Compressor Stator Vane Forward Sweep for Flow Loss Reduction
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Solution Overview
Problem
In gas turbine engines, non-uniform inlet pressure and high flow turning lead to secondary flow structures along stator vanes, causing momentum deficits and increased losses, which negatively affect compressor performance and stall range.
Innovation Solution
The compressor design features a radially inner portion with aerodynamic dihedral at the leading edge, a maximum forward sweep between 5% and 30% of the span, and a trailing edge belly region extending forwardly, redistributing high energy flow to improve airflow and reduce losses.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Loss of energy
If conventional stator vane design is used, then manufacturing is simpler, but secondary flow extends along radial length causing momentum deficits and higher losses
Solution Approach 1:
The patent applies local quality by implementing forward sweep specifically in the radially inner portion of the stator vane (between 5% and 40% of span) while keeping other regions conventional. This localized modification targets the specific region where secondary flow causes momentum deficits, improving flow energy in that critical area without unnecessarily complicating the entire vane structure.
Solution Approach 2:
The stator vane is segmented into distinct radial regions: a radially inner portion (5-40% span) with forward sweep and aerodynamic dihedral, and a radially outer portion with conventional geometry. This segmentation allows different geometric features to address different flow problems in different radial zones, reducing overall flow losses while managing complexity through regional specialization.
2Reliability
If forward sweep is increased to reduce secondary flow, then flow energy is improved, but vane geometry becomes more complex
Solution Approach 1:
Forward sweep is applied locally to the radially inner portion (5-40% span) rather than the entire vane span. This localized application improves reliability by addressing the specific region where secondary flow causes momentum deficits, while limiting geometric complexity to only the necessary portion of the vane.
Solution Approach 2:
The vane is divided into segments with different sweep characteristics: the radially inner portion has forward sweep (10-30 degrees) while the radially outer portion maintains conventional geometry. This segmentation improves compressor reliability in the critical inner region without unnecessarily increasing overall geometric complexity.
3Loss of energy
If aerodynamic dihedral is added at leading edge, then flow redistribution is improved, but manufacturing precision requirements increase
Solution Approach 1:
Aerodynamic dihedral is implemented locally at the leading edge of the radially inner portion, concentrated near the root area. This local application improves flow redistribution where it is most needed while limiting the extent of precision requirements to a specific region rather than the entire vane.
Solution Approach 2:
The leading edge is segmented into a radially inner portion with aerodynamic dihedral and a radially outer portion without dihedral. This segmentation improves flow deviation in the critical inner region while reducing overall manufacturing precision requirements compared to applying dihedral across the entire span.
4Loss of energy
If belly region is extended forwardly, then trailing edge flow is improved, but chord length effectively increases
Solution Approach 1:
The belly region is extended forwardly locally at the trailing edge of the radially inner portion (5-30% of chord length from trailing edge). This localized extension improves trailing edge flow quality where momentum deficits occur without significantly increasing the overall chord length of the airfoil.
Solution Approach 2:
The airfoil is segmented such that the radially inner portion has an extended belly region at the trailing edge while the radially outer portion maintains conventional geometry. This segmentation improves trailing edge flow and reduces energy loss without substantially increasing the overall chord length.
Data Source
Figure 1
Figure 2A~2C
Figure 3A~3B
AI summary
The present compressor (13) of a gas turbine engine (10) includes a rotor (18) and a stator (17) located immediately downstream of the rotor (18). The stator (17) comprises a plurality of stator vanes (20; 120; 320) each having an airfoil (21; 121) extending along a generally radial span (S) between a root (22; 122; 322) and a tip (23; 123) of the airfoil (21; 121). The radially inner portion (22A) has aerodynamic forward sweep at the leading edge (24; 124; 324), the forward sweep having a maximum forward sweep that is more forward than a sweep at the leading edge in the intermediate portion (22C).