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No description available.

No description available.

No description available.

No description available.

No description available.

Steady state, inter row measurements in multistage axial compressors are relevant to the current design process. The objective in obtaining such data is for evaluation of compressor blading as well as validation for the computer programmes used in compressor design. Multi-hole pressure probes are a reliable and economical method of collecting detailed flowfield data in compressors for these purposes. These probes are calibrated in a uniform flow in a wind tunnel prior to use, to determine their response to a range of flow angles and speeds. When the probe is subsequently used for measurements in the compressor, often the small inter row spacing means that the probe has to be close to the downstream stator passage and upstream rotor. The result is that the probe is no longer situated in the uniform flow in which it was calibrated, in terms of influences from both the upstream rotor and downstream stator. This project presents the investigation of these two effects on steady state pressure probe measurements. The effects of blockage on a probe positioned in front of a stator row in a high speed compressor were studied using CFD. This was also carried out on a large scale probe in a low speed compressor. It was found that the blockage effect caused a reduced mass flow in the downstream stator passage which in turn lead to an altered flow angle and a small reduction in measured total pressure. Experimental in rig calibrations showed that the change in flow angle was due to an angular offset of the pressure distribution about the probe. These calibrations also showed that the wind tunnel calibration was valid in the compressor within a small angular range. The influence of the upstream rotor passing was studied using an unsteady CFD model. Responses of the individual probe ports and the deduced flow angle and total pressure indicated that the steady state blockage effect is present throughout the wake passing. The wake passing was found to be a largely two dimensional effect in that the radial flow component changes in the low speed compressor wakes had little influence. The Total Technology thesis incorporates a management project on the relevant topic of project selection within companies. An existing project selection model was applied to a sample group of projects to determine the applicability of such models. The main findings were that these models can generate useful information for further selection decisions and that the applicability is towards lower budget projects where a structured approach is often not used.

A computational study has been performed to investigate the effects of blade row interaction on the performance of radial turbomachines, which was motivated by the need to improve our understanding of the blade row interaction phenomena for further improvement in the performance. High-speed centrifugal compressor stages with three settings of radial gap are configured and simulated using a three-dimensional Navier-Stokes flow method in order to investigate the impact of blade row interaction on stage efficiency. The performance predictions show that the efficiency deteriorates if the gap between blade rows is reduced to intensify blade row interaction, which is in contradiction to the general trend for stage axial compressors, hi the compressors tested, the wake chopping by diffuser vanes, which usually benefits efficiency in axial compressor stages, causes unfavourable wake compression through the diffuser passages to deteriorate the efficiency. Similarly, hydraulic turbine stages with three settings of radial gap are simulated numerically. A new three-dimensional Navier-Stokes flow method based upon the dual-time stepping technique combined with the pseudo-compressibility method has been developed for hydraulic flow simulations. This method is validated extensively with several test cases where analytical and experimental data are available, including a centrifugal pump case with blade row interaction. Some numerical tests are conducted to examine the dependency of the flow solutions on several numerical parameters, which serve to justify the sensitivity of the solutions. Then, the method is applied to performance predictions of the hydraulic turbine stages. The numerical performance predictions for the turbines show that, by reducing the radial gap, the loss generation in the nozzle increases, which has a decisive influence on stage efficiency. The blade surface boundary layer loss and wake flow mixing loss, enhanced with a higher level of flow velocity around blading and the potential flow disturbances, are responsible for the observed trend.

An experimental investigation has been carried out to determine the aerodynamic performance of an annular S-shaped duct representative of that used to connect the compressor spools of aircraft gas turbine engines. Measurements of both the mean flow and turbulent structure have been obtained using both 5 hole pressure probes and a3 component Laser Doppler Anemometry (LDA) system. The measurements indicate that development of the flow within the duct is complex and significantly influenced by the combined effects of streamwise pressure gradients and flow curvature. For inlet conditions in which boundary layers are developed along an upstream entry length the static pressure, shear stress and velocity distributions are presented. The data shows that as a result of flow curvature significant streamwise pressure gradients exist within the duct, with this curvature also affecting the generation and suppression of turbulence. The stagnation pressure loss within the duct is also assessed and is consistent with the measured distributions of shear stress. More engine representative conditions are provided by locating a single stage compressor at inlet to the duct. Relative to the naturally developed inlet conditions the flow within the duct is less likely to separate, but mixing out of the compressor blade wakes increases the measured duct loss. With both types of inlet conditions the effect of a radial strut, such as that used for carrying loads and engine services, is also described both in terms of the static pressure distribution along the strut and its contribution to overall loss. The effects of inlet swirl on the flow field that develops within an annular S-shaped duct have also been investigated. By removing the outlet guide vanes from an upstream single stage compressor swirl angles in excess of 30° were generated. Results show that within the S-shaped duct tangential momentum is conserved, leading to increasing swirl velocities through the duct as its radius decreases. Furthermore, this component influences the streamwise velocity as pressure gradients are established to ensure the mean flow follows the duct curvature. Consequently in the critical region adjacent to the inner casing, where separation is most likely to occur, higher streamwise velocities are observed. Within the duct substantial changes also occur to the turbulence field which results in an increased stagnation pressure loss between duct inlet and exit. Data is also presented showing the increasing swirl angles through the duct which has consequences both for the design of the downstream compressor spool and of any radial struts which may be located within the duct.

Core engine icing has been recognized to affect a wide variety of engines since the 1990's. This previously unrecognized form of icing occurs in flights through high altitude convective regions and vicinity of thunderstorms. Engine icing events involve power loss or damage associated to the engine core, namely instabilities such as compressor surge, stall, engine rollback and even combustor flameout events. The effects on compressor performance are significant in understanding the response of the engine to atmospheric ice ingestion. A one-dimensional axisymmetric flow model is used to simulate the continuous phase through the compressor. The steady state operation of dry air is validated with an industrial database. By changing an exit throttle, the point where the dry compressor mass flow rate slowly starts to drop, is predicted. The stage that is the first to locally collapse, causing the remaining stages and eventually the complete compressor failure, is determined. The continuous flow model is then coupled with a Lagrangian model for the discrete phase in a framework that conserves mass, momentum and energy. From numerical simulations of the coupled, continuous-discrete phase flow model, it is observed that a rematching of the stages across the compressor occurs with increasing ice flow rates to accommodate loss of energy to the ice flow. The migration of the operating point towards the stall point at the rear stage eventually causes the compressor to stall. The onset of stall is characterized by initial oscillations followed by a rapid decay of pressures of the last stage with the instability traveling quickly towards the front of the compressor. Effectively, a reduction in the compressor stall margin is observed as the ice flow rate increases. Further, the relevance of factors such as blockage due to discrete particles and break/splash semi-empirical models in the icing physics, are analyzed through parametric studies. Conclusions are drawn that underscore the influence of the assumptions and models in prediction of the flow behavior in the presence of ice ingestion. Smaller ice crystal diameters have a greater influence on the gas flow dynamics in terms of a higher reduction in surge margin. The break empirical model for ice crystals and splash model for the droplets that are used to calculate the secondary particle size upon impact with rotor blades have a significant influence on the gas flow predictions.