An experimental and numerical investigation of the flow and convective heat transfer distribution around a high turning angle gas turbine rotor blade has been carried out at the University of Kwa-Zulu, Durban campus. This study in gas turbine blade aerothermodynamics was done to meet the research and development requirements of the CSIR and ARMSCOR. The experimental results were generated using an existing continuously running supersonic cascade facility which offers realistic engine conditions at low operating costs. These results were then used to develop and validate a 2-D model created using the commercially available Computational Fluid Dynamics (CFD) software package, FLUENT. An initial phase of the study entailed a restoration of what was an unoperational experimental facility to a state capable of producing test simulation conditions. In the analysis, a 4-blade cascade system with provisions for an interchangeable, test blade was subjected to the steady state conditions set up by the facility. Firstly, the flow was characterised by evaluating the static pressures around the midspan of a pressure measurement test blade. This was done using two pressure transducers, a scanivalve, an upgraded data acquisition system and LABview software. The method for measuring the heat transfer distributions made use of a transient measuring technique, whereby a pre-chilled Macor test blade, instrumented with thin film heat flux gauges was rapidly introduced into the hot cascade flow conditions by displacing an aluminum dummy blade while still maintaining the flow conditions. Measurement of the heat flux and generation of the isothermal heat transfer co-efficient distributions entailed re-instrumentation of the test blade section with gauges of increased temperature sensitivity along with modifications of the associated electrical circuitry to improve on the quality of experimental data. Both the experimental flow and heat transfer data were used to validate the CFD model developed in FLUENT. An investigation into different meshing strategies and turbulence models placed emphasis on the choice of model upon correlation. The outcome of which showed the k -co model's superiority in predicting the flow at transonic conditions. A feasibility study regarding a new means of implementing a film cooled turbine test blade at the supersonic cascade facility was also successfully investigated. The study comprised of experimental facility modifications as well as cascade and blade redesigns, all of which were to account for the requirements of film cooling. The implementation of this project, however, demanded the resources of both time and money of which neither commodity was available. / Thesis (M.Sc.Eng.)-University of KwaZulu-Natal, Durban, 2006.

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Thesis (M. App. Sc.)--Carleton University, 2004. / Includes bibliographical references (p. 131-139). Also available in electronic format on the Internet.

Thesis (Ph.D.) - Carleton University, 2005. / Includes bibliographical references (p. 295-301). Also available in electronic format on the Internet.

Thesis (M. App. Sc.)--Carleton University, 2004. / Includes bibliographical references (p. 221-230). Also available in electronic format on the Internet.

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1993. / Vita. Abstract. Includes bibliographical references (leaves 113-115). Also available via the Internet.

In current design methods for gas turbines there are important features of the flow which are not yet within the scope of the available prediction methods for both the calculation of surface pressures and heat transfer rates. Such features include the prediction of three-dimensional viscous flowfields, the accurate location and strengths of the secondary flow regimes in a turbine passage, and allowance for time-dependent variations. It is the understanding of the time-varying phenomena which is the subject of this study. Such phenomena occur due to the periodic interaction between stages in a turbine, either that of a nozzle guide vane on a rotor downstream or vice-versa. In most contemporary designs of turbines the effects are due primarily to the wakes from the trailing-edge of the upstream airfoil, and to any associated shock structures resulting from transonic exit flow Mach numbers. The present investigation is concerned with furthering knowedge of these wake and shock interactions, using a method of simulation established in the Isentropic Light Piston Tunnel and Oxford. Measurements of heat transfer rates and pressures are presented, supported by flow visualisation methods such as surface oil-dots and schlieren photography, for two examples of high-pressure turbine rotor blades. The majority of analysis deals with the first of these (a highty-loaded transonic profile) whilst the second blade (designed for use in a large civil engine) is included for investigation of the effects of flow unsteadiness on the film cooling process The transition process is examined in detail by use of wide bandwith heat transfer measurements, and a new method derived for modelling this process. It has been possible to observe the effect of the enhanced turbulence in the simulated nozzle guide vane wake and effects due a shock-boundary layer interaction. The reaction of the blade boundary layers to these disturbances is identified, and trajectories of disturbed events tracked along the blade surfaces. The measurements which have been taken allow for some aspects of wake and shock interactions to be included in the design process for turbine blading. A better understanding has been obtained of how these types of transient flow regimes affect the boundary layers on the blade surfaces.

