ADDITIVELY DEPOSITED GAS TURBINE ENGINE COOLING COMPONENT

• US 20200131931A1
• Filed: 12/27/2019
• Published: 04/30/2020
• Est. Priority Date: 09/13/2016
• Status: Active Grant

• Current Assignee: Rolls-Royce Corporation (Rolls-Royce Holdings Plc)
• Original Assignee: Rolls-Royce Corporation (Rolls-Royce Holdings Plc)
• Inventors: Varney, Bruce Edward
• Granted Patent: US 11,248,491 B2
• Time in Patent Office: Days
• CPC Class Codes: B23P 15/04 turbine or like blades from several pieces
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 5/182 Transpiration cooling
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F05D 2220/32 in gas turbines
• CPC Class Codes: F05D 2230/20 essentially without removing material
• CPC Class Codes: F05D 2230/31 Layer deposition
• CPC Class Codes: F05D 2240/12 Fluid guiding means, e.g. vanes
• CPC Class Codes: F05D 2240/30 Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
• CPC Class Codes: F05D 2240/35 Combustors or associated equipment
• CPC Class Codes: F05D 2260/201 by impingement of a fluid
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/203 by transpiration cooling
• CPC Class Codes: F05D 2260/94 Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
• CPC Class Codes: F23R 3/002 Wall structures F23R3/02 and F23R3/007 take precedence
• CPC Class Codes: F23R 3/005 Combined with pressure or heat exchangers
• CPC Class Codes: F23R 3/06 Arrangement of apertures along the flame tube

Abstract

An example gas turbine engine component includes a component configured to separate a cooling air plenum from a heated gas environment. The component includes a substrate defining a surface, and a unitary structure. The unitary structure includes a cooling region and a cover layer. The cover layer defines a hot wall surface configured to face the heated gas environment. The cooling region is disposed between the cover surface and the substrate and includes a plurality of support structures extending between the cover layer and the surface of the substrate. At least some of the support structures define a respective bond surface bonded to the substrate at the surface of the substrate. An example technique for fabricating the gas turbine engine component includes additively depositing the unitary structure on the surface of the substrate.

Claims

1-7. -7. (canceled)

8. A method of fabricating a dual walled component configured to separate a cooling air plenum from a heated gas environment, the method comprising:

additively depositing a unitary structure on a surface of a substrate, wherein the substrate comprises a cold section part of the dual walled component, wherein the unitary structure comprises a cooling region and a cover layer and comprises a hot section part of the dual walled component, wherein the cover layer defines a hot wall surface configured to face the heated gas environment, wherein the cooling region is disposed between the cover layer and the substrate and comprises a plurality of support structures extending between the cover layer and the surface of the substrate, and wherein at least some of the support structures of the plurality of support structures define a respective bond surface bonded to the substrate at the surface of the substrate.

9. The method of claim 8, wherein the plurality of support structures defines a plurality of cooling channels, respective cooling channels of the plurality of cooling channels being between adjacent support structures of the plurality of support structures.

10. The method of claim 8, further comprising installing the dual walled component in a gas turbine engine.

11. The method of claim 10, wherein installing the dual walled component comprises bonding the dual walled component to a gas turbine engine surface.

12. The method of claim 11, wherein the bonding comprises diffusion bonding.

13. The method of claim 10, wherein the installing the dual walled component comprises connecting the dual walled component to an air-cooling system of the gas turbine engine.

15. The method of claim 8, wherein the dual walled component comprises a flame tube, a combustion ring, a combustor casing, a combustor guide vane, a turbine vane, a turbine disc, or a turbine blade.

16. The method of claim 9, wherein the hot wall surface defines a plurality of cooling apertures fluidly connected to the plurality of support structures, wherein respective cooling apertures of the plurality of cooling apertures are fluidly connected to respective cooling channels of the plurality of cooling channels for fluidly connecting the heated gas environment to the cooling region.

17. The method of claim 8, wherein additively depositing the unitary structure comprises:

directing a material stream and an energy stream at a focal region on the surface of the substrate, and

moving the focal region along a predetermined path.

18. The method of claim 17, wherein the energy beam comprises a laser beam and wherein the material stream comprises one or more of metal or alloy powder, wire, or ribbon.

19. The method of claim 8, wherein the bond surface inhibits transfer of heat across the bond surface from the unitary structure to the substrate.

20. The method of claim 8, wherein the unitary structure comprises a turbine vane defining an exterior surface comprising the hot wall surface and defining an internal hollow chamber comprising the cooling air plenum, wherein the cover layer comprises a coversheet, and wherein the substrate comprises a spar.

21. The method of claim 20, wherein the hot wall surface defines a plurality of cooling apertures fluidly connected to the plurality of support structures, wherein respective cooling apertures of the plurality of cooling apertures are fluidly connected to respective cooling channels of the plurality of cooling channels for fluidly connecting the heated gas environment to the cooling region, and wherein the plurality of cooling apertures are positioned along a leading edge of the turbine vane.

22. The method of claim 8, wherein the unitary component comprises a combustor component configured to separate the cooling air plenum from a combustion chamber comprising the heated gas environment.

23. The method of claim 8, wherein the cold section part defines a plurality of impingement apertures that extend through a thickness of the cold section part, wherein the plurality of support structures comprises a plurality of pedestals that connect the cold section part to the hot section part to define a plurality of cooling channels between the cold section part and the hot section part, and wherein the plurality of impingement apertures, the plurality of cooling channels, and the cooling apertures fluidly connect the cooling air plenum to the heated gas environment.

24. The method of claim 8, further comprising, before depositing the unitary structure on the surface of the substrate, depositing a filler composition on the surface of the substrate in a predetermined pattern that defines a plurality of cooling channels between the plurality of support structures.

25. The method of claim 24, further comprising removing the filler composition from the plurality of cooling channels by subjecting the filler composition to at least one of a thermal treatment, a leaching composition, or an oxidizing environment.

26. The method of claim 17, wherein the unitary structure is additively deposited by depositing successive layers that define successive cross-sections of predetermined channels to form a plurality of cooling apertures fluidly connected to the plurality of support structures, wherein respective cooling apertures of the plurality of cooling apertures are fluidly connected to respective cooling channels of the plurality of cooling channels for fluidly connecting the heated gas environment to the cooling region.

27. The method of claim 16, wherein the plurality of cooling apertures defines an angle of incidence to the hot wall surface that is between about 10 degrees and about 75 degrees.

28. The method of claim 8, wherein the dual walled component comprises a gas turbine engine component.

14. (canceled)