Gas turbine engines with improved leading edge airfoil cooling

Patent metadata

Field	Value
Publication number	US-2016108740-A1
Application number	US-201414515131-A
Country	US
Kind code	A1
Filing date	Oct 15, 2014
Priority date	Oct 15, 2014
Publication date	Apr 21, 2016
Grant date	—

Abstract

An airfoil for a gas turbine engine includes a body with a first side wall and a second side wall joined at a leading edge and a trailing edge, the first side wall having a first interior surface and the second side wall having a second interior surface. The airfoil further includes an internal wall disposed within of the body and extending between the first interior surface and the second interior surface to define a supply passage and a leading edge passage. The internal wall defines a plurality of cooling holes to direct cooling air from the supply passage to the leading edge passage such that the cooling air impinges upon the leading edge. The airfoil further includes a first plurality of grooves formed in the first interior surface, each the first plurality of grooves extending in a chordwise direction within the leading edge passage.

First claim

What is claimed is:

1 . An airfoil for a gas turbine engine, comprising:
a body comprising a first side wall and a second side wall joined at a leading edge and a trailing edge, the first side wall having a first interior surface and the second side wall having a second interior surface; an internal wall disposed within of the body and extending between the first interior surface and the second interior surface to define a supply passage and a leading edge passage, wherein the internal wall defines a plurality of cooling holes to direct cooling air from the supply passage to the leading edge passage such that the cooling air impinges upon the leading edge; and a first plurality of grooves formed in the first interior surface, each the first plurality of grooves extending in a chordwise direction within the leading edge passage.

2 . The airfoil of claim 1 , further comprising a second plurality of grooves formed in the second interior surface, each of the second plurality of grooves extending in the chordwise direction within the leading edge chamber.

3 . The airfoil of claim 2 , wherein the first plurality of grooves and the second plurality of grooves are aligned in the chordwise direction.

4 . The airfoil of claim 2 , wherein the first plurality of grooves and the second plurality of grooves are radially offset relative to one another.

5 . The airfoil of claim 2 , each of the plurality of cooling holes is associated with one of the first plurality of grooves and one of the second plurality of grooves.

6 . The airfoil of claim 2 , wherein the each of the first plurality of grooves and each of the second plurality of grooves have a first end at the internal wall.

7 . The airfoil of claim 2 , wherein each of the first plurality of grooves and each of the second plurality of grooves have approximately the same length.

8 . The airfoil of claim 2 , wherein each of the first plurality of grooves and each of the second plurality of grooves are parallel to one another.

9 . The airfoil of claim 2 , wherein the first plurality of grooves and the second plurality of grooves are non-parallel to one another.

10 . The airfoil of claim 1 , wherein the first interior surface is generally flat between adjacent grooves of the first plurality of grooves.

11 . The airfoil of claim 1 , wherein each of the first plurality of grooves is formed by a curved side wall.

12 . The airfoil of claim 1 , wherein each of the first plurality of grooves has a width and a depth, wherein the width is at least three times the depth.

13 . The airfoil of claim 1 , wherein the width is approximately six times the depth.

14 . A gas turbine engine, comprising:
a compressor section configured to receive and compress air; a combustion section coupled to the compressor section and configured to receive the compressed air, mix the compressed air with fuel, and ignite the compressed air and fuel mixture to produce combustion gases; and a turbine section coupled to the combustion section and configured to receive the combustion gases, the turbine section defining a combustion gas path and comprising a turbine rotor positioned within the combustion gas path, the turbine rotor comprising
a platform at least partially defining the combustion gas path; and
an airfoil extending from the platform, the airfoil including
a body comprising a first side wall and a second side wall joined at a leading edge and a trailing edge, the first side wall having a first interior surface and the second side wall having a second interior surface;
an internal wall disposed within of the body and extending between the first interior surface and the second interior surface to define a supply passage and a leading edge passage,
wherein the internal wall defines a plurality of cooling holes to direct cooling air from the supply passage to the leading edge passage such that the cooling air impinges upon the leading edge;
a first plurality of grooves formed in the first interior surface, each the first plurality of grooves extending in a chordwise direction within the leading edge passage; and
a second plurality of grooves formed in the second interior surface, each of the second plurality of grooves extending in the chordwise direction within the leading edge chamber.

15 . The gas turbine engine of claim 14 , wherein the first plurality of grooves and the second plurality of grooves are aligned in the chordwise direction.

16 . The gas turbine engine of claim 14 , each of the plurality of cooling holes is associated with one of the first plurality of grooves and one of the second plurality of grooves.

17 . The gas turbine engine of claim 14 , wherein the each of the first plurality of grooves and each of the second plurality of grooves have a first end at the internal wall.

18 . The gas turbine engine of claim 14 , wherein each of the first plurality of grooves and each of the second plurality of grooves have approximately the same length.

19 . The gas turbine engine of claim 14 , wherein each of the first plurality of grooves is formed by a curved side wall.

20 . The gas turbine engine of claim 14 , wherein each of the first plurality of grooves has a width and a depth, wherein the width is at least three times the depth.

Assignees

• Honeywell Int Inc

Inventors

• Srinivasan Balamurugan
• Narayanan Sridharan Reghupathy
• Paramanandam Karthikeyan
• Jayamurugan Chandiran
• Malak Malak Fouad

Classifications

• F01D5/187 Primary

  Convection cooling · CPC title

• F05D2240/303

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

• F05D2260/201

  by impingement of a fluid · CPC title

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F05D2220/32

  in gas turbines · CPC title