Geared Turbofan Layout for Lower Drag and Closer Wing Mounting
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Solution Overview
Problem
Scaling up gas turbine engine components does not proportionally increase power or efficiency, leading to issues like increased drag and installation challenges, particularly in reducing drag and fitting the engine beneath an aircraft wing without adjusting dimensions.
Innovation Solution
A gas turbine engine design with a specific engine area ratio of the fan face area to turbine diameter and core length, ranging from 1.7 to 3, and a bypass to core exhaust nozzle pressure ratio between 1.1 and 2, which allows for improved propulsive efficiency and closer wing mounting, reducing drag and installation difficulties.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Power
If the overall size of a gas turbine engine is increased, then power output is improved, but drag increases and installation becomes more difficult
Solution Approach 1:
The patent applies parameter changes by optimizing the engine area ratio (fan face area to turbine diameter and core length) to a specific range of 1.7 to 3.0, and the bypass to core exhaust nozzle pressure ratio to 1.1 to 2.0. These parameter optimizations enable the engine to achieve improved propulsive efficiency and reduced drag while maintaining high power output, resolving the contradiction between power increase and drag reduction.
2Power
If the overall size of a gas turbine engine is increased, then power output is improved, but installation difficulty increases
Solution Approach 1:
The patent resolves installation difficulty by optimizing key dimensional parameters, specifically the engine area ratio and bypass to core pressure ratio. These optimized parameters enable closer wing mounting, which reduces the moment applied to the wing by the engine mass and facilitates installation on aircraft with constrained wing space, thereby improving ease of installation while maintaining high power output.
3Loss of energy
If the fan size is increased, then propulsive efficiency is improved, but engine area increases making installation difficult
Solution Approach 1:
The patent applies parameter changes by defining an optimized engine area ratio range of 1.7 to 3.0, which establishes a specific relationship between fan face area and turbine dimensions. This ratio optimization allows the engine to achieve improved propulsive efficiency through larger fan size while controlling the overall engine area to facilitate installation, resolving the contradiction between propulsive efficiency improvement and installation feasibility.
Data Source
AI summary
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub (66); and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine (10) has an engine length (110) and a gearbox location (112) relative to a forward region of the fan (23), and a gearbox location ratio of:gearbox location/engine lengthis in a range from 0.19 to 0.45.