Turbine shroud segment feather seal located in radial shroud legs

• US 7,374,395 B2
• Filed: 07/19/2005
• Issued: 05/20/2008
• Est. Priority Date: 07/19/2005
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Rtx Corporation)
• Original Assignee: Pratt & Whitney Canada Corporation (Rtx Corporation)
• Inventors: Clermont, Martin, Durocher, Eric
• Primary Examiner(s): Look; Edward K.
• Assistant Examiner(s): White; Dwayne J
• Application Number:
  US11/183,922
• Publication Number:
  US 20070020087A1
• Time in Patent Office: 1,036 Days
• Field of Search: 415/115, 415/116, 415/182.1, 415/108, 415/196, 415/173.1, 416/191, 416/93.R, 416/96.R, 416/97.R
• US Class Current: 415/115
• CPC Class Codes: F01D 11/005 Sealing means between non relatively rotating elements
• CPC Class Codes: F01D 11/08 for sealing space between rotor blade tips and stator specially-shaped blade tips therefor F01D5/20
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 9/04 forming ring or sector
• CPC Class Codes: F05D 2240/11 Shroud seal segments
• CPC Class Codes: F05D 2240/57 Leaf seals
• CPC Class Codes: F05D 2260/203 by transpiration cooling
• CPC Class Codes: F05D 2260/205 Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes

Abstract

A turbine shroud assembly is configured to adequately adjust a distribution of cooling air flow such that air leakage between radial shroud legs of adjacent shroud segments is minimized, while permitting cooling air to leak between platforms of adjacent shroud segments in order to cool sides of the platforms thereof.

Claims

1. A turbine shroud assembly of a gas turbine engine comprising a plurality of shroud segments disposed circumferentially one adjacent to another, an annular support structure supporting the shroud segments together within an engine casing, and seals provided between adjacent shroud segments, each of the shroud segments including a platform collectively with platforms of adjacent shroud segments forming a shroud ring, and also including front and rear legs integrated with the platform and extending radially and outwardly therefrom for connection with the annular support structure, thereby supporting the platform radially and inwardly spaced apart from the annular support structure to define an annular cavity between the front and rear legs, the seals being disposed between the radial legs of adjacent shroud segments while radial air passages are defined substantially by clearances between mating side surfaces of adjacent platforms to permit cooling of substantially an entire axial length of sides of the platforms of the respective shroud segments.

2. The turbine shroud assembly as claimed in claim 1 wherein the seals comprise feather seals disposed between each pair of adjacent front legs and between each pair of adjacent rear legs.

3. The turbine shroud assembly as claimed in claim 2 wherein each of the shroud segments comprises radial slots defined in opposite sides of the respective front and rear legs thereof, each for receiving a portion of one feather seal.

4. The turbine shroud assembly as claimed in claim 1 wherein each of the shroud segments comprises a cooling passage extending within and through the platform and having at least one inlet thereof defined on an outer surface between the front and rear legs.

5. The turbine shroud assembly as claimed in claim 4 wherein the cooling passage comprises at least one outlet defined in a trailing end of the platform.

6. A cooling arrangement in a turbine shroud assembly of a gas turbine engine, the turbine shroud assembly having a plurality of shroud segments, the shroud segments including platforms disposed circumferentially adjacent one to another collectively to form a shroud ring, and including front and rear legs extending radially from an outer surface of the platforms thereby defining a cavity therebetween, the cooling arrangement comprising a first means for substantially preventing cooling air within the cavity from leakage between the front legs and between the rear legs of adjacent shroud segments and a second means for permitting use of cooling air within the cavity to cool substantially entire axial edges joining an inner surface and respective opposite sides of the platforms of the respective shroud segments.

7. The cooling arrangement as claimed in claim 6 further comprising a third means for transpiration cooling of the platforms of the shroud segments.

8. The cooling arrangement as claimed in claim 7 wherein the third means comprises a plurality of axial passages extending through the platform of each shroud segment, the axial passages being in fluid communication with the annular cavity between the front and rear legs for intake of the cooling air therein and for discharging same at a trailing end of the platform.

9. The cooling arrangement as claimed in claim 6 wherein the first means comprises a plurality of radially extending feather seals, disposed to substantially block an axial passage between adjacent front legs and between adjacent rear legs, respectively.

10. The cooling arrangement as claimed in claim 9 wherein each of the shroud segments comprises a cavity in opposite sides of the respective front and rear legs, each pair of the cavities defined in mating sides of adjacent legs, in combination accommodating one of the feather seals.

11. The cooling arrangement as claimed in claim 6 wherein the second means comprises a clearance between mating sides of each pair of adjacent shroud segments.

12. A method for cooling shroud segments of a turbine shroud assembly of a gas turbine engine, comprising steps of:

(a) continuously introducing cooling air into a cavity defined axially between radial front legs and radial rear legs of the shroud segments and radially between platforms of the shroud segments and an annular support structure;

(b) substantially preventing air leakage between the radial front legs and between the radial rear legs of the shroud segments for maintaining a predetermined pressure of the cooling air within the cavity; and

(c) cooling substantially entire axial edges joining an inner surface and respective opposite sides of the platforms by continuously directing the cooling air from the cavity through radial passages between platforms of adjacent shroud segments into a gas path defined by the platforms of the shroud segments.

13. The method as claimed in claim 12 comprising a step of (d) continuously directing the cooling air from the cavity through a passage extending within and through the individual shroud segments for transpiration cooling of the platforms of the shroud segments.

14. The method as claimed in claim 13 wherein step (d) is practiced by use of at least one inlet of the passage defined on an outer surface and positioned between the front and rear legs of the individual shroud segments for intake of the cooling air.

15. The method as claimed in claim 14 wherein step (d) is practiced by use of at least one outlet of the passage defined in a trailing end of the platform of the individual shroud segments for discharging the cooling air from the passage to cool a part of the engine before entering into the gas path.

16. The method as claimed in claim 12 wherein step (b) is practiced by use of feather seals provided between the radial front legs and between the radial rear legs of the shroud segments.

17. The method as claimed in claim 12 wherein step (e) is practiced by use of clearances between mating sides of adjacent platforms to form the radial passages.