Turbine blade cooling hole arrangement
AssigneeRaytheon Technologies Corporation
InventorsAlex J. Schneider, Timothy J. Jennings, David A. Niezelski, Jeffrey T. Morton, Nicholas M. LoRicco, Bret M. Teller
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Abstract
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.
Claims20
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.
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.
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.
Dependencies are inferred from the claim text.