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An aircraft in low-level flight experiences a large increase in lift and a marked reduction in drag, compared with flight at altitude. This phenomenon is termed the 'wing-in-ground' effect. In these circumstances a region of high pressure is created beneath the aerofoil, and a pressure difference is set up between its upper and lower surfaces. A pressure difference is not permitted at the trailing edge and therefore a mechanism must exist, which allows the pressures above and below to adjust themselves to produce a continuous pressure field in the wake. It is the study of this mechanism and its role in the aerodynamics of low-level flight that forms the basis of our investigation

MEMS actuators and their effect on boundary layers is investigated using numerical simulation. The thesis is specifically focussed on jet actuators and their application to the targeted control of turbulent boundary layers. A complete numerical model of jet-type actuators, including the popular synthetic-jet actuator, is developed. The assumed input is the voltage signal to the piezocermic driver and the calculated output is the exit jet velocity. Thorough validation of the numerical code is presented and simulations performed to highlight the key issues in MEMS-actuator design. The three-dimensional boundary-layer disturbance created by the MEMS actuator is modelled using a velocity-vorticity formulation of the Navier-Stokes equations. The parallel code is rigorously validated against results from linear stability theory and transitional-streak measurements. The boundary-layer code is used to determine a performance criterion for MEMS jets; it is shown that the net mass flow from a jet best determines its effectiveness. The code is also used to demonstrate the macro-scale capabilities of MEMS-scale actuators; a grid-scaling method is described and employed to facilitate this calculation. A method is presented that enables high- and low-speed streaks to be modelled economically in otherwise undisturbed mean flows. Using this model, the fundamental principles of targeted control using MEMS actuation are explored. The MEMS-actuator and boundary-layer models are coupled, and an investigation into the interactive effects of the two systems is described. Using the coupled code, disturbances in the boundary layer are shown to induce velocities in inactive devices. One special case occurs when an oscillating pressure field creates Helmhotz resonance within the cavity of a MEMS actuator, thus causing large mass flow rates in and out of the device. It is also suggested that the MEMS device could strongly interact with the random fluctuations of a turbulent boundary layer, leading to highly unpredictable actuator responses.

A description of the experimental investigation of the aerodynamic characteristics of variable geometry of an aircraft model is presented. Aerodynamically, the model is tested-for a sweep range of 0°,12.5° 32.5°, and 52.5° and incidence range of 0° to 200° in 4° intervals. All the pressure distributions on the wing, glove, and body are recorded for each wind tunnel test. Aerodynamic forces and moments were also taken through a balance mechanism system which is attached to the model. This is connected to an independent computer terminal and a Teletype printer. Initially, a flow visualization to test the flow separation on the wing model was carried out. A three-dimensional subsonic program, which was already developed by Hawker Siddeley Aviation Limited, was modified for our purposes in order to carry out numerical calculation of the aerodynamic characteristics and investigate the interference of wing and body. This programme has also been developed to include the compressibility effects and compare these results with those for incompressible flow. The three-dimensional numerical solution was a Panel method for the subsonic case. This investigates the three-dimensional flow-field using a distribution of quadrilateral vortex panels, the effects of which are summed to calculate the aerodynamic characteristics of the model. This subsonic theory was applied to calculate the characteristics of the wind tunnel model over a similar range of sweep and incidence to those tested, for Mach numbers of 0 and 0.5. As the only input data required is the configuration geometry and the flight condition, however, the program can be used to calculate the aerodynamics of any wing-body arrangement specified by the user. The program includes the capability of analysing both fixed-wing and variable sweep-wing configurations. This computational method is capable of being applied to general arbitrary subsonic three-dimensional potential flows, including inlet flow fields. In panel methods, the velocity potential at any point in a flow field is expressed in terms of the induced effects of source and doublet (or vortex) sheet distributed on the boundary surfaces. The configuration surfaces are divided into panels, and essentially, this is a general three-dimensional boundary value problem solver that is capable of being applied to most problems that can be modelled within the limitations of potential flow. Compressibility effects are approximated by the Göthert rule. Comparisons were made between the subsonic calculations and the experimental results and some other theoretical results. Hence, an indication of agreement and accuracy among them is seen, which is good up to a certain degree of incidence (about 10°). Owing to viscous effects, the experimental results for lift coefficient show a significant decline in size with respect to subsonic calculated results. Wing-body interference was calculated for subsonic flows and found to be favourable. Similarly, a general supersonic program was developed for numerical analysis of the aerodynamic characteristics of a thin wing. The theory was extended to include wing-body interferences. This extended treatment consists of slender body theory combined with a thin wing solution using a "characteristic box" method for supersonic analysis. Streamwise pressure-distributions on an aircraft wing are presented, and also-some aerodynamic force and moment coefficients of this wing are-presented. Finally, for wing body interaction analysis, the Nielsen method was used. All the relevant computations including centre of pressure position and interferences of wing and body for a combined model are presented. Comparisons of the supersonic results with some theoretical and experimental results shows good agreement. The interference calculations in this case showed favourable effects, which very broadly tend to be lower than those calculated for subsonic flow.

