Sri Sreekanth

Influence of Surface Roughness on the Aerodynamic Losses of a Turbine Vane

The effects of surface roughness on the aerodynamic performance of a turbine vane are investigated for three Mach number distributions, one of which results in transonic flow. Four turbine vanes, each with the same shape and exterior dimensions, are employed with different rough surfaces. The nonuniform, irregular, three-dimensional roughness on the tested vanes is employed to match the roughness which exists on operating turbine vanes subject to extended operating times with significant particulate deposition on the surfaces. Wake profiles are measured for two different positions downstream the vane trailing edge. The contributions of varying surface roughness to aerodynamic losses, Mach number profiles, normalized kinetic energy profiles, Integrated Aerodynamics Losses (IAL), area-averaged loss coefficients, and mass-averaged loss coefficients are quantified. Total pressure losses, Mach number deficits, and deficits of kinetic energy all increase at each profile location within the wake as the size of equivalent sandgrain roughness increases, provided the roughness on the surfaces is uniform. Corresponding Integrated Aerodynamic Loss IAL magnitudes increase either as Mach numbers along the airfoil are higher, or as the size of surface roughness increases. Data are also provided which illustrate the larger loss magnitudes which are present with flow turning and cambered airfoils, than with symmetric airfoils. Also described are wake broadening, profile asymmetry, and effects of increased turbulent diffusion, variable surface roughness, and streamwise development.

Suction-Side Gill Region Film Cooling: Effects of Hole Shape and Orientation on Adiabatic Effectiveness and Heat Transfer Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on thermal film cooling characteristics, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, the exit Mach number is 0.35, and the tests are conducted using the first row of holes only, second row of holes only, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.73 to 1.92 similar to values present in operating gas turbine engines. A mesh grid is used to give a magnitude of longitudinal turbulence intensity of 5.7% at the inlet of the test section. Results show that the best overall protection over the widest range of blowing ratios and streamwise locations is provided by either the RC holes or the RR holes. This result is particularly significant because the RR hole arrangement, which has lower manufacturing costs compared with the RC and SA arrangements, produces better or equivalent levels of performance in terms of magnitudes of adiabatic film cooling effectiveness and heat transfer coefficient. Such improved performance (relative to RA and SA holes) is most likely a result of compound angles, which increases lateral spreading. As such, the present results indicate that compound angles appear to be more effective than hole shaping in improving thermal protection relative to that given by RA holes.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77-1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient C p magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak C p magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

Effect of Input Variability on the Performance of Turbine Blade Thermal Design Using Monte Carlo Simulation: An Exploratory Study

A probabilistic approach to the thermal design and analysis of cooled turbine blades is presented. Various factors that affect the probabilistic performance of the blade thermal design are grouped into categories and a select number of factors known to be significant, for which the variability could be assessed are modeled as random variables. The variability data for these random variables were generated from separate Monte Carlo simulations (MCS) of the combustor and the upstream stator and secondary air system. The oxidation life of the blade is used as a measure to evaluate the thermal design as well as to evaluate validity of the methods. Two approaches have been explored to simulate blade row life variability and compare it with the field data. Field data from several engine removals are used for investigating the approach. Additionally a response surface approximation technique has been explored to expedite the simulation process. The results indicate that the conventional approach of a worst-case analysis is overly conservative and analysis based on nominal values could be very optimistic. The potential of a probabilistic approach in predicting the actual variability of the blade row life is clearly evident in the results. However, the results show that, in order to predict the blade row life variability adequately, it is important to model the operating condition variability. The probabilistic techniques such as MCS could become very practical when approximation techniques such as response surface modeling are used to represent the analytical model.

Suction-Side Gill-Region Film Cooling: Effects of Hole Shape and Orientation on Adiabatic Effectiveness and Heat Transfer Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah Transonic Wind Tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on thermal film cooling characteristics are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes only, second row of holes only, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.73 to 1.92 similar to values present in operating gas turbine engines. A mesh grid is used to give a magnitude of longitudinal turbulence intensity of 5.7 percent at the inlet of the test section. Results show that the best overall protection over the widest range of blowing ratios and streamwise locations is provided by either the RC holes, or the RR holes. This result is particularly significant because the RR hole arrangement, which has lower manufacturing costs compared with the RC and SA arrangements, produces better or equivalent levels of performance in terms of magnitudes of adiabatic film cooling effectiveness and heat transfer coefficient. Such improved performance (relative to RA and SA holes) is likely mostly a result of compound angles, which increases lateral spreading. As such, the present results indicate that compound angle appears to be more effective than hole shaping in improving thermal protection relative to that given by RA holes.Copyright

Blade Air Cooling Feed System CFD Analysis and Validation

A cold flow static rig test and computational fluid dynamics (CFD) analyses have been performed to verify the benefits of a modified “deflector” design to the blade air cooling feed system. The area of interest is the broach passage which is a space between the bottom of the blade fixing and the blade disk. Cooling flow enters into this region from the disk cavity and then exits upward into the blade internal cooling passages. Steady Navier-Stokes analyses of the cold flow static rig geometries were performed using the in-house code NS3D and the commercial code CFX v5.6. Due to the setup of the experimental rig, the numerical domain model with appropriate boundary conditions had to be carefully selected. Unstructured grids including wall prism layers combined with mesh adaptation were employed for both CFD codes on the baseline geometry. The commercial adaptation code OPTIMESH was used for NS3D while CFX employed its internal grid adaptation tool. No grid adaptation was necessary for the modified geometry. The pressure and mass flow measurements for the baseline and modified blade broach geometries demonstrated a reduction in losses for the new design. The CFD flow visualization showed the presence of a strong vortex in the broach for the baseline case with an accompanying low pressure zone. The modified broach design deflects the flow thus avoiding the formation of the strong vortex and low pressure zone. The CFD pressure distributions predicted the trends qualitatively. Quantitative comparisons of the CFD pressure distributions along broach walls were reasonable for the new design. For the baseline geometry, the CFD pressure drop is higher than measurements indicate. Modeling issues associated with a screen in the rig set-up is thought to be the source of the difference.Copyright