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An oblique shock generated by a variable-angle wedge on the side wall of a wind tunnel, has been used to investigate the three-dimensional glancing interaction problem. The shock interacts with the turbulent boundary layer growing along the side wall. Two related test programmes have been completed using a 2.5 x 2.5 inch intermittent tunnel and a 9 x 9 inch continuous-running tunnel. For both the test programmes, the Mach number was approximately 2.5 and the Reynolds number relative to the wall boundary-layer thickness 5 x 10 4 . The experimental results include oil-flow pictures, vapour-screen and smoke photographs,wall pressure distributions, local heat transfers, wall surface temperatures and viscous layer surveys. The experimental results suggest that the interaction reg10n consists of two different viscous layers between which an ordinary separation can take place, (the double viscous layer flow-field model). The three- dimensional separation is found to depend significantly on the pressure rise in the direction normal to the swept shock. In this sense the separation is similar to the two-dimensional case.

The effects of leading edge bluntness and sweep angle on the three dimensional glancing shock wave - boundary layer interaction have been investigated. A large number of hemi-cylindrically blunted fins with leading edge diameter ranging from 0 to l.0 in, with leading edge sweep angles between 0° and 75° were tested. The incidence angle was varied from 0° to 21°. The shock wave from each configuration interacted with a fully developed turbulent boundary layer growing along the tunnel side wall. The free stream Mach number in the 9in x 9in continuous flow supersonic wind tunnel was 2.4 and the Reynolds number based on boundary layer thickness was 5 x 10^. Experimental investigations included oil smear tests, surface pressure surveys, schlieren pictures of the inviscid shock envelopes and shock structure in the plane of symmetry. The study highlighted the significant effects of bluntness and sweep on the scale and character of the interaction. While bluntness intensified the interaction, sweep alleviated its intensity. The most dramatic effect of sweep angle was observed when the leading edge was swept from 0° to 30°. Sufficiently outboard of the plane of symmetry, the features of blunt and sharp fins became similar. The boundary between the inner "bluntness dominated" and the outer "viscous dominated" regions shifted inboard at the higher incidence and sweep angles. The characteristic surface oil flow patterns showed little change for sweep angles up to A = 60°. Leading edge bluntness increased the scale of the interaction almost linearly while leaving its character unchanged. The multiplicity of the separation and attachment lines on the side wall and the fin surface, suggested a system of vortices in the interaction region. Flow field models have been proposed over the range of sweep angles considered in the present study. The number and strength of the vortices is seen to depend on the leading edge bluntness, sweep and the incidence angle. The important parameters governing the primary separation distance and the peak pressure in the plane of symmetry have been identified. Correlation formulae suggest a strong interdependence of the various parameters concerned.

The use of the estímation-before-modellíng (EBM) two step identification procedure for the determination of aircraft aerodynamic derivatives from flight test data is analysed and illustrated. In the first step of the identification procedure the usual Extended Kalman Filter (EKF) associated with the Modified Bryson-Frazíer (MBF) smoother is compared with a new alterative filtering and smoothing process. The new smoother is simpler and less computationally demanding than the MBF smoother. However, its main advantage is that it enables simultaneous data smoothing with state derivative estimation, thereby avoiding the need for a separate differentiation algorithm. The new smoother differentiator has an important feature that is the determination of the noise characteristics of the measurement signal under analysis prior to the smoothing process. This is done by variance matching between the theoretical and measured autocorrelation of the innovation process generated by a Kalman filter. The new technique is compared with the old one by determining the aerodynamic models for a EMB-312 Tucano dutch roll manoeuvre. It is demonstrated that the new smoother may be used to replace the MBF. Otherwise the new technique is used in the analysis of the Handley Page Jetstream-100 aircraft low speed controls free phugoid trying to identify the contribution of the power Variation observed during the phugoid to the stability of the oscillation. Finally the models obtained from the phugoid analysis are reprocessed using the Total Least Square regression and the results are compared with those from the ordinary Least Square formulation.

Using the design requirements of Remotely Piloted Vehicles (RPV's), selected for wind tunnel testing over the Reynolds number range 3 x 105 to 1 x 106. The first aerofoil, NACA 643-418, showed a degradation of performance in terms of lift-to-drag ratio as the Reynolds number was reduced. There was also a laminar separation bubble of notable extent on both the upper and lower surfaces at most incidences throughout the Reynolds number range. The second aerofoil, Göttingen 797, had good performance in terms of lift-to-drag ratio and maximum lift coefficient, even at the lowest Reynolds number. This was attributed to the flat bottom of the aerofoil, which allowed the formation of extensive laminar flow on the lower surface without the formation of a laminar separation bubble. The third aerofoil, Wortmann FX63-137, generally exhibited the best aerodynamic performance in terms of maximum values of both lift-to-drag ratio and lift coefficient, throughout the Reynolds number range considered. Four alternative lower surface geometries for this aerofoil were also tested. The modifications reduced the maximum values of both the lift coefficient and lift-to- drag ratio of the original aerofoil throughout the Reynolds number range, but generally improved the lift-to-drag ratios at low values of lift coefficient. The notable exception was the modification which resulted in a flat bottomed section. This had maximum values of lift-to-drag ratio which were within a few percent of those of the original aerofoil throughout the Reynolds number range. Wind tunnel results were used to evaluate low-speed aerofoil analysis computer programs written by Eppler and Somers (13) and Van Ingen (18). The results were disappointing. However, using the same wind tunnel results it was noted that computer programs using semi-inverse viscous methods show great promise.