This study develops analytical methods for the analysis of harmonically forced vibrations of multi-supported beams and stiffened plates. The methods are based on the response of infinite, uninterrupted structures to a single external excitation from which families of waves propagate outwards in all directions. The structures considered are uniform in thickness. The beams have a single-point force. The plates are finite in width with the two opposite parallel edges being simply-supported along the length. A line forcing varies sinusoidally between those edges. The first part of the thesis analyses the dynamics of infinite, uninterrupted structures. Special attention is subsequently paid to three-layered sandwich structures and their dispersion characteristics are also investigated. The second part considers finite structures. An analytical approach is presented for single-bay and multi-bay structures with arbitrary support spacing. Influences of support elastic/inertial characteristics are investigated in detail. Effects of end reflections are fully dealt with. The third part deals with the free and forced vibrations of infinite, periodic structures. Particular attention is focused on the effects of stiffness characteristics and cross-sectional distortion of the stiffeners. Part four outlines experimental work undertaken to validate the above theories. The experimental and theoretical results are compared.

This thesis investigates three elements necessary for the prediction of the broadband noise from a turbofan engine due to the interaction between the turbulent rotor wakes with the Outlet Guide Vanes (OGVs). These are (i) the sound radiation from a cascade of closely spaced blades interacting with rotor wake turbulence, (ii) an analysis of the behaviour of hotwire velocity data from a Large Scale Fan Rig (LSFR), (iii) the development of a scheme for the prediction of the blockage due to the transmission of multi mode sound across the rotor necessary for the prediction of noise in the forward-arc. (i) Cascade noise model A noise model is presented for the prediction of rotor wake turbulence with a cascade of OGVs. Similar to other approaches of this kind, computation time becomes excessive at high frequencies as the number of modes required increases. This thesis shows that at sufficiently high frequencies, when at least two modes are cut-on between adjacent blades, the acoustic blade coupling is weak and the cascade sound radiation closely approximates to that of an isolated aerofoil whose radiation can be computed efficiently using single airfoil theory, thereby greatly reducing computation time. (ii) Characteristics of rotor wake turbulence One factor currently limiting accurate fan broadband noise predictions is an understanding of rotor wake turbulence at the OGV leading edge. This thesis analyses in detail recent hotwire velocity data measured in the interstage of an LSFR. The focus here is on assessing the extent of self-preservation in the rotor wake, whereby the mean and turbulent wake characteristics can be deduced at any position downstream of the rotor and at any operating condition from a limited number of measurements. Unlike as previously assumed, this analysis demonstrates insufficient self-preserving behaviour to justify further pursuit of this approach. Rotor wake turbulence must therefore be measured or predicted at each operating condition separately. An analysis procedure is developed by which the characteristics of individual wakes, necessary for broadband noise predictions, may be inferred from rotor wake velocity measurements in situations in which there is significant overlap between adjacent wakes. (iii) Multi mode rotor blockage Noise generated by the OGV propagates to the forward arc by passing upstream through the spinning rotor. This thesis presents a model for the sound power transmission loss associated with crossing the rotor that includes modal frequency scattering effects. It is shown that the results obtained using exact cascade scattering closely agree at low and high frequencies with the results from a relatively simple prediction scheme that assumes that only plane waves propagate through the cascade, thereby ignoring modal scattering effects. The advantage of making this approximation is that the computation is considerably more efficient than a full cascade calculation. At low frequencies, where only plane waves propagate in the gap, exact agreement is obtained between the exact and plane wave models. Close agreement is also observed in the high frequency limit where a large number of cascade modes are cut-on, most of which are well cut-on and hence whose behaviour tends that of the plane wave mode. The three components of the prediction procedure outlined above are combined to perform a prediction of the rear-arc and forward-arc broadband noise from an LSFR. Comparison of the measured and predicted noise spectra are in reasonable agreement with variations with working line and fan speed being reasonably well captured.

