GAS TURBINE ENGINE COMPONENT WITH COMBINED MATE FACE AND PLATFORM COOLING

• US 20150377032A1
• Filed: 02/07/2014
• Published: 12/31/2015
• Est. Priority Date: 02/15/2013
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: RAPP, Brandon M., HOUGH, Matthew Andrew, LEWIS, Scott D., BEATTIE, Jeffrey S., TELLER, Bret M., JACQUES, Jeffrey Michael, ASTERLIN, Max
• Granted Patent: US 10,227,875 B2
• Time in Patent Office: Days
• US Class Current: 1/1
• CPC Class Codes: F01D 11/08 for sealing space between rotor blade tips and stator specially-shaped blade tips therefor F01D5/20
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 5/18 Hollow blades, i.e. blades with cooling or heating channels or cavities structure of hollow blades in general F01D5/147; Heating, heat-insulating or cooling means on blades
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 9/023 Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
• CPC Class Codes: F01D 9/065 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 Class Codes: F05D 2240/11 Shroud seal segments
• CPC Class Codes: F05D 2240/35 Combustors or associated equipment
• CPC Class Codes: F05D 2240/81 Cooled platforms
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F23R 2900/03041 Effusion cooled combustion chamber walls or domes
• CPC Class Codes: F23R 2900/03042 Film cooled combustion chamber walls or domes
• CPC Class Codes: F23R 3/002 Wall structures F23R3/02 and F23R3/007 take precedence
• CPC Class Codes: F23R 3/02 characterised by the air-flow or gas-flow configuration reverse- flow combustion chambers F23R3/54; cyclone or vortex type combustion chambers F23R3/58
• CPC Class Codes: F23R 3/04 Air inlet arrangements
• CPC Class Codes: F23R 3/06 Arrangement of apertures along the flame tube
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes a gas path surface and a non-gas path surface. The component defines at least one cavity that extends at least partially inside of the component. A first plurality of cooling holes extends from the at least one cavity to at least one of the first mate face and the second mate face and a second plurality of cooling holes extends from either the at least one cavity or the non-gas path surface to the gas path surface.

Claims

1. A component for a gas turbine engine, comprising:

a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes at least one gas path surface and at least one non-gas path surface;

said component defining at least one cavity that extends at least partially inside of the component;

a first plurality of cooling holes extending from said at least one cavity to at least one of said first mate face and said second mate face; and

a second plurality of cooling holes extending from either said at least one cavity or said at least one non-gas path surface to said at least one gas path surface.

2. The component as recited in claim 1, wherein said platform is part of a blade.

3. The component as recited in claim 1, wherein said platform is part of a vane.

4. The component as recited in claim 1, wherein said platform is part of a blade outer air seal (BOAS).

5. The component as recited in claim 1, wherein said first plurality of cooling holes are round holes.

6. The component as recited in claim 1, comprising a first cavity and a second cavity that extend inside of the component, wherein a first portion of each of said first plurality of cooling holes and said second plurality of cooling holes are in fluid communication with said first cavity and a second portion of each of said first plurality of cooling holes and said second plurality of cooling holes are in fluid communication with either said second cavity or said at least one non-gas path surface.

7. The component as recited in claim 1, wherein a first portion of said first plurality of cooling holes is disposed at a first axial region of a pressure side mate face and a second portion of said first plurality of cooling holes is disposed at a second axial region of said pressure side mate face.

8. The component as recited in claim 7, wherein said second plurality of cooling holes are disposed between said first portion and said second portion of said first plurality of cooling holes.

9. The component as recited in claim 1, wherein each of said first plurality of cooling holes and said second plurality of cooling holes are located adjacent a pressure side of said platform.

10. The component as recited in claim 1, wherein at least one cooling hole of said first plurality of cooling holes includes a greater diameter than the remaining cooling holes of said first plurality of cooling holes.

11. The component as recited in claim 1, wherein an outlet portion of said second plurality of cooling holes include a diffusion angle of between 5 and 15 degrees.

12. The component as recited in claim 1, wherein at least one cooling hole of said first plurality of cooling holes is disposed at a position that is about 20% to 30% of an axial length of said first mate face.

13. A gas turbine engine, comprising:

a compressor section;

a combustor section in fluid communication with said compressor section;

a turbine section in fluid communication said combustor section;

a component positioned within at least one of said compressor section and said turbine section, wherein said component includes;

a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes at least one gas path surface and at least one non-gas path surface;

said component defining at least one cavity that extends at least partially inside of said component;

a first plurality of cooling holes extending through at least one of said first mate face and said second mate face; and

a second plurality of cooling holes extending from either said at least one cavity or said at least one non-gas path surface to said at least one gas path surface.

14. The gas turbine engine as recited in claim 13, wherein said component is a turbine blade.

15. The gas turbine engine as recited in claim 13, wherein said second plurality of cooling holes are disposed between a first portion and a second portion of said first plurality of cooling holes.

16. The gas turbine engine as recited in claim 13, wherein said at least one cavity includes a first cavity and a second cavity and a first portion of said first plurality of cooling holes extend from said first cavity and a second portion of said first plurality of cooling holes extend from said second cavity.

17. A method of cooling a component of a gas turbine engine, comprising the steps of:

circulating a cooling airflow through a cavity of the component;

directing a first portion of the cooling airflow into a first plurality of cooling holes that extend from the cavity to a mate face of the component; and

directing a second portion of the cooling airflow into a second plurality of cooling holes that extend from either the cavity or a non-gas path surface to a gas path surface of the component.

18. The method as recited in claim 17, comprising the step of:

expelling the first portion of the cooling airflow from the first plurality of cooling holes such that the first portion of the cooling airflow impinges on a second component.

19. The method as recited in claim 17, comprising a step of:

blocking hot combustion gases from a core flow path of the gas turbine engine from ingestion into the mate face with the second portion of the cooling airflow.

20. The method as recited in claim 17, wherein the second plurality of cooling holes are disposed between a first portion and a second portion of the first plurality of cooling holes.