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The topic of this thesis is the construction of a formula to approximate stress-strain responses at notches under thermomechanical fatigue (TMF) loading. The understanding of material behavior of the V-notched component which experiences TMF is important to the mechanical industries where V-notched structures are often utilized. In such applications, it is crucial that the designers be able to predict the material behavior; therefore, the purpose of this research is to examine and to model the precise effects a stress concentration will have on a specimen made of a generic Ni-base superalloy. The effects of non-isothermal loading will be studied, and it is the goal of this research to formulate an extension of Neuber's rule appropriate for TMF which is to approximate the temperature range with a single value, T'. One strategy to extend Neuber's rule, which relies on Finite Element Modeling (FEM), Bilinear Kinetic Hardening Model (BKIN), and test data, will be used to predict the stress-strain behavior at the notch of a thin plate subjected to axial loading. In addition, the CHABOCHE model will be utilized in the FEA to have the highest fidelity to material response at high temperatures. Parametric study of the FEA simulations will be employed to determine the correlation between the Neuber hyperbola, temperature range, stress concentration, the nominal stress, and the temperature cycling. Using the Neuber hyperbola and simplified constitutive model (i.e., bilinear kinematic strain hardening), the stress-strain solutions of the specimen will be calculated and compared to analytical results.

Air velocity and fuel concentration data have been collected throughout the flow fields of two gas turbine mixers in an effort to better understand the mixing of fuel and air in gas turbine mixers. The two gas turbine mixers consisted of an annular flow profile and incorporated swirl vanes to produce a swirling flow to promote fuel/air mixing. The fuel was injected into the bulk flow from the pressure side of the swirl vanes. The first mixer had a swirl angle of 45°, while the second had a swirl angle of 55°. In order to examine the effect of the swirl angle on the mixing of fuel and air as the flow progressed through gas turbine mixers, axial and tangential air velocity data was taken using a laser Doppler velocimeter (LDV). Also, fuel concentration data was taken separately using a hydrocarbon concentration probe with methane diluted with air as the fuel. The data were taken at varying axial and varying angular locations in an effort to capture the spatial development of the fuel and velocity profiles. The spectra of the data were analyzed as well in an effort to understand the turbulence of the flow. It was found that the 55° swirler exhibited smaller variations in both velocity and fuel concentration values and that the fuel reached a uniform concentration at axial locations further upstream in the 55° degree mixer than in the 45° mixer. The RMS values of the velocity, which were influenced by the swirl vanes, were higher in the 55° mixer and likely contributed to the better mixing performance of the 55° mixer. The fuel concentration spectrum data showed that the spectra of the two mixers were similar, and that the fluctuations in fuel concentration due to flow emanating from the swirl vanes were seen throughout the length of the two mixers.

Luminescent coating measurement system is a relatively new technology for quantitative pressure and temperature measurement. Usually referred to as Pressure Sensitive Paint (PSP) and Temperature Sensitive Paint (TSP), luminescent coatings contain sensor molecules, which undergoes a luminescent transition when excited with light of proper wavelength. The reaction is pressure and/or temperature sensitive. The image of TSP or PSP coated model surface can be captured with a scientific grade camera and then processed to obtain full field temperature and pressure distribution with very high fidelity. The preparation time of the technique is short. The measurement system offers an economic alternative to conventional testing methods using large number of pressure taps and thermocouples. The purpose of the experiment in this thesis is to take the benefits of the TSP and PSP technique, develop a well-controlled process and then apply the technique for a fundamental study on jet impingement heat transfer. First, Uni-Coat TSP and Binary-FIB PSP purchased from ISSI Inc. are calibrated to high accuracy. The calibration uncertainty of TSP and PSP are found to be ±0.93 °C and ±0.12 psi over temperature and pressure ranges of 22 to 90 ° C and 5 to 14.7 psia, respectively. The photodegradation of TSP is then investigated with the same calibration system. The photodegradation refers to the phenomenon of decreasing emission intensity as the luminescent paint is exposed to the illumination light during testing. It was found that photodegradation rate is a strong function of temperature and the optical power of illumination lighting. The correlation developed in this work is expected to compensate the degradation of TSP to achieve high measurement accuracy. Both TSP and PSP were then applied in the flow and heat transfer measurement of single round impinging air jet. Various separation distance (Z/D) and jet Reynolds number are tested. Pressure measurement on the jet impinged target surface using PSP clearly shows the boundary of jet impingement zone, which broadens with separation distance. In heat transfer experiment using TSP, the "second peak" in local heat transfer occurring at radial distance r/D around 2 is clearly observed when the separation distance Z/D is shorter than the length of jet potential core. The slight variation in radial location and the amplitude of the "second peak" are captured as Z/D and jet Reynolds number change. The optimum Z/D of stagnation point heat transfer is found to be around 5. The effect of jet nozzle configuration is investigated. It is found that the heat transfer rate associated with "tube jet" is generally higher than that of "plate jet". The difference in heat transfer between the two jet configurations is related to the weaker entrainment effect associated with "plate jet", where the entrainment of surrounding air is confined by the injection plate, especially under small Z/D circumstances. When compared with the benchmark data in the literature, the averaged heat transfer data of "tube jet" matches the empirical data better than those of "plate jet". The maximum difference is 3.3% for tube jet versus 15.4% for plate jet at Reynolds number of 60000 and Z/D of 5. The effect of surface roughness on jet impingement heat transfer is also studied. Heat transfer can be significantly increased by the enhanced roughness of the target surface. The largest roughness effect is achieved near stagnation point at high jet Reynolds number. Compared to the heat transfer to a smooth plate, as high as 30.9% increase in area-averaged Nusselt number is observed over a rough surface at r/D=1.5 and jet Reynolds number of 60000. The most significant advance of the present work is that both temperature and pressure measurement be obtained with the same measurement system and with accuracy comparable to traditional testing methods. The procedures that were employed in this work should be easy to apply in any university or industrial testing facility. It provides a rapid testing tool that can help solve complex problems in aerodynamics and heat transfer

