Integrated Compressor Diffuser and Turbine Vane Assembly
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Solution Overview
Problem
Conventional gas turbine engines face challenges in reducing axial length due to the axial alignment of the combustor and compressor sections, which also leads to increased compressor size to compensate for bleed cooling air diverted to cool turbine vanes, reducing the air available for combustion.
Innovation Solution
The configuration integrates a compressor diffuser with a high-pressure turbine blade assembly, allowing the compressor air to be used for backface cooling of the turbine vanes, and positioning at least the final compressor stages radially inward from the combustor, thereby reducing the need for bleed cooling and shortening the engine length.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Device complexity
If the combustor is axially aligned with the compressor section, then the engine structure is simplified, but the axial length of the engine increases
Solution Approach 1:
The patent transitions from a conventional axial arrangement to a radial arrangement by positioning the compressor radially inward from the combustor. This dimensional change allows the compressor and combustor to occupy different radial zones rather than requiring sequential axial space, thereby reducing the overall axial length of the engine while maintaining structural simplicity.
2Temperature
If bleed air is diverted to cool turbine vanes, then the turbine vanes are effectively cooled, but the air available for combustion is reduced
Solution Approach 1:
The compressor serves a dual function by simultaneously cooling the turbine vanes through its radially outward flow and supplying air to the combustor. The compressed air flows radially outward through the turbine vane assembly, providing self-cooling to the vanes while the remaining air continues to the combustor, eliminating the need for separate bleed air extraction for cooling purposes.
3Quantity of substance
If more compressor size is used to compensate for bleed cooling air, then sufficient air is available for both cooling and combustion, but the compressor size and weight increase
Solution Approach 1:
The compressor is designed to perform multiple functions: compressing air for combustion, cooling the turbine vanes through radial flow, and potentially driving the fan. By making the compressor multi-functional, the system eliminates the need for additional compressors or larger compressor capacity that would otherwise be required to compensate for bleed air extraction, thereby reducing overall compressor size and weight.
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 reduces the axial length of the engine by approximately 6 inches, allows for a smaller compressor, and improves combustion efficiency by utilizing the heat picked up during cooling, while also minimizing the need for bleed air.
Implementation Method 1
a compressor section having an impeller configured to compress a core flow
Implementation Method 2
allowing the compressor air to be used for backface cooling of the turbine vanes
Implementation Method 3
the internal vane path has an increasing cross-sectional area extending in a direction of flow of the core flow
Data Source
Figure 1
Figure 2A
Figure 2B
AI summary
Gas turbine engines (200) include a compressor section (202) having an impeller (212) configured to compress a core flow (208). A combustor section (204) having a combustor (220) is arranged downstream from the compressor section (202) along a path of the core flow (208). A turbine section (206) is arranged downstream from the combustor section (204) along the path of the core flow (208), with the turbine section (206) having a plurality of first vanes (214) arranged at an outlet of the combustor (220) and at least one first vane (214) of the plurality of first vanes (214) includes an internal vane path. The path of the core flow (208), in a flow direction, passes through the impeller (212) of the compressor section (202), the internal vane path internal to the first vane (214), the combustor (220), and then between external surfaces of the plurality of first vanes (214) to enter the turbine section (206).