Axial-radial cooling slots on inner air seal

Patent metadata

Field	Value
Publication number	US-11041396-B2
Application number	US-201916528238-A
Country	US
Kind code	B2
Filing date	Jul 31, 2019
Priority date	Oct 6, 2016
Publication date	Jun 22, 2021
Grant date	Jun 22, 2021

Abstract

An air seal may comprise an annular ring defined by at least a proximal surface, a distal surface, an aft side and a forward side. A channel may be disposed in the forward side of the air seal and/or the aft side of the air seal and may extend between the proximal surface and the distal surface. An additional channel extending from at least one of the forward side or the aft side may be disposed in the distal surface. The channel and the additional channel may be circumferentially in line. The channels may define a flow path for direction cooling air from a proximal side of the air seal to a distal side of the air seal. The radial channel may interface with the axial channel at an edge of the air seal.

First claim

What is claimed is:

1. An air seal for a gas turbine engine comprising:
an annular ring defined by at least a proximal surface, a distal surface, an aft side and a forward side;
a radial channel disposed in the air seal, the radial channel disposed in at least one of the forward side or the aft side, the radial channel extending between the proximal surface and the distal surface; and
an axial channel disposed in the distal surface, the axial channel extending from at least one of the forward side or the aft side and circumferentially in line with the radial channel;
wherein the radial channel and the axial channel interface at an edge,
wherein the radial channel and the axial channel are configured to direct a cooling air from a proximal side of the air seal to a distal side of the air seal for cooling the air seal.

2. The air seal of claim 1 , wherein the radial channel is disposed on the aft side and the axial channel extends from the aft side.

3. The air seal of claim 1 , wherein the radial channel is disposed on the forward side and the axial channel extends from the forward side.

4. The air seal of claim 1 , wherein the air seal is configured to receive the cooling air from an aperture disposed in a rotor disk leg, the rotor disk leg being located radially inward from the air seal.

5. The air seal of claim 1 , wherein the air seal is configured to be coupled between a forward rotor disk and an aft rotor disk.

6. The air seal of claim 1 , wherein the air seal comprises knife edges extending from the distal surface, the knife edges configured to interface with a proximal surface of a vane platform.

7. The air seal of claim 1 , wherein the air seal comprises a nickel-based alloy.

8. The air seal of claim 1 , wherein a cross-section area of the radial channel is greater than a cross-section area of the axial channel.

9. A gas turbine engine comprising:
a compressor section;
a combustor section;
a turbine section;
an aft blade disk;
a forward blade disk; and
an air seal coupled between the aft blade disk and the forward blade disk comprising:
an annular ring defined by at least a proximal surface, a distal surface, an aft side and a forward side;
a radial channel disposed in at least one of the forward side or the aft side and extending between the proximal surface and the distal surface; and
an axial channel disposed in the distal surface and extending from at least one of the forward side or the aft side and circumferentially in line with the radial channel;

wherein the radial channel and the axial channel interface at an edge,
wherein the radial channel and the axial channel are configured to direct a cooling air from a proximal side of the air seal to a distal side of the air seal for cooling the air seal.

10. The gas turbine engine of claim 9 , wherein the radial channel is disposed on the aft side and the axial channel extends from the aft side.

11. The gas turbine engine of claim 9 , wherein the radial channel is disposed on the forward side and the axial channel extends from the forward side.

12. The gas turbine engine of claim 9 , wherein the air seal is configured to receive the cooling air from an aperture disposed in a rotor disk leg, the rotor disk leg being located radially inward from the air seal.

13. The gas turbine engine of claim 9 , wherein the air seal is configured to be coupled between a forward rotor disk and an aft rotor disk.

14. The gas turbine engine of claim 9 , wherein the air seal comprises knife edges extending from the distal surface, the knife edges configured to interface with a proximal surface of a vane platform.

15. The gas turbine engine of claim 9 , wherein the air seal comprises a nickel-based alloy.

16. A method of manufacturing an air seal for a gas turbine engine comprising:
forming a radial channel in at least one of a forward side or an aft side of the air seal, the radial channel extending between a proximal surface and a distal surface;
forming an axial channel in a distal surface of the air seal, the axial channel extending from at least one of the forward side or the aft side and circumferentially in line with the radial channel,
wherein the radial channel and the axial channel interface at an edge,
wherein the radial channel and the axial channel are configured to direct a cooling air from a proximal side of the air seal to a distal side of the air seal for cooling the air seal.

17. The method of claim 16 , wherein the forming the radial channel is performed by milling the at least one of the forward side or the aft side of the air seal.

18. The method of claim 16 , wherein the forming the radial channel and the forming the axial channel provides the radial channel having a cross-section area which is greater than a cross-section area of the axial channel.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Taubler James W
• Gunderson Peter E
• Patel Sunny Anant
• Boursy Andrew P

Classifications

• F05D2230/31

  Layer deposition · CPC title

• F05D2300/17

  Alloys · CPC title

• F01D5/06

  Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections {(F01D5/022, F01D5/023 take precedence)} · CPC title

• F05D2230/13

  using lasers · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title