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The present study analyses the effects of a number of geometric parameters on the performance of a trailing edge control flap on a hypersonic body. The tests were conducted in a gun tunnel at Mach 8.2 and Mach 12.3. The study revealed that flap deflection promoted separation lengthscales and boundary layer transition. The latter significantly increased the local aerothermal loads on the flap. For well separated flows, flap heat transfer rates were successfully predicted by reference temperature theory. The promotion of transition caused a progressive reduction in the lengthscales of separated flows. In a free-flight environment, vehicle incidence varies considerably. Incidence was found to promote transition on both flat plates and control flaps. The latter resulted in a considerable increase in flap heat transfer. A modified version of reference temperature theory successfully predicted the aerothermal loads on the flap. For laminar and transitional interactions, the separated flow lengthscale was found to have a complex variation with incidence. A number of relevant flow parameters were identified. The intense heat loads on a vehicle in hypersonic flight dictates the blunting of the leading edge. This strengthens the leading edge shock structure and generates an entropy layer. Bluntness was found to significantly decrease the separation interaction scales on the flap. This was due to a reduction in the pressure recovered on the flap. The latter adverse affects control effectiveness. The aerothermal loads on the control flap was successfully predicted by reference temperature theory. An investigation into the efficiency of an under-expanded transverse jet controls was conducted on an axi-symmetric slender blunt cone. Force measurements found that the interaction augmented the jet reaction force by 70% at zero incidence. This increased to 110% at low incidence. The experiments found that the scale of the interaction region was determined by Poj/pes. Using this parameter, a closed loop algorithm for the shape of the separation front was developed. The latter can be used to predict jet reaction control effectiveness.

A computational investigation was undertaken to examine techniques for predicting supersonic dynamic flows, involving unsteadiness over fixed and moving surfaces. The fixed geometries examined were cylinder-flares and compression ramps, and the moving body geometries a pitching aerofoil and a rapidly deployed flap. Investigation into the characteristics of incipient separation of a supersonic cylinder-flare flow revealed that the separated length varied with a power of the flare angle and that the variation in height of the separated region varies in a bi-modal manner with flare angle. For small-scale separations (flare angles less than those which would traditionally have been expected to induce separation) the height of the separated region was seen to vary slowly with flare angle. For larger flare angles, the separation bubble was found to grow rapidly in height and length with increasing flare angle and produce significant deflection of the external flow. Computations of a Mach 5, compression ramp induced unsteady shock boundary layer interaction exhibited self-sustained oscillations at frequencies and amplitudes consistent with experimental data. Large dynamic structures (up to 1.7 boundary layer thicknesses in extent) were observed, and their production, propagation and deformation illustrated. By modifying the turbulent viscosities produced by a non-dimensional implementation of the Baldwin-Lomax turbulence model (using under- relaxation) a turbulence model was produced which accurately predicted separation lengths for a series of Mach 6.85 compression ramp flows encompassing laminar, transitional and turbulent flow regimes (dependent on ramp angle). A technique was developed to enable efficient computation of dynamically moving and/or deforming body flows. This technique was based on hierarchical, adaptive mesh refinement coupled with automatic generation of body surfaces, in which mesh adaption was used to capture the body geometry to within a specified accuracy. This, in conjunction with automatic cell creation and destruction, enabled the derivation of steady and unsteady, time accurate, conservative boundary conditions. This algorithm was used to compute a quasi-steady laminar supersonic pitching aerofoil flow, and an unsteady turbulent supersonic flap deployment. In both cases agreement with experiment was found to be good.

