Geared Turbofan Engine Modular Efficiency Design
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Solution Overview
Problem
Current gas turbine engines, despite improvements in geared architectures, continue to seek enhancements in thermal, transfer, and propulsive efficiencies to optimize engine performance.
Innovation Solution
A turbine engine design featuring a fan section with a leading edge to hub ratio of less than 0.34 and a speed change mechanism with a gear ratio greater than 2.6:1, along with a bypass ratio greater than 8, and a multi-stage compression system, including a low fan pressure ratio and a fan drive turbine with multiple stages, optimized for high-bypass geared aircraft engines.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Productivity
If a geared architecture with epicyclical gear assembly is used to drive the fan section at reduced speed, then propulsive efficiency is improved, but device complexity increases
Solution Approach 1:
The patent replaces the traditional mechanical geared architecture with a direct-drive configuration where the fan section is driven directly by the low pressure turbine through a common shaft. This eliminates the epicyclical gear assembly and its associated complexity while maintaining the ability to optimize fan speed for propulsive efficiency through alternative design approaches such as optimized fan blade geometry and hub-to-tip ratio
Solution Approach 2:
The patent segments the compression system into distinct low pressure and high pressure compressor sections with independent shafts, allowing each section to operate at its optimal speed. The fan section is integrated with the low pressure turbine on a common shaft, creating a modular architecture that simplifies the overall system while maintaining performance
2Productivity
If fan diameter is increased and fan pressure rise is reduced to improve propulsive efficiency, then thrust specific fuel consumption improves, but device complexity increases
Solution Approach 1:
The patent optimizes the fan design by changing key parameters including hub-to-tip ratio (less than 0.34), fan blade geometry, and operating speed. These parameter changes allow the fan to operate at reduced pressure rise (less than 1.6) while maintaining high bypass ratio (greater than 8), achieving improved thrust specific fuel consumption without requiring complex geared mechanisms
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 achieves a unique combination of component efficiencies leading to high engine efficiency, particularly in thrust specific fuel consumption, benefiting commercial engines and single-aisle aircraft with reduced engine weight and increased power density.
Implementation Method 1
a speed change mechanism with gear ratio greater than about 2.6 to 1
Implementation Method 2
the fan section provides a bypass ratio greater than about 8
Implementation Method 3
A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section
Implementation Method 4
The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section
Data Source
AI summary
A turbine engine includes a first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section.