Turbine Vane Cooling Channel Insert for Gas Turbine
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Solution Overview
Problem
Turbine vanes in gas turbines face challenges in efficiently guiding cold air flow through internal cooling channels to prevent deformation due to high-temperature combustion gases, leading to suboptimal cooling and potential damage.
Innovation Solution
The internal cooling channel of the turbine vane is divided by an insert with a partition and throttle plate, creating separate pressure side and suction side passages, allowing for optimal flow rates of cold air through strategically sized throttle holes and communication holes, ensuring efficient cooling.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Device complexity
If a single cooling channel is used in the turbine vane, then the structure is simple, but the cold air flow rate to each passage cannot be optimized, leading to insufficient cooling efficiency
Solution Approach 1:
The single cooling channel is divided into multiple separate cooling passages (first cooling passage and second cooling passage) using a partition structure. This segmentation allows independent flow control for each passage, enabling optimized cold air distribution to different regions of the turbine vane, thereby improving cooling efficiency while maintaining relatively simple overall structure.
2Reliability
If cold air flow rate is increased to improve cooling, then cooling efficiency improves, but the structure cannot guide cold air to flow at optimal rates to each passage, causing uneven cooling distribution
Solution Approach 1:
Different throttle plates are installed in different cooling passages, with each throttle plate having specifically designed opening areas to control the cold air flow rate for that particular passage. This local differentiation allows each cooling passage to receive the optimal amount of cold air needed for its specific cooling requirements, achieving uniform and efficient cooling distribution across the turbine vane.
3Device complexity
If the turbine vane is exposed to high-temperature combustion gases without adequate cooling, then the structure remains simple, but deformation occurs leading to potential damage
Solution Approach 1:
The cooling system is segmented into multiple independent passages that can be strategically positioned to cover different high-temperature exposure zones on the turbine vane. This segmentation enables targeted cooling of specific areas most susceptible to thermal deformation, providing comprehensive protection against heat-related damage while maintaining structural simplicity.
Solution Approach 2:
Cold air is introduced into the cooling passages before the hot combustion gases contact the turbine vane surfaces. This preliminary cooling action establishes a protective temperature gradient and thermal barrier in advance, preventing the turbine vane material from reaching deformation temperatures when exposed to combustion gases.
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 configuration enables efficient cooling of the turbine vane by optimizing the flow rate of cold air through the passages, reducing the risk of deformation and enhancing the operational efficiency of the gas turbine.
Implementation Method 1
a throttle plate coupled to the partition to cover an inlet of the at least one cooling channel. The throttle plate may include a first throttle hole for introducing cold air into the pressure side passage and a second throttle hole for introducing cold air into the suction side passage
Implementation Method 2
at least one cooling channel formed radially in the airfoil... guiding cold air flow through internal cooling channels to prevent deformation due to high-temperature combustion gases
Data Source
AI summary
A turbine vane and a gas turbine including the same are provided. The turbine vane includes an airfoil having a pressure side and a suction side, at least one cooling channel formed radially in the airfoil, and an insert inserted into the at least one cooling channel to divide the cooling channel into a pressure side passage and a suction side passage.