Turbine Blade Trailing Edge Cooling with Segmented Ribs
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Solution Overview
Problem
In gas turbine engines, existing turbine blade cooling technologies face challenges in achieving efficient heat transfer at the trailing edge while maintaining a strong casting core, often resulting in excessive coolant flow at the airfoil root and tip, which can lead to inefficient coolant consumption and structural issues.
Innovation Solution
The implementation of an array of elongated pins at the trailing edge of the turbine blade, arranged in radial rows with staggered configurations, along with a ceramic casting core design featuring indentations and perforations, creates a longer coolant flow path and tortuous passages to enhance heat transfer and restrict coolant flow, thereby improving cooling efficiency and core strength.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Strength
If larger coolant exit aperture dimensions are used at the airfoil trailing edge near the root and tip, then the casting core structural strength is improved, but the coolant flow rate increases excessively
Solution Approach 1:
The coolant exit aperture is segmented into multiple smaller sub-apertures arranged in a matrix pattern rather than using a single large aperture. This segmentation maintains the overall exit area for structural strength while dividing the flow into multiple restricted paths, reducing the total coolant flow rate through the trailing edge.
Solution Approach 2:
Different regions of the trailing edge are given different aperture characteristics. The matrix pattern of sub-apertures creates local variations in flow resistance and distribution, allowing optimization of both structural integrity and coolant flow control at specific locations along the airfoil span.
2Strength
If a picture frame like configuration is used at the airfoil trailing edge, then the casting core strength is improved, but coolant flow is increased near the airfoil root and tip
Solution Approach 1:
The picture frame configuration is modified by segmenting the trailing edge structure into a matrix of smaller features rather than a continuous frame. This maintains the structural reinforcement benefits while breaking up the coolant flow path to reduce overall consumption and improve distribution efficiency.
Solution Approach 2:
The trailing edge structure combines the picture frame configuration with a matrix pattern of sub-apertures, creating a composite structural-cooling feature that provides both mechanical strength and optimized coolant flow characteristics.
3Speed
If the trailing edge is made thin for aerodynamic efficiency, then the aerodynamic performance is improved, but the structural strength and cooling efficiency are compromised
Solution Approach 1:
The trailing edge features localized matrix patterns of cooling sub-apertures and framing elements at specific positions along the span, allowing the overall trailing edge to remain thin for aerodynamics while providing localized structural reinforcement and cooling functionality where needed.
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 increases convective heat transfer and pressure drop, reducing the coolant flow rate while maintaining structural integrity of the casting core, thus optimizing cooling efficiency and minimizing coolant consumption.
Implementation Method 1
This design increases convective heat transfer and pressure drop, reducing the coolant flow rate while maintaining structural integrity
Data Source
Figure 1
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Figure 3~4
AI summary
A turbine airfoil (10) includes a trailing edge coolant cavity (41f) located in an airfoil interior (11) between a pressure sidewall (14) and a suction sidewall (16). The trailing edge coolant cavity (41f) is positioned adjacent to a trailing edge (20) of the turbine airfoil (10) and is in fluid communication with a plurality of coolant exit slots (28) positioned along the trailing edge (20). At least one framing passage (70, 80) is formed at a span-wise end of the trailing edge coolant cavity (41f). The airfoil (10) further includes framing features (72A-B, 82A-B) located in the framing passage (70, 80). The framing features are configured as ribs (72A-B, 82A-B) protruding from the pressure sidewall (14) and/or the suction sidewall (16). The ribs (72A-B, 82A-B) extend partially between the pressure sidewall (14) and the suction sidewall (16).