Abstract
Thermal barrier coatings (TBC) found on turbine blades are a key element in the performance and reliability of modern gas turbines. TBC reduces the heat transfer into turbine blades by introducing an additional surface thermal resistance; consequently allowing for higher gas temperatures. During the service life of the blades, the TBC surface may be damaged due to manufacturing imperfections, handling damage, service spalling, or service impact damage, producing chips in the coating. While an increase in aerofoil temperature is expected, it is unknown to what degree the blade will be affected and what parameters of the chip shape affect this result. During routine inspections, the severity of the chipping will often fall to the discretion of the inspecting engineer. Without a quantitative understanding of the flow and heat transfer around these chips, there is potential for premature removal or possible blade failure if left to operate. The goal of this preliminary study is to identify the major driving parameters that lead to the increase in metal temperature when TBC is damaged, such that more quantitative estimates of blade life and refurbishing needs can be made.
A two-dimensional computational Conjugate Heat Transfer model was developed; fully resolving the hot gas path and TBC, bond-coat, and super alloy solids. Representative convective conditions were applied to the cold side to emulate the characteristics of a cooled turbine blade. The hot gas path properties included an inlet temperature of 1600 K with varying Mach numbers of 0.30, 0.59, and 0.80 and Reynolds number of 5.1×105, 7.0×105, and 9.0×105 as referenced from the leading edge of the model. The cold side was given a coolant temperature of 750 K and a heat transfer coefficient of 1500 W/m2*K. The assigned thermal conductivities of the TBC, bond-coat, and metal alloys were 0.7 W/m*K, 7.0 W/m*K, and 11.0 W/m*K, respectively, and layer thicknesses of 0.50 mm, 0.25 mm, and 1.50 mm, respectively. A flat plate model without the presence of the chip was first evaluated to provide a basis of validation by comparison to existing correlations. Comparing heat transfer coefficients, the flat plate model matched within uncertainty to the Chilton-Colburn analogy. In addition, flat plate results captured the boundary layer thickness when compared with Prandtl’s 1/7th power-law. A chip was then introduced into the model, varying the chip width and the edge geometry. The most sensitive driving parameters were identified to be the chip width and Mach number. In cases where the chip width reached 16 times the TBC thickness, temperatures increased by almost 30% when compared to the undamaged equivalents. Additionally, increasing the Mach number of the incoming flow also increased metal temperatures. While the Reynolds number based on the leading edge of the model was deemed negligible, the Reynolds number based on the chip width was found to have a noticeable impact on the blade temperature. In conclusion, this study found that chip edge geometry was a negligible factor, while the Mach number, chip width, and Reynolds number based on the chip width had a significant effect on the total metal temperature.
Geometric Effects of Thermal Barrier Coating Damage on Turbine Blade Temperatures
