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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.

At low flow rates, the flowfield of axial compressors becomes unstable. Two examples of this are rotating stall (a compression system level instability) and so-called "Rotating Instabilities" (a localized fluctuation in the tip flow region). This thesis presents full-annulus URANS computations demonstrating newly discovered properties of spike-type stall inception, Rotating Instabilities, and "multi-spikes" (a phenomenon related to Rotating Instabilities that has not been discussed in previous literature). It is shown, via the series of computations on a low speed rotor designed by Mitsubishi Heavy Industries (MHI) with nominal tip clearance (1.2%ctip), that spike-type stall inception is preceded by exponentially growing and circumferentially propagating long wavelength harmonic perturbations upstream of the rotor tip leading edge. These harmonic perturbations directly influence the circumferential variation of incidence and hence the blade which first exhibits leading edge separation that characterizes spike-type stall inception. The harmonic perturbations are supported by the rotor tip's rolled over local instantaneous axisymmetric total-to-static pressure rise characteristic, which reaches peak pressure rise at a stable operating point prior to spike-type stall inception. The rolling over of the tip axisymmetric characteristic is caused by blockage due to increased tip leakage jet strength as the rotor mass flow rate is reduced, a phenomenon common to other rotors. These findings provide connections between spike-type and modal stall inception, in that both mechanisms involve pre-stall exponentially growing long wavelength perturbations and total-to-static pressure rise characteristics, at least over part of the span, rolling over prior to stall inception. In the series of computations of the MHI rotor with large tip clearance (4.5%ctip), Rotating Instabilities are captured with their characteristic frequency spectral "hump" (a property of Rotating Instabilities widely reported in the literature) reproduced. Rotating Instabilities are shown to begin to develop when the circumferentially averaged axial velocity above the blade tip becomes negative. The computations allow the structure of the Rotating Instabilities to be identified. During the development process, the tip leakage flow starts to oscillate, and rolls up into circumferentially propagating disturbances of a vortex tube nature with a circumferential spacing of between 1.16 to 1.65 blade pitches. This range in circumferential spacing of the Rotating Instabilities vortex tubes gives rise to the characteristic "hump" in the frequency spectra. These disturbances are confined within the top 15%span, exist downstream of the leading edge, and do not trigger rotating stall inception. Previously unreported "multi-spikes", which are found to be adjacent spike-like disturbances measured in casing static pressure traces, occur when the mass flow rate is reduced from the Rotating Instabilities operating point. The computations show that "multi-spikes" consist of Rotating Instabilities disturbances (the vortex tubes) superposed with non-growing, circumferentially-propagating, long wavelength perturbations. The long wavelength perturbations create a pressure trough with a width of several blade passages (one-eighth of circumference in the case presented), which causes adjacent Rotating Instabilities disturbances to move upstream of the leading edge. Thus, the now-upstream vortex tubes are registered as adjacent spike-like disturbances by the upstream static pressure traces. The rest of the annulus, meanwhile, operates with Rotating Instabilities disturbances still contained within the blade passage downstream of the leading edge.

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.