Gas Turbine Blade Tip Recess with Thermal Barrier Coating
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Solution Overview
Problem
Turbine blade tips in gas turbine engines operate in extreme temperatures, leading to erosion and oxidation, which reduces efficiency due to leakage air and existing cooling techniques alter the aerodynamic shape, affecting blade efficiency.
Innovation Solution
A recess is created at the terminal end surface of the blade tip, filled with a thermal barrier coating that forms a planar surface, extending beyond the edge, without cooling holes, to provide thermal protection without altering the aerodynamic shape.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If cooling holes are provided in the tip shelf or squealer pocket to reduce tip temperatures, then thermal protection is improved, but the aerodynamic shape is altered and blade efficiency is reduced
Solution Approach 1:
The invention removes cooling holes from the tip surface and extracts the cooling function to the blade interior. The interior cooling passages cool the blade body, and the thermal barrier coating on the tip surface provides thermal protection without requiring surface cooling holes, thus preserving the aerodynamic shape while achieving thermal protection.
Solution Approach 2:
The invention introduces a thermal barrier coating as an intermediary layer between the hot tip surface and the blade metal. This coating provides thermal protection and allows the tip surface to maintain its aerodynamic shape without requiring cooling holes, mediating between thermal protection requirements and aerodynamic efficiency.
2Temperature
If a large airfoil tip shelf is provided to reduce tip temperatures, then thermal protection is improved, but the tip aerodynamic shape is changed and blade efficiency is reduced
Solution Approach 1:
The invention applies local quality by providing thermal barrier coating specifically on the tip surface where thermal protection is needed, while maintaining the original aerodynamic shape. The cooling passages are positioned in the blade interior rather than altering the tip geometry, allowing localized thermal protection without global shape modification.
3Temperature
If the recess depth is increased to provide better thermal protection, then temperature reduction is improved, but the aerodynamic shape is more significantly altered
Solution Approach 1:
The invention applies preliminary action by pre-cooling the blade interior through internal cooling passages before heat reaches the tip surface. The thermal barrier coating is also applied in advance to the tip surface, providing thermal protection before thermal exposure occurs, thereby reducing the need for deep recesses that would alter aerodynamic shape.
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
The thermal barrier coating effectively reduces temperature at the blade tip without relying on cooling fluids, maintaining aerodynamic efficiency and preventing erosion, while being flush with the terminal end surface to preserve the tip shape.
Implementation Method 1
the recess is completely filled with a thermal barrier coating
Implementation Method 2
which results in erosion and oxidation at the tip
Data Source
Figure 1
Figure 2A~2C
Figure 3A~6
AI summary
A blade (64) for a gas turbine engine (20) includes an airfoil (78) that has a tip (80;180;280;380) with a terminal end surface (102). The terminal end surface (102) includes a recess (100;200;300;400) that has a depth (104) of less than 40 mils (1.016 mm). The recess (100;200;300;400) is filled with a thermal barrier coating (106). The recess (100;200;300;400) is without any cooling holes.