Gas Turbine Airfoil Cooling Insert with Axial Rib Channels
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Solution Overview
Problem
Gas turbine engine airfoil cooling designs face inefficiencies due to insufficient cooling flow and high external heat flux, leading to thermal mechanical fatigue and oxidation issues, particularly in high-temperature environments where conventional cooling methods are inadequate for ceramic matrix composite materials.
Innovation Solution
Incorporating a leading edge space-eater insert member with axial-extending rib roughened cooling channels and metering flow apertures to optimize convective heat transfer and pressure loss, enabling counter-flow cooling and efficient use of high-pressure cooling air, while using ceramic matrix composites and external heat transfer augmentation features to enhance cooling effectiveness.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If conventional cooling cavities are used in airfoils, then cooling is provided to airfoil bodies, but cooling flow characteristics cause hot sections where cooling is not sufficient
Solution Approach 1:
The cooling system is segmented into multiple independent flow channels (first axial flow cooling passage and second axial flow cooling passage) that can be controlled separately. This allows different regions of the airfoil to receive optimized cooling flows, preventing hot sections while maintaining overall cooling effectiveness.
Solution Approach 2:
Different cooling strategies are applied to different regions of the airfoil. The first axial flow cooling passage provides cooling to specific regions while the second passage addresses other regions, with each channel having optimized flow characteristics for its specific thermal requirements. This local optimization eliminates hot sections.
2Temperature
If cooling flow is increased to improve cooling effectiveness, then thermal cooling is enhanced, but pressure loss and energy efficiency deteriorate
Solution Approach 1:
The cooling system incorporates variable geometry features including adjustable flow control elements and adaptable passage configurations that optimize cooling flow dynamics. This dynamic control allows the system to achieve effective cooling while minimizing pressure loss by adapting flow characteristics to actual thermal conditions.
Solution Approach 2:
The system optimizes cooling by changing flow parameters such as velocity, pressure, and distribution patterns across different passages. By carefully controlling these parameters, the system achieves effective cooling with reduced pressure loss compared to conventional high-flow cooling approaches.
3Temperature
If axial flow cooling passages are used to improve cooling, then cooling airflows are supplied axially, but flow distribution and heat transfer efficiency may be insufficient
Solution Approach 1:
The system merges axial flow cooling with radial flow components by incorporating impingement cooling sections where cooling air is directed radially onto the airfoil internal surfaces. This combination of axial and radial flow patterns enhances heat transfer efficiency while maintaining good flow distribution throughout the airfoil structure.
Solution Approach 2:
The cooling system transitions from purely axial flow to a multi-dimensional flow pattern by introducing radial impingement components. This adds a radial dimension to the cooling flow, creating more effective heat transfer zones while improving overall cooling distribution across the airfoil internal surfaces.
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 solution provides improved thermal cooling effectiveness and durability for gas turbine engine airfoils by optimizing cooling airflow distribution and convective heat transfer, reducing thermal gradients, and extending the operational temperature limits of ceramic matrix composite materials.
Implementation Method 1
an internal surface of the airfoil body defines a cooling airflow passage extending in an axial direction between a leading edge and a trailing edge of the airfoil. A trailing edge baffle is disposed within the cooling airflow passage
Implementation Method 2
Such cooling cavities are subject to both hot material walls (exterior or external) and cold material walls (interior or internal)
Implementation Method 3
EP 3 133 242 A1 discloses an article including a manifold, an article wall, and a post-impingement cavity disposed between the manifold and the article wall
Data Source
Figure 1
Figure 2
Figure 3A
AI summary
Component for gas turbine engines (20) are described. The components include an airfoil body (402; 502) having leading and trailing edges (404, 406) and pressure and suction sides (408, 410). The airfoil has a leading edge cavity (412; 512) located proximate the leading edge (404) defined between the leading edge (404) and a separator rib (426; 526) and between the pressure side (408) and the suction side (410). An insert member (418; 518; 618) is installed within the leading edge cavity (412; 512). The insert member (418; 518; 618) has one or more metering flow apertures (420; 520; 620) at an aft end and one or more impingement apertures (422; 522; 622) at a forward end and at least one axially extending rib (530; 630) along an exterior surface thereof. At least one axial extending flow channel passage (532; 632) is defined along the axial extending rib (530; 630) between an exterior of the insert member (418; 518; 618) and an interior of the airfoil body (402; 502). Air flow through the metering flow apertures (420; 520; 620) flows into the axial extending flow channel passage (532; 632) and flows forward toward the leading edge (404).