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A novel experimental facility for the testing of modern high pressure nozzle guide vanes, the Annular Sector Heat Transfer Facility, is described in this thesis. Non- dimensionally similar conditions to a thermal paint test are reproduced, in a warm flow field, by the use of actual engine hardware, contoured sidewalls, and an innova- tive system of deswirl vanes in a five-passage annular sector cascade. External Mach and Reynolds numbers, inlet turbulence intensity, and coolant-to-mainstream pres- sure ratio are all matched to engine conditions. The test vanes are heavily cooled both internally (by convection and impingement) and externally (by film cooling). Detailed aerodynamic measurements are discussed, which demonstrate that a peri- odic, transonic, and highly engine-realistic flow is established in the cascade. High resolution full coverage maps of overall cooling effectiveness are presented, acquired on the vane surfaces at steady state conditions by wide-band liquid crys- tals and infrared thermography. Experimental measurements are then scaled to en- gine conditions by a new theoretical procedure, argued from first principles, which extends the principle of superposition to fully-cooled compressible flows. A newly- defined recovery temperature is proposed, which accounts for the redistribution of heat between the internal and external vane flows in a fully-integrated manner. This technique makes the results analogous to those of a thermal paint test, but allows for fundamental research and early and inexpensive cooling system validation. Overall cooling effectiveness measurements are complemented by those of the re- quired cooling flow capacity to achieve them, conducted in a second test rig commis- sioned during this research: the Flow Testing Facility. To conclude, the approach developed is applied to the global thermal assessment of the dendritic geometry, an innovative turbine cooling system. Experimental results show promising benefits over the baseline vane, especially in regions of low coolant-to-mainstream pressure margin.

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This thesis presents CFD predictions of gas turbine film and effusion cooling. The dearth of detailed experimental adiabatic effusion cooling data led to the validation of the computational procedures against the experimental adiabatic cooling effectiveness data for a single row of inclined round film cooling holes. This showed that the overall best agreement of the CFD predictions with experimental data was for the realizable k-e turbulence model with enhanced wall function. This was also shown to give good predictions of experimental results for trench outlet film cooling. This film cooling CFD work was extended .to demonstrate trench outlet lip geometries that could further improve the cooling effectiveness. The limitation of the CFD model was at higher blowing rates, M, when the film jet lifted off from the surface, where the CFD did not accurately predict the adiabatic cooling effectiveness close to the hole. For attached jets at lower M the agreement was good. The same CFD procedures were used for all the effusion cooling conjugate heat transfer (CHT) predictions. The hot metal wall effusion cooling experimental data base of Andrews and co-workers (1983-1995) was used to validate the CHT effusion cooling predictions. This database was for combustor flat wall cooling with mainly 90° injection holes. The overall effusion cooling effectiveness was measured and this required conjugate heat transfer CFD predictions. The adiabatic film cooling effectiveness was also predicted, by using a gas tracer in the cooling air and predicting its concentration at the effusion wall. For each effusion hole configuration, the coolant mass flow rate, G kg/srrr2bar, was varied from 0.1 to 1.5 and each G required a separate computation. The influence of the number of holes at a constant X!D of 4.6 and the hole size at fixed X were investigated. The agreement between the predictions and experimental data was good. Finally, the influence of the effusion coolant jets flow direction to the hot-gas crossflow on effusion cooling performance was investigated. This included 30° inclined opposed-flow jets effusion wall, which was predicted to be the best effusion jets flow pattern. The addition of the filleted shape trench outlet to effusion cooling was predicted to improve the cooling performance with reduced coolant mass flow rate, due to the improved adiabatic film cooling.

Near surface compressive residual stresses are known to improve the fatigue performance and damage tolerance of materials. In the aerospace gas turbine industry there is a compressive residual stress fields to extend the safe life time of a component is developing as a new field of research. Global near surface compressive stresses such as those produced by shot or glass bead peening have been used in service for a long time. The benefits of these stresses are not normally considered in the safety analysis of components; they normally act as an additional safety margin. The development of new surface treatment techniques capable introducing compressive residual stresses into specific locations has lead to increased industrial interest in the use of engineered compressive residual stresses for quantifiable life extension.

The experimental investigation of the combustion dynamics in a full scale industrial gas turbine combustor has been conducted. The successful capture of different combustion observables; including chemiluminescence emissions from C2* and CH* radicals, acoustic pressure oscillations emanating from the combustor, sound pressure level, fuel pressure oscillations before entry to the combustor, differential pressure oscillations across orifice plates for fuel flow measurement has yielded a huge database.

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No description available.