Cooling hole for a gas turbine engine component

Patent metadata

Field	Value
Publication number	US-10215030-B2
Application number	US-201414766475-A
Country	US
Kind code	B2
Filing date	Feb 7, 2014
Priority date	Feb 15, 2013
Publication date	Feb 26, 2019
Grant date	Feb 26, 2019

Abstract

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall having an internal surface, an outer skin and a cooling hole having an inlet extending from the internal surface and merging into a metering section, and a diffusion section downstream of the metering section that extends to an outlet located at the outer skin. At least two lobes are embedded within the diffusion section of the cooling hole. At least one surface of each of the at least two lobes is at least partially cylindrical.

First claim

What is claimed is:

1. A component for a gas turbine engine, comprising:
a wall having an internal surface and an outer skin;
a cooling hole having an inlet extending from said internal surface and merging into a metering section, and a diffusion section downstream of said metering section that extends to an outlet located at said outer skin;
at least two lobes embedded within said diffusion section of said cooling hole such that a trailing edge of said at least two lobes is upstream from a trailing edge of said diffusion section, wherein at least one surface of each of said at least two lobes is at least partially cylindrical;
wherein said at least two lobes each include a leading edge, said trailing edge, a first side surface that extends between said leading edge and said trailing edge of said at least two lobes along a first curved edge, and a second side surface that extends between said trailing edge of said at least two lobes and said leading edge along a second curved edge; and
a curved transition portion extending between said second curved edge of a first lobe of said at least two lobes and said first curved edge of a second lobe of said at least two lobes,
wherein each of said first curved edges is curved along an entire length of said first side surfaces and each of said second curved edges is curved along an entire length of said second side surfaces.

2. The component as recited in claim 1 , wherein said wall is part of one of an airfoil, a turbine vane, a turbine blade, a blade outer air seal (BOAS), a combustor liner and a platform.

3. The component as recited in claim 1 , wherein said trailing edge of said diffusion section is linear.

4. The component as recited in claim 1 , wherein said first lobe and said second lobe diverge longitudinally and laterally from said metering section.

5. The component as recited in claim 1 , wherein said curved transition portion extends to said outer skin.

6. The component as recited in claim 1 , wherein said curved transition portion is below said outer skin.

7. The component as recited in claim 1 , comprising a coating layer at said outer skin, wherein said diffusion section extends into said coating layer.

8. The component as recited in claim 7 , wherein an entirety of said diffusion section is formed within said coating layer and said metering section is formed entirely within a substrate of said wall.

9. The component as recited in claim 7 , wherein a first portion of said diffusion section extends into said coating layer and a second portion of said diffusion section extends within a substrate of said wall.

10. The component as recited in claim 1 , wherein said curved transition portion extends between said first lobe and said second lobe at a position that is upstream from a downstream portion of said diffusion section.

11. The component as recited in claim 1 , wherein said first curved edge diverges laterally from said leading edge and converges laterally before reaching said trailing edge.

12. The component as recited in claim 11 , wherein said second curved edge diverges proximally.

13. The component as recited in claim 1 , wherein said at least two lobes extend at an angle that is between 10° and 60° relative to an axis of said metering section.

14. The component as recited in claim 1 , wherein said diffusion section defines an asymmetric design.

15. The component as recited in claim 1 , wherein said diffusion section includes a downstream surface that extends at an angle between 135° and 180° relative to an axis of said metering section.

16. The component as recited in claim 1 , wherein said at least two lobes include different radii.

17. A method of forming a cooling hole in a component of a gas turbine engine, comprising the step of:
forming a cooling hole in a wall of the component including an inlet extending from an internal surface of the wall toward an outer skin of the wall, the inlet merging into a metering section; and
providing the cooling hole with a diffusion section downstream of the metering section, the diffusion section including at least two lobes that are embedded within the diffusion section of the cooling hole, the at least two lobes having a surface that is at least partially cylindrical, wherein the step of providing the cooling hole with the diffusion section includes forming a trailing edge of the at least two lobes at an upstream position from a trailing edge of the diffusion section;
wherein the at least two lobes each include a leading edge, the trailing edge, a first side surface that extends between the leading edge and the trailing edge of the at least two lobes along a first curved edge, and a second side surface that extends between the trailing edge of the at least two lobes and the leading edge along a second curved edge,
wherein each of the first curved edges is curved along an entire length of the first side surfaces and each of the second curved edges is curved along an entire length of the second side surfaces.

18. The method as recited in claim 17 , comprising the step of providing a coating layer at the outer skin of the wall.

19. The method as recited in claim 18 , wherein the step of providing the cooling hole with the diffusion section includes forming the diffusion section entirely within the coating layer.

20. The component as recited in claim 1 , wherein said trailing edge of said at least two lobes is longitudinally offset in a direction toward said internal surface from said trailing edge of said diffusion section.

21. The method as recited in claim 17 , wherein the trailing edge of the at least two lobes is longitudinally offset in a direction toward the internal surface from the trailing edge of the diffusion section.

22. A component for a gas turbine engine, comprising:
a wall having an internal surface and an outer skin;
a cooling hole having an inlet extending from said internal surface and merging into a metering section, and a diffusion section downstream from said metering section that extends to an outlet located at said outer skin;
a first lobe and a second lobe embedded within said diffusion section of said cooling hole such that a first trailing edge of each of said first lobe and said second lobe is upstream from a second trailing edge of said diffusion section;
wherein said first lobe and said second lobe each include a leading edge, said first trailing edge, a first side surface that extends between said leading edge and said first trailing edge along a first curved edge, and a second side surface that extends between said first trailing edge and said leading edge along a second curved edge; and
a third lobe embedded within said first lobe.

Assignees

• United Technologies Corp

Inventors

• Xu Jinquan

Classifications

• F01D5/186 Primary

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

• F01D5/288

  Protective coatings for blades · CPC title

• F05D2260/202

  by film cooling · CPC title

• F23R3/002

  Wall structures (F23R3/02 and F23R3/007 take precedence) · CPC title

• F01D9/065

  Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids (see also F01D25/16, F01D25/24 and F01D25/26) · CPC title