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Gas turbine engine performance improvement has been requested continuously for both military and commercial applications due t various reasons. One of the issues is to save fuel and/or to increase the engine life to meet the multi-mission and operation cost economics requirements. I order to satisfy the customers' requirements, the engine manufacturers invested a lot of money and time if the gas turbine performance improvement. The most straight forward and simple approach is to trade the excess remained surge margins for performance. NASA has demonstrated the feasibility of this concept in their F-15 Highly Integrated Digital Electronic Control and Performance Seeking Control programs. It offers not only obvious benefits if the overall system performance improvement but also cost effective operations such a fuel saving and extended component life. Those were carried out with traditional control approaches which have to face the modelling difficulties. ' Due to successful control implementations of fuzzy logic if various environment of uncertainties, a proportional plus integral z logic controller if proposed. The fuzzy logic control system simulation results prove that the fuzzy logic controller is appropriate for gas turbine engine control. Basic fuzzy logic control concept is used with new approaches to simplifying the fuzzy logic controller. I order to enhance the engine performance, fuzzy logic control concept is used to optimize the engine performance parameters. A time function linear control scheme is proposed to the engine to a new operation location System simulation results prove the new methodology. It has to be understood that the engine model used if this research is not representative of a gas turbine, but it `is appropriate for the fuzzy logic control design analysis and simulation.

This thesis studies the application of sensitivity analysis and optimization methods to the design of turbomachinery components. Basic design issues and a survey of current design trends are presented. The redesign of outlet guide vanes (OGV's) in an aircraft high bypass turbofan engine is attempted. The redesign is necessitated by the interaction of the pylon induced static pressure field with the OGV's and the fan, leading to reduced OGV efficiency and shortened fan life. The concept of cyclically varying camber is used to redesign the OGV row to achieve suppression of the downstream disturbance in the domain upstream of the OGV row. The redesign is performed using (a) a linear perturbation CFD analysis and (b) a minimisation of the pressure mismatch integral by using a Newton method. In method (a) the sensitivity of the upstream flow field to changes in blade geometry is acquired from the linear perturbation CFD analysis, while in method (b) it is calculated by perturbing the blade geometry and differencing the resulting flow fields. Method (a) leads to a reduction in the pylon induced pressure variation at the fan by more than 70% while method (b) achieves up to 86%. An OGV row with only 3 different blade shapes is designed using the above method and is found to suppress the pressure perturbation by more than 73%. Results from these calculations are presented and discussed. The quasi-Newton design method is also used to redesign a three dimensional OGV row and achieves considerable reduction of upstream pressure variation. A concluding discussion summarises the experiences and suggests possible avenues for further work.

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Gas turbine engine NOX, CO2 and H20 exhaust emissions from civil subsonic fleets are potentially in sufficient amounts to affect atmospheric ozone and climate, particularly with the projected growth in air traffic. Because, it may be that the future envisaged low-NOX combustor technologies may not keep up with the industry requirements for increased engine thermal efficiency, the potential benefits from optimising aero engine cycles and flight operations for low mission emissions deserve to be thoroughly investigated. The SKY computer simulation system developed to examine such alternative routes integrates flight route performance, aero engine performance and the formation of pollutants within the combustor. Based on Turbomatch, Cranfield Gas Turbine Simulation System, SKY can be used to optimise mission/aircraft/engine/combustor combinations with respect to landing and take-off (LTO) as well as mission emissions. A model of the high capacity Boeing 747-400 powered by Turbomatch high bypass ratio turbofan models and simulated on long range routes such as London-Tokyo is selected for this work. On the one hand, aero engine cycles can be designed at a optimum bypass ratio and deliver mission NOX reductions of up to 10% over designs optimised for LTO NOX, indicating that the current ICAO regulatory regime is a inadequate parameter to control mission NOX. On the other hand, operational measures such as speed reductions could bring further reductions of the order of 10%, but some of the improvement would be made at the expense of fuel burn, CO2 and H20 emissions, payload-range capability and direct operating costs. The benefits from such alternative routes are not negligible but smaller in comparison to the 30 to 80% potential cuts from future low-NOX technology, as well as to the 30% reduction due to expected improvements in the next 20 years or so in airframe weight and aerodynamics and more efficient navigation practices.

