Compressor Rotor Hub Blade Geometry for Surge Margin
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Solution Overview
Problem
Gas turbines face challenges in balancing efficiency, emissions, noise levels, and stability across varying operating speeds, particularly in maintaining stability and operability margins when increasing aerodynamic duty or work, which can lead to issues like stall/surge phenomena and reduced operating margins.
Innovation Solution
A rotor design for a compressor with thicker blades at the hub and a unique endwall profile that increases the hub-thickness to chord ratio, allowing for higher aerodynamic duty while maintaining stability and operability margins, achieved through a specific blade geometry and endwall profile that counteracts increased boundary layer diffusion.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Power
If the aerodynamic duty or work of the rotor is increased, then the efficiency and power output are improved, but the stability and operability margins deteriorate, leading to stall/surge phenomena
Solution Approach 1:
The endwall profile is specifically modified at the hub region with a radius change between 15-30%, creating localized flow acceleration where it is most needed to counteract the adverse effects of increased blade thickness. This local modification allows the rotor to maintain stability margins while operating at higher aerodynamic duty.
Solution Approach 2:
The invention changes the geometric parameters of the endwall profile, specifically the hub radius change between 15-30%, to optimize the flow characteristics. This parameter modification creates relative acceleration that compensates for the boundary layer diffusion caused by thicker blades, enabling higher aerodynamic duty without sacrificing stability.
2Strength
If thicker blades are used at the hub to increase structural strength, then the mechanical strength and durability are improved, but the aerodynamic performance deteriorates due to increased boundary layer diffusion
Solution Approach 1:
The invention converts the harmful effect of increased boundary layer diffusion caused by thicker blades into a benefit by pairing it with an endwall profile that generates relative acceleration. This acceleration compensates for the diffusion, allowing the thick blades to provide structural strength without sacrificing aerodynamic performance.
Solution Approach 2:
The endwall profile is made asymmetric with a specific radius change concentrated at the hub region (15-30% change), creating non-uniform flow acceleration that specifically addresses the aerodynamic penalty at the thickened blade section while maintaining overall rotor performance.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
The design enables increased aerodynamic duty and work without compromising stability and operability, achieving a balance between efficiency and performance by offsetting the aerodynamic penalty of thicker blades with relative acceleration from the endwall profile, thus enhancing the compressor's performance and surge margin.
Implementation Method 1
offsetting the aerodynamic penalty of thicker blades with relative acceleration from the endwall profile
Implementation Method 2
relative acceleration from the endwall profile
Data Source
Figure 1A~1C
Figure 2A~4
Figure 5
AI summary
A rotor, for a compressor of a gas turbine, comprising a rotatable support for rotation about an axis of rotation and a plurality of blades. Each blade comprising a hub, a leading edge and a trailing edge and a chord is defined between the leading edge and the trailing edge. Each of the blades extends from its hub away from the rotatable support and at least one of the blades has a hub-thickness to chord ratio greater than 10 percent. The leading edge of the at least one of the blades at its hub is positioned at a leading-edge-hub-radius from a position of the axis of rotation and the trailing edge of the at least one of the blades is positioned at its hub at a trailing-edge-hub-radius from the position of the axis of rotation. The trailing-edge-hub-radius is greater than the leading-edge-hub-radius.