No description available.

No description available.

No description available.

No description available.

The performance requirements of turbofan engines demands a stability and transient capability beyond that associated with the past generations of gas turbine engines. The axial flow fan unit is most vulnerable to loading limitations due to the primary problems associated with the compression process, its sensitivity to inlet distortion and the difficulty to design for an overall optimum blade duty in a machine of wide radial blade loading distribution. The development of mathematical models with some capability of predicting the stable operating range of an axial flow fan has to overcome the difficulties associated with the modelling of the radially distinct flow regions and their dynamic interaction. ' The current investigation combined the available knowledge of one-dimensional models (based on the principles of conservation of mass, linear momentum and energy) with the assumptions of the parallel compressor theory, in order to develop a linearized system of equations for stability analysis (surge prediction). The stability conditions which emerged from this approach were applied on the experimentally derived characteristics of a low hub to tip ratio split flow fan in a manner which involved the modelling of the dynamic interaction of the inner and outer flow region of the fan. The development of the governing equations was achieved by applying one-dimensional flow analysis to the inner and outer section of the fan. Their interaction was modelled on the experimentally obtained radial movement of the splitter streamline and the discharge ,static pressure 'radial distribution. The inner and outer region were treated as a lumped volume element search operating on a local masflow averaged total pressure rise characteristic and alternatively acting in conjunction with a common nozzle and separate nozzles. The experimental investigation was carried out on a low hub totipratio two-stage split flow fan(with the facility of independent bypass and core throttles)in order to examine the localised and overall performance of such a fan(and the staling processes involved)and to enable the application of the stability analysis. The influence of reducing the distance between the fan flow spliter and the last bladerowasal so investigated, «The mathematical mode1s predicted the point of dynamic instability within 4.52 of the experimental observed mas flow rate and pressure is value.

The ingestion of hot mainstream gas into the wheel-space between the rotor and staler discs is one of the most important internal cooling problems for gas turbine designers. To solve this problem, engineers design a rim seal at the periphery of wheel-space and direct a sealing flow from the internal cooling system to prevent ingress. The main aim of this thesis is to build a simple computational model to predict the scaling effectiveness of externally-induced ingress for engine designers. The axisymmetric model represents a gas turbine wheel-space and provides useful information related to the fluid dynamics and heat transfer in the wheel-space. At the same time, this model saves much computation time and cost for engine designers who currently use complex and time-consuming 3D models. The- computational model in this -thesis is called the prescribed ingestion model. Steady simulations are carried out using the commercial CFD code, ANSYS CFX with meshes built using ICEM CFD. Boundary conditions are applied at the ingress inlet of the model using experimental measurements and a mass-based averaging procedure. Computational parameters such as rotational Reynolds number, non-dimensional sealing flow rate and thermal conditions on the rotor are selected to investigate the fluid dynamics and heat transfer at typical experimental rig operating conditions. Different rim seal geometries arc investigated and results are compared with experimental data. In addition to the prescribed ingestion model, two typical axisymmetric rotor-stator system models without ingress arc established. The aim of these rotor-stator models is to investigate the fluid dynamics and heat transfer of the wheel-space in the situation without ingress. The effects of geometry and turbulence model also arc studied in these simulations. Most results from these simulations are in good agreement with experimental data from the literature, which enhances confidence in the prescribed Ingestion model.

