Conformal seal and vane bow wave cooling

Patent metadata

Field	Value
Publication number	US-11041391-B2
Application number	US-202016852728-A
Country	US
Kind code	B2
Filing date	Apr 20, 2020
Priority date	Aug 30, 2017
Publication date	Jun 22, 2021
Grant date	Jun 22, 2021

Abstract

A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A plurality of vane cooling passages extending through a forward flange of the first vane stage to the radial inner surface for communicating cooling airflow into the core flow path forward of the leading edge. A seal assembly is disposed between the combustor and the forward flange of the vane stage. The seal includes a plurality of circumferentially spaced apart slots and a plurality of seal cooling passages that extend from the radially outer surface to the radially inner surface at each of the plurality of circumferentially spaced apart slots.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a combustor including an aft end;
a vane stage including a leading edge, a forward flange, a radial outer surface and a radial inner surface, the radial inner surface defining a portion of a core flow path aft of the combustor, the vane stage including a plurality of vane cooling passages extending through the forward flange to the radial inner surface for communicating cooling airflow into the core flow path forward of the leading edge; and
a seal assembly disposed between the combustor and the forward flange of the vane stage, the seal assembly including a radially outer surface, a radially inner surface, a plurality of circumferentially spaced apart slots and a plurality of seal cooling passages, the plurality of seal cooling passages extend from the radially outer surface to the radially inner surface at each of the plurality of circumferentially spaced apart slots for communicating cooling airflow into a gap between the aft end of the combustor and the forward face of the vane stage.

2. The gas turbine engine as recited in claim 1 , wherein the vane stage includes a plurality of vanes and the plurality of circumferentially spaced apart slots are circumferentially aligned with the leading edge of each of the plurality of vanes.

3. The gas turbine engine as recited in claim 2 , wherein the each of the vane cooling passages includes an inner opening on the radially inner surface of the vane and a centerline through the inner opening is spaced a distance from a forward face of the forward flange that is no less than a diameter of the inner opening.

4. The gas turbine engine as recited in claim 3 , wherein the centerline of the inner opening of each the plurality of vane cooling passages is spaced a distance from a forward face of the forward flange that is no greater than three times the diameter of the opening.

5. The gas turbine engine as recited in claim 4 , wherein the plurality of vane cooling passages are grouped in annular sectors about a complete circumference of the vane stage, wherein a circumferential spacing between the inner openings within each of the annular sectors is no more than six hole diameters and no less than four hole diameters.

6. The gas turbine engine as recited in claim 5 , wherein each of the annular sectors comprises an arc length between one percent and two percent of the complete circumference of the vane stage.

7. The gas turbine engine as recited in claim 6 , wherein each of the annular sectors includes an annular area on the radial inner surface and the inner openings for each of the plurality of vane passages comprise an open area within the arc length that is between 18 percent and 22 percent of the annular area.

8. The gas turbine engine as recited in claim 7 , wherein each of the annular sectors is aligned circumferentially with a leading edge of one of the plurality of vanes.

9. The gas turbine engine as recited in claim 8 , wherein the seal assembly includes an aft face abutting the forward face of each of the plurality of vanes, the aft face is transverse to the radially inner surface and the radially outer surface.

10. The gas turbine engine as recited in claim 9 , wherein the aft face includes a wearing end portion extending axially aft from the seal assembly and configured to wear down during initial operation to provide a seal against the forward face.

11. The gas turbine engine as recited in claim 10 , wherein the seal assembly includes an alignment slot that aligns the seal assembly circumferentially with the first vane stage.

12. A first vane stage assembly for a gas turbine engine comprising:
a plurality of vanes spaced circumferentially about an engine longitudinal axis, each of the plurality of vane include a leading edge, a forward flange radially outward of the leading edge, a radial outer surface and a radial inner surface, the radial inner surface defining portion of a core flow path and a plurality of vane cooling passages extend through the forward flange to the radial inner surface for communicating cooling airflow into the core flow path forward of the leading edge;
a seal abutted against the forward flange for containing core gas flow within the core flow path, the seal including a radially outer surface, a radially inner surface, a plurality of circumferentially spaced apart slots and a plurality of seal cooling passages, the plurality of circumferentially spaced apart slots are each aligned with the leading edge of one of the plurality of vanes and the plurality of seal cooling passages extend from the radially outer surface to the radially inner surface at each of the plurality of circumferentially spaced apart slots for communicating cooling airflow into a gap forward of the plurality of vanes.

13. The first vane stage assembly as recited in claim 12 , wherein the each of the vane cooling passages includes an inner opening on the radially inner surface of the vane and a centerline through the inner opening is spaced a distance from a forward face of the forward flange that is no less than a diameter of the inner opening and no more than three times the diameter of the inner opening.

14. The first vane stage assembly as recited in claim 13 , wherein the plurality of vane cooling passages are grouped in annular sectors about a complete circumference of the vane stage, wherein a circumferential spacing between the inner openings within each of the annular sectors is no more than six hole diameters and no less than four hole diameters.

15. The first vane stage assembly as recited in claim 14 , wherein each of the annular sectors comprises an arc length between one percent and two percent of the complete circumference of the vane stage.

16. The first vane stage assembly as recited in claim 15 , wherein each of the annular sectors includes an annular area on the radial inner surface and the inner openings for each of the plurality of vane passages comprise an open area within the arc length that is between 18 percent and 22 percent of the annular area.

17. The first vane stage assembly as recited in claim 16 , wherein the seal includes an aft face abutting the forward face of each of the plurality of vanes, the aft face is transverse to the radially inner surface and the radially outer surface and includes a wearing end portion extending axially aft from the seal assembly and configured to wear down during initial operation to provide a seal against the forward face.

18. A gas turbine engine comprising:
a combustor including a forward end and an aft end;
a first turbine vane stage including an forward flange with a forward face, the first vane stage defining a portion of a core flow path; and
a seal abutting the forward face of the first vane stage and the aft end of the combustor and extending axially across a gap between the combustor and the first vane stage, wherein the seal includes a plurality of seal openings that extend from a radially outer surface to a radially inner surface and the first vane stage includes a plurality of vane openings through the axial face for communicating cooling airflow into the core flow path.

19. The gas turbine engine as recited in claim 18 , wherein the seal includes a plurality of slots disposed at spaced apart circumferential positions corresponding with a leading edge of each vane of the first turbine vane stage.

20. The gas turbine engine as recited in claim 19 , wherein the seal includes an aft face that seals against a forward face of the first turbine vane stage, the af

Assignees

• Raytheon Tech Corp

Inventors

• Porter Steven D
• Mcmahon Shawn M
• Zacchera Kevin
• Wadsworth Noah
• Whitfield Christopher
• Bradshaw Sean D
• Moura Dennis M
• Bitzko David
• Schneider Alex J

Classifications

• F05D2240/121

  related to the leading edge of a stator vane · CPC title

• F05D2260/20

  Heat transfer, e.g. cooling · CPC title

• F01D9/023 Primary

  Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings · CPC title

• F05D2240/55

  Seals · CPC title

• F01D25/24

  Casings (modified for heating or cooling F01D25/14); Casing parts, e.g. diaphragms, casing fastenings (casings for rotary machines or engines in general F16M {; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045}) · CPC title