Outer rim seal assembly in a turbine engine

Patent metadata

Field	Value
Publication number	US-9260979-B2
Application number	US-201414546309-A
Country	US
Kind code	B2
Filing date	Nov 18, 2014
Priority date	Feb 15, 2013
Publication date	Feb 16, 2016
Grant date	Feb 16, 2016

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 axially 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 each include a portion that is curved as the passage extends radially outwardly to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path.

First claim

What is claimed is:

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 axially adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, the blades extending from a platform of the blade assembly; 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 each include a portion that is curved against the direction of rotation of the turbine rotor as the passage extends radially outwardly to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path during operation of the engine.

2. The seal assembly according to claim 1 , wherein outlets of the flow passages are located axially between a downstream end of the inner shroud and an upstream end of the platform.

3. The seal assembly according to claim 2 , wherein the flow passages are entirely located axially between the downstream end of the inner shroud and the upstream end of the platform.

4. 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.

5. The seal assembly according to claim 4 , wherein the seal member is located radially outwardly from the wing member and overlaps the wing member in the axial direction, and wherein the outlets of the flow passages are located axially between a downstream axial end of the seal member and the upstream end of the platform.

6. The seal assembly according to claim 5 , wherein the wing member includes an annular radially outwardly extending flange that is in close proximity to the seal surface of the seal member.

7. The seal assembly according to claim 6 , 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.

8. 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; and
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.

9. The seal assembly according to claim 1 , wherein the scooping 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.

10. 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 axially adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, the blades extending from a platform of the blade assembly;
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 the seal member 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 outlets of the flow passages are located axially between a downstream axial end of the seal member and an upstream end of the platform, and wherein the flow passages each include a portion that is curved in the circumferential direction against the direction of rotation of the turbine rotor as it extends radially outwardly through the wing member to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path during operation of the engine.

11. The seal assembly according to claim 10 , wherein the seal member axially overlaps the wing member.

12. The seal assembly according to claim 11 , 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, and 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.

13. The seal assembly according to claim 10 , wherein:
the 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; and
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.

14. The seal assembly according to claim 10 , wherein the flow passages are entirely located axially between the upstream end of the platform and each of a downstream end of the inner shroud and the downstream axial end of the seal member.

15. 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 axially adjacent to the vane assembly and including a row of blades and a turbine disc that forms a part of a turbine rotor, the blades extending from a platform of the blade assembly; 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 outlets of the flow passages are located axially between a downstream end of the inner shroud and an upstream end of the platform, and wherein the flow passages each include a portion that is curved against the direction of rotation of the turbine rotor as the passage extends radially outwardly to effect a scooping of cooling fluid from the disc cavity into the flow passages and toward the hot gas path during operation of the engine.

16. The seal assembly according to claim 15 , 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, wherein the seal member is located radially outwardly from the wing member and overlaps the wing member in the axial direction, and wherein the outlets o

Assignees

• Siemens Ag

Inventors

• Lee Ching-Pang
• Tham Kok-Mun
• Shivanand Manjit
• Laurello Vincent Paul
• Azad Gm Salam
• Martin Nicholas F

Classifications

• F01D5/081

  Cooling fluid being directed on the side of the rotor disc or at the roots of the blades (F01D5/087 takes precedence) · CPC title

• F01D11/001 Primary

  for sealing space between stator blade and rotor · CPC title

• F01D11/04

  using sealing fluid, e.g. steam · CPC title

• F05D2260/202

  by film cooling · CPC title

• F01D25/12 Primary

  Cooling · CPC title