Ceramic Matrix Composite Turbine Components in Gas Turbine Engines
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Solution Overview
Problem
Conventional gas turbine engines face challenges in minimizing fuel burn and weight due to mature design limitations, with cooling systems adding complexity, cost, and mass, while ceramic matrix composites offer thermal benefits but require careful optimization for efficient use.
Innovation Solution
Incorporating ceramic matrix composites in specific components of the gas turbine engine, such as the fan and turbines, with a reduction gearbox, to achieve a higher thrust-to-core efficiency ratio, optimizing their use between 1% to 15% of the total mass to balance thermal capability and cooling efficiency.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If ceramic matrix composites are used in turbine components, then thermal capability and temperature resistance are improved, but manufacturing complexity and cost increase
Solution Approach 1:
The patent applies ceramic matrix composites selectively in specific turbine components (such as turbine blades and vanes) where high temperature resistance is most critical, rather than throughout the entire turbine assembly. This localized application provides thermal capability where needed while avoiding unnecessary manufacturing complexity in other components.
Solution Approach 2:
The patent utilizes ceramic matrix composite materials that combine ceramic fibers in a matrix structure, creating a material with superior high-temperature properties compared to conventional metals. This composite structure enables the turbine components to withstand higher temperatures while managing the manufacturing challenges through specialized fabrication processes.
2Reliability
If cooling flow is increased to protect turbine components, then component durability is improved, but engine efficiency deteriorates
Solution Approach 1:
The patent changes the material parameter from conventional metal to ceramic matrix composite, which has inherently higher temperature resistance. This parameter change allows the turbine components to operate at higher temperatures without requiring proportional increases in cooling flow, thereby reducing the energy penalty associated with cooling while maintaining component durability.
3Loss of energy
If turbine entry temperature is increased to improve efficiency, then fuel burn reduction is improved, but component thermal stress increases
Solution Approach 1:
The patent employs ceramic matrix composite materials in turbine components to enable operation at higher turbine entry temperatures. These composite materials resist thermal stress and degradation better than conventional metals, allowing the engine to burn less fuel by operating at higher temperatures without compromising component integrity.
4Loss of energy
If cooling passages are made more intricate to improve cooling efficiency, then cooling effectiveness is improved, but manufacturing cost and complexity increase
Solution Approach 1:
The patent focuses cooling passages and thermal management features on specific high-stress regions of turbine components rather than uniformly throughout. This localized cooling approach improves cooling effectiveness where it is most needed while reducing the overall complexity and manufacturing cost compared to comprehensive intricate cooling throughout the entire component.
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 enhances engine efficiency by reducing cooling flow requirements, minimizing engine mass, and optimizing thermal management, leading to improved thrust-to-core efficiency and reduced fuel burn, while maintaining component durability.
Implementation Method 1
the reduced thermal conductivity of CMCs (compared to an equivalent metallic component) means that they may not be suitable in some other areas
Implementation Method 2
a fan driven by the power turbine through a geared architecture at a second speed lower than the first speed
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 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 the maximum net thrust of the engine at sea level is in the range of from 160 kN to 550kN and the normalized thrust is in the range of from 0.25 to 0.5 kN/kg.