TURBINE VANE FOR GAS TURBINE ENGINE

• US 20190323362A1
• Filed: 04/19/2018
• Published: 10/24/2019
• Est. Priority Date: 04/19/2018
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Arisi, Allan N., Teller, Bret M.
• Granted Patent: US 10,677,071 B2
• Time in Patent Office: Days
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 5/189 the insert having a tubular cross-section, e.g. airfoil shape
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F05D 2220/3212 the first stage of a turbine
• CPC Class Codes: F05D 2220/323 for aircraft propulsion, e.g. jet engines
• CPC Class Codes: F05D 2230/54 by sheet metal manufacturing
• CPC Class Codes: F05D 2230/60 Assembly methods
• CPC Class Codes: F05D 2240/121 related to the leading edge of a stator vane
• CPC Class Codes: F05D 2250/74 given by a set or table of xyz-coordinates
• CPC Class Codes: F05D 2260/201 by impingement of a fluid
• 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 having a plurality of cooling holes defined therein, at least some of the plurality of cooling holes being located on a leading edge of an airfoil of the turbine vane and in fluid communication with an internal cavity of the turbine vane; and a baffle insert located in the internal cavity, the baffle insert having a plurality of holes formed therein at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane, the baffle insert being formed from a flat sheet of metal wherein the plurality of holes of the baffle insert are formed in the flat sheet of metal prior to the baffle insert being formed from the flat sheet of metal, the plurality of holes of the baffle insert being formed in the flat sheet of metal according to the coordinates of Table 1.

Claims

1. A turbine vane for a gas turbine engine having a plurality of cooling holes defined therein, at least some of the plurality of cooling holes being located on a leading edge of an airfoil of the turbine vane and in fluid communication with an internal cavity of the turbine vane;

and

a baffle insert located in the internal cavity, the baffle insert having a plurality of holes formed therein at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane, the baffle insert being formed from a flat sheet of metal wherein the plurality of holes of the baffle insert are formed in the flat sheet of metal prior to the baffle insert being formed from the flat sheet of metal, the plurality of holes of the baffle insert being formed in the flat sheet of metal according to the coordinates of Table 1.

2. The turbine vane of claim 1, wherein the turbine vane is a first stage turbine vane of a high pressure turbine of the gas turbine engine.

3. The turbine vane of claim 2, wherein the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane have a hole diameter in a range of 0.015 inches to 0.030 inches.

4. The turbine vane of claim 3, wherein the plurality of holes of the baffle insert other than the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane have a hole diameter that is less than the hole diameter of the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane.

5. The turbine vane of claim 1, wherein the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane have a hole diameter in a range of 0.015 inches to 0.030 inches.

6. The turbine vane of claim 5, wherein the plurality of holes of the baffle insert other than the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane have a hole diameter that is less than the hole diameter of the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane.

7. The turbine vane of claim 1, wherein the plurality of holes of the baffle insert other than the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane have a hole diameter that is less than the hole diameter of the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane.

8. A turbine stator assembly for a gas turbine engine comprising a plurality of vanes, each vane having a plurality of cooling holes defined therein, at least some of the plurality of cooling holes of each of the plurality of vanes being located on a leading edge of an airfoil of each of the plurality of vanes and in fluid communication with an internal cavity of a respective one of the plurality of vanes;

and

a baffle insert located in the internal cavity of each of the plurality of vanes, the baffle insert having a plurality of holes formed therein, at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of a respective one of each of the plurality of vanes, the baffle insert being formed from a flat sheet of metal wherein the plurality of holes of the baffle insert are formed in the flat sheet of metal prior to the baffle insert being formed from the flat sheet of metal, the plurality of holes of the baffle insert being formed in the flat sheet of metal according to the coordinates of Table 1.

9. The turbine stator assembly of claim 8, wherein each vane of the plurality of vanes is a first stage vane of a high pressure turbine of the gas turbine engine.

10. The turbine stator assembly of claim 9, wherein the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of a respective one of the plurality of vanes have a hole diameter in a range of 0.015 inches to 0.030 inches.

11. The turbine stator assembly of claim 10, wherein the plurality of holes of the baffle insert other than the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of each vane of the plurality of vanes have a hole diameter that is less than the hole diameter of the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane.

12. The turbine stator assembly of claim 8, wherein the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of a respective one of the plurality of vanes have a hole diameter in a range of 0.015 inches to 0.030 inches.

13. The turbine stator assembly of claim 12, wherein the plurality of holes of the baffle insert other than the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of a respective one of the plurality of vanes have a hole diameter that is less than the hole diameter of the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane.

14. The turbine stator assembly of claim 8, wherein the plurality of holes of the baffle insert other than the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of a respective one of the plurality of vanes have a hole diameter that is less than the hole diameter of the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane.

15. A method of cooling a leading edge of a turbine vane of a gas turbine engine, comprising:

forming a plurality of cooling holes in the turbine vane, at least some of the plurality of cooling holes being located on a leading edge of an airfoil of the turbine vane and in fluid communication with an internal cavity of the turbine vane;

forming a baffle insert from a flat sheet of metal wherein a plurality of holes are formed in the flat sheet of metal prior to the baffle insert being formed from the flat sheet of metal, the plurality of holes of the baffle insert being formed in the flat sheet of metal according to the coordinates of Table 1; and

inserting the formed baffle insert into the internal cavity of the turbine vane.

16. The method of claim 15, wherein at least some of the plurality of holes of the baffle insert correspond to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane.

17. The method of claim 15, wherein the turbine vane is a first stage vane of a high pressure turbine of the gas turbine engine.

18. The method of claim 15, wherein the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane have a hole diameter in a range of 0.015 inches to 0.030 inches.

19. The method of claim 18, wherein the plurality of holes of the baffle insert other than the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane have a hole diameter that is less than the hole diameter of the at least some of the plurality of holes of the baffle insert corresponding to the at least some of the plurality of cooling holes located in the leading edge of the turbine vane.

20. The method of claim 15, wherein the turbine vane is a first stage vane of a high pressure turbine of the gas turbine engine, and wherein all of the plurality of holes of the baffle insert have a hole diameter in the range of 0.015 inches to 0.030 inches.