Gas turbine engines with turbine airfoil cooling

Patent metadata

Field	Value
Publication number	US-9394798-B2
Application number	US-201313855417-A
Country	US
Kind code	B2
Filing date	Apr 2, 2013
Priority date	Apr 2, 2013
Publication date	Jul 19, 2016
Grant date	Jul 19, 2016

Abstract

An airfoil for a gas turbine engine is provided. The airfoil includes a body with a leading edge, a trailing edge, a first side wall extending between the leading edge and the trailing edge, and a second side wall extending between the leading edge and the trailing edge. The body defines an interior cavity. The airfoil includes an interior wall disposed within the interior cavity of the body and extending between the first wall and the second wall to define a supply chamber and a leading edge chamber. The interior wall defines a cooling hole with a base portion and a locally extended portion to direct cooling air from the supply chamber to the leading edge chamber such that the cooling air impinges upon the leading edge.

First claim

What is claimed is:

1. An airfoil for a gas turbine engine, comprising:
a body comprising a leading edge, a trailing edge, a pressure side wall extending between the leading edge and the trailing edge of the airfoil, and a suction side wall extending between the leading edge and the trailing edge of the airfoil, wherein the body defines an interior cavity; and
an interior wall disposed within the interior cavity of the body and extending between the first wall and the second wall to define a supply chamber and a leading edge chamber,
wherein the interior wall defines a first cooling hole with a first base portion and a first locally extended portion completely surrounding the first cooling hole and a second cooling hole being adjacent to the first cooling hole in a radial row of cooling holes in the interior wall, the second cooling hole having a second base portion and a second locally extended portion completely surrounding the second cooling hole, the first and second cooling holes configured to direct cooling air from the supply chamber to the leading edge chamber such that the cooling air impinges upon the leading edge, wherein the first and second locally extended portions are separated at a respective distance from each of the pressure and suction side walls and are separated from one another.

2. The airfoil of claim 1 , wherein the first cooling hole has a generally oval cross-sectional shape.

3. The airfoil of claim 1 , wherein the first cooling hole has a non-circular cross-sectional shape.

4. The airfoil of claim 1 , wherein the interior wall has a first surface facing the leading edge chamber and a second surface facing the supply chamber, and wherein the first locally extended portion extends from the first surface.

5. The airfoil of claim 1 , wherein the interior wall has a first surface facing the leading edge chamber and a second surface facing the supply chamber, and wherein the first locally extended portion extends from the second surface.

6. The airfoil of claim 1 , wherein the first locally extended portion is generally cylindrical.

7. The airfoil of claim 1 , wherein the interior wall has a first surface facing the leading edge chamber and a second surface facing the supply chamber, and wherein the interior wall further comprises a scoop that extends from the second surface adjacent to the first cooling hole.

8. The airfoil of claim 1 , wherein the first cooling hole has a nonperpendicular central axis.

9. The airfoil of claim 1 , wherein the first cooling hole has a varying cross sectional area.

10. An airfoil for a gas turbine engine, comprising:
a body comprising a leading edge, a trailing edge, a pressure side wall extending between the leading edge and the trailing edge of the airfoil, and a suction side wall extending between the leading edge and the trailing edge of the airfoil, wherein the body defines an interior cavity; and
an interior wall disposed within the interior cavity of the body and extending between the first wall and the second wall to define a supply chamber and a leading edge chamber,
wherein the interior wall defines a cooling hole with a base portion and a locally extended portion to direct cooling air from the supply chamber to the leading edge chamber such that the cooling air impinges upon the leading edge, and
wherein the locally extended portion is generally conical, and wherein the locally extended portion has a generally circular perimeter.

11. 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 leading edge, a trailing edge, and side walls defining a body with an interior cavity, the airfoil further comprising an interior wall disposed within the interior cavity of the body and extending between the side walls to define a supply chamber and a leading edge chamber,

wherein the interior wall defines a first cooling hole with a first base portion and a first locally extended portion completely surrounding the first cooling hole and a second cooling hole being adjacent to the first cooling hole in a radial row of cooling holes in the interior wall, the second cooling hole having a second base portion and a second locally extended portion completely surrounding the second cooling hole, the first and second cooling holes configured to direct cooling air from the supply chamber to the leading edge chamber such that the cooling air impinges upon the leading edge, wherein the first and second locally extended portions are separated at a respective distance from each of the first and second side walls and are separated from one another.

12. The gas turbine engine of claim 11 , wherein the first locally extended portion is generally conical, and wherein the first locally extended portion has a generally circular perimeter.

13. The gas turbine engine of claim 11 , wherein the interior wall has a first surface facing the leading edge chamber and a second surface facing the supply chamber, and wherein the first locally extended portion extends from the first surface.

14. The gas turbine engine of claim 11 , wherein the interior wall has a first surface facing the leading edge chamber and a second surface facing the supply chamber, and wherein the interior wall further comprises a scoop that extends from the second surface adjacent to the first cooling hole.

Assignees

• Honeywell Int Inc

Inventors

• Crites Daniel Cregg
• Morris Mark C
• Crosatti Lorenzo
• Riahi Ardeshir

Classifications

• F05D2260/201

  by impingement of a fluid · CPC title

• F01D5/186

  Film cooling (F01D5/187 takes precedence) · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title

• F05D2250/232

  conical · CPC title

• F05D2250/231

  cylindrical · CPC title