Conformal seal bow wave cooling

Patent metadata

Field	Value
Publication number	US-10738701-B2
Application number	US-201715690555-A
Country	US
Kind code	B2
Filing date	Aug 30, 2017
Priority date	Aug 30, 2017
Publication date	Aug 11, 2020
Grant date	Aug 11, 2020

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 seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A combustor assembly and method are also disclosed.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a combustor including a radially outward extending rib;
a turbine section in fluid communication with the combustor, the turbine section including a first vane stage with a forward face aft of the combustor; and
a seal disposed between an axial surface of the radially extending rib of the combustor and the forward face of the first vane stage, the seal including a forward surface abutting the axial surface of the radially extending rib of the combustor, an aft surface, that is parallel to the forward surface, abutting the forward face of the first vane stage, a radially outer surface disposed parallel to a radially inner surface, a plurality of circumferentially spaced apart slots facing the forward face of the first vane stage and openings that extend through the seal and are disposed in groups spaced circumferentially to align with a corresponding one of the circumferentially spaced apart slots, the openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage.

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

3. The gas turbine engine as recited in claim 2 , wherein the plurality of openings are angled relative to the radially outer surface.

4. The gas turbine engine as recited in claim 1 , wherein the seal assembly includes an alignment slot that aligns the seal assembly circumferentially with the first vane stage.

5. An assembly for a gas turbine engine comprising:
a combustor including an aft end and a radially outward extending rib; and
a seal abutting an axial surface of the radially outward extending rib from the aft end across a gap between the combustor and a first turbine vane stage, the seal including a forward surface abutting the axial surface of the radially outward extending rib of the combustor, an aft surface parallel to the forward surface and abutting the first turbine vane stage, a radially outer surface disposed parallel to a radially inner surface, a plurality of circumferentially spaced apart slots facing a forward face of the first vane stage and a plurality of openings that extend through the seal and are disposed in groups spaced circumferentially to align with a corresponding one of the circumferentially spaced apart slots for communicating cooling airflow into the gap.

6. The combustor assembly as recited in claim 5 , wherein the plurality of slots disposed at circumferential positions that align with a leading edge of a respective vane from a plurality of vanes of the first turbine stage.

7. The combustor assembly as recited in claim 6 , wherein the plurality of openings are angled relative to the radially outer surface.

8. The combustor assembly as recited in claim 5 , wherein the seal includes an alignment slot that aligns the seal assembly circumferentially with the first vane stage.

9. A method of cooling an assembly comprising:
providing a combustor including an aft end and a radially outward extending rib;
providing a seal abutting an axial surface of the radially outward extending rib from the aft end across a gap between the combustor and a first turbine vane stage, the seal including a forward surface abutting the axial surface of the radially outward extending rib of the combustor, an aft surface parallel to the forward surface and abutting the first turbine vane stage, a radially outer surface disposed parallel to a radially inner surface, a plurality of circumferentially spaced apart slots facing a forward face of the first vane stage and a plurality of openings that extend through the seal and are disposed in groups spaced circumferentially to align with a corresponding one of the circumferentially spaced apart slots for communicating cooling airflow into the gap; and
communicating cooling air flow into the gap through the plurality of openings in the seal.

10. The method as recited in claim 9 , wherein the plurality of circumferential slots align with a leading edge of a turbine vanes within the turbine vane stage.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Porter Steven D
• Mcmahon Shawn M
• Zacchera Kevin
• Wadsworth Noah
• Whitfield Christopher
• Bradshaw Sean D
• Moura Dennis M

Classifications

• F01D11/005

  Sealing means between non relatively rotating elements · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F02C7/28

  Arrangement of seals · CPC title

• F23R3/42

  characterised by the arrangement or form of the flame tubes or combustion chambers · CPC title

• F02C7/18 Primary

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title