Turbine blade cooling hole arrangement
A turbine blade for a gas turbine engine having a plurality of cooling holes defined therein, the plurality of cooling holes are located in an airfoil of the turbine blade according to the coordinates of Table 1.
1. A turbine blade for a gas turbine engine having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in an airfoil of the turbine blade according to coordinates of Table 1, wherein the coordinates of Table 1 are distances in inches from a point of origin O on the turbine blade, the point of origin being located at a center point of an inner diameter edge of a forward root face of a root of the turbine blade and wherein the coordinates of Table 1 have a tolerance of ± 0.060 inches.
2. The turbine blade of claim 1 , wherein the turbine blade is a second stage turbine blade of a high pressure turbine of the gas turbine engine.
3. The turbine blade of claim 2 , wherein the at least some of the plurality of holes have a hole diameter in a range of 0.010 inches to 0.020 inches.
4. The turbine blade of claim 3 , further comprising a platform, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.
5. The turbine blade of claim 1 , wherein the at least some of the plurality of holes have a hole diameter in a range of 0.010 inches to 0.020 inches.
6. The turbine blade of claim 5 , further comprising a platform, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.
7. The turbine blade of claim 1 , further comprising a platform, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.
8. A turbine rotor assembly for a gas turbine engine, comprising: a rotor disk; a plurality of turbine blades secured to the rotor disk, each turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in an airfoil of each turbine blade according to the coordinates of Table 1, wherein coordinates of Table 1 are distances in inches from a point of origin O on each turbine blade, the point of origin being located at a center point of an inner diameter edge of a forward root face of a root of each turbine blade and wherein the coordinates of Table 1 have a tolerance of ± 0.060 inches.
9. The turbine rotor assembly of claim 8 , wherein the turbine rotor assembly is a second stage turbine rotor assembly of a high pressure turbine of the gas turbine engine.
10. The turbine rotor assembly of claim 9 , wherein the at least some of the plurality of holes have a hole diameter in a range of 0.010 inches to 0.020 inches.
11. The turbine rotor assembly of claim 10 , wherein each of the plurality of turbine blades further comprise a platform, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.
12. The turbine rotor assembly of claim 8 , wherein the at least some of the plurality of holes have a hole diameter in a range of 0.010 inches to 0.020 inches.
13. The turbine rotor assembly of claim 12 , wherein each of the plurality of turbine blades further comprise a platform, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.
14. The turbine rotor assembly of claim 8 , wherein each of the plurality of turbine blades further comprise a platform, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.
15. A method of cooling an airfoil of a turbine blade of a gas turbine engine, comprising: forming a plurality of cooling holes in an airfoil of the turbine blade, wherein the plurality of cooling holes are located in the turbine blade according to coordinates of Table 1, wherein the coordinates of Table 1 are distances in inches from a point of origin O on the turbine blade, the point of origin being located at a center point of an inner diameter edge of a forward root face of a root of the turbine blade and wherein the coordinates of Table 1 have a tolerance of ± 0.060 inches.
16. The method of claim 15 , wherein the turbine blade is a second stage turbine blade of a high pressure turbine of the gas turbine engine.
17. The method of claim 16 , wherein the at least some of the plurality of holes have a hole diameter in a range of 0.010 inches to 0.020 inches.
18. The method of claim 17 , wherein the turbine blade further comprises a platform, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.
19. The method of claim 15 , wherein the at least some of the plurality of holes have a hole diameter in a range of 0.010 inches to 0.020 inches.
20. The method of claim 15 , wherein the turbine blade further comprises a platform, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.