Gas turbine engine airfoil

Patent metadata

Field	Value
Publication number	US-2016348691-A1
Application number	US-201515114922-A
Country	US
Kind code	A1
Filing date	Feb 12, 2015
Priority date	Feb 19, 2014
Publication date	Dec 1, 2016
Grant date	—

Abstract

An airfoil of a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between a stagger angle and span position that defines a curve with a stagger angle that is greater than 35° from 0% span to at least 50% span.

First claim

What is claimed is:

1 . An airfoil of a turbine engine comprising:
pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position, wherein the airfoil has a relationship between a stagger angle and span position that defines a curve with a stagger angle that is greater than 35° from 0% span to at least 50% span.

2 . The airfoil according to claim 1 , wherein the curve includes at least one of a non-increasing positive slope and a negative slope in a range of 80% span to 100% span.

3 . The airfoil according to claim 1 , wherein the stagger angle is generally linear from 0% span to 100% span.

4 . The airfoil according to claim 1 , wherein the stagger angle is between 35° and 45° at 0% span.

5 . The airfoil according to claim 1 , wherein the stagger angle is between 55° and 65° at 100% span.

6 . A gas turbine engine comprising:
a combustor section arranged between a compressor section and a turbine section; a fan section having an array of twenty-six or fewer fan blades, wherein the fan section has a low fan pressure ratio of less than 1.55; a geared architecture coupling the fan section to the turbine section or the compressor section; and an airfoil including pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position, wherein the airfoil has a relationship between a stagger angle and span position that defines a curve with a stagger angle that is greater than 35° from 0% span to at least 50% span.

7 . The gas turbine engine according to claim 6 , wherein the airfoil is arranged in the compressor section.

8 . The gas turbine engine according to claim 7 , wherein the compressor section includes at least a low pressure compressor and a high pressure compressor, the high pressure compressor arranged immediately upstream of the combustor section, the airfoil provided in a compressor outside the high pressure compressor section.

9 . The gas turbine engine according to claim 8 , wherein the low pressure compressor is counter-rotating relative to the fan blades.

10 . The gas turbine engine according to claim 8 , wherein the gas turbine engine is a two-spool configuration.

11 . The gas turbine engine according to claim 8 , wherein the low pressure compressor is immediately downstream from the fan section.

12 . The gas turbine engine according to claim 7 , wherein the airfoil is rotatable relative to an engine static structure.

13 . The gas turbine engine according to claim 6 , wherein the curve includes at least one of a non-increasing positive slope and a negative slope in a range of 80% span to 100% span.

14 . The gas turbine engine according to claim 6 , wherein the stagger angle is generally linear from 0% span to 100% span.

15 . The gas turbine engine according to claim 14 , wherein the stagger angle is between 35° and 45° at 0% span.

16 . The gas turbine engine according to claim 14 , wherein the stagger angle is between 55° and 65° at 100% span.

Assignees

• United Technologies Corp

Inventors

• Gallagher Edward J
• Brilliant Lisa I
• Straccia Joseph C
• Balamucki Stanley J
• Stephens Mark A
• Hudon Kate

Classifications

• F01D5/141 Primary

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

• F05D2220/36

  specially adapted for the fan of turbofan engines · CPC title

• F01D17/162

  for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line (F01D17/167 takes precedence) · CPC title

• F04D29/38 Primary

  Blades {(for axial flow compressors F04D29/324)} · CPC title

• F04D29/325

  for axial flow fans (blade mountings F04D29/34, blades F04D29/38) · CPC title