Airfoil cooling apparatus

Patent metadata

Field	Value
Publication number	US-10012090-B2
Application number	US-201514804865-A
Country	US
Kind code	B2
Filing date	Jul 21, 2015
Priority date	Jul 25, 2014
Publication date	Jul 3, 2018
Grant date	Jul 3, 2018

Abstract

An airfoil for a gas turbine engine, the airfoil includes a wall that has a leading edge and a trailing edge and at least partially defining a boundary of a leading edge cavity radially along the leading edge. A cooling jet structure is operatively associated with a portion of the wall proximate the leading edge and is configured to direct a cooling fluid tangent to the portion of the wall.

First claim

What is claimed is:

1. An airfoil for a gas turbine engine, the airfoil comprising:
a wall having a leading edge and a trailing edge, and at least partially defining a boundary of a leading edge cavity radially along the leading edge, and a radially extending feed cavity;
a cooling jet operatively associated with a portion of the wall proximate the leading edge and configured to direct a cooling fluid from the radially extending feed cavity to the leading edge cavity tangent to the portion of the wall; and
at least one dividing wall separating the leading edge cavity into multiple segments and configured to minimize radial flow of the cooling fluid through the leading edge cavity, wherein the at least one dividing wall at least partially defines the cooling jet.

2. The airfoil of claim 1 , wherein the at least one dividing wall is generally perpendicular to the radial direction of the airfoil.

3. The airfoil of claim 1 , wherein the cooling jet is at least partially defined by a protrusion operatively disposed opposite the portion of the wall to form a channel therebetween configured to accelerate flow of the cooling fluid.

4. The airfoil of claim 3 wherein the portion of the wall includes a concave portion about the leading edge.

5. The airfoil of claim 4 , wherein the concave portion follows a profile of the leading edge.

6. The airfoil of claim 3 , wherein the radially extending feed cavity is in fluid communication with the channel.

7. The airfoil of claim 6 , wherein the cooling jet includes an inlet adjacent the radially extending feed cavity and an outlet adjacent the leading edge cavity and the at least one dividing wall extends from the radially extending feed cavity to the leading edge.

8. The airfoil of claim 1 , wherein the wall defines at least one film hole therethrough fluidly coupled to the leading edge cavity.

9. The airfoil of claim 1 , further comprising a pressure side cooling circuit extending along a pressure side of the airfoil, a suction side cooling circuit extending along a suction side of the airfoil, and a trailing edge cooling circuit extending along a trailing edge of the airfoil.

10. The airfoil of claim 9 , further comprising a central wall dividing the pressure side cooling circuit from the suction side cooling circuit.

11. A gas turbine engine assembly, comprising:
a rotor configured to rotate about an axis of rotation; and
an airfoil including:
a wall having a leading edge and a trailing edge, and at least partially defining a boundary of a leading edge cavity radially along the leading edge, and a radially extending feed cavity;
a cooling jet operatively associated with a portion of the wall proximate the leading edge and configured to direct a cooling fluid from the radially extending feed cavity to the leading edge cavity tangent to the portion of the wall; and
at least one dividing wall separating the leading edge cavity into multiple segments and configured to minimize radial flow of the cooling fluid through the leading edge cavity, wherein the at least one dividing wall at least partially defines the cooling jet.

12. The assembly of claim 11 , wherein the at least one dividing wall is generally perpendicular to the radial direction of the airfoil.

13. The assembly of claim 11 , wherein the cooling jet is at least partially defined by a protrusion operatively disposed opposite the portion of the wall to form a channel therebetween configured to accelerate flow of the cooling fluid.

14. The assembly of claim 13 , wherein the radially extending feed cavity is in fluid communication with the channel.

15. The assembly of claim 14 , wherein the cooling jet includes an inlet adjacent the radial feed cavity and an outlet adjacent the leading edge cavity and the at least one dividing wall extends from the radially extending feed cavity to the leading edge.

16. A method of cooling an airfoil in a gas turbine engine, comprising:
directing a cooling fluid through a cooling jet from a radially extending feed cavity to a leading edge cavity in a direction tangential to a portion of a wall of the leading edge cavity extending along a leading edge of the airfoil; and
separating the leading edge cavity into multiple segments configured to minimize radial flow of the cooling fluid through the leading edge cavity with at least one dividing wall, wherein the at least one dividing wall at least partially defines the cooling jet.

17. The method of claim 16 , wherein the at least one dividing wall extends generally perpendicular to the radial direction of the airfoil.

18. The method of claim 16 , wherein the cooling jet is at least partially defined by a protrusion operatively disposed opposite the portion of the wall to form a channel therebetween configured to accelerate flow of the cooling fluid.

19. The method of claim 16 , wherein the at least one dividing wall extends from the radially extending feed cavity to the leading edge.

Assignees

• United Technologies Corp

Inventors

• Slavens Thomas N
• Snyder Brooks E
• Burdick Andrew D

Classifications

• F05D2240/303

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

• F05D2250/322

  tangential · CPC title

• F05D2260/2212

  by creating turbulence · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title