OUTER RIM SEAL ASSEMBLY IN A TURBINE ENGINE

• US 20140234076A1
• Filed: 02/15/2013
• Published: 08/21/2014
• Est. Priority Date: 02/15/2013
• Status: Active Grant

• Current Assignee: Siemens Energy Global GmbH & Co. KG (Siemens AG)
• Original Assignee: Nicholas F. Martin Jr
• Inventors: Lee, Ching-Pang, Tham, Kok-Mun, Shivanand, Manjit, Laurello, Vincent P., Azad, Gm Salam, Martin, Nicholas F. Jr.
• Granted Patent: US 8,939,711 B2
• Time in Patent Office: Days
• US Class Current: 415/116
• CPC Class Codes: F01D 11/001 for sealing space between stator blade and rotor
• CPC Class Codes: F01D 11/04 using sealing fluid, e.g. steam
• CPC Class Codes: F01D 11/122 with erodable or abradable material blades having cutting or grinding tips F01D5/20
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 5/081 Cooling fluid being directed on the side of the rotor disc or at the roots of the blades F01D5/087 takes precedence
• CPC Class Codes: F05D 2260/202 by film cooling

Abstract

A seal assembly between a hot gas path and a disc cavity in a turbine engine includes a non-rotatable vane assembly including a row of vanes and an inner shroud, a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, and an annular wing member located radially between the hot gas path and the disc cavity. The wing member extends generally axially from the blade assembly toward the vane assembly and includes a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof. The flow passages effect a pumping of cooling fluid from the disc cavity toward the hot gas path during operation of the engine.

Claims

1. A seal assembly between a hot gas path and a disc cavity in a turbine engine comprising:

a non-rotatable vane assembly including a row of vanes and an inner shroud;

a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor; and

an annular wing member located radially between the hot gas path and the disc cavity and extending generally axially from the blade assembly toward the vane assembly, the wing member including a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof, wherein the flow passages effect a pumping of cooling fluid from the disc cavity toward the hot gas path during operation of the engine.

2. The seal assembly according to claim 1, wherein the flow passages extend generally straight through the wing member in the radial direction.

3. The seal assembly according to claim 1, wherein the flow passages are at least one of curved and angled in the circumferential direction as they extend through the wing member.

4. The seal assembly according to claim 3, wherein the flow passages are at least one of curved and angled against a direction of rotation of the turbine rotor to effect a scooping of cooling fluid from the disc cavity into the flow passages.

5. The seal assembly according to claim 1, further comprising an annular seal member that extends axially from the vane assembly toward the blade assembly, the seal member including a seal surface that is in close proximity to a portion of the wing member.

6. The seal assembly according to claim 5, wherein the seal member is located radially outwardly from the wing member and overlaps the wing member.

7. The seal assembly according to claim 6, wherein the wing member includes an annular radially outwardly extending flange that is in close proximity to the seal surface of the seal member.

8. The seal assembly according to claim 7, wherein the seal surface of the seal member comprises an abradable material that is sacrificed in the case of contact between the flange and the seal surface.

9. The seal assembly according to claim 1, wherein outlets of the flow passages are positioned near known areas of ingestion of hot gas from the hot gas path into the disc cavity such that the cooling fluid exiting the flow passages through the outlets forces the hot gas away from the known areas of ingestion.

10. The seal assembly according to claim 9, wherein the known areas of ingestion are located between the vane assembly and the blade assembly at an upstream side of the blade assembly with reference to a flow direction of the hot gas through the hot gas path.

11. The seal assembly according to claim 1, wherein the pumping of cooling fluid from the disc cavity toward the hot gas path is effected by rotation of the turbine rotor and the blade assembly to limit hot gas ingestion from the hot gas path to the disc cavity by forcing hot gas in the hot gas path away from the seal assembly.

12. A seal assembly between a hot gas path and a disc cavity in a turbine engine comprising:

a non-rotatable vane assembly including a row of vanes and an inner shroud;

a rotatable blade assembly adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor;

an annular seal member that extends axially from the vane assembly toward the blade assembly and includes a seal surface; and

an annular wing member located radially inwardly from the hot gas path and radially outwardly from the disc cavity, the wing member extending generally axially from an axially facing side of the blade assembly toward the vane assembly and including;

a portion in close proximity to the seal surface of the seal member; and

a plurality of circumferentially spaced apart flow passages extending therethrough from a radially inner surface thereof to a radially outer surface thereof;

wherein a pumping of cooling fluid from the disc cavity toward the hot gas path is effected through the flow passages during operation of the engine by rotation of the turbine rotor and the blade assembly to limit hot gas ingestion from the hot gas path to the disc cavity by forcing the hot gas away from the seal assembly.

13. The seal assembly according to claim 12, wherein the flow passages extend generally straight through the wing member in the radial direction.

14. The seal assembly according to claim 12, wherein the flow passages are at least one of curved and angled in the circumferential direction as they extend through the wing member.

15. The seal assembly according to claim 14, wherein the flow passages are at least one of curved and angled against a direction of rotation of the turbine rotor to effect a scooping of cooling fluid from the disc cavity into the flow passages.

16. The seal assembly according to claim 12, wherein the seal member is located radially outwardly from the wing member and overlaps the wing member.

17. The seal assembly according to claim 12, wherein the wing member includes an annular radially outwardly extending flange that comprises the portion of the wing member in close proximity to the seal surface of the seal member.

18. The seal assembly according to claim 17, wherein the seal surface of the seal member comprises an abradable material that is sacrificed in the case of contact between the flange and the seal surface.

19. The seal assembly according to claim 12, wherein outlets of the flow passages are positioned near known areas of ingestion of the hot gas from the hot gas path into the disc cavity such that the cooling fluid exiting the flow passages through the outlets forces the hot gas away from the known areas of ingestion.

20. The seal assembly according to claim 19, wherein the known areas of ingestion are located between the vane assembly and the blade assembly at an upstream side of the blade assembly with reference to a flow direction of the hot gas through the hot gas path.