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 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.

No description available.

No description available.

Rocket plume flowfields have an importance due to their influence on the signature of the rocket and also on the distribution of the plume gases around the vehicle. Little information on the co-flowing situation exists other than a previous study at Oxford. This thesis thus represents a significant database for co-flowing rocket plumes of this form. The work presented deals with two new aspects of co- flowing rocket plumes in that detailed flowfield measurements have been made and plumes from twin nozzle have been investigated for the first time in this thesis. This study on twinjet rocket plumes was carried out using the University of Oxford Gun Tunnel. Twinjet rockets with nozzle exit Mach numbers of 3 and 5 were tested in quiescent and in co-flow at Mach 7 using nitrogen and hydrogen injections. A major feature of the twinjet case was the so-called impingement shock between the flows from the two nozzles. It was discovered that this shock was insensitive to the freestream and scaling parameters are suggested for its geometry. Comparisons with single equivalent thrust nozzles are made at downstream locations and similar Pitot pressure profiles were observed for nitrogen injection in a nitrogen freestream after approximately 3 nozzle diameters downstream. Shear layers were studied and fluctuations in this region were measured by fast-response Pitot pressure and heat transfer probes sampled at 1.1 MHz. The extent of the shear layer was deduced using a new Oxford Total Temperature Probe. With the freestream stagnation temperature at approximately 650 K and injected gas at 350 K, a linear variation for the deduced total temperature across the shear layer was obtained. This was consistent with the Pitot pressure variations across this region. Convective heat transfer coefficient fluctuations and flow total temperature fluctuations across rocket flowiields were obtained using three thin-film heat transfer probes and found to be closely correlated. Experimental results for the twinjet and the single jet were compared with CFD simulations and good overall agreements were achieved. Instrumentation for the hypersonic experiments was investigated and a fast-response (~ 20 kHz) Pitot probe suited for flows heavily contaminated with particulate was developed and tested.

No description available.

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.

The design and development of gas turbine combustors is a crucial but uncertain part of an engine development process. Combustion within a gas turbine is a complex interaction of, among other things, fluid dynamics, heat and mass transfer and chemical kinetics. At present, the design process relies upon a wealth of experimental data and correlations. The proper use of this information requires experienced combustion engineers and even for them the design process is very time consuming. Some major engine manufacturers have attempted to address the above problem by developing one dimensional computer programs based on the above test and empirical data to assist combustor designers. Such programs are usually proprietary. The present work, based on this approach has yielded DEPTH, a combustor design program. DEPTH ( Design and Evaluation of Pressure, Temperature and Heat transfer in combustors) is developed in Fortran-77 to assist in preliminary design and evaluation of conventional gas turbine combustion chambers. DEPTH can be used to carry out a preliminary design along with prediction of the cooling slots for a given metal temperature limit or to evaluate heat transfer and temperatures for an existing combustion chamber. Analysis of performance parameters such as efficiency, stability and NOx based on stirred reactor theories is also coupled. DEPTH is made sufficiently interactive/user-friendly such that no prior expertise is required as far as computer operation is concerned. The range of variables such as operating conditions, geometry, hardware, fuel type can all be effectively examined and their contribution towards the combustor performance studied. Such comprehensive study should provide ample opportunity for the designer to make the right decisions. It should also be an effective study aid. Returns in terms of higher thermal efficiencies is an incentive to go for combined cycles and cogeneration. In such cases, opting for higher cycle pressures together with a second or reheat combustor promise higher thermal efficiencies and exhaust temperatures and hence such designs are likely to be of interest. The concepts that are needed for understanding a double or reheat combustor are also addressed using the programme. A specific application of the programme is demonstrated through the design of a double combustor.

No description available.

No description available.