Gas Turbine Rotor Blade Cooling Cavity Design
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Solution Overview
Problem
Serpentine cooling passages in gas turbine engine rotor blades result in high pressure losses due to repeated coolant flow direction changes, increasing parasitic losses and reducing engine efficiency.
Innovation Solution
The implementation of a low pressure loss cooling scheme in rotor blades, featuring discrete, generally upward-flowing cooling cavities separated by ribs and connected by crossover holes, which reduces pressure losses by minimizing coolant flow direction changes.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If serpentine cooling passages are used to ensure sufficient convective cooling, then cooling effectiveness is improved, but pressure losses increase due to repeated flow direction changes
Solution Approach 1:
The cooling passage is divided into multiple discrete cooling cavities (first cooling cavity, second cooling cavity, third cooling cavity) separated by ribs. Each cavity receives coolant through dedicated inlet passages, allowing independent flow paths that reduce the need for sharp direction changes while maintaining comprehensive cooling coverage across the airfoil section.
Solution Approach 2:
The cooling system transitions from a single serpentine path to a multi-dimensional array of discrete cavities arranged spanwise. Coolant is distributed across multiple parallel flow paths through separate inlet passages, enabling cooling in multiple spatial dimensions simultaneously without requiring the coolant to traverse the entire airfoil length in a single winding path.
2Reliability
If serpentine passages are used to provide full downward pass cooling, then convective cooling is improved, but parasitic losses increase due to flow direction changes
Solution Approach 1:
The airfoil is divided into multiple cooling zones with discrete cavities separated by ribs. Each cavity is fed by dedicated inlet passages, creating segmented cooling zones that can be independently optimized. This segmentation eliminates the need for a single long serpentine path, reducing flow direction changes and associated parasitic losses while maintaining reliable cooling in each zone.
Solution Approach 2:
Different regions of the airfoil are provided with locally-optimized cooling cavities positioned at specific locations (leading edge, intermediate, trailing edge). Each local cooling cavity is tailored to the thermal requirements of its specific region, allowing effective convective cooling where needed without forcing coolant through unnecessary path length and direction changes in other regions.
3Object-affected harmful factors
If serpentine cooling passages are used to cool hot combustion products, then thermal protection is improved, but engine efficiency decreases due to increased pressure losses
Solution Approach 1:
The cooling system uses multiple parallel coolant flow paths distributed spanwise across the airfoil, rather than a single serpentine path. This multi-dimensional arrangement allows coolant to be delivered simultaneously to multiple cooling zones, providing comprehensive thermal protection while minimizing the total path length and direction changes, thereby reducing pressure losses and improving engine efficiency.
Solution Approach 2:
The cooling system is segmented into multiple independent cooling cavities with dedicated inlet passages. This segmentation allows each cavity to be optimized for its specific thermal protection role while using shorter, more direct coolant paths. The overall system provides complete thermal protection across the airfoil without the excessive pressure losses associated with a single long serpentine passage.
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 cooling scheme effectively maintains convective cooling while significantly reducing pressure losses, thereby enhancing engine efficiency by optimizing coolant flow and reducing parasitic losses.
Implementation Method 1
provide sufficient convective cooling, the internal cooling cavities often include a serpentine portion through which the coolant is intended to make at least one full downward pass through the airfoil
Data Source
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AI summary
A rotor blade comprises a root section, an airfoil section, a leading edge cooling cavity, an intermediate cooling cavity, and a trailing edge cooling cavity. The leading edge, intermediate, and trailing edge cooling cavities each extend spanwise through the airfoil section from a coolant inlet passage in the root section, and each terminate proximate the airfoil tip.