An efficient three-dimensional structured solver for the Euler and Navier-Stokes equations is developed based on a finite volume upwind algorithm using Roe fluxes. Multigrid and optimal smoothing multi-stage time stepping accelerate convergence. The accuracy of the new solver is demonstrated for inviscid flows in the range 0.675 :5M :5 25. A comparative grid convergence study for transonic turbulent flow about a wing is conducted with the present solver and a scalar dissipation central difference industrial design solver. The upwind solver demonstrates faster grid convergence than the central scheme, producing more consistent estimates of lift, drag and boundary layer parameters. In transonic viscous computations, the upwind scheme with convergence acceleration is over 20 times more efficient than without it. The ability of the upwind solver to compute viscous flows of comparable accuracy to scalar dissipation central schemes on grids of one-quarter the density make it a more accurate, cost effective alternative. In addition, an original convergencea cceleration method termed shock acceleration is proposed. The method is designed to reduce the errors caused by the shock wave singularity M -+ 1, based on a localized treatment of discontinuities. Acceleration models are formulated for an inhomogeneous PDE in one variable. Results for the Roe and Engquist-Osher schemes demonstrate an order of magnitude improvement in the rate of convergence. One of the acceleration models is extended to the quasi one-dimensiona Euler equations for duct flow. Results for this case d monstrate a marked increase in convergence with negligible loss in accuracy when the acceleration procedure is applied after the shock has settled in its final cell. Typically, the method saves up to 60% in computational expense. Significantly, the performance gain is entirely at the expense of the error modes associated with discrete shock structure. In view of the success achieved, further development of the method is proposed.

An attempt to minimize interference drag in a wing-fuselage junction by means of inserting a fillet is presented in this thesis. The case of a low-wing com- mercial transport aicraft at cruise conditions is examined. Due to the highly three dimensional behaviour of the flow field around the junction, a thin-layer Navier-Stokes code was implemented to estimate the drag forces at the junc- tion. Carefully selected design variable combinations based on-the theory of Design of Experiments constituted the initial group of feasible cases for which the flow solver had to be run. The drag values of these feasible cases were then used to create a second order response surface which could predict with rea- sonable accuracy the interference drag given the value of the design variables within the feasible region. A further optimization isolated the minimum in- terference drag combination of design variable values within the design space. The minimurn interference drag combination of design variable values was eval- uated numerically by the flow solver. The prediction of the response surface and the numerical value obtained by the flow solver for the interference drag of the optimal wing-fuselage combination differed by less than five percent. To demonstrate the ability of the method to be used in an interdisciplinary analysis and optimization program, a landing gear design module is included which provides volume constraints on the fillet geometry during the fillet sur- face definition phase. The Navier Stokes flow analyses were performed on the Cranfield Cray su- percomputer. Each analysis required between eight to twelve CPU hours, and the total CPU time required for the optimization of the six variable model described in the thesis required thirty Navier Stokes runs implementing the Design of Experimens and Surface Response Methodology implementation. For comparison, a typical optimization implementing a classical conjugate di- rections optimizer with no derivative information available would probably require more than forty iterations. Both the optimization and the flow solver results are discussed and some recommendations for improving the efficiency of the code and for further ap- plications of the method are given.

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A major addition to European research facilities is the Oxford University Engineering Laboratory (O.U.E.L.) blowdown tunnel which can provide full-scale Reynolds and Mach number simulations on large and small models of turbine stage components. The facility was designed to provide extended aerodynamic capabilities to complement the existing heat transfer research in the Isentropic Light Piston Tunnel (ILPT) at O.U.E.L. The blowdown tunnel will be used for fundamental investigations of the boundary layers and flow fields around turbine blades in a linear cascade. The study of these flow fields is necessary for the prediction of heat transfer rates and for the optimisation of materials and cooling schemes required to improve gas turbine efficiencies. As a commissioning exercise measurements were made on cascades of similar geometry to those which had been previously tested in the ILPT and in other European facilities in order to compare results and analyse differences which occur due to the influence of tunnel geometry. Measurements made on various rotor profiles identified regions on the suction surface where surface pressure data is sensitive to the various types of exit plenums and exit pressure gradients. A second phase of work included measurements and a theoretical study of the boundary layer on a large-chord turbine rotor profile. Measurements on the pressure surface of the blade suggested the presence of secondary longitudinal vortices which rapidly lose an identifiable structure towards the trailing edge. On the suction surface, boundary layer measurements were compared with theoretical models, and it was shown that current numerical models of compressible turbulent boundary layers approximately correspond with the data. An adjunct to the boundary layer work was research on the use of a hot-wire anemometer, intended for future boundary layer measurements, and for which calibration laws and temperature effects were studied.

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To provide power on a national scale, a large number of windmills will have to be deployed in wind farms or arrays because the output of individual machines is relatively small. Within an array, some windmills will be faced with the wakes generated by others. This interaction leads to a loss of power relative to upwind turbines and changes in the wind loading across the turbine blades.

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