Internally Cooled Airfoil Segmented Cooling Passage Design
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Solution Overview
Problem
Gas turbine engines face challenges in efficiently cooling turbine airfoils due to high temperatures, leading to increased manufacturing costs and cycle penalties, as existing internal cooling features are complex and costly to produce.
Innovation Solution
A casting core design with specific geometries and orientations of pedestals, including racetrack and spear pedestals, is used to create internal cooling cavities in airfoils, allowing for effective cooling fluid distribution and reducing manufacturing complexity.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If numerous film cooling holes and high volumes of cooling fluid are used to cool turbine airfoils, then cooling effectiveness is improved, but manufacturing complexity and cost increase
Solution Approach 1:
The cooling system is segmented into multiple functional zones along the airfoil passage: impingement cooling section with perpendicular jets, transition section with angled jets, and film cooling section with surface-parallel jets. This segmentation allows each zone to address specific thermal challenges independently, achieving comprehensive cooling without requiring excessive cooling fluid or complex manufacturing
Solution Approach 2:
The patent employs parameter changes by varying the injection angle of cooling fluid throughout the passage length. The injection angle transitions from perpendicular (90 degrees) at the leading end to progressively smaller angles (45 degrees, then parallel to surface) toward the trailing end. This continuous parameter change optimizes heat transfer efficiency at different locations while maintaining manageable manufacturing complexity
2Reliability
If internal cooling features such as pedestals, air jet impingement, and turbulator treatments are implemented, then cooling performance is enhanced, but manufacturing cost increases due to intricate cores
Solution Approach 1:
The patent merges multiple cooling functions into a single integrated passage structure. The same passage performs impingement cooling, transition cooling, and film cooling functions sequentially, eliminating the need for separate cooling systems or multiple intricate cores. This consolidation maintains high cooling performance while significantly reducing manufacturing complexity and cost
Solution Approach 2:
The cooling passage is designed as a multi-functional component that performs various cooling operations along its length. Different sections of the passage deliver cooling fluid at different angles and positions to address diverse thermal conditions, allowing a single structure to replace multiple specialized components and reduce overall manufacturing burden
3Productivity
If higher turbine inlet temperatures are used to boost engine performance, then efficiency is improved, but the risk of burn-through or creep rupture increases
Solution Approach 1:
The cooling system applies preliminary cooling action by introducing cooling fluid at the leading end of the passage where temperatures are highest. The impingement cooling section immediately counteracts thermal loading before heat can accumulate to dangerous levels, enabling the airfoil to withstand higher inlet temperatures without burn-through or creep rupture
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 design enhances the robustness of airfoil tips, maintains airflow, and reduces mixing losses, thereby improving stage efficiency and reducing the risk of burn-through or creep rupture, while being cost-effective to manufacture.
Implementation Method 1
cooling fluid may be directed to flow through the internal cavities and over a number of pedestals
Implementation Method 2
the cooling fluid temperature may be lowered as it flows across the pedestals
Implementation Method 3
turbulator treatments for the walls
Data Source
AI summary
A casting core may comprise a tip flag material having a forward cavity and a first spear cavity disposed aft of the forward cavity. A trailing edge discharge material may be separated from the tip flag material and include a first row of cavities. The first row of cavities may comprise a first racetrack cavity. A second row of cavities may be disposed aft of the first row of cavities and include a second racetrack cavity. A third row of cavities may be disposed aft of the second row of cavities and include a circular cavity. A fourth row of cavities may be disposed aft of the third row of cavities and include a second spear cavity.