Two-Portion Cooling Passages for Turbine Airfoil Leading Edge
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Solution Overview
Problem
Turbine airfoils face challenges in cooling due to excessive heat loads, with traditional cooling passages experiencing reduced effectiveness at the leading edge due to the small radius of curvature, and there is a need to reduce overall coolant usage while maintaining film cooling efficiency.
Innovation Solution
The implementation of two portion cooling passages with a first portion extending from the exterior surface to a coolant chamber and a second portion intersecting the mid-portion of the first portion, featuring a diffuser outlet to enhance coolant coverage and effectiveness, which can be manufactured in separate steps and applied to the leading edge of turbine airfoils.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Ease of manufacture
If traditional round or conical shaped exit holes are used at the leading edge with radial orientation, then the cooling passages can be easily manufactured, but the cooling flow blows off the airfoil surface reducing coolant coverage and cooling effectiveness
Solution Approach 1:
The patent changes the geometric parameters of the cooling passage by using shaped diffusion exit holes instead of traditional round or conical holes. The shaped exit holes have specific angle ranges (30-60 degrees from perpendicular to surface) that optimize the cooling flow direction to prevent blow-off while maintaining manufacturability through controlled drilling angles.
Solution Approach 2:
The cooling passages are oriented at asymmetric angles relative to the airfoil surface rather than perpendicular. The exit holes are drilled at specific angles (30-60 degrees from perpendicular) to match the local surface curvature and flow direction, creating an asymmetric configuration that improves coolant attachment to the surface.
2Reliability
If shaped diffusion exit holes are used in other regions on the airfoil, then cooling effectiveness is relatively high, but they cannot be successfully used in the leading edge due to the small radius of curvature requiring drilling nearly perpendicular to the surface
Solution Approach 1:
The patent adapts the shaped diffusion exit hole geometry to the leading edge region by modifying the drilling angle parameters. Instead of drilling perpendicular to the surface, the exit holes are drilled at specific angles (30-60 degrees from perpendicular) that accommodate the small radius of curvature while maintaining the diffusion shape for effective cooling.
Solution Approach 2:
The patent applies different drilling angle parameters to different regions of the airfoil. At the leading edge with small radius of curvature, exit holes are drilled at 30-60 degrees from perpendicular, while other regions can use different angles optimized for their local geometry, allowing shaped diffusion holes to be successfully implemented throughout the airfoil.
3Loss of substance
If fewer cooling passages are used to reduce overall coolant usage, then coolant consumption decreases, but spacing between holes increases reducing film cooling effectiveness at the leading edge
Solution Approach 1:
The patent changes the parameters of individual cooling passages by implementing shaped diffusion exit holes with optimized angles and geometries. This increases the laterally-averaged cooling effectiveness of each passage, allowing fewer passages to be used while maintaining or improving overall film cooling effectiveness at the leading edge.
Solution Approach 2:
The patent divides the cooling function into multiple discrete cooling passages with optimized individual performance. Each passage is designed as a separate segment with specific shaped diffusion geometry, allowing the system to achieve high overall effectiveness with reduced total passage count compared to traditional uniform hole arrangements.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This design improves averaged film cooling effectiveness and coolant coverage, increasing combined cycle efficiency and part durability, reducing metal temperature, and decreasing the likelihood of thermal barrier coating spallation, thus enhancing turbomachine performance and reducing unplanned outages.
Implementation Method 1
a coolant is typically introduced through cooling passages from an interior chamber of the airfoil through holes to an exterior surface of the airfoil... the coolant creates a cooling film, i.e., a flow across and close to the surface of the airfoil
Implementation Method 2
Airfoils are typically covered with a high concentration of a thermal barrier coating (TBC)
Data Source
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AI summary
A turbine airfoil 150, 176, 200 includes a body 210 having an airfoil wall, and a coolant chamber 216 within the body 210. A plurality of cooling passages 204 are within the airfoil wall, each cooling passage 204 includes: a first portion 218 extending from a first point 220 on an exterior surface 222 of the airfoil wall to the coolant chamber 216, and a second portion 230 extending from a second point 232 on the exterior surface 222 of the airfoil wall distal from the first point 220 to intersect a mid-portion 234 of the first portion 218. A cap element 250 closes the first portion 218 at the first point 220 but leaving the second portion 230 open. Each cooling passage 204 has a single inlet 252 in fluid communication with the coolant chamber 216 and a single outlet 254 at the second point 232 of the second portion 230.