Compressor Blade Tip Solidity Reduction
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Solution Overview
Problem
The sensitivity of compressor blades in gas turbine engines to the clearance between the blade tip and the shroud reduces efficiency and overall performance due to inefficient airflow through this small volume of air.
Innovation Solution
The design of the compressor blade airfoil features a reduction in local chord from 75% span to the tip, decreasing leading and trailing edge sweep and dihedral angles from 50% span to the tip, resulting in a local reduction in solidity and improved airflow.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Productivity
If the compressor blade tip clearance is reduced to improve efficiency, then airflow efficiency improves, but the blade becomes more sensitive to clearance variations and manufacturing precision requirements increase
Solution Approach 1:
The patent applies local quality by varying the airfoil geometry parameters (chord length, sweep angle, dihedral angle) specifically in the tip region (from 75% span to tip) while maintaining conventional geometry in the root region. This localized modification optimizes the tip clearance flow without affecting the overall blade structure, addressing the efficiency improvement while concentrating the design changes where they are most needed.
Solution Approach 2:
The patent employs parameter changes by systematically modifying geometric parameters of the airfoil in the tip region: reducing local chord from 75% span to tip, decreasing leading and trailing edge sweep angles from 50% span to tip, and decreasing leading and trailing edge dihedral angles from 50% span to tip. These parameter variations optimize airflow through the tip clearance region.
2Productivity
If the local chord is reduced from 75% span to tip to reduce solidity, then airflow efficiency improves, but the blade structural strength may be compromised
Solution Approach 1:
The chord reduction is applied locally only in the tip region (75% span to tip) where it optimizes airflow and reduces adverse effects from tip clearance, while the root region (0% to 75% span) maintains conventional chord dimensions to preserve structural strength and loading characteristics.
Solution Approach 2:
The blade is effectively segmented into two regions: a root region (0% to 75% span) with conventional geometry for structural integrity, and a tip region (75% to 100% span) with modified geometry for airflow optimization. This segmentation allows each region to be optimized for its primary function.
3Productivity
If the sweep and dihedral angles are decreased from 50% span to tip, then airfoil performance improves, but the manufacturing complexity increases
Solution Approach 1:
The sweep and dihedral angle modifications are applied locally in the tip region (50% span to tip) where they improve airfoil performance, while the root region (0% to 50% span) maintains conventional angles. This localized approach minimizes the overall geometric complexity while achieving performance benefits.
Solution Approach 2:
Instead of modifying the entire blade span, the patent applies sweep and dihedral angle reductions partially, only in the tip region where they are most beneficial for airflow. This partial action reduces manufacturing complexity compared to a full-span modification while maintaining performance improvements.
Data Source
AI summary
An airfoil for a compressor blade of a gas turbine engine has a chord between a leading edge and a trailing edge and a span between a root and a tip. The airfoil can further include a reduction in local chord from about 75% span to the tip for about a 5% reduction in local solidity. The airfoil can have decreasing sweep angles for the leading edge and the trailing edge from 50% span to the tip and can have decreasing leading and trailing edge dihedral angles from 50% span to the tip.