sequence can be easily solved when the form of variation in wall thickness is specified. A complete solution is obtained in a particular case when the wall thickness varies linearly and the tank is subjected to hydrostatic pressure. For this particular case the numerical results are compared with the ones available in the literature which seem to be in good agreement.

Measurements made at a Mach number of 0.18 and a chord-based Reynolds number of 4·2 x 106 on a constant-chord model having a NACA 4412 aerofoil section are described and compared with the results of flow field calculations.

Both the experimental arrangement and the difficulties initially experienced in achieving an adequate approximation to two-dimensional flow above the wing are briefly outlined.

The measurements include static pressure distributions on the wing surface and on the wind-tunnel walls above and below the mid-span section of the wing. The main emphasis in the experiment was, however, on defining the development of the upper surface boundary layer through separation (at about 20% chord ahead of the trailing-edge) and on into the wake, making extensive use of laser anemometry to measure mean velocities. In addition, Reynolds stresses were measured in certain parts of the flow field by hot-wire anemometry.

The flow field calculations are of the semi-inverse kind in which an inverse momentum-integral treatment of the shear flow, used to avoid difficulties at separation, is coupled to a direct solution of the inviscid flow problem. The main features of the method are outlined.

Large deflection problem of a spring loaded hinged nonuniform cantilever beam subjected to a rotational distributed loading is formulated by means of a second-order non-linear integro-differential equation. The problem is examined by considering the beam of rectangular cross-section with linear depth taper subjected to a uniform rotational distributed load. The elastic curves of a beam for this special case are presented.