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The flow at the trailing edge of a turbine blade at transonic air velocities can be extremely complex. Solid trailing edges can shed vortices in a manner known as transonic vortex shedding, where vortices form very close to the trailing edge and cause large trailing edge shear layer deflection. This, in turn, results in shockwaves that can propagate upstream of the trailing edge. When this flow regime occurs, it is known to be the loss dominating mechanism for trailing edge flows. In this dissertation experiments and LES simulations are performed to increase the understanding of the flow mechanisms at the trailing edge, both for solid and cooled trailing edges. It was found that transonic vortex shedding is not the only flow regime possible behind round trailing edges at transonic air velocities. Under certain conditions the vortices are found to form approximately one trailing edge diameter downstream of the trailing edge, both the shear layer deflection and shockwave formation are dramatically reduced. This flow regime has been referred to as detached vortex shedding. The switch from detached to transonic vortex shedding is found to be the result of the transition of the pressure surface boundary layer and is characterised by a close to doubling in the mixed-out loss for the trailing edge tested. The effect of trailing edge wedge angle on the performance of solid, round trailing edges is also investigated. It is found that wedge angle is an important parameter governing trailing edge performance, with higher wedge angles offering reductions in loss. Switching from an 8◦ to a 14◦ wedge angle plate was found to reduce the loss by up to 29% when the plates were in the same flow regime. The effect of geometry variation on blown, through trailing edge holes geometries is investigated experimentally. These geometries undergo transonic vortex shedding when there is no trailing edge blowing, but the addition of coolant to the base region is able to suppress the vortex shedding and reduce the loss by up to 39%. In order to disrupt the vortex shedding, a threshold coolant mass flow ratio must be reached; the value of this threshold depends on Reynolds number, since this governs the strength of the vortex shedding. Holes of half the area, or holes drilled obliquely to the trailing edge are able to reduce the coolant mass flow for a given value of coolant stagnation pressure coefficient. Finally, the effects of manufacturing variation on trailing edge loss is investigated with the aid of GOM scans from real trailing edge geometries. For solid trailing edges, variations in geometry away from an ideal round trailing edge are found to have significant effects on the loss. This is almost always beneficial, since the performance is governed by the ability of the geometry to suppress transonic vortex shedding. Some variations are able to reduce the loss by over 50% from the baseline round geometry. For blown, cutback geometries, the differences in loss due to manufacturing variations are smaller (at most 5% reduction in loss). But, over the majority of operating points tested, there was no disadvantage and indeed sometimes a small advantage of real geometry.

In gas turbines, film cooling is required to protect metal parts from hot combustion gases. Reduction in coolant mass flow increases cycle efficiency, and hence reduces greenhouse gas emissions. However, the lifespan of a cooled component is sensitive to the metal temperature within the part. A designer requires predictions of cooling performance to make this compromise, yet present design methods are subject to uncertainty and are not viable without empirical input. Flow through a turbine is inherently unsteady due to relative motion of stators and rotors, termed blade row interaction. Blade row interaction is not captured in flat-plate and cascade testing, or present design methods, contributing to uncertainty in predicted cooling performance. The aims of this thesis are to establish the mechanisms by which blade row interaction affects rotor film cooling, and quantify their influence on cooling performance in a representative case. A new experimental rig is developed to facilitate aerodynamic and heat transfer measurements of cooling holes subject to unsteady main-stream boundary conditions. The effect of unsteadiness is set by non-linearity in the hole response. Unsteadiness reduces film effectiveness by up to 31% with cylindrical holes at a low momentum flux ratio, because the response to perturbations is non-linear. Cylindrical holes at a high momentum flux ratio, and fan-shaped holes, are robust to unsteadiness because they respond linearly. Non-film-resolved computations are used to identify the blade row interaction mechanisms generating unsteady main-stream boundary conditions in a turbine rotor. A quasi-steady model is used to predict instantaneous excursions in cooling hole momentum flux ratio. Fluctuations of at least ±30% are present for all hole locations, due to both upstream vane wake and potential field interaction. A hybrid URANS-LES computational approach is implemented, validated against experimental data, and applied to a turbine stage cascade model. Compared to steady conditions, blade row interaction reduces rotor film effectiveness: by up to 18% on the pressure side, due to migration of vane coolant across the passage; and by up to 30% on the suction side, due to wake interactions increasing the film mixing rate.