SHROUD FOR A GAS TURBINE ENGINE

• US 20180016924A1
• Filed: 07/18/2016
• Published: 01/18/2018
• Est. Priority Date: 07/18/2016
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Gallier, Kirk Douglas, Craig, III, Charles William
• Granted Patent: US 10,450,897 B2
• Time in Patent Office: Days
• 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.

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, 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 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 inner end of the nozzle to the 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, further comprising a radially inner surface and a radially outer surface, wherein a clearance gap is defined between each of a plurality of blades of the rotor blade stage and the radially inner surface.

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

8. The shroud segment of claim 1, further comprising a mounting rail near the forward end and a pin mount near the aft end for mounting the shroud segment within the gas turbine engine.

9. 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; 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 near a leading edge of each of the plurality of rotor blades and the aft end positioned near a trailing edge of each nozzle of the plurality of nozzles.

10. The gas turbine engine of claim 9, 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.

11. The gas turbine engine of claim 9, 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 and a second side extending axially from the forward end to the aft end, the first side of one shroud segment positioned adjacent the second side of another shroud segment to form the shroud.

13. The gas turbine engine of claim 9, 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.

14. The gas turbine engine of claim 13, 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 includes a radially outer nozzle end, each nozzle extending through an opening in a shroud segment such that the outer nozzle end of the nozzle extends radially outward beyond the outer surface of the shroud segment.

15. The gas turbine engine of claim 13, wherein each nozzle includes a radially inner nozzle end, the inner nozzle end received within the inner band of the nozzle stage.

16. The gas turbine engine of claim 15, 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 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.

12. The gas turbine engine of claim 12, 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.

17. 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.

18. The method of claim 17, 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.

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

20. The method of claim 17, 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.