The research focuses on testing the film cooling effectiveness on a gas turbine blade suction side surface. The test is performed on a five bladed cascade with a blow down facility. Four different blowing ratios are used in this study, which are 0.5, 1.0, 1.6, and 2.0; mainstream flow conditions are maintained at exit Mach number of 0.7, 1.1 and 1.3. Nitrogen is injected as the coolant so that the oxygen concentration levels can be obtained for the test surface. Based on mass transfer analogy, film cooling effectiveness can be computed with pressure sensitive paint (PSP) technique. The effect of blowing ratio on film cooling effectiveness is presented for each testing condition. The spanwise averaged effectiveness for each case is also presented to compare the blowing ratio and mainstream effect on film cooling effectiveness. Results show that due to effects of shock, the optimum blowing ratio is 1.6 for exit Mach number of 1.1 and 1.3; however; without the effects of shock, the optimum blowing ratio is 1.0 for exit Mach number of 0.7.

High turbine inlet temperature becomes necessary for increasing thermal efficiency of modern gas turbines. To prevent failure of turbine components, advance cooling technologies have been applied to different portions of turbine blades. The detailed film cooling effectiveness distributions along a rotor blade has been studied under combined effects of upstream trailing edge unsteady wake with coolant ejection by the pressure sensitive paint (PSP). The experiment is conducted in a low speed wind tunnel with a five blade linear cascade and exit Reynolds number is 370,000. The density ratios for both blade and trailing edge coolant ejection range from 1.5 to 2.0. Blade blowing ratios are 0.5 and 1.0 on suction surface and 1.0 and 2.0 on pressure surface. Trailing edge jet blowing ratio and Strouhal number are 1.0 and 0.12, respectively. Results show the unsteady wake reduces overall effectiveness. However, the unsteady wake with trailing edge coolant ejection enhances overall effectiveness. Results also show that the overall effectiveness increases by using heavier coolant for ejection and blade film cooling. Leading edge film cooling has been investigated using PSP. There are two test models: seven and three-row of film holes for simulating vane and blade, respectively. Four film holes’ configurations are used for both models: radial angle cylindrical holes, compound angle cylindrical holes, radial angle shaped holes, and compound angle shaped holes. Density ratios are 1.0 to 2.0 while blowing ratios are 0.5 to 1.5. Experiments were conducted in a low speed wind tunnel with Reynolds number 100,900. The turbulence intensity near test model is about 7%. The results show the shaped holes have overall higher effectiveness than cylindrical holes for both designs. As increasing density ratio, density effect on shaped holes becomes evident. Radial angle holes perform better than compound angle holes as increasing blowing and density ratios. Increasing density ratio generally increases overall effectiveness for all configurations and blowing ratios. One exception occurs for compound angle and radial angle shaped hole of three-row design at lower blowing ratio. Effectiveness along stagnation row reduces as increasing density ratio due to coolant jet with insufficient momentum caused by heavier density coolant, shaped hole, and stagnation row.

No description available.

Cooling combustor chambers for gas turbine engines is challenging because of the complex flow fields inherent to this engine component. This complexity, in part, arises from the interaction of high momentum dilution jets required to mix the fuel with effusion film cooling jets that are intended to cool the combustor walls. The dilution and film cooling flow have different performance criteria, often resulting in conflicting flow mechanisms. The purpose of this study is to evaluate the influence that the dilution jets have on the film cooling effectiveness and how the flow and thermal patterns in the cooling layer are affected by both the dilution flow and the closely spaced film cooling holes. This study also intends to characterize the development of the flow field created by effusion cooling injection without dilution injection. This work is unique because it allows insight into how the full-coverage discrete film cooling layer is interrupted by high momentum dilution jets and how the surface cooling is affected. The film cooling flow was disrupted along the combustor walls in the vicinity of the high momentum dilution jets and the surface cooling effectiveness was reduced with increased dilution jet momentum. This was due to the secondary flows that were intensified by the increased jet momentum. High turbulence levels were generated at the dilution jet shear layer resulting in efficient mixing. The film cooling flow field was affected by the freestream turbulence and complex flow fields created by the combined dilution and effusion cooling flows both in the near dilution jet region as well as downstream of the jets. Effusion cooling holes inclined at 20Ë created lower coolant layer turbulence levels and higher surface cooling effectiveness than 30Ë cooling holes. Results showed an insensitivity of the coolant penetration height to the diameter and angle of the cooling hole in the region downstream of the dilution mixing jets. When high momentum dilution jets were injected into crossflow, a localized region in the flow of high vorticity and high streamwise velocity was created. When film cooling air was injected the inlet flow field and the dilution jet wake were fundamentally changed and the vortex diminished significantly. The temperature field downstream of the dilution jet showed evidence of a hot region which was moderated appreciably by film cooling flow. Differences in the temperature fields were nominal compared to the large mass flow increase of the coolant. A study of streamwise oriented effusion film cooling flow without dilution injection revealed unique and scaleable velocity profiles created by the closely spaced effusion holes. The effusion cooling considered in these tests resulted in streamwise velocity and turbulence level profiles that scaled well with blowing ratio which is a finding that allows the profile shape and magnitude to be readily determined at these test conditions. Results from a study of compound angle effusion cooling injection showed significant differences between the flow field created with and without crossflow. It was found from the angle of the flow field velocity vectors that the cooling film layer grew nearly linearly in the streamwise direction. The absence of crossflow resulted in higher turbulence levels because there was a larger shear stress due to a larger velocity difference between the coolant and crossflow. The penetration height of the coolant was relatively independent of the film cooling momentum flux ratio for both streamwise oriented and compound angle cooling jets. / Ph. D.

