Rotary airfoil

Patent metadata

Field	Value
Publication number	US-10443390-B2
Application number	US-201414469948-A
Country	US
Kind code	B2
Filing date	Aug 27, 2014
Priority date	Aug 27, 2014
Publication date	Oct 15, 2019
Grant date	Oct 15, 2019

Abstract

A rotary airfoil in a gas turbine engine is provided. The airfoil includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. An axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil. A method for forming such blade is also presented.

First claim

The invention claimed is:

1. A rotary airfoil ofa compressor in a gas turbine engine, the airfoil comprising: opposed pressure and suction sides joined together at chordally opposite leading and trailing edges, the pressure and suction sides extending in a span direction from a root to a tip, a chord of the airfoil extending between the leading and trailing edges;
an axial sweep of centers of gravity of the airfoil having a maximum value at the tip, the axial sweep decreasing from the maximum value at the tip to a minimum value at an intermediate position along the span between the tip and the root, and increasing from the minimum value at the intermediate position toward the root, the axial sweep increasing in a direction opposite to a direction of airflow across the airfoil;
a continuous tangential lean of the centers ofgravity of the airfoil along the span from the root to the tip in a direction of rotation ofthe airfoil; and
a first center of gravity of a first cross-section ofthe airfoil at the tip upstream of a second center of gravity of a second cross-section of the airfoil at the root relative to the direction of airflow across the_rotary airfoil.

2. The rotary airfoil of claim 1 , wherein a thickness is defined across the airfoil between the pressure and suction sides, the thickness having a chord-wise distribution and a maximum thickness, a thick region of a given cross-section being defined by a thickness of at least 85% of the maximum thickness and extending from the maximum thickness towards the leading edge by a length of at most 15% of the chord and extending towards the trailing edge by a length of at most 15% of the chord, a second thick region ofthe second cross-section being chordally longer than a first thick region ofthe first cross-section.

3. The rotary airfoil of claim 2 , wherein the second thick region is comprised between 30% and 60% ofthe chord taken from the leading edge.

4. The rotary airfoil of claim 3 , wherein the first thick region is comprised between 30% and 45% ofthe chord taken from the leading edge.

5. The rotary airfoil of claim 1 , wherein a leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical, the leading edge dihedral angle having a span-wise distribution, the distribution having at least one change ofdirection.

6. The rotary airfoil of claim 5 , wherein the at least one change of direction in the distribution of the leading edge dihedral angle is toward the tip ofthe blade.

7. The rotary airfoil of claim 5 , wherein the distribution of the leading edge dihedral angle generally decreases from the root to the tip.

8. The rotary of claim 1 , wherein a leading edge sweep angle is defined between a tangent to the airfoil and flow velocity vector at a point on the leading edge;
a leading edge dihedral angle defined between the tangent to the airfoil and a vertical at the point on the leading edge; and
a ratio ofthe leading edge sweep angle to the leading edge dihedral angle being smaller than 1 .

9. The rotary airfoil of claim 1 , wherein the first center of gravity is upstream of any other center of gravity of a cross-section of the airfoil relative to the direction of airflow across the rotary airfoil.

10. A gas turbine engine comprising:
a compressor section including a plurality of rotors, each of the plurality of rotors including a hub and a plurality of blades extending radially from the hub, the hubs being aligned axially, each one of the blades including an airfoil portion, the airfoil portion comprising:
opposed pressure and suction sides joined together at chordally opposite leading and trailing edges, the pressure and suction sides extending in a span direction from a root to a tip, a chord of the airfoil extending between the leading and trailing edges;
an axial sweep of centers of gravity of the airfoil having a maximum value at the tip, the axial sweep decreasing from the maximum value at the tip to a minimum value at an intermediate position along the span between the tip and the root, and increasing from the minimum value at the intermediate position toward the root, the axial sweep increasing in a direction opposite to a direction of airflow across the airfoil;
a continuous tangential lean of the centers ofgravity of the airfoil along the span from the root to the tip in a direction of rotation ofthe airfoil; and
a first center ofgravity of a first cross-section of the airfoil at the tip upstream of a second center of gravity of a second cross-section of the airfoil at the root relative to the direction of airflow in the compressor section.

11. The gas turbine engine of claim 10 , wherein a thickness is defined across the airfoil between the pressure and suction sides, the thickness having a chord-wise distribution and a maximum thickness, a thick region of a given cross-section being defined by a thickness at least 85% of the maximum thickness and extending from the maximum thickness towards the leading edge by a length of at most 15% of the chord and extending towards the trailing edge by a length of at most 15% of the chord, a second thick region of the second cross-section being chordally longer than a first thick region of the first cross-section.

12. The gas turbine engine of claim 11 , wherein the second thick region is comprised between 30% and 60% ofthe chord taken from the leading edge.

13. The gas turbine engine of claim 12 , wherein the first thick region is comprised between 30% and 45% ofthe chord taken from the leading edge.

14. The gas turbine engine of claim 10 , wherein a leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical, the leading edge dihedral angle having a span-wise distribution, the distribution having at least one change of direction.

15. The gas turbine engine of claim 10 , wherein a leading edge sweep angle is defined between a tangent to the airfoil and flow velocity vector at a point on the leading edge;
a leading edge dihedral angle defined between the tangent to the airfoil and a vertical at the point on the leading edge; and
a ratio ofthe leading edge sweep angle to the leading edge dihedral angle being smaller than 1.

16. The gas turbine engine of claim 10 , wherein the first center of gravity is upstream of any other center of gravity of a cross-section of the airfoil relative to the direction of airflow in the compressor section.

17. A method of forming a rotary blade of a compressor having opposed pressure and suction sides joined together at chordally opposite leading and trailing edges, a chord of the airfoil extending between the leading and trailing edges, the pressure and suction side extending in a span direction from a root to a tip, the method comprising:
forming an airfoil having a first center ofgravity of a first cross-section of the airfoil at the tip upstream of a second center of gravity of a second cross-section of the airfoil at the root relative to a direction of airflow across the airfoil, an axial sweep of centers of gravity of the airfoil having a maximum value at the tip, the axial sweep decreasing from the maximum value at the tip to a minimum value at an intermediate position along the span between the tip and the root, and increasing from the minimum value at the intermediate position toward the root, the axial sweep increasing in a direction opposite to a direction of airflow across
the airfoil, and a continuous tangential lean of the centers of gravity of the airfoil along the span from the root to the tip in a direction of rota

Assignees

• Pratt & Whitney Canada

Inventors

• Duong Hien
• Balike Krishna Prasad
• Veitch Thomas
• Warikoo Raman
• Lobo Keegan

Classifications

• F05D2250/71

  curved · CPC title

• F01D5/141 Primary

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

• F05D2240/303

  related to the leading edge of a rotor blade · CPC title

• F01D5/145

  Means for influencing boundary layers or secondary circulations (for compressors F04D29/68) · CPC title

• F04D29/324

  Blades · CPC title