Turbomachine Blade Stacking Axis Curvature
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Solution Overview
Problem
The geometric contour of rotating blades in turbomachines, particularly gas turbines, influences loading and stressing, and existing designs fail to effectively manage fluid forces and reduce stress concentrations, especially at the leading edge, which can lead to operational robustness issues and vulnerability to Foreign Object Damage (FOD).
Innovation Solution
The rotating blades are designed with a specific radial stacking axis course that combines a positive linear x-lean (sweep) and negative y-lean, incorporating a superimposed sinusoidal component, which deviates from a purely linear course by a controlled amount, reducing stress concentrations and improving robustness by optimizing the blade's geometric profile.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If a conventional linear stacking axis course is used, then the blade design is simple, but the stress concentration at the leading edge increases and robustness against FOD decreases
Solution Approach 1:
The patent applies curvature to the stacking axis course by superimposing a sinusoidal function on the linear progression, creating a curved rather than straight path for the blade element centers. This curved geometry redistributes stresses along the blade span, reducing stress concentration at the leading edge and improving robustness against foreign object damage, while maintaining manufacturability through defined mathematical functions.
2Strength
If the stacking axis follows a purely linear course, then manufacturing is easier, but the loading and stressing of the blade root and leading edge cannot be optimized
Solution Approach 1:
The patent modifies the geometric parameters of the blade by introducing a sinusoidal component to the stacking axis course, changing the linear parameter progression into a non-linear function. This parameter change optimizes the stress distribution along the blade span, reducing peak stresses at critical locations while the mathematical definition ensures the profile remains manufacturable.
Data Source
AI summary
The present invention relates to a rotating blade for a turbomachine, in particular a compressor stage or a turbine stage of a gas turbine, particularly of an aircraft engine, having a blade element for deflecting the flow, with a pressure side and a suction side, these sides being joined by a leading edge and a trailing edge, wherein a stacking axis of the blade element, in the radial direction over a radius r from a root of a blade element at r=0 to a tip of a blade element at r=H, has a course x(r) in a first downstream direction perpendicular to the radial direction and parallel to a principal axis of the turbomachine and has a course y(r) in a second direction perpendicular to the radial direction and to the first direction.