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The specific power (or specific thrust) of modern gas turbines is much influenced by the gas temperature at turbine inlet. Even with the use of the best superalloy available and the most advanced cooling configurations, there are competitive pressures to operate engines at even higher gas temperatures. Ceramic coatings operate as thermal barriers and can allow the gas temperature to be increased by 50 to 220 K over the operating gas temperature for an uncoated turbine . It is important that the surface temperature of the blade be determined as accurately as possible. Large uncertainties as to the surface temperature require significant margins for safe operation . Blade surface temperatures can be determined with an accuracy of 10 K using radiation pyrometry and about'30 to 40 K by calculating the blade temperature based on---gas temperature measurement of the exhaust gas plane. This'- makes pyrometry an attractive option for advanced high temperature gas turbines . However, there is little experience in measuring surface temperatures of blades coated with ceramic coatings. There is evidence that the. radiation signal picked up by the pyrometer will not only depend on the surface temperature but also on a number of optical properties of the coating. Important among these are the emissivity of the coating and whether the coating is translucent. Parameters affecting this are the coating material, coating surface finish, coating thickness and whether or not a bond coat is used . This work explores these variables in a rig that simulates the conditions within a turbine stage of a gas turbine engine. In which six thermal barrier coating systems were tested. These systems are of current interest to gas turbine manufacturers and users. They include the latest advances in coating technology. Four stabilized zirconia systems and two alumina based systems were tested. It was found experimentally that the surface emissivity of these coating systems was invariant over the range 873 to 1023 K surface temperature. It was found that the use of different stabilizers did not affect the surface spectral emissivity. In further experiments six turbine wheels were coated with these systems and tested at turbine entry temperatures of 973, 1073, and 1173 K. It was found that the blade surface temperature was function of the coating material, coating thickness and turbine entry temperature. The blade surface temperature was also function of the blade height being maximum at the blade tip and minimum at the blade root . It was found that the C-YPSZ was better insulator than the rest of the systems. Whilst the blades coated with zirconia based systems suffered minor loss near the edges, the two alumina based systems were lost from more than a blade during the test. This coating loss was picked up by. the pyrometer . Analysis shows that the measured blade surface temperature was within 10 K of that calculated. The use of 0.3 mm of C-YPSZ on air cooled turbine blades caused 250 K surface temperature increase and 270 K metal temperature decrease for turbine entry temperature of 1673 K. The metal temperature reduction was as high as 310 K for coating thickness of 0.5 mm.

This thesis presents a detailed investigation of the capability of a modern three-dimensional Navier-Stokes solver to predict the secondary flows and losses in a linear cascade of high turning turbine rotor blades. Three codes were initially tested, to permit selection of the best of the available numerical solvers for this case. This program was then tested in more detail. Results showed that although very accurate prediction of the effects of inviscid fluid mechanics is now possible, the Reynolds stress modelling can have profound effects upon the quality of the solutions obtained. Solutions using two different calculation meshes, have shown that the results are not significantly grid dependent. The flowfield of the cascade was traversed with hot-wires to obtain measurements of the turbulent Reynolds stresses. A turbulence generating grid was placed upstream of the cascade, to produce a more realistic inlet turbulence intensity. Results showed that regions of high turbulent kinetic energy are associated with regions of high total pressure loss. Calculation of eddy viscosities from the Reynolds stresses showed that downstream of the -cascade the eddy viscosity is fairly isotropic. Evaluation of terms in the kinetic energy equation, also indicated that both the normal and shear Reynolds stresses are important as loss producing mechanisms in the downstream flow. The experimental Reynolds stresses have been compared with those calculated from the eddy viscosity and velocity fields of Navier-Stokes predictions using a mixing length turbulence model, a one equation model, and K - ϵ model. It was found that in the separated, shear flows, agreement was poor, although the K - ϵ model performed best. Further experimental work is suggested to obtain data with which to determine the accuracy of the models within the blade and endwall boundary layers.

The loss of performance of a gas turbine engine at off-design is primarily due to the rapid drop of the major cycle performance parameters with decrease in power and this may be aggravated by poor component performance. More and more stringent requirements are being put on the performance demanded from gas turbines and if future engines are to exhibit performances superior to those of present day: engines, then a means must be found of controlling engine cycle such that the lapse rate of the major cycle parameters with power is reduced. In certain applications, it may be desirable to vary engine cycle with operating conditions in an attempt to re-optimize performance. Variable geometry in key engine components offers the advantage of either improving the internal performance of a component or re-matching engine cycle to alter the flow-temperature-pressure relationships. Either method has the potential to improve engine performance. Future gas turbines, more so those for aeronautical applications, will extensively use variable geometry components and therefore, a tool must exist which is capable of evaluating the off-design performance of such engines right from the conceptual stage. With this in mind, a computer program was developed which can simulate the steady state performance of arbitrary gas turbines with or without variable geometry in the gas path components. The program is a thermodynamic component-matching analysis program which uses component performance maps to evaluate the conditions of the gas at the various engine stations. The program was used to study the performance of a number of cycles incorporating variable geometry and it was concluded that variable geometry can significantly improve the off-design performance of gas turbines.