Gas Turbine Engine Area Ratio Optimization
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Solution Overview
Problem
Scaling up gas turbine engine components to increase power and efficiency often results in increased drag and installation challenges, such as larger nacelle drag and difficulty in mounting beneath an aircraft wing, due to proportional scaling issues.
Innovation Solution
The design focuses on specific relative component dimensions, including an engine area ratio of 1.7 to 3, a bypass to core nozzle pressure ratio of 1.1 to 2, and a gearbox location ratio of 0.19 to 0.45, to optimize engine size and mounting efficiency while reducing drag and improving installation.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Productivity
If the overall size of a gas turbine engine is increased to provide improved propulsive efficiency, then the fan face area increases, but the overall drag produced by the larger nacelle increases and installation becomes more difficult
Solution Approach 1:
The patent applies parameter changes by optimizing the engine area ratio (fan face area to product of turbine diameter and core length) to fall within a specific range of 1.7 to 3. This dimensional parameter optimization allows the engine to achieve improved propulsive efficiency through increased fan face area while controlling the overall drag by maintaining appropriate proportions between fan size and core dimensions, thereby resolving the contradiction between efficiency improvement and drag reduction.
2Power
If the components of the engine are scaled proportionally to increase power output, then the power and thrust increase, but the nacelle drag increases and the engine may not fit to be mounted beneath the wing
Solution Approach 1:
The patent employs parameter changes by defining specific dimensional relationships through the engine area ratio parameter. Rather than simple proportional scaling, the patent optimizes the ratio of fan face area to the product of turbine diameter and core length within the range of 1.7 to 3. This allows power output to be increased through larger fan face area while maintaining installation feasibility by controlling the relative dimensions of core components, thus resolving the contradiction between power increase and installation ease.
Solution Approach 2:
The patent transitions from one-dimensional proportional scaling to multi-dimensional optimization by introducing the engine area ratio as a composite parameter that relates fan face area (two-dimensional) to core dimensions (linear measurements). This dimensional transformation enables independent optimization of power-generating components (fan) from installation-critical components (core dimensions), allowing power output to increase while maintaining compatibility with wing mounting constraints.
3Power
If the turbine diameter and core length are increased to provide higher power output, then the power increases, but the effective engine area in axial plane increases making close-coupled installation more difficult
Solution Approach 1:
The patent applies parameter changes by optimizing the engine area ratio to fall within the range of 1.7 to 3. This optimization allows the fan face area to be increased for higher power output while controlling the product of turbine diameter and core length to limit the engine's axial plane footprint. By balancing these dimensional parameters, the patent enables close-coupled installation (reducing the moment applied to the wing) while maintaining power output improvements, thereby resolving the contradiction between power increase and reduced engine area.
Data Source
AI summary
A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.