Gas Turbine Core Flange Layout for Higher Bending Resistance
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Solution Overview
Problem
Larger gas turbine engines face structural challenges due to increased bending loads, particularly exacerbated by fan outlet guide vanes moving rearward, which can lead to core bending and deformation, necessitating improved stiffness and load management.
Innovation Solution
The gas turbine engine design incorporates specific relative component positions and configurations, including a gearbox-driven fan with a reduction ratio, optimized flange and core casing arrangements, and strategically positioned fan outlet guide vanes to enhance engine core stiffness and manage bending loads effectively.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Productivity
If the fan size of a gas turbine engine is increased, then the thrust and efficiency of the engine is improved, but bending loads on the engine core are deleteriously increased
Solution Approach 1:
The patent changes the axial position parameter of the first flange connection along the engine core. By positioning the flange at a specific axial location (downstream of the low-pressure compressor and upstream of the high-pressure compressor), the bending moment distribution is optimized to reduce peak bending loads while maintaining the large fan diameter configuration
Solution Approach 2:
The patent transitions from considering only radial support to incorporating axial positioning of the flange connection. By adding the axial dimension to the support structure design, the bending loads are distributed more effectively along the engine core length, reducing the harmful bending moments caused by large fan diameter
2Adaptability or versatility
If the fan outlet guide vanes are moved rearward, then the engine design flexibility is improved, but core bending is exacerbated
Solution Approach 1:
The patent changes the axial position parameter of the first flange connection to compensate for the rearward movement of fan outlet guide vanes. By adjusting the flange position downstream, the support point is optimized to counteract the increased bending moment caused by the rearward OGV positioning
Solution Approach 2:
The patent applies preliminary anti-action by positioning the flange connection to preemptively counteract the bending loads generated by rearward OGV placement. The flange is positioned upstream of the OGVs to provide early structural support that offsets the bending effects before they propagate through the engine core
3Strength
If the engine core stiffness is increased, then the bending load resistance is improved, but the device complexity increases
Solution Approach 1:
The patent applies local quality by concentrating structural reinforcement at the specific location of the first flange connection rather than uniformly increasing core stiffness throughout. The flange and its supporting structure are locally optimized to provide maximum bending resistance at the critical section where bending moments are highest
Solution Approach 2:
The patent segments the engine core into distinct sections with different structural characteristics. The region around the first flange connection is reinforced with a segmented structure comprising the flange, support ribs, and connecting elements, while other regions maintain their original design, thus increasing stiffness only where needed
Data Source
Figure 1~2
Figure 3A~3B
Figure 4A~4C
AI summary
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising: a first compressor (14), a second compressor (15) and an outer core casing (70) surrounding the compressors. The gas turbine engine further comprises a fan (23), the fan comprising a plurality of fan blades. The outer core casing comprises: a first flange connection (60) having a first flange radius (104), wherein the first flange connection (60) is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor (14) and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor (15) and a ratio first flange radius/mass of each blade is equal to or less than 19.0 mm/lb.