Strong blast waves, such as those associated with gun fire or rocket exhausts, can cause serious physiological and/or structural damage. It is necessary, therefore, to develop ways in which to minimise this damage whilst still allowing the system that produces the blast waves to function normally. In an effort to develop such a system this study has examined, both theoretically and experimentally, the interaction that occurs when a shock wave is passed along a free unconstrained liquid tube before emerging into the surrounding atmosphere. In the theoretical analysis the problem was confined to the two-dimensional case and involved dividing the liquid sheet up into infinitesimal sections that were then regarded as small piston-cylinder systems, which were driven by the high pressure shocked gas behind the shock wave. The results from these one-dimensional systems were then used as input into a two-dimensional solution of the wave equation which predicted the gross changes in the remainder of the space. The experimental investigation involved laboratory experiments that examined, visually and with pressure transducers, the result of the shock/liquid interactions in two-dimensional and axi-symmetric cases, both between and external to the liquid sheets. The experimental investigation also included field trials that examined the pressure profiles of blast waves that were produced when a shock wave, which resulted from the ignition of a rocket motor, was passed along a liquid tube. From this work it was found that the high pressure gas, behind the shock wave, caused the liquid sheets to move perpendicularly from the line of travel of the shock wave which in turn caused expansion waves and compression waves to propagate out from the face of the water sheets, into the shocked gas and the surrounding atmosphere, respectively. These compression waves were then found to interact with the weak blast (produced when the shock, which had been weakened by the expansion waves, emerged into the atmosphere) in such a way that they produced a weaker blast field than would have been the case had the shock wave emerged directly into the atmosphere; the maximum observed reduction in the strength of the blast wave was 16.4 dB. Experiments were also performed that examined the effect of using rigid sheets in place of liquid sheets. From these experiments it was found that the differences between the liquid and rigid sheet cases was a function of the size of the inertial barrier (i.e. the mass of the water sheet) that the water presented to the shocked gas. Consequently, it was noted that, in terms of attenuating the blast wave, the rigid sheets proved to be inferior to the thicker water sheets and superior to the thinner water sheets. However, when the spectra of the pressure disturbances were examined it was found that, with regard to the attenuation of the 2-4 kHz region of the spectra, all the liquid sheet results showed an improvement in relation to the rigid sheet results.

