Ceramic Matrix Composite Turbine Cooling Flow Reduction
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Solution Overview
Problem
Conventional gas turbine engines face challenges in reducing fuel burn and weight due to mature design, with cooling systems adding complexity and cost, and ceramic matrix composites (CMCs) not being suitable for all temperature ranges, requiring intricate cooling passages and additional mass.
Innovation Solution
Optimizing the use of ceramic matrix composites (CMCs) in specific ranges within the turbine, combining with metallic components for effective thermal management, and employing a geared architecture to enhance efficiency and reduce cooling flow requirements.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If cooling systems are added to protect metal turbine components from high temperatures, then component reliability is improved, but device complexity and weight increase
Solution Approach 1:
The patent changes the material parameter from metal to ceramic matrix composite, which fundamentally alters the temperature resistance capability. This material parameter change eliminates the need for cooling systems while maintaining component reliability at high temperatures
Solution Approach 2:
The patent extracts and removes the cooling system from the turbine component design. By using CMC materials that inherently resist high temperatures, the cooling system is completely taken out of the design, reducing device complexity and weight
2Ease of manufacture
If metal turbine components are used in high temperature zones, then manufacturing maturity is maintained, but fuel burn increases due to cooling requirements
Solution Approach 1:
The patent changes the material parameter from metal to ceramic matrix composite, enabling the components to operate at higher temperatures without cooling. This parameter change directly reduces fuel burn by eliminating the energy loss associated with cooling metal components
3Temperature
If ceramic matrix composites are used throughout the entire turbine, then temperature resistance is improved, but manufacturing difficulty increases due to material limitations in certain temperature ranges
Solution Approach 1:
The patent applies local quality by using CMC materials selectively in specific high-temperature zones (first turbine and second turbine) while using other materials in regions where CMC manufacturing is not yet成熟. This localized application optimizes both temperature resistance and manufacturing feasibility
4Temperature
If cooling passages are added to turbine components, then thermal management is improved, but mass increases due to additional structure
Solution Approach 1:
The patent extracts and removes the cooling passages from the turbine component design. By using CMC materials that inherently resist high temperatures, the cooling passages are completely taken out, reducing component mass while maintaining thermal management through material properties rather than structural cooling features
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 approach reduces the cooling flow requirements, increases engine efficiency, and minimizes mass while allowing for higher turbine entry temperatures, leading to improved performance and reduced engine weight.
Implementation Method 1
the first and second turbines may comprise ceramic matrix composites... allowing for higher turbine entry temperatures
Implementation Method 2
ceramic matrix composites (CMCs) not being suitable for all temperature ranges
Data Source
Figure 1~2
Figure 3
Figure 4
AI summary
A gas turbine engine (10) for an aircraft with an engine core (11) comprising a first turbine (19), a first compressor (14), and a first core shaft (26) connecting the first turbine to the first compressor, a second turbine (17), a combustor, a second compressor (15), and a second core shaft (27) connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft, the gas turbine engine further comprising a fan (23) comprising a plurality of fan blades and a gearbox (30) that receives an input from the first core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, wherein part (C) of the flow (A) that enters the engine core bypasses the combustor and is used as turbine cooling flow to cool the turbine, a cooling flow requirement is defined as the ratio of the mass flow rate of the turbine cooling flow (C) to the mass flow rate of the flow entering the engine core (A) at cruise conditions,a turbine entry temperature is defined as the temperature (K) at the inlet to the most axially upstream turbine rotor in the gas turbine engine at a maximum power condition of the gas turbine engine and the cooling efficiency ratio, defined as the ratio between the turbine entry temperature and the cooling flow requirement, is in the range of from 8000 to 20000 K.