Gas Turbine Flow Path Surface for Uniform Cooling
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Solution Overview
Problem
The non-uniform static pressure distribution around turbine blades in a gas turbine engine leads to uneven cooling air flow and temperature distribution, resulting in reduced cooling efficiency and increased cooling air supply, which deteriorates the engine's performance.
Innovation Solution
The engine features a flow path surface with depression and protrusion portions that balance the static pressure distribution, ensuring uniform cooling air flow and efficiency by adjusting the flow path area and aerodynamic characteristics.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If cooling air is supplied to the upstream side of the turbine blades, then the heat-resistance performance of the turbine blades and end wall is improved, but the static pressure distribution at the surface of the end wall becomes non-uniform
Solution Approach 1:
The flow path surface is designed with local variations in shape, including depression portions in areas overlapping with turbine blade front ends and protrusion portions in areas between blade front ends. This local geometric differentiation modifies the flow characteristics and pressure distribution in specific zones to achieve more uniform cooling air flow while maintaining heat-resistance performance.
Solution Approach 2:
The flow path surface introduces asymmetric geometric features (depression and protrusion portions) that break the symmetry of the flow path. This asymmetric design creates controlled variations in flow direction and pressure distribution to prevent localized high-pressure zones that would otherwise cause non-uniform cooling air flow.
2Reliability
If the supply flow rate of cooling air is increased to satisfy heat-resistance performance in areas with poor cooling air flow, then the heat-resistance performance is improved, but the entire supply of cooling air is increased and engine performance deteriorates
Solution Approach 1:
Instead of uniformly increasing cooling air supply throughout the entire flow path, the invention applies partial action by concentrating cooling air flow through the depression portions in specific areas where it is most needed. This targeted approach ensures adequate cooling in critical zones without unnecessarily increasing the overall cooling air supply, thereby maintaining engine performance.
Solution Approach 2:
The invention changes the geometric parameters of the flow path surface by introducing depression and protrusion portions with specific dimensions and positions. These parameter changes modify the flow characteristics and pressure distribution to optimize cooling air flow patterns, ensuring efficient heat-resistance performance without excessive cooling air consumption.
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 uniformizes the static pressure and cooling film thickness, preventing high-temperature areas and reducing the overall cooling air supply, thereby enhancing the gas turbine engine's performance.
Implementation Method 1
the static pressure distribution at the surface of the end wall in the upstream side of the turbine blades becomes non-uniform due to the existence of the turbine blades
Implementation Method 2
cooling air is supplied to the upstream side of the turbine blades, the cooling air flows along the surface of the end wall, and thereby the film cooling is performed on the surface of the end wall
Data Source
AI summary
The gas turbine engine (S1) includes: turbine blades (7b); and a cooling air supply unit (11) to supply cooling air to the turbine blades (7b). A flow path surface (31) is formed so as to be positioned in an upstream side of the turbine blades (7b) and so as to be connected to a base surface (32) in which the turbine blades (7b) are provided. The flow path surface (31) includes: depression portions (31a) depressed relative to the base surface (32), each depression portion (31a) including at least an area overlapping with a front end (7b1) of each turbine blade (7b), when viewed from a direction of the turbine axis (L); and protrusion portions (31b) protruding relative to the base surface (32), each protrusion portion (31b) being at least part of each area positioned between front ends (7b1) of the turbine blades (7b), when viewed from the above direction.