Compressor rotor airfoil

Patent metadata

Field	Value
Publication number	US-10760424-B2
Application number	US-201715674865-A
Country	US
Kind code	B2
Filing date	Aug 11, 2017
Priority date	Aug 27, 2014
Publication date	Sep 1, 2020
Grant date	Sep 1, 2020

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 blade for a gas turbine engine, the compressor blade comprising:
an airfoil having opposed pressure and suction sides joined together at chordally opposite leading and trailing edges, the pressure and suction sides extending along a span in a spanwise direction from a root to a tip;
a leading edge dihedral angle defined at a point on the leading edge between a tangent to the airfoil and a vertical in a radial direction relative to an axial axis of the gas turbine engine, a positive dihedral angle leans the airfoil in a direction of rotation; and
first, second and third regions of the airfoil, the first region extending from the root toward the tip, the third region extending from the tip toward the root, and the second region extending between the first and third regions, the leading edge dihedral angle generally decreases within the first region in the spanwise direction, generally increases within the second region in the spanwise direction, and generally decreases within the third region in the spanwise direction, the leading edge dihedral angle remains positive along the span from the root to the tip, wherein the leading edge dihedral angle decreases within the first region between 20 degrees to 30 degrees.

2. The compressor blade as defined in claim 1 , wherein the leading edge dihedral angle has at most two changes of direction between a decrease and an increase thereof.

3. The compressor blade as defined in claim 1 , including a leading edge sweep angle defined relative to the tangent to the airfoil and a flow velocity vector, a ratio of the leading edge sweep angle to the leading edge dihedral angle being smaller than 1.

4. The compressor blade as defined in claim 1 , wherein a first center of gravity of a first cross-section of the airfoil at the tip is 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.

5. The compressor blade as defined in claim 1 , wherein the leading edge dihedral angle decreases within the first region between 24 degrees to 26 degrees.

6. The compressor blade as defined in claim 1 , wherein the leading edge dihedral angle increases within the second region between 5 degrees to 10 degrees.

7. The compressor blade as defined in claim 1 , wherein the leading edge dihedral angle decreases within the third region between 10 degrees and 15 degrees.

8. 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 blade of claim 1 .

9. A compressor blade for a gas turbine engine, the compressor blade comprising:
an airfoil having opposed pressure and suction sides joined together at chordally opposite leading and trailing edges, the pressure and suction sides extending along a span in a spanwise direction from a root to a tip;
a leading edge dihedral angle defined at a point on the leading edge between a tangent to the airfoil and a vertical in a radial direction relative to an axial axis of the gas turbine engine, a positive dihedral angle leans the airfoil in a direction of rotation; and

first, second and third regions of the airfoil, the first region extending from the root toward the tip, the third region extending from the tip toward the root, and the second region extending between the first and third regions, the leading edge dihedral angle generally decreases within the first region in the spanwise direction, generally increases within the second region in the spanwise direction, and generally decreases within the third region in the spanwise direction, the leading edge dihedral angle remains positive along the span from the root to the tip, wherein the leading edge dihedral angle decreases within the third region between 10 degrees and 15 degrees.

Assignees

• Pratt & Whitney Canada

Inventors

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

Classifications

• Y02T50/60

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

• F01D5/141 Primary

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

• F04D19/00

  Axial-flow pumps (F04D21/00 takes precedence; {pump comprising axial flow and radial flow stages F04D17/025}) · CPC title

• F01D5/145

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

• F04D29/324

  Blades · CPC title