We quantify the sensitivity of turbine acoustic impedance to aerodynamic design parameters. Impedance boundary conditions are an influential yet uncertain parameter in predicting the thermoacoustic stability of gas turbine combustors. We extend the semi-actuator disk model to cambered blades, using non-linear time-domain computations of turbine vane and stage cascades with acoustic forcing for validation data. Discretising cambered aerofoils into multiple disks improves reflection coefficient predictions, reducing error by up to an order of magnitude compared to a flat plate assumption. A parametric study of turbine stage designs using the analytical model shows acoustic impedance is a weak function of degree of reaction and polytropic efficiency. The design parameter with the strongest influence is flow coefficient, followed by axial velocity ratio and Mach number. We provide the combustion engineer with improved tools to predict impedance boundary conditions, and suggest thermoacoustic stability is most likely to be compromised by change in turbine flow coefficient.

Loss in axial compressor bleed systems is quantified, and the loss mechanisms identified, in order to determine how efficiency can be improved. For a given bleed air pressure requirement, reducing bleed system loss allows air to be bled from further upstream in the compressor, with benefits for the thermodynamic cycle. A definition of isentropic efficiency which includes bleed flow is used to account for this. Two cases with similar bleed systems are studied: a low-speed, single-stage research compressor and a large industrial gas turbine high-pressure compressor. A new method for characterising bleed system loss is introduced, using research compressor test results as a demonstration case. A loss coefficient is defined for a control volume including only flow passing through the bleed system. The coefficient takes a measured value of 95% bleed system inlet dynamic head, and is shown to be a weak function of compressor operating point and bleed rate, varying by ±2.2% over all tested conditions. This loss coefficient is the correct non-dimensional metric for quantifying and comparing bleed system performance. Computations of the research compressor and industrial gas turbine compressor identify the loss mechanisms in the bleed system flow. In both cases, approximately two-thirds of total loss is due to shearing of a high-velocity jet at the rear face of the bleed slot, one quarter is due to mixing in the plenum chamber and the remainder occurs in the off-take duct. Therefore, the main objective of a designer should be to diffuse the flow within the bleed slot. A redesigned bleed slot geometry is presented that achieves this objective and reduces the loss coefficient by 31%.

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The mechanisms of blade row interaction affecting rotor film cooling are identified in order to make recommendations for the design of film cooling in the real, unsteady turbine environment. Present design practice makes the simplifying assumption of steady boundary conditions, despite intrinsic unsteadiness due to blade row interaction; we argue that if film cooling responds non-linearly to unsteadiness, the time-averaged performance will then be in error. Non-linear behaviour is confirmed using experimental measurements of flat-plate cylindrical film cooling holes, main-stream unsteadiness causing a reduction in film effectiveness of up to 31% at constant time-averaged boundary condition. Unsteady computations are used to identify the blade row interaction mechanisms in a high-pressure turbine rotor: a ‘negative jet’ associated with the upstream vane wake, and frozen and propagating vane potential field interactions. A quasi-steady model is used to predict unsteady excursions in momentum flux ratio of rotor cooling holes, with fluctuations of at least ±30% observed for all hole locations. Computations with modified upstream vanes are used to vary the relative strength of wake and potential field interactions. In general, both mechanisms contribute to rotor film cooling unsteadiness. It is recommended that the designer should choose a cooling configuration which behaves linearly over the expected unsteady excursions in momentum flux ratio as predicted by a quasi-steady hole model.

Loss in axial compressor bleed systems is quantified, and the loss mechanisms identified, in order to determine how efficiency can be improved. For a given bleed pressure requirement, reducing loss allows air to be bled from further upstream in the compressor, benefiting the thermodynamic cycle. A definition of efficiency which includes bleed is used to account for this. Two cases with similar bleed systems are studied: a low-speed research compressor and an industrial gas turbine high-pressure compressor. A new method for characterising bleed loss is introduced, using research compressor test results as a demonstration case. A loss coefficient is defined for a control volume including only flow passing through the bleed system. The coefficient takes a measured value of 95% bleed system inlet dynamic head, and is shown to be a weak function of operating point and bleed rate, varying +2.2% over all tested conditions. This loss coefficient is the correct non-dimensional metric for quantifying and comparing bleed system performance. Computations of the research compressor and industrial gas turbine compressor identify loss mechanisms in the bleed flow. In both cases, approximately two-thirds of total loss is due to shearing of a high-velocity jet at the rear face of the bleed slot, one quarter is due to mixing in the plenum chamber and the remainder occurs in the off-take duct. A redesigned bleed slot geometry is presented which increases diffusion in the bleed slot and hence reduces the loss coefficient by 31%.