The aerodynamic performance of the low solidity curved-bladed Darrieus vertical axis wind turbine has been studied both theoretically and experimentally. Initial studies showed the need for an engineering prediction scheme sufficiently accurate to give blade forces as functions of rotational position which did not require excessive computational time. The scheme proposed here develops a suggestion first made by Lapin in 1975, to treat the turbine as two actuator discs, one upwind and the other down-wind. This suggestion, combined with a multiple streamtube approach, the momentum equations in the freestream direction and blade element theory enables the system of equations to be solved. A continuity argument connects the flow between the two discs. The theory does not rely on analytic formulations for aerofoil force coefficients and can therefore use data obtained from experiment, tabulated for a range of Reynolds numbers, thereby including the effects of stall and drag. Comparison with the power coefficients obtained from experiments, using a two-bladed wind tunnel turbine at a Reynolds number of 28,000 (based on free wind speed and blade chord) shows that the theory is accurate enough to detect the effect of dynamic stall. It is also shown that the continuity argument is essential for improved power output predictions, over earlier single actuator disc theories. The new theory also indicates large differences between air speeds on the upwind and downwind sides of the turbine. These were also confirmed by experiment. Comparison of blade force predictions with those obtained using a computationally expensive time-marching discrete vortex theory shows that good estimates are obtained over the normal turbine operating range. At present only a uniform freestream is treated and turbulence is not accounted for.

Two dimensional subsonic wind tunnel tests have been conducted on a 20% thickness: chord ratio circulation controlled elliptic aerofoil section equipped with forward and reverse blowing slots. Overall performance measurements were made over a range of trailing edge blowing momentum coefficients from 0 to 0.04; some included the effect of leading edge blowing. The effective incidence was determined experimentally and lift augmentations of 70 were obtained at low blowing rates. A detailed investigation of the trailing edge wall jet, using split film probes, hot wire probes and total head tubes, provided measurements of mean velocity components, Reynolds normal and shear stresses, and radial static pressure. Corrections for the effects of ambient temperature variation, flow angle and shear flow gradient upon the various probes were examined and some corrections for the low bandwidth of the split film probes proposed. In some cases, the effects of slot height and slot lip thickness were investigated. The results were mostly taken at a geometric incidence of 0°. The closure of the two dimensional angular momentum and continuity equations was examined using the measured data, with and without correction, and the difficulty of obtaining a satisfactory solution illustrated. The experimental results have led to some suggestions regarding the nature of the flow field which should aid the understanding of Coanda effect and the theoretical solution of highly curved wall jet flows.

A supersonic aircraft's intake has been simulated from just upstream of the throat down to the engine face, with a wide bleed slot employed for boundary layer removal and mass flow trimming. A comprehensive experimental survey of the aerodynamic characteristics of the simulated intake has been made for several slot configurations, slot length, rear lip planform and profile all being changed. Rear lip scoop has been employed with one configuration and flow; unsteadiness in the form of pressure fluctuations has also been assessed. The general flow mechanisms prevailing have been identified and in many instances explained either qualitatively or quantitatively using basic gas dynamic relations. For optimum total pressure recovery at the engine face, irrespective of slot configuration, the terminal shock should be located at, or upstream of, the front lip. If limited amounts of scoop are applied when using a bluff rear lip improvements in bleed total pressure recovery may be obtained with no deterioration in the optimum engine face pressure recovery. The bluff lip also suppresses shear layer unsteadiness, as does a reduced void depth or reduced slot width.

Current trends in modem combat aircraft design have seen a move towards canard configurations with all moving foreplanes, providing a manoeuvre advantage with reduced stability. At the same time, with rapid advances in the field of assisted flight control and emphasis now placed on computer controlled, fly-by-wire aircraft, there is an unprecedented requirement for detailed knowledge of motion dependent aerodynamics, such as may be experienced on a foreplane undergoing rapid corrective motions. In this study, investigations have been carried out into the rigid body, motion dependent aerodynamics of a 55* delta wing, undergoing small amplitude pitching oscillations. Steady and unsteady surface pressures have been measured on the wing under low speed, pre-stalled conditions, for a range of mean incidence and oscillation frequencies, up to frequencies approximating a full scale foreplane under low speed conditions, such as landing approach. Relationships between the motion of the wing and the unsteady pressures have been identified, and it has been shown that they may not be approximated by a simple quasi-steady model due to significant phase shifts in specific regions of the flow. The lower surface flow is shown to be highly dependent on the effective incidence of the wing. The vortical flow of the upper surface has a more complex response to the pitching motion, with the shear layer and burst motion reacting at different rates. There is also significant attenuation/overshoot and phase in the unsteady loads and moments (obtained by integrating the pressure data) relative to the quasi-steady. These are shown to be highly dependent on the pitching oscillation frequency and the location of the pitch axis. It is suggested that there may be a pitch axis location such that quasi-steady loading may be obtained under oscillatory conditions. Application of the key findings to a simple all moving control surface shows that the stability of the control system is strongly influenced by the pitch axis location.

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