INNER SHROUD COOLING ARRANGEMENT IN A GAS TURBINE ENGINE

• US 20120177479A1
• Filed: 01/06/2011
• Published: 07/12/2012
• Est. Priority Date: 01/06/2011
• Status: Abandoned Application

• Current Assignee: Siemens Energy Incorporated (Siemens AG)
• Original Assignee: Siemens Energy Incorporated (Siemens AG)
• Inventors: Lee, Ching-Pang, Gao, Zhihong, Azad, Gm Salam
• Application Number:
  US12/985,571
• Publication Number:
  US 20120177479A1
• Time in Patent Office: Days
• US Class Current: 415/115
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F05D 2240/59 Lamellar seals
• CPC Class Codes: F05D 2240/81 Cooled platforms
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/60 Fluid transfer

Abstract

A component in a gas turbine engine includes an airfoil and a shroud. The shroud has an outer surface supporting an end of the airfoil and defines a portion of an annular gas path. The shroud includes axial edges extending between upstream and downstream edges thereof. Each of the axial edges includes a seal slot that receives a seal member extending between the shroud and an adjacent shroud. A cooling air channel extends between the upstream and downstream edges of the shroud. A cooling air supply passage extends from a cooling air chamber at an inner surface of the shroud to the cooling air channel. At least one cooling air exit passage extends from the cooling air channel to one of the axial edges. The cooling air channel is located radially between the outer surface of the shroud and the seal slot.

Claims

1. A component in a gas turbine engine, said component comprising:

an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge;

a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream edge and said downstream edge;

each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud;

a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream edge and said downstream edge;

a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel;

at least one cooling air exit passage extending from said cooling air channel to said one of said axial edges; and

said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.

2. The component of claim 1, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges.

3. The component of claim 1, wherein said cooling air exit passage includes an exit opening located between said outer surface of said shroud and said seal slot.

4. The component of claim 1, wherein said cooling air exit passage comprises a purge passage having an exit opening providing a volume of air for purging hot gas from said seal slot and said seal, said exit opening being located at a downstream end of said cooling air channel adjacent said downstream edge of said shroud.

5. The component of claim 4, further comprising a replenishing cooling air supply passage extending from a mid-chord cooling air chamber at said inner surface of said shroud to said cooling air channel, said replenishing cooling air supply passage located between said cooling air supply passage and said shroud downstream edge.

6. The component of claim 5, wherein said cooling air supply passage supplies a first portion of cooling air from a leading edge cooling air chamber to a location adjacent an upstream end of said cooling air channel and said replenishing cooling air supply passage supplies a second portion of cooling air from said mid-chord cooling air chamber to a location proximate to said purge passage.

7. The component of claim 5, wherein said component comprises a vane in a first row of vanes within said gas turbine engine.

8. The component of claim 1, wherein said cooling air exit passage comprises a plurality of impingement passages having exit openings providing a flow of cooling air impinging on an axial edge of said adjacent shroud, said exit openings being located at an upstream end of said cooling air channel adjacent said upstream edge of said shroud.

9. The component of claim 8, wherein said upstream end of said cooling air channel receives cooling air from a leading edge cooling air chamber at said inner surface of said shroud.

10. The component of claim 9, wherein said cooling air supply passage comprises a replenishing cooling air supply passage that supplies cooling air from a mid-shroud impingement cavity located at an axial midpoint between said upstream edge and said downstream edge at said inner surface of said shroud, said replenishing cooling air supply passage providing replenishing cooling air to said cooling air channel.

11. The component of claim 9, wherein said component comprises a vane in a second row of vanes within said gas turbine engine.

12. A vane in a gas turbine engine, said vane comprising:

an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge;

a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream edge and said downstream edge;

each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud;

a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream edge and said downstream edge;

a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel;

at least one cooling air exit passage extending from said cooling air channel to said one of said axial edges, said cooling air exit passage comprising a purge passage having an exit opening providing a volume of air for purging hot gas from said seal slot and said seal, said exit opening being located at a downstream end of said cooling air channel adjacent said downstream edge of said shroud; and

said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.

13. The vane of claim 12, further comprising a replenishing cooling air supply passage extending from a mid-chord cooling air chamber at said inner surface of said shroud to said cooling air channel, said replenishing cooling air supply passage located between said cooling air supply passage and said shroud downstream edge.

14. The vane of claim 13, wherein said cooling air supply passage supplies a first portion of cooling air from a leading edge cooling air chamber to a location adjacent an upstream end of said cooling air channel and said replenishing cooling air supply passage supplies a second portion of cooling air from said mid-chord cooling air chamber to a location proximate to said purge passage.

15. The vane of claim 12, wherein said cooling air channel is without openings for discharge of air along a length of said cooling air channel from said cooling air supply passage to within close proximity of said replenishing cooling air supply passage.

16. The vane of claim 12, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges.

17. A vane in a gas turbine engine, said vane comprising:

an airfoil adapted to extend radially through an annular hot gas path extending in a generally axial direction through said turbine engine, said airfoil including a pressure side and a suction side, an upstream leading edge and a downstream trailing edge;

a shroud having an outer surface supporting an end of said airfoil, said shroud defining a portion of said annular gas path through said gas turbine engine and including an upstream edge and a downstream edge, and opposing axial edges extending between said upstream and downstream edges;

each of said axial edges including a generally axially extending seal slot adapted to receive a seal member extending between said shroud and an adjacent shroud;

a cooling air channel extending generally axially substantially parallel to at least one of said axial edges between said upstream and said downstream edges;

a cooling air supply passage extending from a cooling air chamber at an inner surface of said shroud to said cooling air channel;

a plurality of impingement passages extending from said cooling air channel to said one of said axial edges, said impingement passages having exit openings providing a flow of cooling air impinging on an axial edge of said adjacent shroud, said exit openings being located at an upstream end of said cooling air channel and adjacent said upstream edge of said shroud; and

said cooling air channel being located radially between said outer surface of said shroud and said seal slot at said one of said axial edges for effecting convective cooling of a corner defined at an intersection of said outer surface and said one of said axial edges.

18. The vane of claim 17, wherein said upstream end of said cooling air channel receives cooling air from a leading edge cooling air chamber at said inner surface of said shroud.

19. The vane of claim 18, wherein said cooling air supply passage comprises a replenishing cooling air supply passage that supplies cooling air from a mid-shroud impingement cavity located at an axial midpoint between said upstream edge and said downstream edge at said inner surface of said shroud, said replenishing cooling air supply passage providing replenishing cooling air to said cooling air channel.

20. The vane of claim 17, wherein said cooling air channel is located on a radial plane passing through said seal slot at said one of said axial edges.