Gas Turbine Airfoil Throughflow Cooling Segregation
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Solution Overview
Problem
Conventional cooled airfoil designs for gas turbine engines, such as single wall cooling, do not effectively segregate cooling flow, leading to inefficient cooling and increased metal temperatures, which can reduce part life and engine performance.
Innovation Solution
The design incorporates a vane with a serpentine flow path and multiple skin core cavities, including throughflow and non-throughflow passages, between external hot and internal cold walls, allowing for optimized cooling by segregating the flow and reducing the amount of cooling flow required.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If conventional single wall cooling is used, then the structure is simple, but the cooling flow is not effectively segregated leading to inefficient cooling and increased metal temperatures
Solution Approach 1:
The airfoil is divided into multiple discrete cooling zones with separate cooling passages for the leading edge, pressure side, and suction side. Each zone has its own cooling flow path, allowing independent temperature control and effective cooling flow segregation, which directly addresses the issue of inefficient cooling in conventional single wall designs.
Solution Approach 2:
Different cooling schemes are implemented in different regions of the airfoil based on local thermal requirements. The leading edge has dedicated cooling passages, while the pressure and suction sides have separate skin-core cooling zones, allowing each region to be cooled according to its specific thermal demands, thereby reducing overall metal temperatures effectively.
2Reliability
If multiple discrete cooling passages are implemented, then cooling efficiency is improved, but the device complexity increases
Solution Approach 1:
The cooling passages are integrated into the skin-core structure of the airfoil, where the cooling channels are formed within the thickness of the airfoil walls themselves. This merging of cooling functionality with the structural skin reduces overall device complexity while maintaining multiple discrete cooling zones for improved reliability.
3Temperature
If more cooling flow is used, then metal temperatures can be reduced, but the amount of cooling flow required increases
Solution Approach 1:
By segmenting the cooling flow into separate zones (leading edge, pressure side, suction side), each zone receives cooling flow only where needed. This prevents the mixing and loss of cooling effectiveness that occurs in single wall designs, allowing metal temperatures to be reduced with less total cooling flow quantity.
Solution Approach 2:
The cooling flow is maintained continuously through dedicated passages that ensure cooling air reaches the hot sections without interruption or mixing with hot gas paths. This continuous, directed cooling action maximizes the useful cooling effect, reducing the total quantity of cooling flow needed to achieve target metal temperatures.
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 approach reduces metal temperatures and improves engine performance by optimizing cooling schemes, allowing for reduced cooling flow while maintaining constant metal temperatures and extending part life.
Implementation Method 1
cooling fluid from opposing sides flow through the passages in a substantially parallel arrangement
Implementation Method 2
Each passage is in thermal communication with a respective external hot wall of the airfoil body
Data Source
Figure 1
Figure 2
Figure 3A~3B
AI summary
Vanes for gas turbine engines are described. The vanes (300) include an airfoil body (302) extending between a first platform and a second platform, the airfoil body having a leading edge (310), a trailing edge (312), a pressure side (314), and a suction side (316) and at least one skin core cavity (320, 322, 324) comprising a first skin core cavity, wherein the first skin core cavity defines a throughflow passage through the vane between the first platform and the second platform, wherein the at least one skin core cavity is defined between an external hot wall (338) of the vane and an internal cold wall (340) of the vane.