Serpentine Cavity Cooling Passage for Gas Turbine Components
Find Innovative SolutionsGenerate Solutions
Solution Overview
Problem
Gas turbine engine components, particularly in the turbine section, face challenges in effectively managing high temperatures due to the lack of efficient cooling mechanisms, which can lead to reduced performance and lifespan.
Innovation Solution
The implementation of a serpentine cavity cooling passage with multiple inlets and cooling augmentation features, such as trip strips and pin fins, within turbine vanes to enhance heat transfer and utilize bleed air from compressor stages for efficient cooling, where the second inlet is strategically placed downstream of a pressure drop to reintroduce cooling fluid and maintain optimal heat transfer.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If a single inlet cooling passage is used in turbine components, then the structure is simple, but the cooling efficiency is insufficient to manage high temperatures
Solution Approach 1:
The cooling passage is divided into multiple segments with separate inlets (first inlet receiving first cooling fluid, second inlet receiving second cooling fluid) rather than a single inlet system. This segmentation allows different cooling fluids from different compressor stages to be introduced at different locations, improving temperature management across the turbine component while maintaining reasonable structural complexity
Solution Approach 2:
Different regions of the turbine component receive cooling fluids with different properties (pressure, temperature) from different compressor stages through the segmented inlet system. The first cooling fluid from the first compressor stage and the second cooling fluid from the second compressor stage are directed to specific locations within the cooling passage, providing locally optimized cooling where needed most
2Reliability
If cooling air is extracted from the compressor section, then turbine components can be cooled, but the overall engine efficiency is reduced
Solution Approach 1:
The system utilizes bleed air from multiple compressor stages with different pressure and temperature parameters. By tapping cooling fluids at different stages (first compressor stage providing first cooling fluid, second compressor stage providing second cooling fluid), the system optimizes the thermal parameters of the cooling air to achieve effective component cooling while minimizing the impact on overall engine efficiency
Solution Approach 2:
The turbine component's own cooling requirements are met by utilizing the engine's existing compressor output. The compressor section naturally produces bleed air that is redirected through the multi-inlet cooling passage system, allowing the turbine component to self-cool using resources already available within the engine system without requiring external cooling mechanisms
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 effectively reduces temperature and pressure gradients within the turbine section, enhancing the cooling efficiency and extending the lifespan of components by utilizing high-pressure bleed air from compressor stages, thereby improving the overall performance of the gas turbine engine.
Implementation Method 1
an internal cooling passage (e.g., a serpentine cavity) that is fed by first and second inlets that respectively receive fluid from the first and second cooling sources
Implementation Method 2
cooling augmentation features, such as trip strips and pin fins
Implementation Method 3
The first compressor stage has a higher pressure than the second compressor stage
Data Source
AI summary
A gas turbine engine includes a compressor section that provides first and second compressor stages that are configured to respectively provide first and second cooling fluids. The first compressor stage has a higher pressure than the second compressor stage. The gas turbine engine further includes a component that has platform with an internal cooling passage fed by first and second inlets that respectively receive fluid from the first and second cooling sources. The second inlet is downstream from the first inlet.