Turbine blade for gas turbine engine and method of cooling same

• US 6,402,471 B1
• Filed: 11/03/2000
• Issued: 06/11/2002
• Est. Priority Date: 11/03/2000
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Demers, Daniel Edward, Acquaviva, Paul Joseph, Manning, Robert Francis
• Primary Examiner(s): Look, Edward K.
• Assistant Examiner(s): WOO, RICHARD SUKYOON
• Application Number:
  US09/706,387
• Time in Patent Office: 585 Days
• Field of Search: 416/92, 416/95, 416/96.R, 416/97.R, 416/96.A, 416/97.A, 416/193.A, 415/115, 415/116
• US Class Current: 416/97.00R
• CPC Class Codes: F01D 11/04 using sealing fluid, e.g. steam
• 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: F01D 5/187 Convection cooling
• CPC Class Codes: F05D 2240/81 Cooled platforms
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/22141 using fins or ribs
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A turbine blade includes a platform having an internal cavity formed therein and an airfoil extending radially from the platform. An internal cooling circuit is formed in the airfoil for circulating a coolant therethrough, and at least one supply passage extends between the internal cooling circuit and the internal platform cavity for diverting coolant to the internal platform cavity. The coolant is expelled from holes located in the forward and aft edges of the platform for purging the forward and aft disk wheel spaces and impingement cooling adjacent nozzle bands.

Claims

1. A turbine blade comprising:

a platform having an internal cavity formed therein;

an airfoil extending radially from said platform;

an internal cooling circuit formed in said airfoil for circulating a coolant therethrough, said internal cooling circuit including a plurality of radially extending cavities fluidly connected in series in a serpentine arrangement with two of said radially extending cavities being connected at a turn; and

at least one supply passage extending between said turn and said internal cavity.

2. The turbine blade of claim 1 further comprising a second supply passage extending between said internal cooling circuit and said internal cavity.

3. The turbine blade of claim 1 further comprising a plurality of holes formed in said platform so as to provide egress for coolant from said internal cavity.

4. The turbine blade of claim 3 wherein said platform includes a forward angel wing and an aft angel wing, said plurality of holes including a first group of outlet holes formed in said forward angel wing and a second group of outlet holes formed in said aft angel wing.

5. The turbine blade of claim 3 wherein said platform includes a radially outer surface, said plurality of holes including film cooling holes formed in said radially outer surface.

6. The turbine blade of claim 1 further comprising a plurality of convective cooling enhancement features disposed in said internal cavity.

7. The turbine blade of claim 6 wherein said convective cooling enhancement features comprise pins extending from a radially outer wall of said platform to a radially inner wall of said platform.

8. A turbine blade for use in a gas turbine engine having a turbine rotor disk, said turbine blade comprising:

a dovetail for mounting said turbine blade to said rotor disk;

a shank extending from said dovetail;

a platform joined to said shank and having an internal cavity formed therein;

an airfoil extending radially from said platform;

an internal cooling circuit formed in said airfoil for circulating a coolant therethrough, said internal cooling circuit including a plurality of radially extending cavities fluidly connected in series in a serpentine arrangement with two of said radially extending cavities being connected at a turn; and

at least one supply passage extending between said turn and said internal cavity.

9. The turbine blade of claim 8 further comprising a second supply passage extending between said internal cooling circuit and said internal cavity.

10. The turbine blade of claim 8 further comprising a plurality of holes formed in said platform so as to provide egress for coolant from said internal cavity.

11. The turbine blade of claim 10 wherein said platform includes a forward angel wing and an aft angel wing, said plurality of holes including a first group of outlet holes formed in said forward angel wing and a second group of outlet holes formed in said aft angel wing.

12. The turbine blade of claim 11 wherein said outlet holes are angled away from the direction of rotation of said rotor disk.

13. The turbine blade of claim 10 wherein said platform includes a radially outer surface, said plurality of holes including film cooling holes formed in said radially outer surface.

14. The turbine blade of claim 8 further comprising a plurality of convective cooling enhancement features disposed in said internal cavity.

15. The turbine blade of claim 14 wherein said convective cooling enhancement features comprise pins extending from a radially outer wall of said platform to a radially inner wall of said platform.

16. In a gas turbine engine having a turbine rotor including at least one blade mounted to a rotor disk, said blade having a platform and an airfoil extending radially from said platform, a method of cooling said blade comprising the steps of:

providing said airfoil with an internal cooling circuit having a plurality of radially extending cavities fluidly connected in series in a serpentine arrangement with two of said radially extending cavities being connected at a turn;

providing said platform with an internal cavity;

introducing coolant into said internal cooling circuit; and

diverting a portion of the coolant from said turn to said internal cavity.

17. The method of claim 16 wherein said coolant is expelled from said internal cavity in a series of jets that impinge on adjacent structure.

18. The method of claim 16 wherein said coolant is expelled from said internal cavity into disk wheel spaces adjacent to said turbine rotor so as to block ingestion of hot gases into said disk wheel spaces.

19. The method of claim 16 wherein said coolant is expelled from said internal cavity into disk wheel spaces adjacent to said turbine rotor so as to purge said disk wheel spaces.

20. The method of claim 16 wherein said coolant is expelled from said internal cavity via film cooling holes formed in said platform.