Turbine blade cooling hole arrangement
A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.
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 the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.
2. The turbine blade of claim 1 , wherein the turbine blade is a first 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 cooling 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 and a root, 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 cooling 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 and a root, 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 and a root, 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 each turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.
9. The turbine rotor assembly of claim 8 , wherein the turbine rotor assembly is a first 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 cooling 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 and a root, 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 cooling 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 and a root, 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 and a root, 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 a suction side of an airfoil of a turbine blade of a gas turbine engine, comprising:
forming a plurality of cooling holes in the turbine blade, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.
16. The method of claim 15 , wherein the turbine blade is a first 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 cooling 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 and a root, 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 cooling 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 and a root, the airfoil extending from the platform, wherein the platform, the root, and the airfoil are cast as a single part.