Turbine Blade Cooling Path with Radial Diffuser
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Solution Overview
Problem
Gas turbine engine turbine components face inefficiencies and degradation due to mechanical and thermal stresses, with existing cooling methods being insufficient to fully reduce thermal loads, limiting the ability to operate at higher temperatures and generate additional power.
Innovation Solution
A turbine blade with a unique cooling path that routes coolant through radially spaced flow dividers proximate the trailing edge, minimizing coolant flow separation and increasing film cooling coverage, formed through machining or casting processes with a diffuser region biased towards the radial direction.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If existing cooling methods are used in turbine components, then some cooling effect is achieved, but the thermal load on the airfoil is not sufficiently reduced
Solution Approach 1:
The cooling cavity is divided into multiple sections by radially spaced flow dividers, creating separate cooling passages. This segmentation allows coolant to be distributed more effectively across different regions of the airfoil, increasing overall cooling coverage and reducing thermal load more uniformly throughout the component structure.
Solution Approach 2:
Flow dividers are positioned at specific locations within the cooling cavity to create localized cooling zones. The diffuser region with biased openings provides targeted cooling in critical areas near the trailing edge, where thermal loads are highest, thereby improving cooling effectiveness in the most vulnerable regions.
2Reliability
If coolant flow is increased to reduce thermal load, then cooling capacity improves, but coolant flow separation increases reducing efficiency
Solution Approach 1:
The diffuser region is designed with a specific geometry that preliminarily conditions the coolant flow before it exits through the biased openings. This preliminary action aligns the flow direction with the desired cooling path, preventing flow separation and maintaining high cooling efficiency even at increased flow rates.
Solution Approach 2:
The flow dividers act as intermediary structures that guide and condition the coolant flow between the cooling cavity and the external environment. These intermediaries ensure smooth flow transition and prevent separation, maintaining cooling efficiency while allowing increased coolant flow capacity.
3Reliability
If conventional cooling passages are used, then simple structure is maintained, but film cooling coverage is insufficient
Solution Approach 1:
The cooling structure is segmented into multiple passages defined by flow dividers, with each passage contributing to the overall film cooling coverage. This segmentation approach systematically increases coverage area without requiring a complete redesign of the entire cooling system.
Solution Approach 2:
The flow dividers extend in the radial dimension, creating three-dimensional cooling passages that utilize the depth of the cooling cavity more effectively. This dimensional approach increases film cooling coverage by utilizing the radial spacing between flow dividers to create multiple parallel cooling streams.
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 enhanced cooling capacity allows the turbine to operate at higher temperatures, increasing power generation and extending the life cycle of the turbine blades by maintaining convective cooling efficiency.
Implementation Method 1
As the coolant moves through the internal cavity of the airfoil it cools the exposed surfaces within the internal cavity through convection
Implementation Method 2
by providing a diffuser region favoring the radial direction of the turbine, coolant flow separation from the flow dividers is minimized and the cooling effectiveness is thereby increased
Data Source
AI summary
A turbine blade for a gas turbine engine. The turbine blade includes a cooling path for a coolant, routing the coolant through an internal cooling cavity and out through a plurality of cooling holes formed proximate a trailing edge of the turbine blade. Each of the cooling holes including a diffusing region designed so that a coolant does not separate from a radially inward sidewall of the diffusing region.