Early stages of axial flow turbine design require a relatively simple prediction technique for estimating both blade row exit angle and loss profiles produced by secondary flows. Detailed experimental investigation of the flow field in a large scale linear cascade of high turning turbine rotor blades has been made. This gave improved understanding of cascade secondary flow phenomena. and a physical basis for secondary flow angle and loss predictions. Data suitable for comparison with three dimensional flow calculations is presented. Experimental data was obtained utilizing cobra probes throughout the flow field. and hot wire probes at cascade inlet and exit. Results are presented graphically on various planes through the flow field using both contour and vector plots. The developing passage and leading edge horseshoe vortic 3S are traced. and their interactions with the cascade inlet boundary layer are clearly visible. At cascade exit two major secondary loss components were identified: a loss core shed from the suction surface formed largely of inlet boundary layer fluid. and an area consisting of new endwall boundary layer fluid swept towards the suction surface. Highly turbulent flows were also evident close to these regions. Secondary losses were predicted using three discrete loss components: the loss core. a non skewed new endwall boundary layer. and an extra secondary loss related to the classical secondary flow kinetic energy. Experimental data from several sources was compared with secondary loss predictions with some success. Some modifications are clearly desirable to enhance the loss prediction technique. but the relatively simple method gives encouraging results.

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This thesis investigates the concept of thrust rating as a means towards reducing the life cycle costs of engine ownership. Towards this end, this thesis has discussed the concept of thrust rating, developed computer programs for mechanical load type failures, which include creep, LCF, and combinations thereof, and conducted simulations of improving life usage and reducing life cycle costs. A study was performed on a military engine, under an original design mission mix, that showed significant gains in creep-LCF life of the HPT blade could be achieved, especially With the recently proposed and presumably more accurate criterion- ductility exhaustion, by thrust rating. The savings were expressed in terms of an approximate reduced life accumulation rates and life cycle costs. The net result was a 50% increase in creep-LCF life with a savings of $50.4 million. These calculations were based on a Feet of 300 engines having the designed lifetime of 8,000 operating hours per engine. Throughout the thesis, mention is also made of employing the thrust rating concept on other engines. To this end, the thesis will also give a blueprint for conducting a feasibility study to employ thrust rating as a maintenance tool. In addition to the technical aspects, the role of maintenance and aircraft operations policy will also be studied to determine the interrelationships that exist between thrust rating technology and its practical application.

The objective of the work described in this thesis is twofold; to elucidate the nature of stall and surge in an axial flow aeroengine compressor, and to improve on current computational stall modelling techniques. Particular attention is paid to the initial stages of the stall/surge transient, and to the possibility of using active control techniques to prevent or delay the onset of stall/surge. A detailed analysis is presented of measurements of the stalling behaviour of a Rolls- Royce VIPER jet engine, showing a wide variety of stall inception and post-stall behaviour. Stall transients are traced from disturbances through to stable rotating stall or axisymmetic surge. The stall inception pattern at nearly all speeds is shown to conform to the short circumferential length scale pattern described by Day [1993a]. A multiple compressors in parallel stall model is developed using conventional stall modelling techniques, but extended to include the effects of the jet engine environment The model is shown to give a good representation of the overall stalling behaviour of the engine, although the details of the stall inception period are not accurately predicted. A system identification technique is applied to the results of the model in order to develop a method of active control of stall/surge. A new stall model is introduced and developed, based on a time-accurate three dimensional (but pitchwise averaged) solution of the viscous flow equations, with bladerow performance represented by body forces. The flow in the annulus boundary layers is calculated directly, and hence this new method is sufficiently complex to model the initial localised disturbances that lead to stall/surge. At the same time the computational power required is compatible with application to long multistage compressors.