Compressor stator vane with leading edge forward sweep

Patent metadata

Field	Value
Publication number	US-10578125-B2
Application number	US-201615361050-A
Country	US
Kind code	B2
Filing date	Nov 24, 2016
Priority date	Nov 24, 2016
Publication date	Mar 3, 2020
Grant date	Mar 3, 2020

Abstract

A compressor of a gas turbine engine is described which includes a rotor and a stator located immediately downstream of the rotor. The stator comprises a plurality of stator vanes each having an airfoil extending along a generally radial span between a root and a tip of the airfoil. The radially inner portion has aerodynamic forward sweep at the leading edge, the forward sweep having a maximum forward sweep that is more forward than a sweep at the leading edge in the intermediate portion.

First claim

The invention claimed is:

1. A compressor of a gas turbine engine, comprising: a rotor rotatable about a center axis and a stator located immediately downstream of the rotor, the stator comprising a plurality of stator vanes each having an airfoil extending along a generally radial span between a root and a tip of the airfoil and having a chord with a chord length defined between a leading edge and a trailing edge of the airfoil, the airfoil having a radially inner portion defined by an inner approximately 40% of the span, a radially outer portion defined by an outer approximately 40% of the span, and an intermediate portion between the inner and outer portions, the leading edge having aerodynamic forward sweep throughout the radially inner portion, the forward sweep having a maximum forward sweep that is more forward than a sweep at the leading edge in the intermediate portion, an axial projection of the stator vane defined in an axial plane including the center axis, the stator vane having in the axial projection a leading edge belly region which extends forwardly of a reference line in the axial projection, the reference line extending between a first point defined at an intersection of the leading edge and the root, and a second point on the leading edge at 50% of the span.

2. The compressor as defined in claim 1 , wherein the maximum forward sweep is disposed between 5% and 30% of the span.

3. The compressor as defined in claim 1 , wherein the radially inner portion has aerodynamic dihedral at the leading edge, the leading edge dihedral being aft at the root of the airfoil.

4. The compressor as defined in claim 3 , wherein the leading edge dihedral has a maximum aft dihedral at the root that is more aft than the dihedral in a remainder of the inner portion.

5. The compressor as defined in claim 1 , wherein the belly region has an axial extent of between 5% and 30% of the chord length of the airfoil.

6. The compressor as defined in claim 1 , wherein the belly region is defined between the root and 40% of the span.

7. The compressor as defined in claim 6 , wherein the belly region is disposed between 10% and 20% of the span.

8. The compressor as defined in claim 1 , wherein the maximum forward sweep is defined at the root.

9. The compressor as defined in claim 1 , wherein the radially inner portion has aerodynamic forward sweep at the trailing edge, the forward sweep having a maximum forward sweep at the root.

10. The compressor as defined in claim 1 , wherein the stator vane has in the axial projection a trailing edge belly region which extends forwardly of a reference line in the axial projection, the reference line extending between a first point defined at an intersection of the trailing edge and the root, and a second point on the trailing edge at 50% of the span.

11. The compressor as defined in claim 10 , wherein the trailing edge belly region has an axial extent of between 5% and 30% of the chord length of the airfoil.

12. A gas turbine engine, comprising: a compressor with a rotor rotatable about a center axis and a stator located immediately downstream of the rotor, the stator comprising a plurality of stator vanes each having an airfoil extending along a generally radial span between a root and a tip of the airfoil and having a chord with a chord length defined between a leading edge and a trailing edge of the airfoil, the airfoil having a radially inner portion defined by an inner approximately 40% of the span, a radially outer portion defined by an outer approximately 40% of the span, and an intermediate portion between the inner and outer portions, the leading edge having aerodynamic forward sweep throughout the radially inner portion, the forward sweep having a maximum forward sweep that is more forward than a sweep at the leading edge in the intermediate portion, an axial projection of the stator vane defined in an axial plane including the center axis, the stator vane having in the axial projection a leading edge belly region which extends forwardly of a reference line in the axial projection, the reference line extending between a first point defined at an intersection of the leading edge and the root, and a second point on the leading edge at 50% of the span.

13. The gas turbine engine as defined in claim 12 , wherein the maximum forward sweep is disposed between 5% and 30% of the span.

14. The gas turbine engine as defined in claim 12 , wherein the radially inner portion has aerodynamic dihedral at the leading edge, the leading edge dihedral being aft at the root of the airfoil.

15. The gas turbine engine as defined in claim 12 , wherein the belly region has an axial extent of between 5% and 30% of the chord length of the airfoil.

16. The gas turbine engine as defined in claim 12 , wherein the belly region is defined between the root and 40% of the span.

17. The gas turbine engine as defined in claim 1 , wherein the radially inner portion has aerodynamic forward sweep at the trailing edge, the forward sweep having a maximum aerodynamic forward sweep at the root.

18. A method of forming a stator vane, comprising:
providing an airfoil having a chord with a chord length and a generally radial span extending over a radially inner portion defined by an inner approximately 40% of the span, a radially outer portion defined by an outer approximately 40% of the span, and an intermediate portion between the inner and outer portions;
forming a leading edge of the airfoil having an aerodynamic forward sweep throughout the radially inner portion, the forward sweep having a maximum forward sweep that is more forward than a sweep at the leading edge in the intermediate portion; and
forming a leading edge belly region in an axial projection of the stator vane in an axial plane including a center axis, the leading edge belly region extending forwardly of a reference line in the axial projection, the reference line extending between a first point defined at an intersection of the leading edge and the root, and a second point on the leading edge at 50% of the span.

Assignees

• Pratt & Whitney Canada

Inventors

• Duong Hien

Classifications

• F01D5/141

  Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147) · CPC title

• F04D29/542 Primary

  Bladed diffusers (fixing blades to stators F01D9/042) · CPC title

• F01D5/14 Primary

  Form or construction (selecting particular materials, measures against erosion or corrosion F01D5/28) · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title