Flutter is an aeroelastic instability phenomenon that can result either in serious damage or complete destruction of a gas turbine blade structure due to high cycle fatigue. Although 90% of potential high cycle fatigue occurrences are uncovered during engine development, the remaining 10% stand for one third of the total engine development costs. Field experience has shown that during the last decades as much as 46% of fighter aircrafts were not mission-capable in certain periods due to high cycle fatigue related mishaps. To assure a reliable and safe operation, potential for blade flutter must be eliminated from the turbomachinery stages. However, even the most computationally intensive higher order models of today are not able to predict flutter accurately. Moreover, there are uncertainties in the operational environment, and gas turbine parts degrade over time due to fouling, erosion and corrosion resulting in parametric uncertainties. Therefore, it is essential to design engines that are robust with respect to the possible uncertainties. In this thesis, the robustness of an axial compressor blade design is studied with respect to parametric uncertainties through the Mu analysis. The nominal flutter model is adopted from [9]. This model was derived by matching a two dimensional incompressible flow field across the flexible rotor and the rigid stator. The aerodynamic load on the blade is derived via the control volume analysis. For use in the Mu analysis, first the model originally described by a set of partial differential equations is reduced to ordinary differential equations by the Fourier series based collocation method. After that, the nominal model is obtained by linearizing the achieved non-linear ordinary differential equations. The uncertainties coming from the modeling assumptions and imperfectly known parameters and coefficients are all modeled as parametric uncertainties through the Monte Carlo simulation. As iv compared with other robustness analysis tools, such as Hinf, the Mu analysis is less conservative and can handle both structured and unstructured perturbations. Finally, Genetic Algorithm is used as an optimization tool to find ideal parameters that will ensure best performance in terms of damping out flutter. Simulation results show that the procedure described in this thesis can be effective in studying the flutter stability margin and can be used to guide the gas turbine blade design.

A one-dimensional, stage-by-stage axial compression system mathematical model has been constructed which can describe system behavior during post-stall events such as surge and rotating stall. The model uses a numerical technique to solve the nonlinear conservation equations of mass, momentum, and energy. Inputs for blade forces and shaft work are provided by a set of quasi-steady stage characteristics modified by a first order lagging equation to simulate dynamic stage characteristics. The model was validated with experimental results for a three-stage, low-speed compressor and a nine-stage, high-pressure compressor. Using these models, a parametric study was conducted to determine the effect of inlet resistance, combustor performance, heat transfer, and stage characteristic changes due to hardware modification on post—stall system behavior. / Ph. D. / incomplete_metadata

The modern gas turbine engine requires innovative cooling techniques to protect its internal components from the harsh operating environment typically seen downstream of the combustor. Much research has been performed on the design of these cooling techniques thus allowing for combustion temperatures higher than the melting point of the parts within the turbine. As turbine inlet temperatures and efficiencies continue to increase, it becomes vitally important to correctly and realistically model all of the turbine's external cooling features so as to provide the most accurate representation of the associated heat transfer to the metal surfaces. This study examines the effect of several realistic endwall features for a turbine vane endwall. The first study addresses the effects of a mid-passage gap, endwall misalignment, and roughness on endwall film-cooling. The second study focuses on the effect of varying the combustor-to-turbine gap width. Both studies were performed in a large-scale low speed wind tunnel with the same vane geometry. Geometric and flow parameters were varied and the variation in endwall cooling effectiveness was evaluated. Results from these studies show that realistic features, such as surface roughness, can reduce the effectiveness of endwall cooling designs while other realistic features, such as varying the combustor-to-turbine gap width, can significantly improve endwall cooling effectiveness. It was found that, for a given coolant mass flowrate, a narrow combustor-turbine gap width greatly increased the coverage area of the leaked coolant, even increasing adiabatic effectiveness upstream of the vane stagnation point. The turbine designer can also more efficiently utilize leaked coolant from the combustor-to-turbine gap by controlling endwall misalignment, thereby reducing the overall amount of film-cooling needed for the first stage. / Master of Science

No description available.