This thesis is an experimental and numerical full-coverage film cooling study. The objective of this work is the quantification of local heat transfer augmentation and adiabatic film cooling effectiveness for two full-coverage film cooling geometries. Experimental data was acquired with a scientific grade CCD camera, where images are taken over the heat transfer surface, which is painted with a temperature sensitive paint. The CFD component of this study served to evaluate how well the v2-f turbulence model predicted film cooling effectiveness throughout the array, as compared with experimental data. The two staggered arrays tested are different from one another through a compound angle shift after 12 rows of holes. The compound angle shifts from ?=-45° to ?=+45° at row 13. Each geometry had 22 rows of cylindrical film cooling holes with identical axial and lateral spacing (X/D=P/D=23). Levels of laterally averaged film cooling effectiveness for the superior geometry approach 0.20, where the compound angle shift causes a decrease in film cooling effectiveness. Levels of heat transfer augmentation maintain values of nearly h/h0=1.2. There is no effect of compound angle shift on heat transfer augmentation observed. The CFD results are used to investigate the detrimental effect of the compound angle shift, while the SST k-? turbulence model shows to provide the best agreement with experimental results.

As modern gas turbines implement more and more complex geometry to increase life and efficiency, attention to unsteady aerodynamic behavior becomes more important. Computational optimization schemes are contributing to advanced geometries in order to reduce aerodynamic losses and increase the life of components. These advanced geometries are less representative of cylinder and backward facing steps which have been used as analogous geometries for most aerodynamic unsteadiness research. One region which contains a high degree of flow unsteadiness and a direct influence on engine performance is that of the MidFrame. The MidFrame (or combustor-diffuser system) is the region encompassing the main gas path from the exit of the compressor to the inlet of the first stage turbine. This region contains myriad flow scenarios including diffusion, bluff bodies, direct impingement, high degree of streamline curvature, separated flow, and recirculation. This represents the most complex and diverse flow field in the entire engine. The role of the MidFrame is to redirect the flow from the compressor into the combustion system with minimal pressure loss while supplying high pressure air to the secondary air system. Various casing geometries, compressor exit diffuser shapes, and flow conditioning equipment have been tested to reduce pressure loss and increase uniformity entering the combustors. Much of the current research in this area focuses on aero propulsion geometries with annular combustors or scaled models of the power generation geometries. Due to the complexity and size of the domain accessibility with physical probe measurements becomes challenging. The current work uses additional measurement techniques to measure flow unsteadiness in the domain. The methodology for identifying and quantifying the sources of unsteadiness are iv developed herein. Sensitivity of MidFrame unsteadiness to compressor exit conditions is shown for three different velocity profiles. The result is an extensive database of measurements which can serve as a benchmark for radical new designs to ensure that the unsteadiness levels do not supersede previous successful levels.

Hydrogen, as an alternative to conventional aviation fuels, has the potential to increase the efficiency of a gas turbine as well as reduce emissions of greenhouse gases. In addition to significantly reducing the number of pollutants due to the absence of carbon, burning hydrogen at low equivalence ratios can significantly reduce emissions of oxides of nitrogen (NOx). Because hydrogen has a wide range of flammability limits, fuel lean combustion can take place at lower equivalence ratios than those with typical hydrocarbon fuels. Numerous efforts have been made to develop gas turbine fuel injectors that premix methane/natural gas and air in fuel lean proportions prior to the reaction zone. Application of this technique to hydrogen combustion has been limited due to hydrogen's high flame rate and the concern of the reaction zone propagating into the premixing injector, commonly referred to as flashback. In this investigation, a lean-premixing hydrogen injector has been developed for application in small gas turbines. The performance of this injector was characterized and predictions about the injector's performance operating under combustor inlet conditions of a PT6-20 Turboprop have been made. / Master of Science

