Thesis (Ph. D.)--University of Washington, 1999. / Vita. Includes bibliographical references (leaves 156-161).

Thesis (Ph. D.)--University of Washington, 2004. / Vita. Includes bibliographical references (leaves 185-189).

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.

A new pressure-measurement technique which employs the tools of molecular spectroscopy has recently received considerable attention in the fluid mechanics community. Measurements are made via oxygen-sensitive molecules attached to the surface of interest as a coating, or paint. The pressure-sensitive-paint (PSP) technique is now commonly used in stationary wind-tunnel tests; this thesis presents the extension of the technique to advanced turbomachinery applications. New pressure- and temperature-sensitive paints (TSPs) have been developed for application to a state-of-the-art transonic compressor where pressures up to 2 atm and surface temperatures up to 140° C are expected for the first-stage rotor. PSP and TSP data has been acquired from the suction surface of the first-stage rotor of a transonic compressor operating at its peak-efficiency condition. The shock structure is clearly visible in the pressure image, and visual comparison to the corresponding computational fluid dynamics (CFD) prediction shows qualitative agreement to the PSP data. / Master of Science

No description available.

The motion of shock waves is important in many fields of engineering and increasingly so with medical applications and applications to inertial confinement fusion technologies. The flow structures that moving shock waves create when they encounter a change in area is complex and can be difficult to understand. Previousresearchers have carried out experimental studies and many numerical studies looking at this problem in more detail. There has been a discrepancy between numerical and experimental work which had remained unanswered. One of the aims of this project is to try and resolve the discrepancy between numerical and experimental work and try to investigate what experimental techniques are suitable for work of this type and the exact way in which they should be applied. Most previous work has focused on sharp changes in geometry which induce immediate flow separation. In this project rounded corners will also be investigated and the complex flow features will be analyzed.Two geometries, namely a sharp 172 degree knife-edge and a 2.8 mm radius rounded corner will be investigated at three experimental pressure ratios of 4, 8 and 12 using air as the driver gas. This yields experimental shock Mach numbers of 1.28, 1.46 and 1.55. High-speed schlieren and shadowgraph photography with varying levels of sensitivity were used to qualitatively investigate the wave structures. Particle image velocimetry (PIV), pressure-sensitive paint (PSP) and traditional pressure transducers were used to quantify the flow field. Numerical simulations were performed using the commercial package Fluent to investigate the effect of numerical schemes on the flow field produced and for comparison with the experimental results. The sharp geometry was simulated successfully using an inviscid simulation while the rounded geometry required the addition of laminar viscosity. Reynolds number effects will be only sparsely referred to in this project as the flows under investigation show largely inviscid characteristics. As the flow is developing in time rather than in space, quotation of a distance-based Reynolds number is not entirely appropriate; however, Reynolds number based on the same spatial location but varying in time will be mentioned. The density-based diagnostics in this project were designed to have a depth of field appropriate to the test under consideration. This approach has been used relatively few times despite its easy setup and significant impact on the results. This project contains the first quantative use of PIV and PSP to shock wave diffraction. Previous studies have almost exclusively used density-based diagnostics which, although give the best impression of the flow field, do not allow for complete analysis and explanation of all of the flow features present. PIV measurements showed a maximum uncertainty of 5% while the PSP measurements showed an uncertainty of approximately 10%.The shock wave diffraction process, vortex formation, shear layer structure, secondary and even tertiary expansions and the shock vortex interaction were investigate. The experimental results have shown that using one experimental technique in isolation can give misleading results. Only by using a combination of experimental techniques can we achieve a complete understanding of the flow field and draw conclusions on the validity of the numerical results. Expanding the range of the experimental techniques currently in use is vital for experimental aerodynamic testing to remain relevant in an industry increasingly dominated by numerical research. To this end, significant research work has been carried out on extending the range of the PSP technique to allow for the capture of shock wave diffraction, one of the fastest transient fluid processes, and for applications to low-speed flow (< 20 ms−1).

Measurement and mapping of the pressure distribution across the surface of a suitably scaled model is an integral step in the design of any aircraft or automobile. For this purpose, the traditional workhorses of the aeronautic and automotive industries have been pressure taps--small orifices that contain electronic pressure transducers. Unfortunately, in addition to the limited spatial resolution achievable with such devices, their technical complexity and cost constitute serious disadvantages. For more than 35 years, researchers have pursued a fundamentally different alternative: indirect measurement of pressure via oxygen-induced quenching of the luminescence emitted by certain chemical species. Porphyrin complexes of dipositive palladium and especially platinum have emerged as one of the principal classes of oxygen-sensitive luminophores; ruthenium(II) polypyridyl complexes comprise another. Various other metals also form luminescent coordination complexes that are susceptible to quenching by O2, however, and these too have contributed to the diversity of luminophores that are now available for incorporation into pressure-sensitive paints and related films and coatings. After treating the photophysics of luminescence quenching by molecular oxygen and quantitative descriptions of this phenomenon in the ideal case and in heterogeneous media, the thesis presents a comprehensive survey of the chemical literature on oxygen-sensitive luminophores. Efforts to prepare and characterize a novel porphyrin-pillared mixed zirconium phosphonate are then detailed. Following complexation of Pt(II) ions by the porphyrin moieties, this material is expected to display oxygen-sensitive luminescence and should ameliorate such difficulties as luminophore aggregation and matrix photodegradation that are associated with many existing pressure-responsive coatings. Its preparation necessitated preliminary formation of a porphyrin functionalized with two phenylphosphonic acid groups, which was obtained by synthesizing dipyrromethane and diethyl 4-formylphenylphosphonate and condensing these two precursors. The mixed phosphonate, a layered material assembled from ZrOCl2 · 8H2O, methylphosphonic acid, and the aforementioned porphyrin, was then prepared in refluxing HF. Solid-state 31P NMR spectra and powder X-ray diffraction patterns were acquired for the final product, its estimated interlayer spacing of 22.8 Å figuring prominently in analysis and discussion of the X-ray data.

No description available.

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 three parts: 1) turbine blade span film-cooling, 2) turbine platform film-cooling, and 3) blade tip film-cooling. Pressure sensitive paint (PSP) technique was used to get the conduction-free film-cooling effectiveness distribution. Film-cooling effectiveness is assessed in terms of cooling hole geometry, blowing ratio, freestream turbulence, and coolant-to-mainstream density ratio. Blade span film-cooling test shows that the compound angle shaped holes offer better film effectiveness than the axial shaped holes. Greater coolant-to-mainstream density ratio prevents coolant to lift-off. Higher freestream turbulence causes effectiveness to drop everywhere except in the region downstream of suction side. Results are also correlated with momentum flux, compound shaped hole has the greatest optimum momentum flux ratio, and then followed by axial shaped hole, compound cylindrical hole, and axial cylindrical hole. For platform purge flow cooling, the stator-rotor gap was simulated by a typical labyrinth-like seal. Two different film-cooling hole geometries, three blowing ratios and density ratios, and two freestream turbulence are examined. Results showed that the shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. Moreover, the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity. The blade tip study was performed in a blow-down flow loop. Results show that a blowing ratio of 2.0 is found to give best results on the tip floor. Lift-off of the coolant jet can be observed for the holes closer to the leading edge as blowing ratio increases from 1.5 to 2.0. A stator vane suction side heat transfer study was conducted in a partial annular cascade. The heat transfer coefficients were measured by using the transient liquid crystal technique. At X/L=0.15, a low heat transfer region where transition occurs. The heat transfer coefficients increase toward the trailing edge as flow accelerates; a spanwise variation can be found at neat tip and hub portions due to passage and horseshoe vortices.