Results are reported from a series of experiments simulating the combustion and expansion processes of a heavy-duty combustion turbine engine burning a heavy residual fuel oil. The tests were carried out in a turbine simulator device, consisting of a combustion chamber and a turbine first-stage nozzle cascade sector. Both film, air-cooled and closed-circuit, water-cooled nozzle sectors were tested. These sectors were four-vane, three-throat sections with throat cross-sectional areas of approximately 50 (10⁻⁴) m². The test fuel was simulated by adding the appropriate contaminants to no. 2 fuel oil. A series of seven full-length tests were performed, ranging in length from 22.5 to 88.2 hours. Four of the tests involved the watercooled nozzle sector and the remaining three used the air-cooled nozzle. The principle objectives of the tests were to assess the rate at which ash accumulates in the turbine nozzle and the relative difficulty in removing these deposits. The variable used to evaluate the extent of the ash deposit on the nozzle was the effective throat area, determined using the calculated gas flow rates, turbine nozzle inlet temperature, and the measured combustion chamber pressure. The parameters varied in the test program, other than the nozzle sectors, were the gas temperature and the gas pressure. The gas pressure variations served to vary the gas path surface temperatures at constant gas temperature. The test conditions were nominal turbine firing (nozzle exit) temperatures of 1283 and 1394 K and combustor pressures of 3 and 6 atmospheres. A 2-to-l pressure ratio was maintained across the nozzle to insure sonic conditions at the throat sections. With the exception of one test, the data show that the deposit rates in the water-cooled turbine nozzle were lower than in the air-cooled nozzle. The effect of increasing the gas temperature was to dramatically increase the ash deposition rates. Decreased gas pressures (and hence surface temperatures) resulted in reduced deposition rates. Ash cleanability was enhanced by water-cooling. Heat transfer data were analyzed from the water-cooled tests and gave significant insight into the ash deposit formation and removal phenomena. One of the more significant conclusions drawn from these data was that the major portion of the effective area decrease observed in a turbine nozzle because of ash deposits is due to the pressure face deposits. A computer simulation of a combustion turbine engine was developed to aid in the evaluation of the turbine simulator test data. Results from field tests of full-sized production engines burning residual oil were used in the simulation to determine the relationship between the extent of ash deposition (throat area reduction) and turbine efficiency. This result was then combined with data from the turbine simulator tests to produce a real-time computer simulation of full-sized combustion turbine engines having air- and water-cooled first-stage turbine nozzles. It was found that water-cooling of the turbine nozzle would result in an increase in engine availability of 27 per cent when operating on heavy residual fuel oil. / Ph. D.

Gas turbine companies are always focused on reducing capital costs and increasing overall efficiency. There are numerous advantages in reducing the number of airfoils per stage in the turbine section. While increased airfoil loading offers great advantages like low cost and weight, they also result in increased aerodynamic losses and associated issues. The strength of secondary flows is influenced by the upstream boundary layer thickness as well as the overall flow turning angle through the blade row. Secondary flows result in stagnation pressure loss which accounts for a considerable portion of the total stagnation pressure loss occurring in a turbine passage. A turbine designer strives to minimize these aerodynamic losses through design changes and geometrical effects. Performance of airfoils with varying loading levels and turning angles at transonic flow conditions are investigated in this study. The pressure difference between the pressure side and suction side of an airfoil gives an indication of the loading level of that airfoil. Secondary loss generation and the 3D flow near the endwalls of turbine blades are studied in detail. Detailed aerodynamic loss measurements, both in the pitchwise as well as spanwise directions, are conducted at 0.1 axial chord and 1.0 axial chord locations downstream of the trailing edge. Static pressure measurements on the airfoil surface and endwall pressure measurements were carried out in addition to downstream loss measurements. The application of endwall contouring to reduce secondary losses is investigated to try and understand when contouring can be beneficial. A detailed study was conducted on the effectiveness of endwall contouring on two different blades with varying airfoil spacing. Heat transfer experiments on the endwall were also conducted to determine the effect of endwall contouring on surface heat transfer distributions. Heat transfer behavior has significant effect on the cooling flow needs and associated aerodynamic problems of coolant-mainstream mixing. One of the primary objectives of this study is to provide data under transonic conditions that can be used to confirm/refine loss predictions for the effect of various Mach numbers and gas turning. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. A published experimental study on the effect of end wall contouring on such high turning blades at high exit Mach numbers is not available in open literature. Hence, the need to understand the parametric effects of endwall contouring on aerodynamic and heat transfer performance under these conditions. / Ph. D.

Airfoil endwall contouring has been widely studied during the past two decades for the reduction of secondary losses in turbine passages. Although many endwall contouring methods have been suggested by researchers, an analytical tool based on the passage design parameters is still not available for designers. Hence, the best endwall contour shape is usually decided through an optimization study. Moreover, a general guideline for the endwall shape variation can be extrapolated from the existing literature. It has not been validated whether the optimum endwall shape for one passage can be fitted to other similar passage geometry to achieve, least of all a non-optimum but a definite, reduction in losses. Most published studies were conducted at low exit Mach numbers and only recently some studies on the effect of endwall contouring on aerodynamics performance of a turbine passage at high exit Mach numbers have been published. There is, however, no study available in the open literature for a very high turning blade with a transonic design exit Mach number and the effect of endwall contouring on the heat transfer performance of a turbine passage. During the present study, a robust, aerodynamic performance based numerical optimization methodology for turbine endwall contouring has been developed. The methodology is also adaptable to a range of geometry optimization problems in turbomachinery. It is also possible to use the same methodology for multi-objective aero-thermal optimization. The methodology was applied to a high turning transonic turbine blade passage to achieve a geometry based on minimum total pressure loss criterion. The geometry was then compared with two other endwall geometries. The first geometry is based on minimum secondary kinetic energy value instead of minimum total pressure loss criterion. The second geometry is based on a curve combination based geometry generation method found in the literature. A normalized contoured surface topology was extracted from a previous study that has similar blade design parameters. This surface was then fitted to the turbine passage under study in order to investigate the effect of such trend based surface fitting. Aerodynamic response of these geometries has been compared in detail with the baseline case without any endwall contouring. A new non-contoured baseline design and two contoured endwall designs were provided by Siemens Energy, Inc. The pitch length for these designs is about 25% higher than the turbine passage used for the endwall optimization study. The aerodynamic performance of these endwalls was studied through numerical simulations. Heat transfer performance of these endwall geometries was experimentally investigated in the transonic turbine cascade facility at Virginia Tech. One of the contoured geometries was based on optimum aerodynamic loss reduction criterion while the other was based on optimum heat transfer performance criterion. All the three geometries were experimentally tested at design and off-design Mach number conditions. The study revealed that endwall contouring results in significant performance benefit from the heat transfer performance point of view. / Ph. D.

