Shroud for a gas turbine engine

• US 10,450,897 B2
• Filed: 07/18/2016
• Issued: 10/22/2019
• Est. Priority Date: 07/18/2016
• Status: Expired due to Fees

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Gallier, Kirk Douglas, Craig, III, Charles William
• Primary Examiner(s): Eastman, Aaron R
• Application Number:
  US15/212,337
• Publication Number:
  US 20180016924A1
• Time in Patent Office: 1,191 Days
• Field of Search: 4151731
• CPC Class Codes: F01D 11/005 Sealing means between non relatively rotating elements
• CPC Class Codes: F01D 25/246 Fastening of diaphragms or stator-rings
• CPC Class Codes: F01D 9/042 fixing blades to stators fixing stator-rings in the casing or to each other F01D25/246
• CPC Class Codes: F05D 2300/6033 Ceramic matrix composites [CMC]
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

Shrouds and shroud segments for gas turbine engines are provided. In one embodiment, a shroud segment for a gas turbine engine having a rotor blade stage and a nozzle stage is provided. The shroud segment comprises a forward end defining an outer wall of the rotor blade stage and an aft end defining an outer wall of the nozzle stage. The aft end defines at least a portion of an opening therethrough for receipt of a nozzle, and the forward end and the aft end form a single, continuous component. In another embodiment, a gas turbine engine is provided, having a shroud with a forward end positioned near a leading edge of a plurality of rotor blades of a rotor blade stage and an aft end positioned near a trailing edge of a plurality of nozzles of a nozzle stage. Methods of assembling a gas turbine engine also are provided.

Claims

1. A shroud segment for a gas turbine engine, the gas turbine engine including a rotor blade stage and a nozzle stage, the shroud segment comprising:

a forward end, the forward end defining an outer wall of the rotor blade stage; and

an aft end, the aft end defining an outer wall of the nozzle stage, the aft end defining at least a portion of an opening therethrough for receipt of a nozzle,

wherein the forward end and the aft end form a single, continuous component, and

wherein the aft end is angled radially outward with respect to the forward end.

2. The shroud segment of claim 1, further comprising a first side extending axially from the forward end to the aft end, wherein the first side defines a portion of the opening and a side of an adjacent shroud segment defines the remainder of the opening.

3. The shroud segment of claim 1, further comprising a first side extending axially from the forward end to the aft end and a second side extending axially from the forward end to the aft end, wherein each of the first side and the second side are positioned adjacent another shroud segment, wherein the aft end defines two openings therethrough, and wherein each opening is configured for receipt of a nozzle.

4. The shroud segment of claim 1, further comprising a radially inner surface and a radially outer surface, and wherein the nozzle comprises a radially outer end and a radially inner end, the radially outer end extending radially beyond the radially outer surface of the shroud segment when the nozzle is received within the opening.

5. The shroud segment of claim 4, wherein the nozzle comprises an airfoil extending from the radially inner end of the nozzle to the radially outer end of the nozzle, the airfoil having an airfoil shape, and wherein the opening has a shape complementary to the airfoil shape.

6. The shroud segment of claim 1, wherein the shroud segment is formed from a ceramic matrix composite material.

7. The shroud segment of claim 1, further comprising a mounting rail at the forward end and a pin mount at the aft end, each of the mounting rail and the pin mount defining an aperture therein, each aperture configured for receipt of a mounting pin to mount the shroud segment within the gas turbine engine.

8. A gas turbine engine, comprising:

a rotor blade stage, the rotor blade stage including a plurality of circumferentially spaced rotor blades, each rotor blade having a tip;

a nozzle stage adjacent the rotor blade stage, the nozzle stage including a plurality of circumferentially spaced nozzles, each nozzle extending radially outward from an inner band, a circumferential stop positioned against a radially outer end of each nozzle

wherein the gas turbine engine defines an axial direction and a circumferential direction, and wherein the circumferential stop is angled with respect to both the axial direction and the circumferential direction; and

a shroud having

a forward end positioned radially outward from the tips of the rotor blades, and

an aft end positioned radially outward from the inner band of the nozzle stage,

wherein the shroud extends axially from the forward end to the aft end, the forward end positioned at a leading edge of each of the plurality of rotor blades and the aft end positioned at a trailing edge of each nozzle of the plurality of nozzles.

9. The gas turbine engine of claim 8, wherein the shroud comprises a plurality of shroud segments, and wherein each shroud segment defines a first side extending axially from the forward end to the aft end, the first side extending at a first non-orthogonal angle with respect to the forward end.

10. The gas turbine engine of claim 9, wherein each shroud segment defines a second side extending axially from the forward end to the aft end, the second side extending at a second non-orthogonal angle with respect to the forward end, and wherein the first side of one shroud segment is positioned adjacent the second side of another shroud segment to form the shroud.

11. The gas turbine engine of claim 10, wherein each first side of the shroud segments defines a portion of an opening and each second side of the shroud segments defines the remainder of the opening such that the adjacent first and second sides together define a plurality of openings in the shroud, each opening receiving a nozzle of the plurality of nozzles of the nozzle stage.

12. The gas turbine engine of claim 8, wherein the aft end of the shroud defines a plurality of openings, each opening receiving a nozzle of the plurality of nozzles of the nozzle stage.

13. The gas turbine engine of claim 12, wherein the shroud comprises a plurality of shroud segments, each shroud segment including a radially inner surface and a radially outer surface, and wherein each nozzle extends through an opening in a shroud segment such that the radially outer nozzle end of the nozzle extends radially outward beyond the outer surface of the shroud segment.

14. The gas turbine engine of claim 12, wherein each nozzle includes a radially inner nozzle end, the radially inner nozzle end received within the inner band of the nozzle stage.

15. The gas turbine engine of claim 14, wherein the inner band of the nozzle stage comprises a plurality of inner band segments, each inner band segment defining a depression, and wherein each nozzle includes a radially inner nozzle end, the radially inner nozzle end of a nozzle received within a depression in an inner band segment such that each nozzle extends from an inner band segment through an opening in the shroud.

16. A method of assembling a gas turbine engine, the gas turbine engine including a rotor blade stage and a nozzle stage, the method comprising:

positioning a shroud to form an outer wall of both the rotor blade stage and the nozzle stage;

inserting a nozzle through an opening defined in the shroud, the nozzle including an airfoil having opposite pressure and suction sides extending radially from an inner nozzle end to an outer nozzle end and extending axially from a leading edge to a trailing edge; and

positioning a circumferential stop against the outer nozzle end at the trailing edge of the airfoil wherein the gas turbine engine defines an axial direction and a circumferential direction, and wherein the circumferential stop is angled with respect to both the axial direction and the circumferential direction.

17. The method of claim 16, further comprising positioning one of the rotor blade stage or the nozzle stage immediately upstream of the other of the rotor blade stage or the nozzle stage such that one stage is an upstream stage and the other stage is a downstream stage.

18. The method of claim 17, wherein the shroud extends axially from leading edges of airfoils of the upstream stage to trailing edges of airfoils of the downstream stage.

19. The method of claim 16, further comprising attaching the shroud to a casing of the gas turbine engine, wherein the shroud comprises a mounting feature at a forward end of the shroud and a mounting feature at an aft end of the shroud for attaching the shroud to the casing.