Fan Blade Trailing Edge Gradient for Gas Turbine Thrust
Find Innovative SolutionsGenerate Solutions
Solution Overview
Problem
Gas turbine engines with larger fan diameters face challenges in delivering thrust at high angles of attack due to distorted inlet flow, leading to reduced performance and increased fuel burn, particularly because existing designs do not effectively manage the radial gradient of the fan trailing edge angle.
Innovation Solution
Implementing a radial gradient of the fan trailing edge angle between 0.4 and 0.8 of the span, with a specific metric M defined by the rate of change of the trailing edge angle, which can be between 65 and 150 degrees, to enhance thrust delivery and allow for shorter intake lengths, thereby reducing fuel burn.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Force
If larger fan diameters are used to increase thrust, then thrust capability is improved, but inlet flow distortion increases leading to reduced performance at high angles of attack
Solution Approach 1:
The patent applies local quality by varying the trailing edge angle specifically in the radial region between 0.4 and 0.8 of the blade span, while other regions maintain different characteristics. This localized modification optimizes flow management in the critical mid-span region where distortion effects are most pronounced, allowing larger fan diameters to deliver thrust without suffering from inlet flow distortion at high angles of attack.
Solution Approach 2:
The patent changes the geometric parameter of the trailing edge angle in a controlled manner across the radial span. By defining a specific radial gradient where the trailing edge angle increases from approximately 15-25 degrees at 0.4 span to 40-55 degrees at 0.8 span, the design transforms the uniform geometry into a variable geometry that adapts to the flow conditions at different radial positions, thereby resolving the contradiction between thrust generation and flow distortion.
2Ease of manufacture
If conventional fan blade designs are used, then manufacturing simplicity is maintained, but thrust delivery at high angles of attack is reduced
Solution Approach 1:
Rather than redesigning the entire blade geometry, the patent applies a localized modification only to the trailing edge angle in the radial region between 0.4 and 0.8 span. This approach maintains manufacturing simplicity for the overall blade structure while introducing a targeted geometric feature that significantly improves thrust delivery at high angles of attack.
Solution Approach 2:
The blade design is segmented into different radial regions with distinct trailing edge angle characteristics. The region between 0.4 and 0.8 span has an increased trailing edge angle gradient, while other regions maintain conventional angles. This segmentation allows the critical mid-span region to be optimized for high angle of attack performance without complicating the entire blade design.
3Object-affected harmful factors
If longer intake lengths are used to reduce flow distortion, then inlet flow quality is improved, but overall engine length and fuel burn increase
Solution Approach 1:
The patent converts the potentially harmful effect of larger fan diameters (which cause flow distortion) into a benefit by designing a trailing edge geometry that actively manages the distorted flow. The increased trailing edge angle in the mid-span region redirects the distorted inlet flow in a way that maintains attachment and thrust generation, allowing shorter intake lengths to achieve the same flow quality that would otherwise require longer intakes.
Solution Approach 2:
By changing the trailing edge angle parameter in the critical radial region, the patent alters the flow turning characteristics to compensate for inlet distortion. This parameter change enables the engine to tolerate shorter intake lengths while still maintaining good flow quality at the fan inlet, thereby reducing overall engine length and associated fuel burn.
Data Source
Figure 1~2
Figure 3~4a
Figure 4b~5
AI summary
The disclosure relates to a gas turbine engine (10) for an aircraft comprising: 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, the fan comprising a plurality of fan blades mounted for rotation about an engine axis, each blade of the plurality of fan blades having a leading edge and a trailing edge extending across a span of an airflow duct from a blade root to a blade tip; and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein the trailing edge of each blade is characterized by a metric M defined as a rate of change of an angle of the trailing edge between 0.4 and 0.8 of the span divided by an area averaged trailing edge angle, M being not less than around 4.