Efficiency and power output of gas turbines improve with an increase in turbine inlet temperatures, and blade designers continually seek out new methods of increasing these temperatures. Increases in turbine inlet temperatures are achieved by utilizing a combination of internal convective cooling and external film-cooling. This study will evaluate several novel cooling schemes for turbine airfoils, called microcircuits. Microcircuits are placed inside the turbine blade wall, and the features turbulate the air and increase heat transfer surface area, thereby augmenting convective cooling. The coolant flow then exits internal cooling passages to the external side of the blade. Here the coolant forms a protective layer along the external surface of the blade to protect the blade from the heated mainstream flow. In the current study, a low-speed large-scale wind tunnel facility was developed to measure internal heat transfer coefficients and external adiabatic effectiveness, using thermal liquid crystallography and infrared thermography. This test facility is unique in that it can be used to test the effects of internal cooling features on external film cooling. Results show that the highest augmentations in internal heat transfer were seen at the lowest Reynolds numbers. Internal features affected the shapes of external film-cooling contours, but the magnitudes of the spanwise averaged values did not change significantly with changes in internal geometry. / Master of Science

In a gas turbine engine, turbine blades are exposed to temperatures above their melting point. Film-cooling and internal cooling techniques can prolong blade life and allow for higher engine temperatures. This study examines a novel cooling technique called a microcircuit, which combines internal convection and pressure side injection on a turbine blade tip. Holes on the tip called dirt purge holes expel dirt from the blade, so other holes are not clogged. Wind tunnel tests are used to observe how effectively dirt purge and microcircuit designs cool the tip. Tip gap size and blowing ratio are varied for different tip cooling configurations. Results show that the dirt purge holes provide significant film cooling on the leading edge with a small tip gap. Coolant injected from these holes impacts the shroud and floods the tip gap reducing tip leakage flow. With the addition of a microcircuit, coolant is delivered to a larger area of the tip. In all cases, cooling levels are higher for a small tip gap than a large tip gap. Increased blowing ratio does not have a dramatic effect on microcircuit film-cooling at the midchord but does improve internal cooling from the microcircuit. While the combined dirt purge holes and microcircuit cool the leading edge and midchord areas, there remains a small portion of the trailing edge that is not cooled. Also, results suggest that blowing from the microcircuit diminishes the tip leakage vortex. Overall, the microcircuit appears to be a feasible method for prolonging blade life. / Master of Science

Conventional understanding of quality of energy suggests that heat is a low grade form of energy. Hence converting this energy into useful form of work was assumed difficult. However, this understanding was challenged by researchers over the last few decades. With advances in solar, thermal and geothermal energy harvesting, they believed that these sources of energy had great potential to operate as dependable avenues for electrical power. In recent times, waste heat from automobiles, oil and gas and manufacturing industries were employed to harness power. Statistics show that US alone has a potential of generating 120,000 GWh/year of electricity from oil , gas and manufacturing industries, while automobiles can contribute upto 15,900 GWh/year. Thermoelectric generators (TEGs) can be employed to capture some of this otherwise wasted heat and to convert this heat into useful electrical energy. This field of research as compared to gas turbine industry has emerged recently over past 30 decades. Researchers have shown that efficiency of these TEGs modules can be improved by integrating heat transfer augmentation features on the hot side of these modules. Gas turbines employ advanced technologies for internal and external cooling. These technologies have applications over wide range of applications, one of which is thermoelectricity. Hence, making use of gas turbine technologies in thermoelectrics would surely improve the efficiency of existing TEGs. This study makes an effort to develop innovative technologies for gas turbine as well as thermoelectric applications. The first part of the study analyzes heat transfer augmentation from four different configurations for low aspect ratio channels and the second part deal with characterizing improvement in efficiency of TEGs due to the heat transfer augmentation techniques. / Master of Science