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No description available.

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.

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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.

Gas Turbines (GTs) are used extensively in pipelines to compress gas at suitable points. The primary objective of this study is to look at CO2 return pipelines and the close coupling of the compression system with advanced prime mover cycles. Adopting a techno-economic and environmental risk analysis (TERA) frame work, this study conducts the modelling and evaluation of CO2 compression power requirements for gas turbine driven equipment (pump and compressor). The author developed and validated subroutines to implement variable stators in an in-house GT simulation code known as Variflow in order to enhance the off-design performance simulation of the code. This modification was achieved by altering the existing compressor maps and main program algorithm of the code. Economic model based on the net present value (NPV) method, CO2 compressibility factor model based on the Peng-Robinson equation of state and pipeline hydraulic analysis model based on fundamental gas flow equation were also developed to facilitate the TERA of selected GT mechanical drives in two case scenarios. These case scenarios were specifically built around Turbomatch simulated GT design and off-design performance which ensure that the CO2 is introduced into the pipeline at the supercritical pressure as well as sustain the CO2 pressure above a minimum designated pressure during transmission along an adapted real life pipeline profile. The required compression duty for the maximum and minimum CO2 throughput as well as the operation site ambient condition, guided the selection of two GTs of 33.9 MW and 9.4 MW capacities. At the site ambient condition, the off design simulations of these GTs give an output of 25.9 MW and 7.6 MW respectively. Given the assumed economic parameters over a plant life of 25 years, the NPV for deploying the 33.9 MW GT is about £13.9M while that of the 9.4 MW GT is about £1.2M. The corresponding payback periods (PBPs) were 3 and 7 years respectively. Thus, a good return on investment is achieved within reasonable risk. The sensitivity analysis results show a NPV of about £19.1M - £24.3M and about £3.1M - £4.9M for the 33.9 MW and 9.4 MW GTs respectively over a 25 - 50% fuel cost reduction. Their PBPs were 3 - 2 years and 5 - 4 years respectively. In addition, as the CO2 throughput drops, the risk becomes higher with less return on investment. In fact, when the CO2 throughput drops to a certain level, the investment becomes highly unattractive and unable to payback itself within the assumed 25 years plant life. The hydraulic analysis results for three different pipe sizes of 24, 14 and 12¾ inch diameters show an increase in pressure drop with increase in CO2 throughput and a decrease in pressure drop with increase in pipe size for a given throughput. Owing to the effect of elevation difference, the 511 km long pipeline profile gives rise to an equivalent length of 511.52 km. Similarly, given the pipeline inlet pressure of 15 MPa and other assumed pipeline data, the 3.70 MTPY (0.27 mmscfd) maximum average CO2 throughput considered in the 12¾ inch diameter pipeline results in a delivery pressure of about 15.06 MPa. Under this condition, points of pressure spikes above the pipeline maximum operating allowable pressure (15.3 MPa) were obtained along the profile. Lowering the pipeline operating pressure to 10.5 MPa gives a delivery pressure of about 10.45 MPa within safe pressure limits. At this 10.5 MPa, over a flat pipeline profile of same length, the delivery pressure is about 10.4 MPa. Thus, given the operating conditions for the dense phase CO2 pipeline transmission and the limit of this study, it is very unlikely that a booster station will be required. So also, compressing the CO2 to 15 MPa may no longer be necessary; which eliminates the need of combining a compressor and pump for the initial pressure boost in order to save power. This is because, irrespective of the saving in energy, the increase in capital cost associated with obtaining a pump and suitable driver far outweighs the extra expense incurred in acquiring a rated GT mechanical drive to meet the compression duty.

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.

This work is devoted to generate knowledge and high quality experimental data of catalytic combustion at operational gas turbine conditions. The initial task of the thesis work was to design and construct a high pressure combustion test facility, where the catalytic combustion experiments can be performed at real gas turbine conditions. With this in mind, a highly advanced combustion test facility has been designed, constructed and tested. This test facility is capable of simulating combustion conditions relevant to a wide range of operating gas turbine conditions and different kinds of fuel gases. The shape of the combustor (test section) is similar to a “can” type gas turbine combustor, but with significant differences in its type of operation. The test combustor is expected to operate at near adiabatic combustion conditions and there will be no additions of cooling, dilution or secondary supply of air into the combustion process. The geometry of the combustor consists of three main zones such as air/fuel mixing zone, catalytic reaction zone and downstream gas phase reaction zone with no difference of the mass flow at inlet and exit. The maximum capacity of the test facility is 100 kW (fuel power) and the maximum air flow rate is 100g/s. The significant features of the test facility are counted as its operational pressure range (1 – 35 atm), air inlet temperatures (100 – 650 °C), fuel flexibility (LHV 4 - 40 MJ/m3) and air humidity (0 – 30% kg/kg of air). Given these features, combustion could be performed at any desired pressure up to 35 bars while controlling other parameters independently. Fuel flexibility of the applications was also taken into consideration in the design phase and proper measures have been taken in order to utilize two types of targeted fuels, methane and gasified biomass. Experimental results presented in this thesis are the operational performances of highly active precious metal catalysts (also called as ignition catalysts) and combinations of precious metal, perovskites and hexaaluminate catalysts (also called as fully catalytic configuration). Experiments were performed on different catalytic combustor configurations of various types of catalysts with methane and simulated gasified biomass over the full range of pressure. The types of catalysts considered on the combustor configurations are palladium on alumina (Pd/AL2O3), palladium lanthanum hexaaluminate (PdLaAl11O19), platinum on alumina (Pt/AL2O3),and palladium:platinum bi-metal on alumina (Pd:Pt/AL2O3). The influence of pressure, inlet temperature, flow velocity and air fuel ratio on the ignition, combustion stability and emission generation on the catalytic system were investigated and presented. Combustion catalysts were developed and provided mainly by the project partner, the Division of Chemical Technology, KTH. Division of Chemical Reaction Technology, KTH and Istituto di Ricerche sulla Combustione (CNR) Italy were also collaborated with some of the experimental investigations by providing specific types of catalysts developed by them for the specific conditions of gas turbine requirements. / <p>QC 20131125</p>

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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.