Granted Patent B2

US 12,371,997 · App. 18/656,892 · Granted Jul 29, 2025

Inventors: Emma J. Place (New Hartford, CT); Christopher J. Peters (Los Lunas, NM); Jeffrey J. DeGray (Somers, CT); Dominic J. Mongillo (West Hartford, CT); Timothy J. Jennings (New Britain, CT); Raymond Surace (Newington, CT); Jeremy B. Fredette (West Hartford, CT)

Assignee: RTX CORPORATION

F01D5/18F05D2220/32F05D2230/60F05D2260/20

Abstract

Claims (55)

1. A casting core assembly for a gas turbine engine component comprising:

a skin core corresponding to a first cooling passage of an airfoil, the first cooling passage including a first section and a tip flag section joined at a first bend, the tip flag section dimensioned to extend from the first bend to a trailing edge of the airfoil, the skin core including a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section, and the tip flag portion including a row of protrusions corresponding to a first row of exit slots along the trailing edge of the airfoil; and

a serpentine core corresponding to a serpentine cooling passage;

wherein the skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil, and the skin core includes a plurality of branched sections corresponding to a plurality of branched paths along the first section, the plurality of branched sections bounding the at least one arcuate slot such that the plurality of branched sections join along the tip flag portion; and

wherein the skin core and the serpentine core are arranged in spaced relationship such that the first cooling passage and the serpentine cooling passage are opposite sides along an internal wall of the airfoil relative to a thickness direction.

2. The casting core assembly as recited in claim 1 , wherein the at least one arcuate slot extends between a first end and a second end, and the first portion includes at least one elongated slot bounded by an adjacent pair of the branched sections, and the first end of the at least one arcuate slot is aligned with the at least one elongated slot relative to a chordwise direction.

3. The casting core assembly as recited in claim 1 , wherein the at least one arcuate slot extends between a first end and a second end, the skin core includes at least one bridge spanning between an adjacent pair of the branched sections such that the at least one arcuate slot is interrupted between the first and second ends, and the at least one bridge corresponds to at least one crossover passage interconnecting an adjacent pair of the branched paths.

4. The casting core assembly as recited in claim 1 , further comprising:

a leading edge core corresponding to a leading edge cooling passage bounded by an external wall along a leading edge of the airfoil; and

a trailing edge core corresponding to a trailing edge cooling passage including a second set of exit ports along the trailing edge of the airfoil, wherein the tip flag portion of the skin core is at least partially aligned with the trailing edge core relative to the thickness direction, and the serpentine core is spaced apart from and forward of the trailing edge core relative to the chordwise direction.

5. The casting core assembly as recited in claim 4 , wherein the trailing edge core includes a second row of protrusions corresponding to a second row of exit ports along the trailing edge of the airfoil.

6. The casting core assembly as recited in claim 4 , further comprising:

at least one connector that joins the leading edge core and the serpentine core, the at least one connector corresponding to at least one crossover passage extending between the leading edge cooling passage and the serpentine cooling passage.

7. A casting core assembly for a gas turbine engine component comprising:

a skin core corresponding to a first cooling passage of an airfoil, the first cooling passage including a first section and a tip flag section joined at a first bend, the tip flag section dimensioned to extend from the first bend to a trailing edge of the airfoil, the skin core including a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section, and the tip flag portion including a row of protrusions corresponding to a first row of exit slots along the trailing edge of the airfoil;

wherein the skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil, and the skin core includes a plurality of branched sections corresponding to a plurality of branched paths along the first section, the plurality of branched sections bounding the at least one arcuate slot such that the plurality of branched sections join along the tip flag portion; and

wherein the skin core includes a protrusion extending from the first bend, the protrusion corresponds to a purge passage interconnecting the first cooling passage and an aperture along an external surface of the airfoil, and the purge passage is dimensioned to eject particulate from the first cooling passage in operation.

8. The casting core assembly as recited in claim 7 , wherein the purge passage is dimensioned to eject particulate from the first cooling passage into a tip pocket of the airfoil.

9. The casting core assembly as recited in claim 1 , wherein the airfoil is a turbine blade.

10. A method of forming an airfoil for a gas turbine engine comprising:

forming a skin core;

forming a trailing edge core;

forming a serpentine core;

assembling the skin core, the serpentine core and the trailing edge core together to establish a core assembly; and

forming an airfoil around the core assembly, including forming an airfoil section including an external wall and an internal wall;

wherein the skin core corresponds to a first cooling passage of the airfoil, the first cooling passage includes a first section and a tip flag section, the first section extends in a radial direction, and the tip flag section extends in a chordwise direction from the first section to a trailing edge of the airfoil;

wherein the trailing edge core corresponds to a trailing edge cooling passage adjacent to the trailing edge of the airfoil;

wherein the serpentine core corresponds to a serpentine cooling passage;

wherein the skin core and the serpentine core are arranged in spaced relationship such that the first cooling passage and the serpentine cooling passage are on opposite sides along the internal wall relative to a thickness direction; and

wherein the skin core includes at least one arcuate slot corresponding to at least one turning vane outward of the trailing edge cooling passage relative to the radial direction, the skin core includes a plurality of branched sections corresponding to a plurality of branched paths along the first section of the first cooling passage that join along the tip flag section.

11. The method as recited in claim 10 , wherein:

the first section and the tip flag section are joined at a first bend; and

the at least one turning vane is established along the first bend.

12. The method as recited in claim 10 , comprising:

forming a leading edge core;

wherein the assembling step includes assembling the skin core, the serpentine core, the leading edge core and the trailing edge core together to establish the core assembly; and

wherein the leading edge core corresponds to a leading edge cooling passage adjacent to a leading edge of the airfoil.

13. The method as recited in claim 12 , wherein:

the external wall defines pressure and suction sides extending in a chordwise direction between the leading edge and the trailing edge, and the pressure and suction sides are spaced apart in the thickness direction between the leading edge and the trailing edge.

14. The method as recited in claim 13 , wherein:

the tip flag section of the first cooling passage is established between the suction side of the airfoil and a first side of the opposite sides of the internal wall relative to the thickness direction, and the serpentine cooling passage is established between the pressure side of the airfoil and a second side of the opposite sides of the internal wall opposed to the first side relative to the thickness direction.

15. The method as recited in claim 14 , wherein:

the step of forming the airfoil includes forming a platform section and a root section, the airfoil section extends outwardly from the platform section to a tip portion relative to the radial direction, the root section extends inwardly from the platform section relative to the radial direction and is dimensioned to mount the airfoil to a rotatable hub;

the internal wall extends inwardly from the tip portion relative to the radial direction;

the tip flag section is established along the tip portion of the airfoil; and

the branched paths are dimensioned to branch outwardly from a trunk of the first section at a position inward of the platform section relative to the radial direction.

16. The method as recited in claim 15 , wherein:

the assembly step includes coupling the skin core, the serpentine core, the leading edge core and the trailing edge core to each other at a position corresponding to the root section.

17. The method as recited in claim 10 , wherein:

the at least one arcuate slot includes a first arcuate slot and a second arcuate slot that extend along opposite sides of one of the branched sections.

18. The method as recited in claim 17 , wherein:

the first section and the tip flag section are joined at a first bend;

the first arcuate slot is dimensioned to establish a respective turning vane that terminates in the tip flag section; and

the second arcuate slot is dimensioned to establish a respective turning vane that terminates in the first bend.

19. The method as recited in claim 10 , wherein the airfoil is a turbine blade.

Continuity (3)

References Cited (56)