Few aircraft engine manufacturers are able to consistently achieve high levels of performance reliability in newly manufactured engines. Much of the variation in performance reliability is due to the combined effect of tolerances of key engine components, including tip clearances of rotating components and flow areas in turbine nozzles. This research presents system analysis methods for determining the maximum possible tolerances of these key components that will allow a turbine engine to pass a number of specified performance constraints at a selected level of reliability. Through the combined use of a state-of-the-art engine performance code, component clearance loss models, and stochastic simulations, regions of feasible design space can be explored that allow for a pre-determined level of engine reliability. As expected, constraints such as spool speed and fuel consumption that are highly sensitive to certain component tolerances can significantly limit the feasible design space of the component in question. Discussed are methods for determining the bounds of any components feasible design space and for selecting the most economical combinations of component tolerances. Unique to this research is the method that determines the tolerances of engine components as a system while maintaining the geometric constraints of individual components. The methods presented in this work allow for any number of component tolerances to be varied or held fixed while providing solutions that satisfy all performance criteria. The algorithms presented in this research also allow for an individual specification of reliability on any number of performance parameters and geometric constraints. This work also serves as a foundation for an even larger algorithm that can include stochastic simulations and reliability prediction of an engine over its entire life cycle. By incorporating information such as time dependent performance data, known mission profiles, and the influence of maintenance into the component models, it would be possible to predict the reliability of an engine over time. Ultimately, a time-variant simulation such as this could help predict the timing and levels of maintenance that will maximize the life of an engine for a minimum cost. / Master of Science

A method for developing transient, predictive models of gas turbine engine performance using system identification techniques in conjunction with test cell data has been successfully demonstrated. Test cell data were obtained for both transient and steady-state operation from two F402-RR-406A USMC AV-8B engines at the Naval Aviation Depot (NADEP), Cherry Point, North Carolina. One engine was run to gather a single set of steady-state data consisting of 24 subsets of five seconds of data. The other engine was run to obtain two sets of transient data, recorded at three different rates of engine acceleration for each set. The acceleration rates of 3, 25, and 100 degrees of throttle per second were preset in the test cell controls. These rates correspond to the angular velocity of the fuel throttle during an acceleration. Each of these six transient test runs consisted of 25 seconds of data. Data were captured at a rate of five Hertz over the engine operating range from idle (26% Low Pressure spool speed) to full military power (105% LP spool speed) for all cases. A BASIC code developed at the NADEP required significant modification before it could be used to reduce the data. The modified code generated engine operating points consisting of mass flow rate, total pressure ratio, spool speed, and rate of acceleration for the inner fan, outer fan, and high pressure compressor. Finally, a multivariate regression technique using the SAS language was developed in cooperation with the Virginia Tech Statistical Consulting Center. This technique was used to generate a closed-form model of each component capable of predicting operating points at spool speeds and acceleration rates intermediate to those measured. / Master of Science

As a result of growing concerns about the carbon emissions associated with the combustion of conventional hydrocarbon fuels, hydrogen is gaining more attention as a clean alternative. The combustion of hydrogen in air produces no carbon emissions. However, hydrogen-air combustion does have the potential to produce oxides of nitrogen (NOx), which are harmful pollutants. The production of NOx can be significantly curbed using lean premixed combustion, wherein hydrogen and air are mixed at an equivalence ratio (the ratio of stoichiometric to actual air in the combustion process) significantly less than 1.0 prior to combustion. Hydrogen is a good candidate for use in lean premixed systems due to its very wide flammability range. The potential for the stable combustion of hydrogen at a wide range of equivalence ratios makes it particularly well-suited to application in gas turbines, where the equivalence ratio is likely to vary significantly over the operating range of the machine. The strong lean combustion stability of hydrogen-air flames is due primarily to high reaction rates and the associated high turbulent burning velocities. While this is advantageous at low equivalence ratios, it presents a significant danger of flashback — the upstream propagation of the flame into the premixing device — at higher equivalence ratios. An investigation has been conducted into the operation of a specific hydrogen-air premixer design in a gas turbine engine. Laboratory tests were first conducted to determine the upper stability limits of a single premixer. Tests were then carried out in which eighteen premixers and a custom-fabricated combustor liner were installed in a modified Pratt and Whitney Canada PT6A-20 turboprop engine. The tests examined the premixer and engine operability as a result of the modifications. A computer cycle analysis model was created to help analyze and predict the behavior of the modified engine and premixers. The model, which uses scaled component maps to predict off-design engine performance, was integral in the analysis of premixer flashback which limited the operation of the modified engine. / Master of Science