Seal arrangement for turbine engine component

Patent metadata

Field	Value
Publication number	US-10458264-B2
Application number	US-201514704278-A
Country	US
Kind code	B2
Filing date	May 5, 2015
Priority date	May 5, 2015
Publication date	Oct 29, 2019
Grant date	Oct 29, 2019

Abstract

A component for a gas turbine engine according to an example of the present disclosure includes, among other things, a body including a cold side surface adjacent to a mate face. A plurality of ridges extends from the cold side surface. A seal member abuts the plurality of ridges to define a plurality of cooling passages. The seal member is configured to move between a first position and a second position relative to the plurality of ridges. Each of the plurality of cooling passages includes a first inlet defined at the first position and a second, different inlet defined at the second position. A method of sealing between adjacent components of a gas turbine engine is also disclosed.

First claim

What is claimed is:

1. A gas turbine engine, comprising:
a first component including a first set of ridges protruding in a radial direction from a first cold side surface adjacent to a first mate face;
a second component including a second set of ridges protruding in the radial direction from a second cold side surface adjacent to a second mate face, the second mate face circumferentially adjacent to the first mate face to define a leakage gap; and
a seal member abutting the first set of ridges to define a first set of cooling channels, and abutting the second set of ridges to define a second set of cooling channels, the seal member spaced apart from the first cold side surface and the second cold side surface;
wherein the seal member includes a sealing surface that abuts the first set and the second set of ridges, and an outer surface spaced apart from the sealing surface;
wherein each ridge of the first and second sets of ridges includes an elongated body extending in a circumferential direction between a proximal end and a distal end, the proximal end adjacent to the leakage gap and the distal end spaced apart from the leakage gap, a first surface of the elongated body abutting the seal member, and the seal member is contained in the circumferential direction within the distal ends of the first and second sets of ridges; and
wherein the distal ends are free of any retention features radially aligned with the seal member in the radial direction such that the seal member is moveable circumferentially past the distal end.

2. The gas turbine engine as recited in claim 1 , wherein each of the first set and the second set of cooling channels includes an inlet spaced apart from the leakage gap and an outlet adjacent to the leakage gap.

3. The gas turbine engine as recited in claim 1 , wherein the seal member defines a first width in the circumferential direction, the distal ends of the first set and the second set of ridges define a second width in the circumferential direction, and a ratio of the first width to the second width is equal to or less than 0.8.

4. The gas turbine engine as recited in claim 1 , wherein each of the first component and the second component is one of an airfoil and a blade outer air seal (BOAS).

5. The gas turbine engine as recited in claim 4 , wherein the first component is an airfoil, the airfoil includes an airfoil section extending from a platform, the platform includes an upper surface bounding a core flow path and an undersurface bounding a cooling cavity, and the first cold side surface is located at the undersurface of the platform such that the first set of ridges are raised directly from the undersurface.

6. The gas turbine engine as recited in claim 5 , wherein:
the seal member bounds the cooling cavity;
each of the first and the second sets of cooling channels includes an inlet spaced apart from the leakage gap and an outlet adjacent to the leakage gap; and
the inlet is unbounded in the circumferential direction such that the inlet is in direct fluid communication with the cooling cavity.

7. The gas turbine engine as recited in claim 2 , wherein:
the first component is an airfoil, the airfoil includes an airfoil section extending from a platform, the platform includes an upper surface bounding a core flow path and an undersurface bounding a cooling cavity, and the first cold side surface is located at the undersurface of the platform such that the first set of ridges are raised directly from the undersurface.

8. The gas turbine engine as recited in claim 7 , wherein the seal member extends between a leading edge and a trailing edge, and the leading and trailing edges are spaced apart from the undersurface of the platform.

9. The gas turbine engine as recited in claim 7 , wherein the airfoil is a turbine blade.

10. A method of sealing between adjacent components of a gas turbine engine, comprising:
providing a first component, including a first set of ridges protruding in a radial direction from a first cold side surface adjacent to a first mate face;
providing, a second component including a second set of ridges protruding in the radial direction from a second cold side surface adjacent to a second mate face, the second mate face circumferentially adjacent to the first mate face to define a leakage gap in a circumferential direction between the first and second mate faces;
positioning a feather seal across the leakage gap and along the first set and the second set of ridges to respectively define first and second sets of cooling passages, wherein the feather seal is spaced apart from the first cold side surface and the second cold side surface, and wherein the feather seal includes a sealing surface that abuts the first set and the second set of ridges and includes an outer surface spaced apart from the scaling surface;
wherein each of the first set and the second set of ridges includes an elongated body extending in the circumferential direction between a proximal end and a distal end, the proximal end adjacent to the leakage gap and the distal end spaced apart from the leakage gap, a first surface of the elongated body abutting the feather seal, and the feather seal is contained in the circumferential direction within the distal ends of the first set and the second set of ridges of the first and second components; and
wherein the distal ends are free of any retention features radially aligned with the feather seal in the radial direction such that the feather seal is moveable circumferentially past the distal end.

11. The method as recited in claim 10 , comprising communicating coolant from a cooling cavity to the first set and the second set of cooling passages when an edge face of the feather seal is substantially aligned with the distal ends of at least some of the first set and the second set of ridges.

12. The method as recited in claim 10 , wherein each of the first set and the second set of cooling passages includes an inlet that is spaced apart from the mate first and second mate faces at each position of the feather seal relative to the first set and the second set of ridges.

13. The method as recited in claim 12 , further comprising:
moving the feather seal between a first circumferential position and a second circumferential position relative to the first set and the second set of ridges when in an installed position, each of the first set and the second set of cooling passages including a first inlet defined at the first circumferential position and a second, different inlet defined at the second circumferential position.

Assignees

• United Technologies Corp

Inventors

• Mongillo Jr Dominic J
• Spangler Brandon W
• Thornton Lane

Classifications

• F05D2240/11

  Shroud seal segments · CPC title

• F05D2240/57

  Leaf seals · CPC title

• F01D11/005

  Sealing means between non relatively rotating elements · CPC title

• F01D11/006 Primary

  Sealing the gap between rotor blades or blades and rotor · CPC title

• F01D25/246

  Fastening of diaphragms or stator-rings · CPC title