Compressor rotor airfoil

Patent metadata

Field	Value
Publication number	US-9765795-B2
Application number	US-201414469938-A
Country	US
Kind code	B2
Filing date	Aug 27, 2014
Priority date	Aug 27, 2014
Publication date	Sep 19, 2017
Grant date	Sep 19, 2017

Abstract

A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are 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. 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 has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

First claim

The invention claimed is:

1. A compressor rotor airfoil for 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; and
a leading edge dihedral angle 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 dihedral angle generally decreases from the root to a first point along the span, generally increases from the first point to a second point along the span, and generally decreases from the second point to the tip, wherein the first point is disposed at 80% of the span, and the second point is disposed at 95% of the span, and wherein the dihedral angle at the root is 20 degrees or more higher than the dihedral angle at the first point.

2. The airfoil of claim 1 , wherein the dihedral angle has at most two changes of direction.

3. The airfoil of claim 1 , wherein the dihedral angle generally decreases from the root to the tip.

4. The airfoil of claim 1 , including a leading edge sweep angle defined relative to a tangent to the airfoil and flow velocity vector; and a ratio of the leading edge sweep angle to the leading edge dihedral angle being smaller than 1.

5. The airfoil of claim 1 , wherein a first center of gravity of a first cross-section of the airfoil at the tip being axially upstream along an axial length taken chordaly relative to a second center of gravity of a second cross-section of the airfoil at the root.

6. A gas turbine engine comprising:
a compressor section including a plurality of rotors, each of the plurality of rotors including a hub, the hubs being aligned axially, each of the rotors including a plurality of blades extending radially from the hub, the blades including the airfoil of claim 1 .

7. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade, the method comprising:
forming an airfoil having 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 leading edge dihedral angle being defined between a tangent to the airfoil and a vertical at a point on the leading edge, the leading edge dihedral angle having a span-wise distribution, the dihedral angle generally decreasing from the root to a first point along the span, generally increasing from the first point to a second point along the span, and generally decreasing from the second point to the tip, wherein the first point disposed at 80% of the span, and the second point disposed at 95% of the span, and wherein the dihedral angle at the root is 20 degrees or more higher than the dihedral angle at the first point.

8. The method of claim 7 , wherein forming the airfoil comprises forming the airfoil with at most two changes of direction in the distribution of the leading edge dihedral angle.

9. The method of claim 7 , wherein forming the airfoil comprises forming the airfoil with the dihedral angle generally decreasing from the root to the tip.

10. The method of claim 7 , wherein forming the airfoil comprises forming the airfoil with a ratio of a leading edge sweep angle to the leading edge dihedral angle being smaller than 1, the leading edge sweep angle being between a tangent to the airfoil and flow velocity vector.

11. The method of claim 7 , wherein forming the airfoil comprises forming the airfoil with a first center of gravity of a first cross-section of the airfoil at the tip being axially upstream along an axial length taken chordaly relative to a second center of gravity of a second cross-section of the airfoil at the root.

Assignees

• Pratt & Whitney Canada

Inventors

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

Classifications

• F04D29/384 Primary

  characterised by form · CPC title

• F01D5/141 Primary

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

• F04D29/324

  Blades · CPC title

• F01D5/145

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

• F05D2240/303

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