Gas turbine engine systems involving turbine blade platforms with cooling holes

• US 8,206,114 B2
• Filed: 04/29/2008
• Issued: 06/26/2012
• Est. Priority Date: 04/29/2008
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Spangler, Brandon W., Paauwe, Corneil S.
• Primary Examiner(s): EDGAR, RICHARD A
• Application Number:
  US12/111,240
• Publication Number:
  US 20090269184A1
• Time in Patent Office: 1,519 Days
• Field of Search: 416/193.A
• US Class Current: 416/193.A
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F05B 2240/801 cooled platforms
• CPC Class Codes: F05D 2240/81 Cooled platforms

Abstract

Gas turbine engine systems involving turbine blade platforms with mateface cooling holes are provided. In this regard, a representative turbine blade for a gas turbine engine includes: an airfoil having a leading edge, a trailing edge, a pressure side and a suction side; and a blade platform on which the airfoil is disposed, the blade platform having a pressure side mateface located adjacent to the pressure side of the airfoil and a suction side mateface located adjacent to the suction side of the airfoil, the blade platform having a cooling hole operative to direct a flow of cooling air toward an adjacent blade platform.

Claims

1. A turbine blade for a gas turbine engine comprising:

an airfoil having a leading edge, a trailing edge, a pressure side and a suction side, an interior cooling passage extends at least partially into the airfoil; and

a blade platform on which the airfoil is disposed, the blade platform having a pressure side mateface located adjacent to the pressure side of the airfoil and a suction side mateface located adjacent to the suction side of the airfoil, the blade platform having a cooling hole that pneumatically communicates with the interior cooling passage, the cooling hole having a cooling hole exit in the suction side mateface to direct a flow of cooling air through the suction side mateface toward a trailing edge of an airfoil disposed on an adjacent blade platform, the cooling hole raked aftward.

2. The turbine blade of claim 1, wherein:

the cooling hole is operative to direct the flow of cooling air toward the pressure side mateface of the adjacent blade platform such that the cooling flow impinges upon the pressure side mateface.

3. The turbine blade of claim 1, wherein the cooling hole exit is a first of multiple cooling hole exits located in the suction side mateface of the blade platform.

4. The turbine blade of claim 3, wherein each cooling hole exit has a corresponding cooling hole in communication with the interior cooling passage.

5. The turbine blade of claim 4, wherein each cooling hole is oriented parallel to an adjacent cooling hole.

6. The turbine blade of claim 1, wherein the cooling hole is oriented parallel to an outer diameter surface of the blade platform.

7. The turbine blade of claim 1, wherein the blade platform is an inner diameter platform.

8. The turbine blade of claim 1, wherein the cooling hole is oriented along the blade platform.

9. The turbine blade of claim 1, wherein the cooling hole is oriented parallel to a blade platform outer surface.

10. A turbine blade assembly for a gas turbine engine comprising:

a first turbine blade having a first blade platform and a first airfoil extending from the first blade platform; and

positioned adjacent to the first turbine blade, the second turbine blade having a second blade platform and a second airfoil extending from the second blade platform;

the second airfoil having a leading edge, a trailing edge, a pressure side and a suction side, an interior cooling passage extends at least partially into the second airfoil;

second blade platform having a first side facing away from the first turbine blade and a second opposing side facing toward the first turbine blade, the second blade platform having a mateface with a cooling hole that pneumatically communicates with the interior cooling passage, the cooling hole raked aftward to a cooling hole exit in the mateface operative to direct a flow of cooling air through the mateface toward a trailing edge of the first airfoil of the first turbine blade.

11. The assembly of claim 10, wherein:

the first side is a pressure side of the platform; and

the second side is a suction side of the platform.

12. The assembly of claim 10, wherein the blade platform of the second turbine blade is operative to direct the flow of cooling air such that the cooling air impinges upon a pressure side mateface of the blade platform of the first turbine blade.

13. A gas turbine engine comprising:

a compressor; and

a turbine operative to drive the compressor, the turbine having a turbine blade assembly, the turbine blade assembly having a first turbine blade and a second turbine blade;

the second blade being positioned adjacent to the first turbine blade, the second turbine blade having a blade platform and an airfoil extending from the blade platform;

the first blade being operative to aftwardly direct a flow of cooling air through a suction side mateface trailing edge of an airfoil disposed on an adjacent blade platform such that the cooling air impinges upon a pressure side mateface toward a trailing edge of the second turbine blade.

14. The engine of claim 13, wherein:

the cooling flow impinges upon the pressure side mateface in a vicinity of a trailing edge of the airfoil of the second blade.

15. The engine of claim 13, wherein the turbine is a high pressure turbine.

16. The engine of claim 13, wherein the engine is a turbofan gas turbine engine.