Rim seal arrangement having pumping feature

Patent metadata

Field	Value
Publication number	US-9631509-B1
Application number	US-201514946893-A
Country	US
Kind code	B1
Filing date	Nov 20, 2015
Priority date	Nov 20, 2015
Publication date	Apr 25, 2017
Grant date	Apr 25, 2017

Abstract

A rim seal arrangement for a gas turbine engine includes a first seal face on a rotor component, and a second seal face on a stationary annular rim centered about a rotation axis of the rotor component. The second seal face is spaced from the first seal face along an axial direction to define a seal gap. The seal gap is located between a radially outer hot gas path and a radially inner rotor cavity. The first seal face has a plurality of circumferentially spaced depressions, each having a depth in an axial direction and extending along a radial extent of the first seal face. The depressions influence flow in the seal gap such that during rotation of the rotor component, fluid in the seal gap is pumped in a radially outward direction to prevent ingestion of a gas path fluid from the hot gas path into the rotor cavity.

First claim

The invention claimed is:

1. A rim seal arrangement for a gas turbine engine, comprising:
a first seal face on a rotor component,
a second seal face on a stationary annular rim centered about a rotation axis of the rotor component, the second seal face facing the first seal face and being spaced from the first seal face along an axial direction to define a seal gap, wherein the seal gap is located between a radially outer hot gas path and a radially inner rotor cavity,
wherein the first seal face comprises a plurality of circumferentially spaced depressions thereon, each depression having a depth in an axial direction and extending along a radial extent of the first seal face between a radially outer end and a radially inner end of the depression,
the plurality of depressions being configured to influence flow in the seal gap such that during rotation of the rotor component, fluid in the seal gap is pumped in a radially outward direction to prevent ingestion of a gas path fluid from the hot gas path into the rotor cavity, and
wherein each depression has a surface that transitions from a generally convex leading surface to a generally concave trailing surface along a circumferential direction.

2. The rim seal arrangement according to claim 1 , further comprising a fillet disposed at a radially inner edge of the first seal face and positioned radially inward with respect to the radially inner end of the depression.

3. The rim seal arrangement according to claim 1 , wherein each depression has a converging cross-sectional area from the radially inner end to the radially outer end.

4. The rim seal arrangement according to claim 3 , wherein each depression has a decreasing circumferential width from the radially inner end to the radially outer end.

5. The rim seal arrangement according to claim 3 , wherein each depression has a maximum axial depth that varies between the radially inner end and the radially outer end.

6. The rim seal arrangement according to claim 1 , wherein each depression comprises a flow axis that is geometrically configured such that the radially inner end is circumferentially offset from the radially outer end in the direction of rotation of the rotor component.

7. The rim seal arrangement according to claim 6 , wherein the flow axis is parallel to the radial direction at the radially outer end, and is angled with respect to the radial direction toward the direction of rotation of the rotor component at the radially inner end.

8. The rim seal arrangement according to claim 1 , wherein the first seal face is an aft face of a base portion of a turbine blade.

9. A turbine blade comprising:
an airfoil body extending span-wise in a radial direction and comprising an internal cooling channel, and
a blade base located at a radially inner end of the airfoil body for attaching the turbine blade to a rotor disc, the blade base comprising an aft face that faces a seal gap between a radially outer hot gas path and a radially inner rotor cavity in fluid communication with the internal cooling channel, the seal gap being defined by a space between the aft face of the blade base and a fore-end of an annular stationary rim centered about a longitudinal axis of the rotor disc,
wherein the aft face of the blade base facing the seal gap comprises a plurality of circumferentially spaced depressions thereon, each depression having a depth in an axial direction, a width in a circumferential direction and extending along a radial extent of the aft face between a radially outer end and a radially inner end of the depression,
the plurality of depressions being configured to influence flow in the seal gap such that during rotation of the turbine blade, fluid in the seal gap is pumped in a radially outward direction to prevent ingestion of a gas path fluid from the hot gas path into the rotor cavity, and
wherein each depression has a surface that transitions from a generally convex leading surface to a generally concave trailing surface along a circumferential direction.

10. The turbine blade according to claim 9 , further comprising a fillet disposed at a radially inner edge of the aft face of the blade base and positioned radially inward with respect to the radially inner end of the depression.

11. The turbine blade according to claim 9 , wherein each depression has converging cross-sectional area from the radially inner end to the radially outer end, the cross-sectional area being defined by the width and the depth of the depression.

12. The turbine blade according to claim 9 , wherein the radially inner end is circumferentially spaced from the radially outer end in the direction of rotation of the turbine blade.

13. The turbine blade according to claim 12 , wherein each depression comprises circumferentially spaced leading and trailing ends that extend from the radially outer end to the radially inner end, a depression centerline being defined extending centrally between the leading end and the trailing end, wherein
the depression centerline is parallel to the radial direction at the radially outer end, and
the depression centerline is angled with respect to the radial direction toward the direction of rotation of the turbine blade at the radially inner end.

14. A gas turbine engine comprising:
a turbine blade comprising a blade base affixed to a rotor disc, and
an annular stationary rim disposed aft of the turbine blade and centered about a longitudinal axis of the rotor disc,
wherein the rim extends fore toward the blade base and comprises a fore-end that is spaced from an aft face of the blade base to define a seal gap between a radially outwardly located hot gas path and a radially inwardly located rotor cavity in fluid communication with an internal cooling channel of the turbine blade,
wherein the aft face of the blade base facing the seal gap comprises a plurality of circumferentially spaced depressions thereon, each depression having a depth in an axial direction and extending along a radial extent of the aft face between a radially outer end and a radially inner end of the depression,
the plurality of depressions being configured to influence flow in the seal gap such that during rotation of the turbine blade, fluid in the seal gap is pumped in a radially outward direction to prevent ingestion of a gas path fluid from the hot gas path into the rotor cavity,
wherein each depression has a surface that transitions from a generally convex leading surface to a generally concave trailing surface along a circumferential direction.

15. The gas turbine engine according to claim 14 , further comprising a cooling fluid supply passage providing fluid communication between the rotor cavity and a source of cooling fluid at atmospheric pressure, wherein the rotation of the turbine blade is effective to draw the cooling fluid from the source into the rotor cavity via the cooling fluid supply passage.

16. The gas turbine engine according to claim 15 , further comprising a pre-swirler disposed downstream of the turbine blade between the rotor cavity and the cooling fluid supply passage.

Assignees

• Siemens Energy Inc

Inventors

• Lee Ching-Pang
• Myers Caleb

Classifications

• F16J15/44

  Free-space packings · CPC title

• F01D11/04 Primary

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

• F05D2260/60

  Fluid transfer · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F01D5/082 Primary

  on the side of the rotor disc · CPC title