Compressor Airfoil Sweep and Dihedral Optimization
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Solution Overview
Problem
In gas turbine engines, the high-speed compressor stages experience transonic flow conditions leading to induced shock losses, and rear stages have small blade heights with large tip clearance, resulting in complex tip flow structures that reduce rotor performance and stall margin due to secondary flow migration and boundary layer buildup.
Innovation Solution
The design of compressor airfoils with forward sweep and lean angles, optimized sweep and dihedral angle distributions, and thickness variations along the span to reduce shock and boundary layer interactions, minimize tip clearance effects, and enhance flow efficiency and stall margin.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Ease of operation
If large tip clearance is used in rear compressor stages, then mechanical limitations are satisfied, but rotor performance and stall margin are reduced due to complex tip flow structure
Solution Approach 1:
The patent applies different geometric characteristics to different regions of the airfoil. The leading edge region has specific sweep and dihedral angles to control shock formation, while the tip region has optimized curvature and thickness distribution to manage tip leakage flow and boundary layer interactions. This local optimization resolves the contradiction by addressing tip clearance effects specifically at the tip region without compromising overall airfoil performance.
Solution Approach 2:
The airfoil geometry incorporates dynamic flow control through its geometric design. The leading edge sweep angle and dihedral angle distribution are optimized to dynamically adjust shock position and strength across different operating conditions. The tip region geometry dynamically manages the interaction between tip clearance vortex and boundary layer, adapting to varying flow conditions to maintain performance.
2Loss of energy
If forward sweep and lean angles are increased, then shock and boundary layer interactions are reduced, but airfoil geometric complexity increases
Solution Approach 1:
The patent systematically varies geometric parameters including leading edge sweep angle, dihedral angle, and tip region curvature along the span. These parameter changes are optimized to reduce shock strength and boundary layer interactions while maintaining manufacturability. The gradual variation of these parameters along the span achieves flow control without excessive geometric complexity.
Solution Approach 2:
The patent addresses two-dimensional shock and boundary layer problems by introducing three-dimensional geometric features. The leading edge sweep and dihedral angles create three-dimensional flow patterns that reduce adverse two-dimensional interactions. The tip region geometry uses spanwise variation to control three-dimensional tip leakage flows, resolving complex interactions through dimensional transition.
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 solution improves rotor performance by reducing shock and boundary layer accumulation, increasing stall margin, and optimizing flow distribution, leading to enhanced efficiency and reduced sensitivity to tip clearance and secondary flow migration.
Implementation Method 1
at high speed conditions the front stages usually have transonic flow conditions and carry large induced shock losses
Implementation Method 2
because of rotor centrifugal effects, there may be a migration of secondary flow along blade surface from the hub to tip section of the blade
Implementation Method 3
resulting in a thick tip boundary layer build up
Implementation Method 4
Interaction between leading edge shocks, tip clearance vortex, blade/shroud surface boundary layer results in complex tip flow structure
Data Source
Figure 1
Figure 2
Figure 3A
AI summary
A compressor airfoil (20) in a gas turbine engine is presented. Opposed pressure and suction sides (32,34) are joined together at chordally opposite leading and trailing edges (36,38). The pressure and suction sides (32,34) extend in a span direction from a root (25) to a tip (27) of the airfoil (20). A leading edge sweep angle (α) is defined relative to a tangent to the airfoil (20) and flow velocity vector at a point on the leading edge (36). A leading edge dihedral angle (β) is defined relative to the tangent to the airfoil (20) and a vertical at the point on the leading edge (36). A ratio of the leading edge sweep angle (α) to the leading edge dihedral angle (β) is smaller than 1. A method of forming such airfoil is also presented.