Gas Turbine Shroud Cooling via Nozzle Plenum Impingement
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Solution Overview
Problem
Gas turbine shrouds face challenges with localized 'hot spots' due to 'bow wave' flow structures, and traditional cooling methods like film cooling are either ineffective or damage thermal barrier coatings, necessitating a cooling method that does not rely on cooling holes.
Innovation Solution
A cooled shroud assembly for gas turbines with arcuate segments and an outer band featuring strategically angled cooling holes in the nozzle plenum that direct cooling air to negate localized heating without damaging thermal barrier coatings.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If film cooling holes are drilled in the shroud to cool localized hot spots, then cooling effectiveness is improved, but the thermal barrier coating integrity is damaged
Solution Approach 1:
The patent introduces an intermediary cooling structure - a cooling element with cooling channels positioned between the hot gas flow and the shroud surface. This mediator transfers heat away from the shroud without requiring direct penetration into the thermal barrier coating, thus maintaining coating integrity while achieving cooling effectiveness.
Solution Approach 2:
The cooling function is segmented from the shroud structure itself and placed in a separate cooling element. This segmentation allows the shroud to maintain its thermal barrier coating without holes, while the cooling element provides the necessary cooling function through its own internal cooling channels.
2Reliability
If a row of film cooling holes is distributed over the entire shroud segment to protect against bow wave heating, then shroud durability is improved, but cooling flow requirement increases and engine performance decreases
Solution Approach 1:
The cooling system is designed with local quality by positioning cooling holes and cooling channels only in specific areas where bow wave induced heating occurs, rather than distributing cooling holes across the entire shroud. This localized approach reduces the total cooling flow requirement while maintaining shroud durability in critical areas.
Solution Approach 2:
The cooling element acts as an intermediary that concentrates cooling flow in specific high-heat areas, reducing the overall cooling flow demand compared to a distributed film cooling approach across the entire shroud surface.
3Temperature
If cooling holes are drilled in the shroud with thermal barrier coating, then cooling is achieved, but manufacturing complexity and coating repair difficulty increase
Solution Approach 1:
The cooling function is separated from the shroud structure into a distinct cooling element. This segmentation allows the shroud to be manufactured with its thermal barrier coating intact, while the cooling element is manufactured separately with its own cooling channels, simplifying both manufacturing processes and avoiding coating damage.
Solution Approach 2:
The cooling element serves as an intermediary component that provides cooling functionality without requiring modification of the shroud structure. This eliminates the need to drill holes through the thermal barrier coating, thereby simplifying manufacturing and preserving coating integrity.
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 solution effectively protects shroud segments from 'bow wave' induced heating without compromising the integrity of thermal barrier coatings and reduces overall cooling flow requirements, enhancing engine performance.
Implementation Method 1
directing a flow of cooling air from the nozzle plenum to the aft edge of the shroud segment so as to substantially negate the localized heating
Implementation Method 2
The hole is positioned so as to direct a flow of cooling air against the shroud segment
Implementation Method 3
gas flow over the airfoil generates a bow wave which causes localized heating of a portion of the aft flange
Data Source
AI summary
A cooled shroud assembly for a gas turbine engine having a longitudinal centerline includes at least one arcuate shroud segment surrounding a row of rotating turbine blades. The shroud segment has a forward flange, an aft flange defining an axially-facing aft edge, and an inwardly-facing flowpath surface, and the shroud segment lacks cooling holes for cooling the aft flange or the aft edge. At least one stationary turbine nozzle is disposed axially next to the shroud segment and includes an airfoil-shaped airfoil and an arcuate outer band disposed at a radially outer end of the airfoil and positioned axially adjacent to the shroud. At least one cooling hole is formed in the outer band in fluid communication with a source of cooling air. The cooling hole is positioned so as to direct a flow of cooling air against the shroud segment.