Turbine Vane Blade Count and Combustion Uniformity
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Solution Overview
Problem
High inlet temperature and chemical species non-uniformities at the exit of a combustor in gas turbine engines, particularly in high fuel-to-air ratio designs, compromise turbine durability, necessitating conservative cooling designs to prevent damage.
Innovation Solution
A gas turbine engine design with a turbine section having a greater number of vanes than blades, where the combustion reaction is completed before the gas flow reaches the downstream blades, utilizing cooling air to react with combustion products and film-cooling the vanes to reduce temperature and chemical species non-uniformities, thereby reducing cooling requirements and enhancing mixing and aerodynamic efficiency.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If conservative cooling designs are used to prevent turbine damage from temperature non-uniformities, then turbine durability is improved, but device complexity and weight increase
Solution Approach 1:
The combustor is designed to complete the combustion reaction before the gas enters the turbine section, eliminating temperature and chemical species non-uniformities in advance. This preliminary action removes the need for complex conservative cooling designs while maintaining turbine durability.
Solution Approach 2:
The harmful temperature and chemical species non-uniformities are extracted and eliminated through optimized combustor design, separating the combustion completion function from the turbine section. This allows the turbine to operate with uniform conditions without requiring complex cooling systems.
2Reliability
If conservative cooling designs are used to prevent turbine damage from temperature non-uniformities, then turbine durability is improved, but weight increases
Solution Approach 1:
Combustion is completed in the combustor before gas enters the turbine, preliminarily eliminating the need for heavy cooling systems. This approach maintains turbine durability while significantly reducing turbine weight by removing unnecessary cooling infrastructure.
Solution Approach 2:
The high fuel-to-air ratio combustion, which initially creates harmful non-uniformities, is optimized to complete combustion before turbine entry. This converts the potential harm into a benefit by allowing higher fuel efficiency while maintaining turbine durability through uniform gas conditions.
3Use of energy by moving object
If high fuel-to-air ratio combustion is used, then energy efficiency is improved, but temperature and chemical species non-uniformities worsen
Solution Approach 1:
The combustor is designed to complete the combustion reaction before gas enters the turbine section. This preliminary combustion completion ensures uniform temperature and chemical species distribution while maintaining high fuel-to-air ratio for improved energy efficiency.
Solution Approach 2:
The combustor design parameters are optimized to achieve complete combustion at high fuel-to-air ratios. By adjusting combustion chamber geometry, mixing characteristics, and residence time, uniform gas composition is achieved while maintaining high energy efficiency.
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 design reduces the need for conservative cooling, achieves weight reduction, and improves aerodynamic efficiency by dissipating temperature and chemical species non-uniformities, leading to a more efficient and lightweight gas turbine engine.
Implementation Method 1
utilizing cooling air to react with combustion products and film-cooling the vanes to reduce temperature and chemical species non-uniformities
Implementation Method 2
completing a combustion reaction along a gas flow path prior to a plane defined by downstream portions of the first set of blades
Data Source
AI summary
An exemplary gas turbine engine includes a turbine section operative to impart rotational energy to a compressor section. The turbine section includes at least a low-pressure turbine and a high-pressure turbine, and a number of stages in the low pressure turbine is from three to five.