Gas turbine engine airfoil

Patent metadata

Field	Value
Publication number	US-11193496-B2
Application number	US-202016740817-A
Country	US
Kind code	B2
Filing date	Jan 13, 2020
Priority date	Feb 19, 2014
Publication date	Dec 7, 2021
Grant date	Dec 7, 2021

Abstract

A gas turbine engine includes a combustor section arranged between a compressor section and a turbine section. The compressor section includes at least a low pressure compressor and a high pressure compressor. The high pressure compressor is arranged upstream of the combustor section. A fan section has an array of twenty-six or fewer fan blades. The low pressure compressor is downstream from the fan section. An airfoil is arranged in the low pressure compressor and 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 camber angle and span position that defines a curve with a decreasing camber angle within the range of 80% span to 100% span. The camber angle is less than 20° within the entire range of 40% span to 100% span.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a combustor section arranged between a compressor section and a turbine section, wherein the compressor section includes at least a low pressure compressor and a high pressure compressor, the high pressure compressor arranged upstream of the combustor section;
a fan section having an array of twenty-six or fewer fan blades, wherein the low pressure compressor is downstream from the fan section, wherein the low pressure compressor is counter-rotating relative to the fan blades; and
an airfoil arranged in the low pressure compressor and 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 camber angle and span position that defines a curve with a decreasing camber angle within the range of 80% span to 100% span, and wherein the camber angle is less than 20° within the entire range of 40% span to 100% span.

2. The gas turbine engine according to claim 1 , wherein the gas turbine engine is a two-spool configuration.

3. The gas turbine engine according to claim 1 , wherein the airfoil is rotatable relative to an engine static structure.

4. The gas turbine engine according to claim 1 , wherein the curve has a decreasing camber angle within the range of 60% span to 100% span.

5. The gas turbine engine according to claim 1 , wherein the camber angle is generally linear within the range of 0% span to 100% span.

6. The gas turbine engine according to claim 1 , wherein the camber angle is less than 15° within the range of 80% span to 100% span.

7. The gas turbine engine according to claim 1 , wherein the fan section has a low fan pressure ratio of less than 1.55.

Assignees

• Raytheon Tech Corp

Inventors

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

Classifications

• F05D2220/32

  in gas turbines · CPC title

• F04D29/38

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

• F02K3/072

  with counter-rotating {, e.g. fan} rotors · CPC title

• F05D2250/70

  Shape · CPC title

• F04D29/325

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