A technique has been developed to enable the measurement of the velocity field in the vicinity of a lifting rotor. This involved the use of a triaxial hot-wire probe coupled to a mini-computer. Using this technique extensive flow measurements in one plane (z/R = -0.075) under the rotor were made at advance ratios of 0.1 and 0.067. In addition the rotor thrust, torque and blade bending moments were also measured. The results of the analysis of the time averaged flow show clear evidence that the wake of the rotor rapidly rolls up into 2 `fixed-wing' type vortices. Also indicated is that the value of K in Glauert's induced velocity equation is underestimated by about 50%. It is also shown that results obtained from tests carried out at the same advance ratio, but different tip speeds, do not produce the same non-dimensionalised velocities at all positions. Analysis of the time varying results has shown that the roll-up of the tip vortex is completed within approximately 1 chord length of the blade and that the strength of the tip vortex is equal to the maximum value of the blade bound vorticity. The time varying data also reveals the influence of the turbulent blade wake and this has been used to obtain estimates of the blade profile drag. It is believed that this is the first time this has been observed in helicopter wake measurements and is potentially of great use in investigating the reductions in drag from using advanced blade sections and tip planforms. Aft of the rotor velocity traces showing the intersection of the tip vortex by the probe has been analysed to provide estimates of the vortex core size. These results also show evidence of axial flow in the vortex core.

In pan A of this thesis, a procedure based on lifting line theory for the design of wind turbines operating in non-uniform, non-axial but axisymmetric flows is presented. This procedure was used for the design of conventional turbines which were compared with turbine designs produced by momentum theory. The overall trends of both theories were found to be similar, although the lifting line procedure was found to produce a more conservative estimate of the turbine performance. The above mentioned design procedure was also used for the turbine blade design of the wind power systems presented in parts B and C of the thesis. Part B of the thesis deals with the development of the delta wing-turbine system: The system was scaled-up using the results of a previously developed design model, and its dimensions were compared with those of equivalent conventional turbines. It was found that the system compares well with conventional turbines up to rated power values equal to 100 kW. Its advantages were found to be the lower turbine diameter required for a given power output and the opportunity it provides for direct connection of the turbines to generators. The cost of this advantage is the relatively large delta wing required. A system prototype with power output in the order of 1 kW was designed for testing. The prototype turbine blades were designed taking into account Reynolds number effects. In order to overcome the detrimental Re effects, use of a low Re aerofoil (GOE 795), reduction of the turbine number of blades to 10, increase of the blade chord, linear blade chord distribution and variable optimum angle of attack were found to be necessary, leading to a reduction of the turbine power coefficient drop to 4.7% below that of the original high Re design. The prototype off design performance was predicted and it was found that increase of the blade chord at the hub region (for strength) and linearisation of the blade optimum twist angle (for ease of manufacture) did not affect the turbine performance significantly. The generator to be used with the prototype turbine was bench-tested. Its model parameters and power losses were identified. For matching the generator with the turbine, an appropriate load for the generator was found. The prototype long-term performance was also estimated using the turbine performance characteristics, the generator test results and the Weibull distribution of wind occurrence probability. It was found that the generator is not ideally suited for the prototype turbine and that a generator of larger ratings would be more suitable. Finally, the effects of yaw on the delta wing vortices were investigated experimentally. This was done in order to determine the feasibility of using the delta wing yaw to regulate the system power output. It was found that the above mentioned regulation technique can be used, provided that undue blade vibrations due to turbine-vortex misalignment and vortex bursting will not occur. In part C, a procedure for the design of the counter-rotating turbine blades was developed. The above mentioned lifting line procedure as well as the existing knowledge of wind turbine wakes and counter-rotating rotor aerodynamics were used for the design of the counter-rotating turbine blades and the semi-empirical modelling of the two rotors' interaction. The optimum axial distance between the two rotors was found to be equal to 1.4 times the rotors' radius. It was demonstrated that proper design of the turbine blades and appropriate axial positioning of the two rotors could increase the turbine performance by 27.4% above that of the original counter-rotating turbine design, called Trimble Mill. It was also found that a considerable increase of the generator effective rotational speed (equal to 58%) can be achieved by the counter-rotating turbine, compared to that of a conventional turbine with the same number of blades, while the two turbines' power output was found to be at the same levels.