This thesis documents an experimental investigation of the base (trailing edge) pressure and its approximate distribution on a transonic turbine blade. Since the base pressure plays an important role in determining the profile loss on blades with thick trailing edges, both the base pressure and the blade losses are presented for a range of transonic exit Mach numbers. The overall objective of this work is to provide experimental data for improving current computer-based models used in designing turbine blades. The two-dimensional cascade was tested in the VPI&SU Transonic Cascade Wind Tunnel, a blow-down type of tunnel facility. The blade design for the cascade was based on the pitchline profile of the high-pressure turbine in a commercial jet engine with a design exit Mach number of approximately 1.2. In order to carefully instrument the thin trailing edge, the blades used in the experiment were made five times the size of the actual engine blade. With this large-scale blade, five static pressure taps were placed around the trailing edge. In addition to these taps, the rearward portion of the suction surface was also instrumented with five static pressure taps. The aerodynamic losses were quantified by a loss coefficient: the mass-averaged total pressure drop divided by the total pressure upstream of the blade row. These measured pressures were taken with a fixed total pressure probe upstream of the cascade and a pitchwise traversing probe in the downstream position. The cascade was tested for an exit Mach number ranging from 0.70 to 1.40. The results of the experiments indicate a decreasing normalized base pressure (p_B/p_t1) with increasing downstream Mach number (M₂) until the minimum value of p_B/p_t1 = 0.30 at M₂ = 1.30. The approximate base pressure distributions for all transonic downstream Mach numbers indicate nearly uniform pressure around the central 90° of the trailing edge. Results for the profile loss are displayed for exit Mach numbers between 0.70 and 1.35; the trend of increasing loss with decreasing base pressure is shown. The shadowgraph pictures taken reveal the trailing edge region of the flow for several downstream transonic Mach numbers. / Master of Science

This thesis documents an experimental investigation into the effect of trailing edge Injection on the aerodynamic performance of turbine blades conducted at Virginia Polytechnic Institute and State University (VPl&SU). A brief description of the arrangement, instrumentation and data acquisition system of the VPl&SU Transonic Cascade Wind Tunnel is given. Testing was conducted under a number of test conditions. Baseline data was obtained for the blades with no trailing edge injection. The blades were then tested for two different blowing rates to test the effect of blowing rate on the total pressure loss coefficient, L. Tests were conducted at a variety of save cascade exit Mach numbers ranging from 0.79 to 1.36. Measurements were taken at three locations downstream of the cascade blade trailing edges. The algorithm used to calculate the L from the measured data is discussed. Results of the testing indicate that trailing edge injection has a negligible effect on the total pressure loss coefficient. Correlations of cascade exit Mach number to L are given. The development of L downstream of the blade trailing edge is discussed. / Master of Science

An investigation of tip leakage flow and its effects on loss production was performed on a large-scale linear turbine cascade having a tip gap measuring 2.1% of the blade height. The Reynolds number based on axial chord and cascade exit velocity was 4.5x10⁵. The experimental work began with a visualization study of the flow in and around the tip gap. The actual flow measurements consisted of two phases, the tip gap exit plane measurements for determination of the losses incurred within the tip gap, and the downstream measurements for determination of the overall cascade losses. The downstream measurements made 140% of an axial chord downstream of the blade leading edges show the development of the leakage flow and its associated losses. Numerical analyses of the data were used to evaluate various flow properties at both the tip gap exit plane and the downstream measurement plane. Using the measured downstream flow, a mixing analysis was performed to estimate the maximum loss of the cascade. Models of the flow were developed to explain and quantify the various factors contributing to the cascade's overall loss. At a particular downstream location, the additional loss due to tip leakage was found to be the sum of the measured loss at the tip gap exit plane and the amount of tip gap secondary kinetic energy which had been dissipated by that downstream location. / M.S.