Turbine Blade Cooling Inserts for Aero-Efficiency
Find Innovative SolutionsGenerate Solutions
Solution Overview
Problem
Conventional internal cooling methods for turbine airfoils, such as film cooling, reduce aero-efficiency and are inefficient due to post-impingement cross-flow degradation effects, which impede the cooling effectiveness and require more coolant to maintain the airfoils at desired temperatures.
Innovation Solution
The design incorporates a hollow airfoil with radially extending chambers and inserts that direct coolant through insert apertures toward the inner surface, forming inward bleed channels and a central collector passage, reducing post-impingement flow and enhancing heat exchange by minimizing cross-flow and optimizing coolant flow paths.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If film cooling is used to cool turbine airfoils, then cooling effectiveness is improved, but aero-efficiency deteriorates
Solution Approach 1:
The cooling system is segmented into multiple functional zones within the airfoil: impingement cooling regions with directed coolant jets, transition regions, and film cooling regions. This segmentation allows different cooling mechanisms to be applied in optimal locations, reducing overall coolant requirements and improving aerodynamic efficiency compared to uniform film cooling.
Solution Approach 2:
Impingement cooling is applied preliminary to film cooling by directing coolant jets at the inner surface of the airfoil before the coolant exits through film cooling holes. This preliminary action pre-cools the airfoil structure, reducing the thermal load on subsequent film cooling and allowing for reduced coolant flow rates, thereby improving aero-efficiency.
2Temperature
If conventional internal cooling circuits are used, then cooling is provided, but post-impingement cross-flow degradation reduces cooling effectiveness
Solution Approach 1:
The cooling circuit is designed with local quality variations through strategically placed impingement zones, transition regions, and film cooling holes at specific locations along the airfoil. Each region has optimized coolant flow characteristics tailored to local thermal conditions, preventing cross-flow degradation and maintaining high cooling effectiveness throughout the airfoil structure.
Solution Approach 2:
The cooling approach transitions from two-dimensional planar film cooling to three-dimensional impingement cooling with directed coolant jets that penetrate deeper into the airfoil structure. This dimensional change allows coolant to reach thermally critical regions more effectively and reduces the impact of cross-flow degradation by creating localized high-velocity cooling zones.
3Temperature
If more coolant is supplied to maintain airfoils at desired temperatures, then cooling effectiveness is improved, but coolant consumption increases
Solution Approach 1:
Instead of applying film cooling uniformly across the entire airfoil surface, the system uses partial action by concentrating impingement cooling in thermally critical regions and applying film cooling only where necessary. This selective approach maintains desired airfoil temperatures while significantly reducing overall coolant consumption compared to full-surface film cooling.
Solution Approach 2:
The system replaces reliance on high-volume film cooling with a hybrid approach that uses impingement cooling (directed coolant jets) and controlled film cooling. This substitution of cooling mechanisms achieves equivalent or superior cooling effectiveness with reduced coolant flow rates, thereby reducing coolant consumption and improving engine efficiency.
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 improves cooling effectiveness, reduces the amount of coolant required, and increases the efficiency of the turbine engine by minimizing coolant usage and enhancing heat transfer, thereby reducing operational stresses on turbine airfoils.
Implementation Method 1
the inserts are configured to initially receive at least a portion of the coolant entering the chamber and direct a substantial portion of the coolant through a plurality of insert apertures toward the inner surface of the outer wall
Implementation Method 2
As the compressed air passes through the airfoil, it convectively cools the airfoil, which may allow the part to withstand firing temperatures that it otherwise could not
Implementation Method 3
the supply of compressed air is released through small holes on the surface of the airfoils. Released in this manner, the supply of air forms a thin layer or film of relatively cool air at the surface of the airfoil, which both cools and insulates the part from the higher temperatures that surround it
Data Source
AI summary
A turbine blade with a generally hollow airfoil having an outer wall that defines at least one radially extending chamber for receiving the flow of a coolant, the airfoil including a leading edge that resides in an upstream or forward direction, a trailing edge that resides in a downstream or aft direction, a convex suction side, and a concave pressure side, the turbine blade comprising: a plurality of inserts disposed within the chamber that are configured to initially receive at least a portion of the coolant entering the chamber and direct a substantial portion of the coolant through a plurality of insert apertures toward the inner surface of the outer wall; wherein the inserts are configured to form at least one inward bleed channel and a central collector passage into which the inward bleed channel flows.