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As the performance required of high pressure turbines continues to increase, there is a need to investigate many details of the flow which occur in a gas turbine stage that were previously overlooked. These include the effects of rotation and three-dimensional flow as well as unsteady effects due to the relative motion of the blade rows. In order to obtain a better understanding of the turbine flowfield a new transient facility has been commissioned in which aerodynamic and heat transfer measurements can be undertaken in a full stage turbine at engine representative conditions. The previously used technique of measuring the heat transfer rate by mounting thin film gauges on models manufactured from machineable glass ceramic was not suitable for use on the rotor blade because of the high stress levels involved. An alternative technique has been developed in which a metal turbine blade is coated with an insulating layer of enamel and thin film gauges painted on top. The developments in signal processing and calibrations which were necessary for the use of this type of thin film gauge are discussed in detail. Signal conditioning electronics have been developed which permit amplification of the thin film gauge output to a higher level within the rotating frame before transmission through a slipring. Extensive tests have been undertaken, in a purpose built spinning rig, to establish the effects of rotation on the performance and mechanical integrity of the instrumentation and associated electronics. The heat transfer measurements recorded in the rotor facility to date are presented and compared with data from a previous two-dimensional simulation of wake passing flow on the mid-height section of the same blade.

The growth of thrust and improved aeroengine efficiency has been gained by increased temperatures throughout the engine. This has been achieved by improved material technology and the continuous cooling of components complemented by the addition of thermal barrier coatings (TBC) to turbine and combustion chamber components. The aggressive nature of the application process of the TBC has previously made the measurement of metal surface temperature and strain exceedingly difficult on components to which it is applied. In the present study magnetron sputter-deposited thin film sensor techniques have been developed specifically for compressor and turbine applications of noble metal thermocouples and strain gauges. The deposition, patterning and evaluation of reactively sputtered aluminium oxide, type R platinum thermocouples as well as PdCr and PtW dynamic strain gauges is reported. A sputtered NiCoCrAlY coating has been developed to replace the vacuum plasma spray process currently used in the TBC system. The most favourable location for the thin film sensor is at the metal/ceramic interface of the TBC system. However, in order to protect the sensor from the aggressive TBC process, the sensor has been deposited in a novel installation between two layers of NiCoCrAlY bond coat. Several trials have been performed to fabricate this package on turbine blade material substrates. This work has demonstrated that the proposed sensor structure is feasible. However there are problems with delamination due to contamination and residual stress and with poor electrical insulation and these have limited the high temperature testing that could be performed. The novel techniques developed are already being utilised in measurement applications on components without TBCs. This work has been performed in an industrial context. The extensive project and risk management activities are reported.

State of the art aircraft afterburners employ spray bars to inject fuel and flameholders to stabilize the combustion process. Such afterburner designs significantly increase the length (and thus weight), pressure losses, and observability of the engine. This thesis presents a feasibility study of a compact prime and trigger (PAT) afterburner concept that eliminates the fuel spray bars and flameholders and, thus, eliminates the above-mentioned problems. In this concept, afterburner fuel is injected just upstream or in between the turbine stages. Downstream of the turbine stages, a low power pilot, or trigger , can be used to control the combustion process. The envisioned trigger for the PAT concept is a jet of product gas from ultra-rich hydrocarbon/air combustion that is injected through the afterburner liner. This partial oxidation (POx) gas, which consists mostly of H2, CO, and diluents, rapidly produces radicals and heat that accelerate the autoignition of the primed mixture and, thus, provide an anchor point for the afterburner combustion process. The objective of this research was to demonstrate the feasibility of the PAT concept by showing that (1) combustion of fuel injected within or upstream of turbine stages can occur only downstream of the turbine stages, and (2) the combustion zone is compact, stable and efficient. This was accomplished using two experimental facilities, a developed theoretical model, and Chemkin simulations. The first facility, termed the Afterburner Facility (AF), simulated the bulk flow temperature, velocity and O2 content through a turbojet combustor, turbine stage and afterburner. The second facility, termed the Propane Autoignition Combustor (PAC), was essentially a scaled-down, simplified version of the AF. The developed model was used to predict and interpret the AF results and to study the feasibility of the PAT concept at pressures outside the AF operating range. Finally, the Chemkin simulations were used to study the effect of several POx gas compositions on the afterburner combustion process.