Gas turbine engine turbine vane airfoil profile

• US 8,707,712 B2
• Filed: 07/02/2012
• Issued: 04/29/2014
• Est. Priority Date: 07/02/2012
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Spangler, Brandon W., Bergman, Russell J.
• Primary Examiner(s): Wongwian, Phutthiwat
• Application Number:
  US13/539,873
• Publication Number:
  US 20140000286A1
• Time in Patent Office: 666 Days
• Field of Search: 608/06, 415115-116, 416/95, 416/96.R, 416/97.A, 416/97.R, 416/193.A
• US Class Current: 60/806
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 9/02 Nozzles; Nozzle boxes; Stator blades; Guide conduits , e.g. individual nozzles nozzle boxes F01D9/047
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F05D 2240/81 Cooled platforms
• CPC Class Codes: F05D 2250/74 given by a set or table of xyz-coordinates
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A turbine vane for a gas turbine engine includes inner and outer platforms joined by a radially extending airfoil. The airfoil includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface. The inner and outer platforms respectively include inner and outer sets of film cooling holes. One of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm).

Claims

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

inner and outer platforms joined by a radially extending airfoil, the airfoil including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface; and

wherein the inner and outer platforms respectively include inner and outer sets of film cooling holes, wherein one of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2, the Cartesian coordinates provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm).

2. The turbine vane according to claim 1, wherein the turbine vane is a first stage turbine vane.

3. The turbine vane according to claim 1, wherein the zero-coordinate corresponds to a tangential on-board injector (TOBI) pin hole.

4. The turbine vane according to claim 1, wherein the inner and outer platforms respectively include inner and outer platform perimeters, and the inner and outer sets of cooling holes respectively lie within the inner and outer platform perimeters regardless of the tolerance.

5. The turbine vane according to claim 1, wherein spacing between edges of adjacent cooling holes is maintained by at least 0.015 inch (0.381 mm).

6. The turbine vane according to claim 1, wherein the film cooling holes have a diameter of 0.010-0.035 inch (0.25-0.89 mm).

7. The turbine vane according to claim 6, wherein some of the film cooling holes are conical and other of the film cooling holes are cylindrical.

8. The turbine vane according to claim 6, wherein the cooling holes are directed away from the airfoil and from the leading edge toward the trailing edge, and the cooling holes are angled relative to the platform surface 10°

-45°

.

9. A gas turbine engine comprising:

a compressor section;

a combustor fluidly connected to the compressor section;

a turbine section fluidly connected to the combustor, the turbine section comprising;

a high pressure turbine coupled to the high pressure compressor via a shaft;

a low pressure turbine; and

wherein the high pressure turbine includes an array of turbine vanes, each turbine vane having inner and outer platforms joined by a radially extending airfoil, the airfoil including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface; and

wherein the inner and outer platforms respectively include inner and outer sets of film cooling holes, wherein one of the inner and outer sets of film cooling holes are formed in substantial conformance with platform cooling hole locations described by one of the sets of Cartesian coordinates set forth in Tables 1 and 2, the Cartesian coordinates provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the cooling holes with Cartesian coordinates in Tables 1 and 2 have a diametrical surface tolerance relative to the specified coordinates of 0.200 inches (5.08 mm).

10. The gas turbine engine according to claim 9, wherein the array is a first stage array of turbine vanes.

11. The gas turbine engine according to claim 10, wherein the high pressure turbine consists of two arrays of turbine blades and two arrays of fixed stator vanes.

12. The turbine vane according to claim 9, wherein the zero-coordinate corresponds to a tangential on-board injector (TOBI) pin hole.

13. The turbine vane according to claim 9, wherein the inner and outer platforms respectively include inner and outer platform perimeters, and the inner and outer sets of cooling holes respectively lie within the inner and outer platform perimeters regardless of the tolerance.

14. The turbine vane according to claim 9, wherein spacing between adjacent cooling holes is maintained by at least 0.015 inch (0.381 mm).

15. The turbine vane according to claim 9, wherein the film cooling holes have a diameter of 0.010-0.035 inch (0.25-0.89 mm).

16. The turbine vane according to claim 15, wherein some of the film cooling holes are conical and other of the film cooling holes are cylindrical.

17. The turbine vane according to claim 15, wherein the cooling holes are directed away from the airfoil and from the leading edge toward the trailing edge, and the cooling holes are angled relative to the platform surface 10°

-45°

.