Blade outer air seal for a gas turbine engine

Patent metadata

Field	Value
Publication number	US-2017306784-A1
Application number	US-201715444971-A
Country	US
Kind code	A1
Filing date	Feb 28, 2017
Priority date	Jul 27, 2012
Publication date	Oct 26, 2017
Grant date	—

Abstract

A blade outer air seal (BOAS) for a gas turbine engine includes, among other things, a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion. The BOAS includes a trough disposed on the radially inner face and an abradable seal received within the trough. The trough is open to expose a leading edge of the abradable seal to a core flow path of the gas turbine engine.

First claim

What is claimed is:

1 . A blade outer air seal (BOAS) for a gas turbine engine, comprising:
a monolithic seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion; a trough disposed on said radially inner face; and an abradable seal received within said trough, wherein said trough is open to expose a leading edge of said abradable seal to a core flow path of the gas turbine engine and a trailing edge of said abradable seal is axially constrained by said trough.

2 . The BOAS as recited in claim 1 , comprising a reverse retention hook that is disposed at one of said leading edge portion and said trailing edge portion, and said reverse retention hook extends in a direction from said leading edge portion toward said trailing edge portion.

3 . The BOAS as recited in claim 1 , wherein said abradable seal is axially constrained at said trailing edge by a radial wall of said trough and is completely unconstrained at said leading edge by any portion of said trough or said body.

4 . The BOAS as recited in claim 1 , comprising an axial retention feature and a radial retention feature disposed at one of said leading edge portion and said trailing edge portion.

5 . The BOAS as recited in claim 4 , wherein said radial retention feature extends from a vertical wall of said axial retention feature.

6 . The BOAS as recited in claim 1 , wherein said abradable seal is axially constrained at said trailing edge by a radial wall of said trough.

7 . The BOAS as recited in claim 1 , wherein said abradable seal is a plasma sprayed seal.

8 . The BOAS as recited in claim 1 , wherein said BOAS is a first stage high pressure compressor BOAS.

9 . The BOAS as recited in claim 1 , wherein said seal body is attached to a casing that includes at least a first surface and a second surface that is different from said first surface, wherein each of said first surface and said second surface include a thermal barrier coating.

10 . The BOAS as recited in claim 9 , wherein said thermal barrier coating of said first surface axially overlaps said thermal barrier coating of said second surface such that a portion of said first surface is axially and radially displaced from a portion of said second surface.

11 . The BOAS as recited in claim 1 , wherein an entirety of said leading edge of said abradable seal is exposed to said core flow path.

12 . The BOAS as recited in claim 1 , wherein said trough is a cavity formed by said radially inner face.

13 . A gas turbine engine, comprising:
a compressor section; a combustor section in fluid communication with said compressor section along a core flow path; a turbine section in fluid communication with said combustor section along said core flow path; and a blade outer air seal (BOAS) that can be received relative to a casing associated with at least one of said compressor section and said turbine section, wherein said BOAS includes:
a seal body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion;
a trough disposed on said radially inner face; and
an abradable seal received within said trough, wherein said trough is open to expose a leading-most edge of said abradable seal to said core flow path.

14 . The gas turbine engine as recited in claim 13 , comprising a reverse retention hook that extends in a direction from one of said leading edge portion and said trailing edge portion toward the other of said leading edge portion and said trailing edge portion.

15 . The gas turbine engine as recited in claim 14 , wherein said reverse retention hook is positioned at said leading edge portion and is received within a groove of said casing.

16 . The gas turbine engine as recited in claim 13 , comprising an axial retention feature and a radial retention feature disposed at one of said leading edge portion and said trailing edge portion.

17 . The gas turbine engine as recited in claim 13 , comprising a thermal barrier coating applied to at least a first surface and a second surface of said casing that is different from said first surface, wherein said thermal barrier coating on said first surface axially overlaps said thermal barrier coating on said second surface.

18 . The gas turbine engine as recited in claim 13 , wherein an entirety of said leading-most edge of said abradable seal is exposed to said core flow path.

19 . The gas turbine engine as recited in claim 13 , wherein said trough is a cavity formed by said radially inner face.

20 . The gas turbine engine as recited in claim 13 , comprising a thermal barrier coating disposed on a first surface and a second surface of said casing, a portion of said first surface being axially and radially displaced from a portion of said second surface such that said thermal barrier coating of said first surface axially overlaps said thermal barrier coating of said second surface.

Assignees

• United Technologies Corp
• MTU Aero Engines AG

Inventors

• Leslie Nicholas R
• Maalouf Fadi S
• Zotz Georg
• Humhauser Werner

Classifications

• F01D11/127

  with a deformable or crushable structure, e.g. honeycomb · CPC title

• F01D11/125

  with a reinforcing structure · CPC title

• F05D2230/312

  by plasma spraying · CPC title

• F01D11/122 Primary

  with erodable or abradable material (blades having cutting or grinding tips F01D5/20) · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title