Computations and experiments were run to study heat transfer and overall effectiveness for a simulated turbine blade leading edge. Computational predictions were run for a film cooled leading edge model using a conjugate numerical method to predict the normalized “metal” temperatures for the model. This computational study was done in conjunction with a parallel effort to experimentally determine normalized metal temperatures, i.e. overall effectiveness, using a specially designed high conductivity model. Predictions of overall effectiveness were higher than experimentally measured values in the stagnation region, but lower along the downstream section of the leading edge. Reasons for the differences between computational predictions and experimental measurements were examined. Also of interest was the validity of Taw as the driving temperature for heat transfer into the blade, and this was examined via computations. Overall, this assumption gave reasonable results except near the stagnation line. Experiments were also conducted on a leading edge with no film cooling to gain a better understanding of the additional cooling provided by film cooling. Heat flux was also measured and external and internal heat transfer coefficients were determined. The results showed roughly constant overall effectiveness on the external surface. / text

No description available.

This study focused on the improvement of film cooling for gas turbine vanes using both computational and experimental techniques. The experimental component used a matched Biot number model to measure scaled surface temperature (overall effectiveness) distributions representative of engine conditions for two new configurations. One configuration consisted of a single row of holes on the pressure surface while the other used numerous film cooling holes over the entire vane including a showerhead. Both configurations used internal impingement cooling representative of a 1st vane. Adiabatic effectiveness was also measured. No previous studies had shown the effect of injection on the mean and fluctuating velocity profiles for the suction surface, so measurements were made at two locations immediately upstream of film cooling holes from the fully cooled cooling configuration. Different blowing conditions were evaluated. Computational tools are increasingly important in the design of advanced gas turbine engines and validation of these tools is required prior to integration into the design process. Two film cooling configurations were simulated and compared to past experimental work. Data from matched Biot number experiments was used to validate the overall effectiveness from conjugate simulations in addition to adiabatic effectiveness. A simulation of a single row of cooling holes on the suction side also gave additional insight into the interaction of film cooling jets with the thermal boundary layer. A showerhead configuration was also simulated. The final portion of this study sought to evaluate the performance of six RANS models (standard, realizable, and renormalization group k-ε; standard k-ω; k-ω SST; and Transition SST) with respect to the prediction of thermal boundary layers. The turbulent Prandtl number was varied to test a simple method for improvement of the thermal boundary layer predictions. / text

Film cooling computations and experiments were performed to study heat transfer and adiabatic effectiveness for several geometries. Various assumptions commonly made in film cooling experiments were computationally simulated to test the validity of using these assumptions to predict the heat flux into conducting walls. The validity of these assumptions was examined via computational simulations of film cooling on adiabatic, heated, and conducting flat plates using the commercial code FLUENT. The assumptions were found to be reasonable overall, but certain regions in the domain suffered from poor predictions. Film cooling adiabatic effectiveness and heat transfer coefficients for axial holes embedded in a 1 [hole diameter] transverse trench on the suction side of a simulated turbine vane were experimentally investigated as well to determine the net heat flux reduction. Heat transfer coefficients were determined with and without upstream heating both with and without a tripped boundary layer approach flow. The net heat flux reduction for the trench was found to be much higher than for the baseline row of holes. Two transverse trench geometries and a baseline row of holes geometry were also simulated using FLUENT and the results were compared to experiments by Waye and Bogard (2006). Trends between simulated trench configurations and baseline cylindrical holes without a trench were found to be largely in agreement with experimental trends, suggesting that FLUENT can be used as a tool for studying new trench configurations. / text

Thesis (Ph.D.)--Cleveland State University, 2010. / Abstract. Title from PDF t.p. (viewed on July 6, 2010). Includes bibliographical references (p. 154-162). Available online via the OhioLINK ETD Center and also available in print.

This research has a small portion focused on interior serpentine channels, with the primary focus on improving the effectiveness of the film cooling technique through the use of a new approach to film cooling. This new approach uses a set of three holes sharing the same inlet and diverging from the central hole to form a three-legged, or tripod, design. The tripod design is examined in depth, in terms of geometric variations, through the use of flat plate and cascade rigs, with both transient and steady-state experiments. The flat plate tests provide a simplified setting in which to test the design in comparison to other geometries, and establish a baseline performance in a simple flow field that does not have the complications of surface curvature or mainstream pressure gradients. Cascade tests allow for testing of the design in a more realistic setting with curved surfaces and mainstream pressure gradients, providing important information about the performance of the design on suction and pressure surfaces of airfoils. Additionally, the cascade tests allow for an investigation into the aerodynamic penalties associated with the injection hole designs at various flow rates. Through this procedure the current state of film cooling technology may be improved, with more effective surface coverage achieved with reduced coolant usage, and with reduced performance penalties for the engine as a whole. This research has developed a new film hole design that is manufacturable and durable, and provides a detailed analysis of its performance under a variety of flow conditions. This cooling hole design provides 40% higher cooling effectiveness while using 50% less coolant mass flow. The interior serpentine channel research provides comparisons between correlations and experiments for internal passages with realistic cross sections. / Ph. D.