Turbine Blade Tip Cooling with Segmented Flow Metering
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Solution Overview
Problem
Current film cooling arrangements for gas turbine engines are inefficient in distributing coolant to the external surfaces of turbine blades, leading to inadequate heat management and potential oxidation, particularly in the radial gaps between the blade outer air seal and the airfoil tip.
Innovation Solution
The design incorporates a turbine blade with a platform, airfoil tip, and airfoil section featuring a cavity and multiple cooling passages between the airfoil tip and the cavity, where internal features within the passages meter the coolant flow to exit ports that eject fluid onto the external surfaces, including a radial gap to form a seal and reduce leakage air.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If traditional film cooling arrangements are used, then the structure is simple, but the coolant distribution efficiency is poor and heat management is inadequate
Solution Approach 1:
The cooling system is segmented into multiple cooling passages (first cooling passage, second cooling passage, third cooling passage) with distinct functions. Each passage delivers coolant to specific regions of the airfoil tip, enabling targeted heat management. The segmentation allows independent optimization of coolant flow to different thermal zones, resolving the contradiction between simple structure and effective heat management.
Solution Approach 2:
Different cooling passages are designed with local quality variations to address specific thermal needs. The first cooling passage targets the radial gap region, the second cooling passage targets the external surface, and the third cooling passage provides additional cooling. Each passage has optimized geometry and flow characteristics suited to its specific cooling zone, achieving effective heat management through localized cooling strategies.
2Temperature
If coolant flow is increased to improve cooling, then heat management improves, but leakage air through the radial gap increases
Solution Approach 1:
The first cooling passage acts as an intermediary system that delivers coolant to the radial gap region, creating a protective coolant barrier. This barrier seals the radial gap between the airfoil tip and blade outer air seal, preventing hot leakage air from entering the radial gap while simultaneously cooling the airfoil tip. The intermediary coolant flow resolves the contradiction by serving dual functions: cooling the surface and sealing the gap.
Solution Approach 2:
The cooling system is designed to self-seal the radial gap through the coolant flow itself. The coolant ejected from the first cooling passage automatically forms a seal in the radial gap without requiring additional sealing components. The same coolant that cools the airfoil tip also performs the sealing function, making the system self-sufficient and resolving the contradiction between cooling and gap sealing.
3Reliability
If multiple cooling passages are added to improve coolant distribution, then cooling efficiency improves, but the device complexity increases
Solution Approach 1:
Each cooling passage is designed with multi-functionality to reduce overall system complexity. The first cooling passage simultaneously cools the radial gap region and seals the gap, preventing leakage air. The second cooling passage cools the external surface while the third cooling passage provides supplemental cooling. This multi-functionality allows effective coolant distribution across multiple zones without proportionally increasing complexity, as each component performs multiple tasks.
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 configuration enhances film cooling, reduces localized heat flux, minimizes oxidation, and improves the durability of the airfoil tip by effectively distributing coolant and sealing the radial gap, thereby increasing the efficiency and performance of the gas turbine engine.
Implementation Method 1
Current film cooling arrangements for gas turbine engines are inefficient in distributing coolant to the external surfaces of turbine blades
Implementation Method 2
An internal feature within each of the plurality of cooling passages is configured to meter flow to the exit port
Implementation Method 3
a plurality of cooling passages radially between the cavity and the airfoil tip
Data Source
AI summary
A turbine blade according to an example of the present disclosure includes, among other things, a platform extending from a root section, an airfoil section extending radially from the platform to an airfoil tip, a plurality of cooling passages defined in an external wall of the airfoil tip, the plurality of cooling passages extending radially between the airfoil tip and a cavity in the airfoil section bounded by the external wall, and each of the plurality of cooling passages defining an inlet port along the cavity and an exit port adjacent the airfoil tip, and at least one internal feature within each of the plurality of cooling passages that meter flow to the respective exit port.