Compressor Rotor Aerodynamic Mistuning for Supersonic Flutter Mitigation
Find Innovative SolutionsGenerate Solutions
Solution Overview
Problem
Gas turbine compressor rotors experience aerodynamic instability issues such as supersonic flutter and resonant stresses, particularly in high-speed regimes, which existing technologies have not adequately addressed, leading to operational barriers and potential damage from excessive stress loads.
Innovation Solution
The implementation of a compressor rotor design featuring alternating blade types with leading edge tip cutbacks and pressure side tip pockets, which generate different shock patterns and aerodynamically mistune the blades, mitigating supersonic flutter and resonant stresses by altering the aerodynamic properties and reducing chord length at the blade tips.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Speed
If compressor rotors operate in high-speed regimes to improve productivity, then speed increases, but supersonic flutter and resonant stresses increase causing aerodynamic instability
Solution Approach 1:
The patent applies local quality by creating different blade types with distinct aerodynamic properties. Specifically, alternate blades have different airfoil cross-sections, leading edge radii, and trailing edge geometries, causing them to generate different shock patterns at supersonic speeds. This local differentiation allows the rotor to maintain aerodynamic stability across the entire blade row while operating at high speeds.
Solution Approach 2:
The patent employs asymmetry through the use of non-uniform blade designs where adjacent blades have asymmetric aerodynamic characteristics. The alternate blades have different airfoil shapes, leading edge radii, and trailing edge angles, creating asymmetric shock wave patterns that prevent resonant coupling between blades and mitigate supersonic flutter.
2Reliability
If blade chord length is reduced at the tip to reduce supersonic flutter, then aerodynamic instability decreases, but blade structural strength may be compromised
Solution Approach 1:
The patent applies local quality by implementing different airfoil cross-sections at different radial positions along the blade span. The blade root maintains a larger chord length for structural strength, while the blade tip features a reduced chord length with optimized airfoil geometry to generate favorable shock patterns and reduce supersonic flutter, thus optimizing both strength and flutter mitigation.
Solution Approach 2:
The patent utilizes parameter changes by varying multiple aerodynamic parameters along the blade span, including airfoil cross-section, leading edge radius, trailing edge angle, and chord length. These parameter variations are designed to optimize the shock wave patterns generated by each blade section, reducing supersonic flutter while maintaining adequate structural strength through careful parameter selection and distribution.
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
This design effectively reduces aerodynamic instability and resonant stresses by creating distinct shock patterns and modifying the aerodynamic damping of the blades, thereby preventing supersonic flutter and associated stress issues within the compressor rotor.
Implementation Method 1
each of the plurality of compressor blades having an airfoil extending radially outward from the hub to a blade tip... generating different shock patterns between adjacent ones of the two or more different blade types when the gas turbine compressor operates in supersonic flow regimes
Data Source
AI summary
The gas turbine compressor for an aircraft gas turbine engine includes a compressor rotor having a plurality of compressor blades circumferentially distributed around a hub. Each of the compressor blades has an airfoil extending radially outward from the hub to a blade tip. A circumferential row of the compressor blades includes two or more different blade types, at least one modified blade of the two or more different blade types having means for generating different shock patterns between adjacent ones of the two or more different blade types when the gas turbine compressor operates in supersonic flow regimes. The means for generating different shock patterns on the modified blade aerodynamically mistune the two or more different blade types.