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In gas turbine development, the direction has been toward higher turbine inlet temperatures to increase the work output and thermal efficiency. This extreme environment can significantly impact component life. One means of preventing component burnout in the turbine is to effectively use film-cooling whereby coolant is extracted from the compressor and injected through component surfaces. One such surface is the endwall of the first stage nozzle guide vane. This thesis details the design, prediction, and testing of two endwall film-cooling hole patterns provided by leading gas turbine engine companies. In addition a flush, two-dimensional slot was included to simulate leakage flow from the combustor-turbine interface. The slot coolant was found to exit in a non-uniform manner leaving a large, uncooled ring around the vane. Film-cooling holes were effective at distributing coolant throughout much of the passage, but at low blowing rates were unable to provide any benefit to the critical vane-endwall junction both at the leading edge and along the pressure side. At high blowing ratios, the increased momentum of the jets induced separation at the leading edge and in the upstream portion of the passage along the pressure side, while the jets near the passage exit remained attached and penetrated completely to the vane surface. Computational fluid dynamics (CFD) was successful at predicting coolant trajectory, but tended to under-predict thermal spreading and jet separation. Superposition was shown to be inaccurate, over-predicting effectiveness levels and thus component life, because the flow field was altered by the coolant injection. / Master of Science

The hot gas temperature in gas turbine engines is far above the permissible metal temperatures. Advanced cooling technologies must be applied to cool the blades, so they can withstand the extreme conditions. Film cooling is widely used in modern high temperature and high pressure blades as an active cooling scheme. In this study, the film cooling effectiveness in different regions of gas turbine blades was investigated with various film hole/slot configurations and mainstream flow conditions. The study consisted of four parts: 1) effect of upstream wake on blade surface film cooling, 2) effect of upstream vortex on platform purge flow cooling, 3) influence of hole shape and angle on leading edge film cooling and 4) slot film cooling on trailing edge. Pressure sensitive paint (PSP) technique was used to get the conduction-free film cooling effectiveness distribution. For the blade surface film cooling, the effectiveness from axial shaped holes and compound angle shaped holes were examined. Results showed that the compound angle shaped holes offer better film effectiveness than the axial shaped holes. The upstream stationary wakes have detrimental effect on film effectiveness in certain wake rod phase positions. For platform purge flow cooling, the stator-rotor gap was simulated by a typical labyrinth-like seal. Delta wings were used to generate vortex and modeled the passage vortex generated by the upstream vanes. Results showed that the upstream vortex reduces the film cooling effectiveness on the platform. For the leading edge film cooling, two film cooling designs, each with four film cooling hole configurations, were investigated. Results showed that the shaped holes provide higher film cooling effectiveness than the cylindrical holes at higher average blowing ratios. In the same range of average blowing ratio, the radial angle holes produce better effectiveness than the compound angle holes. The seven-row design results in much higher effectiveness than the three-row design. For the trailing edge slot cooling, the effect of slot lip thickness on film effectiveness under the two mainstream conditions was investigated. Results showed thinner lips offer higher effectiveness. The film effectiveness on the slots reduces when the incoming mainstream boundary layer thickness decreases.

The hot gas temperature in gas turbine engines is far above the permissible metal temperatures. Advanced cooling technologies must be applied to cool the blades, so they can withstand the extreme conditions. Film cooling is widely used in modern high temperature and high pressure blades as an active cooling scheme. In this study, the film cooling effectiveness in different regions of gas turbine blades was investigated with various film hole/slot configurations and mainstream flow conditions. The study consisted of four parts: 1) effect of upstream wake on blade surface film cooling, 2) effect of upstream vortex on platform purge flow cooling, 3) influence of hole shape and angle on leading edge film cooling and 4) slot film cooling on trailing edge. Pressure sensitive paint (PSP) technique was used to get the conduction-free film cooling effectiveness distribution. For the blade surface film cooling, the effectiveness from axial shaped holes and compound angle shaped holes were examined. Results showed that the compound angle shaped holes offer better film effectiveness than the axial shaped holes. The upstream stationary wakes have detrimental effect on film effectiveness in certain wake rod phase positions. For platform purge flow cooling, the stator-rotor gap was simulated by a typical labyrinth-like seal. Delta wings were used to generate vortex and modeled the passage vortex generated by the upstream vanes. Results showed that the upstream vortex reduces the film cooling effectiveness on the platform. For the leading edge film cooling, two film cooling designs, each with four film cooling hole configurations, were investigated. Results showed that the shaped holes provide higher film cooling effectiveness than the cylindrical holes at higher average blowing ratios. In the same range of average blowing ratio, the radial angle holes produce better effectiveness than the compound angle holes. The seven-row design results in much higher effectiveness than the three-row design. For the trailing edge slot cooling, the effect of slot lip thickness on film effectiveness under the two mainstream conditions was investigated. Results showed thinner lips offer higher effectiveness. The film effectiveness on the slots reduces when the incoming mainstream boundary layer thickness decreases.

