Turbine Blade Trailing Edge Cooling with Zig-Zag Passages
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Solution Overview
Problem
Current cooling designs for gas turbine airfoils, particularly in the trailing edge, face inefficiencies in heat transfer and uniformity due to turbulent environments and high coolant flow rates, leading to suboptimal performance and increased compressor air diversion.
Innovation Solution
The design incorporates a series of cooling chambers and passages within the airfoil with strategically positioned flow paths and textured surfaces to enhance heat transfer, allowing for more efficient cooling with reduced air usage by optimizing the flow of coolant through zig-zag patterns and sloped passages, thereby increasing heat transfer efficiency and uniformity.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Stability of the object's composition
If film cooling is used along exterior wall surfaces, then cooling uniformity is improved, but cooling effectiveness deteriorates in turbulent environment
Solution Approach 1:
The patent uses a thin film cooling approach where cooling air is directed along the exterior wall surface to create a protective film. This film acts as a thermal barrier between the hot combustion gases and the airfoil surface, improving cooling uniformity while maintaining effectiveness through careful film formation and attachment control.
Solution Approach 2:
The patent optimizes film cooling by adjusting parameters such as cooling air temperature, pressure, and injection angle to maintain film stability and effectiveness in the turbulent gas turbine environment. By changing these parameters, the system achieves both uniform cooling distribution and reliable heat transfer protection.
2Use of energy by moving object
If internal cooling channels are used, then heat transfer efficiency is improved, but cooling air volume requirement increases
Solution Approach 1:
The patent divides the internal cooling system into multiple separate channels and chambers within the airfoil structure. This segmentation allows for optimized flow distribution through each channel, improving overall heat transfer efficiency while reducing the total cooling air volume required compared to a single large channel system.
Solution Approach 2:
The patent implements nested cooling channels where smaller channels are positioned within or alongside larger channels. This nested arrangement maximizes the cooling surface area within the limited airfoil thickness, enhancing heat transfer efficiency without proportionally increasing the cooling air volume requirement.
3Power
If trailing edge is made thin for aerodynamic efficiency, then aerodynamic performance is improved, but heat transfer efficiency deteriorates
Solution Approach 1:
The patent applies localized cooling measures specifically at the trailing edge region where heat transfer is most critical. By concentrating cooling resources and optimizing channel geometry in this specific area, the system maintains thin trailing edge geometry for aerodynamic efficiency while compensating for the reduced heat transfer capability through enhanced local cooling design.
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 enhances heat transfer efficiency and uniformity in the trailing edge of turbine airfoils, allowing for higher power output and extended operational temperatures while minimizing the volume of air diverted from the compressor for cooling.
Implementation Method 1
textured surfaces to enhance heat transfer
Implementation Method 2
heat transfer efficiency and uniformity in the trailing edge
Implementation Method 3
flow paths and textured surfaces to enhance heat transfer, allowing for more efficient cooling with reduced air usage by optimizing the flow of coolant
Implementation Method 4
coolant flow rate is desired to provide the requisite rate of heat transfer for maintaining mechanical integrity
Implementation Method 5
internal cooling channels which remove heat from the pressure sidewall and the suction sidewall
Data Source
AI summary
A turbine blade (10) including an airfoil (12) having multiple interior wall portions (70) each separating at least one chamber from another one of multiple chambers (46, 48, 50, 58, 60). In one embodiment a first wall portion (70-2) between first and second chambers (60, 52) includes first and second pluralities of flow paths (86P, 86S) extending through the first wall portion. The first wall portion includes a first region R1 having a first thickness, t, measurable as a distance between the chambers. One of the paths extends a first path distance, d, as measured from an associated path opening (78) in the first chamber (60), through the first region and to an exit opening (82) in the second chamber (52) which path distance is greater than the first thickness.