Geared Gas Turbine Compression Ratio Layout for Rotor Bow Control
Find Innovative SolutionsGenerate Solutions
Solution Overview
Problem
Modern gas turbine engines face challenges in optimizing overall efficiency without compromising engine operability and maintenance requirements, particularly due to issues like rotor bow and inefficient compression systems.
Innovation Solution
The design incorporates a gearbox to drive the fan at a lower rotational speed than the core shaft, with specific compression system blade ratios, core compressor pressure ratios, and compression system radius ratios within defined ranges, optimizing the balance between thermal, propulsive, and installation efficiencies.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Productivity
If the fan blade height is increased to improve propulsive efficiency, then the compression system becomes excessively long, but this leads to adverse impacts on aircraft installation and increased rotor bow risk
Solution Approach 1:
The patent applies parameter changes by optimizing the compression system blade ratio (CSBR) to a specific range (45-95) and core compressor aspect ratio (CCAR) to a specific range (2.0-3.5). These parameter optimizations allow the engine to achieve high propulsive efficiency with a larger fan blade height while maintaining an acceptable compression system length, thus resolving the contradiction between propulsive efficiency and installation flexibility.
2Adaptability or versatility
If the core compressor aspect ratio is increased to reduce compression system length, then installation flexibility improves, but propulsive efficiency decreases
Solution Approach 1:
The patent optimizes the core compressor aspect ratio (CCAR) to a specific range (2.0-3.5) and compression system blade ratio (CSBR) to a specific range (45-95). These parameter changes enable the compression system to be sufficiently compact for flexible aircraft installation while maintaining the fan blade height necessary for high propulsive efficiency, thus resolving the contradiction between installation flexibility and propulsive efficiency.
3Loss of energy
If the fan blade height is increased to improve propulsive efficiency, then thermal efficiency may improve, but the compression system length increases excessively
Solution Approach 1:
The patent applies parameter changes by optimizing both the compression system blade ratio (CSBR) to a range of 45-95 and the core compressor aspect ratio (CCAR) to a range of 2.0-3.5. These coordinated parameter optimizations allow the engine to achieve high thermal efficiency through an optimized compression ratio while maintaining a compact compression system length, thus resolving the contradiction between thermal efficiency and compression system length.
4Ease of manufacture
If the compression system is made more compact to improve installation, then manufacturing becomes easier, but compression efficiency decreases
Solution Approach 1:
The patent optimizes the core compressor aspect ratio (CCAR) to a specific range (2.0-3.5) and compression system blade ratio (CSBR) to a specific range (45-95). These parameter optimizations enable the compression system to achieve a compact design that is easier to install while maintaining sufficient compression efficiency through the optimized geometric parameters, thus resolving the contradiction between installation ease and compression efficiency.
Data Source
AI summary
A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.