In this dissertation, Immersed Boundary Method (IBM) is evaluated in ribbed duct geometries to show the potential of simulating complex geometry with a simple structured grid. IBM is first investigated in well-accepted benchmark cases: channel flow and pipe flow with circular cross-section. IBM captures all the flow features with very good accuracy in these two cases. Then a two side ribbed duct geometry is test using IBM at Reynolds number of 20,000 under fully developed assumption. The IBM results agrees well with body conforming grid predictions. A one side ribbed duct geometry is also tested at a bulk Reynolds number of 1.5⨉10⁴. Three cases have been examined for this geometry: a stationary case; a case of positive rotation at a rotation number (Ro=ΩDₕ/U) of 0.3 (destabilizing); and a case of negative rotation at Ro= -0.3 (stabilizing). Time averaged mean, turbulent quantities are presented, together with heat transfer. The overall good agreement between IBM, BCG and experimental results suggests that IBM is a promising method to apply to complex blade geometries. Due to the disadvantage of IBM that it requires large amount of cells to resolve the boundary near the immersed surface, wall modeled LES (WMLES) is evaluated in the final part of this thesis. WMLES is used for simulating turbulent flow in a developing staggered ribbed U-bend duct. Three cases have been tested at a bulk Reynolds number of 10⁵: a stationary case; a positive rotation case at a rotation number Ro=0.2; and a negative rotation case at Ro=-0.2. Coriolis force effects are included in the calculation to evaluate the wall model under the influence of these effects which are known to affect shear layer turbulence production on the leading and trailing sides of the duct. Wall model LES prediction shows good agreement with experimental data. / Master of Science

The present work has been an experimental investigation to evaluate the applicability of gas turbine cooling technology. With the temperature of the mainstream gas entering the turbine elevated above the melting temperature of the metal components, these components must be cooled, so they can withstand prolonged exposure to the mainstream gas. Both external and internal cooling techniques have been studied as a means to increase the life of turbine components. Detailed film cooling effectiveness distributions have been obtained on the turbine blade platform with a variety of cooling configurations. Because the newly developed pressure sensitive paint (PSP) technique has proven to be the most suitable technique for measuring the film effectiveness, it was applied to a variety of platform seal configurations and discrete film flows. From the measurements it was shown advanced seals provide more uniform protection through the passage with less potential for ingestion of the hot mainstream gases into the engine cavity. In addition to protecting the outer surface of the turbine components, via film cooling, heat can also be removed from the components internally. Because the turbine blades are rotating within the engine, it is important to consider the effect of rotation on the heat transfer enhancement within the airfoil cooling channels. Through this experimental investigation, the heat transfer enhancement has been measured in narrow, rectangular channels with various turbulators. The present experimental investigation has shown the turbulators, coupled with the rotation induced Coriolis and buoyancy forces, result in non-uniform levels of heat transfer enhancement in the cooling channels. Advanced turbulator configurations can be used to provide increased heat transfer enhancement. Although these designs result in increased frictional losses, the benefit of the heat transfer enhancement outweighs the frictional losses.

Gas turbine engines are extensively used in the aviation and power generation industries. They are used as topping cycles in combined cycle power plants, or as stand alone power generation units. Gains in thermodynamic efficiency can be realized by increasing the turbine inlet temperatures. Since modern turbine inlet temperatures exceed the melting point of the constituent superalloys, it is necessary to provide an aggressive cooling system. Relatively cool air, ducted from the compressor of the engine is used to remove heat from the hot turbine blade. This air flows through passages in the hollow blade (internal cooling), and is also ejected onto the surface of the blade to form an insulating film (film cooling). Modern land-based gas turbine engines use high Reynolds number internal flow to cool their internal passages. The first part of this study focuses on experiments pertaining to passages with Reynolds numbers of up to 400,000. Common turbulator designs (45degree parallel sharp-edged and round-edged) ribs are studied. Older correlations are found to require corrections in order to be valid in the high Reynolds number parameter space. The effect of rotation on heat transfer in a typical three-pass serpentine channel is studied using a computational model with near-wall refinement. Results from this computational study indicate that the hub experiences abnormally high heat transfer under rotation. An experimental study is conducted at Buoyancy numbers similar to an actual engine on a wedge shaped model trailing edge, roughened with pin-fins and equipped with slot ejection. Results show an asymmetery between the leading and trailing surfaces due to rotation - a difference which is subdued due to the provision of pin-fins. Film cooling effectiveness is measured by the PSP mass transfer analogy technique in two different configurations: a flat plate and a typical high pressure turbine blade. Parameters studied include a step immediately upstream of a row of holes; the Strouhal number (quantifying rotor-stator interaction) and coolant to mainstream density ratio. Results show a deterioration in film cooling effectiveness with on increasing the Strouhal number. Using a coolant with a higher density results in higher film cooling effectiveness.

High combustion temperatures are required in gas-turbine engines to achieve high cycle efficiencies. With increasing temperature, however, the life span of the turbine components are reduced. The ability to accurately predict engine component temperature as a function of combustion temperature is required to strike this balance correctly. An experimental heat transfer measurement technique is developed in this thesis, which builds on a large body of existing literature. The technique enables a detailed quantification of turbine heat transfer on test rigs which closely represent gas-turbine engine configurations. Fundamental improvements are made to existing methods, in the definition of the ‘semi- infinite limit’ for transient measurement techniques, in Infra-red camera calibration, and in thermal effusivity measurement. The improvements were developed from first principles, verified experimentally, and have been used on a world leading heat transfer rig (the FACTOR combustor-turbine interaction rig, run on the NG-Turb facility at DLR Göttingen). It was found that optimisation of a number of measurement parameters was required to minimise the measurement uncertainty. It is shown that the optimum measurement parameters are dependant, and sensitive to the specific configuration of the test rig. An experimental procedure was developed and tested, which has been ‘tuned’ for measurements on the FACTOR test rig. Despite the challenging measurement environment on the FACTOR rig, it was found that state-of-the-art heat transfer measurement uncertainties of approximately 5%, could nevertheless still be achieved, by using the new methods. General principles and rules are established which can be used to guide the design of future heat transfer measurements, with the aim of minimising measurement uncertainty.

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