Dual-Wall Impingement Cavity for Gas Turbine Airfoil Cooling
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Solution Overview
Problem
Current cooling designs for gas turbine engine airfoils face inefficiencies due to insufficient cooling in certain sections, leading to hot spots and thermal management challenges, particularly in the interaction between hot external and cold internal walls.
Innovation Solution
The implementation of a dual-walled leading edge configuration with impingement holes in the interior impingement wall to create cooling jets on the external surface, along with compartmentalization by ribs to optimize pressure drop splits between impingement and film cooling, and separate feed sources for different impingement cavities to enhance thermal management.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If conventional cooling cavities are used in airfoils, then cooling is provided to airfoil bodies, but hot sections develop where cooling is insufficient
Solution Approach 1:
The cooling cavity is segmented into multiple zones: a feed cavity and multiple impingement cavities (leading edge, pressure side, suction side). This segmentation allows different cooling strategies to be applied to different thermal zones of the airfoil, ensuring adequate cooling in previously problematic hot sections.
Solution Approach 2:
Different cooling methods are applied to different locations: impingement cooling is used in the leading edge cavity where heat transfer coefficients need enhancement, while film cooling is applied to the pressure and suction sides. This local customization of cooling quality addresses the specific thermal requirements of each airfoil section.
2Temperature
If impingement holes are added to create cooling jets, then heat transfer coefficients improve, but device complexity increases
Solution Approach 1:
The impingement cavities are nested within the airfoil structure between the internal and external walls. The feed cavity contains multiple impingement cavities, which in turn contain impingement holes. This nested arrangement integrates the complex cooling functionality within the existing airfoil geometry without requiring external additions.
Solution Approach 2:
The feed cavity and multiple impingement cavities are merged into a single integrated cooling system. Cooling air from a single feed source is distributed to multiple impingement zones through shared cavity space, reducing the number of separate cooling systems needed while maintaining high heat transfer coefficients.
3Temperature
If ribs are used to compartmentalize cavities, then pressure drop splits are optimized, but manufacturing complexity increases
Solution Approach 1:
Ribs are used to segment the feed cavity into distinct regions that feed different impingement cavities. This segmentation allows independent control of pressure drops to each cooling zone, optimizing cooling efficiency while using simple rib structures that are relatively easy to manufacture.
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 design achieves improved heat transfer coefficients and thermal effectiveness by tailoring cooling strategies to meet local thermal and aerodynamic requirements, reducing thermal strain and optimizing cooling performance across the airfoil surfaces.
Implementation Method 1
The impingement wall includes a plurality of impingement holes that fluidly connect the feed cavity to the at least one impingement cavity
Implementation Method 2
cooling jets on the external surface
Implementation Method 3
the external wall includes a plurality of film holes that fluidly connect the at least one impingement cavity to an exterior surface of the external wall
Implementation Method 4
improved heat transfer coefficients and thermal effectiveness
Implementation Method 5
heat transfer coefficients
Data Source
AI summary
Components for gas turbine engines are provided. The components include a hot external wall that is exposed to hot gaspath air when installed within a gas turbine engine, and an interior impingement wall, wherein the interior impingement wall defines a feed cavity and at least one impingement cavity is defined between the impingement wall and the external wall. The impingement wall includes a plurality of impingement holes that fluidly connect the feed cavity to the at least one impingement cavity, the external wall includes a plurality of film holes that fluidly connect the at least one impingement cavity to an exterior surface of the external wall, and wherein the only source of cooling air within the at least one impingement cavity is the feed cavity.