High pressure turbine blade cooling hole distribution
First Claim
1. A turbine blade for a gas turbine engine comprising an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine, the plurality of cooling holes including at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first row extending at least substantially radially and closer to a trailing edge than to a leading edge of the airfoil portion, the second row extending at least approximately axially near a tip of the blade and near the leading edge, the third and fourth rows extending at least substantially radially adjacent the leading edge, and the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8 located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3, wherein the point of origin of the X, Y, Z Cartesian system is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine blade, the X axis being angled with respect to a turbine rotor centerline by an angle corresponding to a restagger of the blade with a positive direction thereof being oriented towards aft of the engine and the Z axis extending generally radially along the stacking line with a positive direction thereof being oriented toward the tip of the blade.
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Accused Products
Abstract
A turbine blade for a gas turbine engine with an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine. The plurality of cooling holes includes at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8 located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3.
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Citations
20 Claims
- 1. A turbine blade for a gas turbine engine comprising an airfoil portion defined by a perimeter wall surrounding at least one enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one enclosure and a gaspath of the gas turbine engine, the plurality of cooling holes including at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first row extending at least substantially radially and closer to a trailing edge than to a leading edge of the airfoil portion, the second row extending at least approximately axially near a tip of the blade and near the leading edge, the third and fourth rows extending at least substantially radially adjacent the leading edge, and the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8 located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3, wherein the point of origin of the X, Y, Z Cartesian system is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine blade, the X axis being angled with respect to a turbine rotor centerline by an angle corresponding to a restagger of the blade with a positive direction thereof being oriented towards aft of the engine and the Z axis extending generally radially along the stacking line with a positive direction thereof being oriented toward the tip of the blade.
- 9. A turbine rotor assembly for a gas turbine engine comprising a blade having an airfoil portion defined by a perimeter wall enclosing at least one cooling cavity, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the at least one cooling cavity and a gaspath of the gas turbine engine, the plurality of cooling holes including at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first row being defined closer to a trailing edge than to a leading edge of the airfoil portion, and the second, third and fourth rows being defined along or in proximity of the leading edge, the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8, located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3, wherein the point of origin of X, Y, Z Cartesian system is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine blade, the X axis being angled with respect to a turbine rotor centerline by an angle corresponding to a restagger of the blade with a positive direction thereof being oriented towards aft of the engine and the Z axis extending generally radially along the stacking line with a positive direction thereof being oriented toward a tip of the blade.
- 16. A high pressure turbine blade comprising an airfoil having a perimeter wall surrounding at least one cooling cavity, the perimeter wall having an outer surface lying substantially on the points of Table 2, the perimeter wall having a plurality of cooling holes defined therethrough in fluid communication with the at least one cooling cavity, the plurality of cooling holes including at least one row of holes selected from the group consisting of a first row, a second row, a third row and a fourth row, with the first, second, third and fourth rows of holes respectively including the holes numbered PA-1 to PA-10, PB-1 to PB-3, HA-1 to HA-9, and SA-1 to SA-8, located such that a central axis thereof extends through the respective point 1 and point 2 having a nominal location in accordance with the X, Y, Z Cartesian coordinate values set forth in Table 3, wherein the point of origin of X, Y, Z Cartesian system is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine blade, the X axis being angled with respect to a turbine rotor centerline by an angle corresponding to a restagger of the blade with a positive direction thereof being oriented towards aft of the engine and the Z axis extending generally radially along the stacking line with a positive direction thereof being oriented toward a tip of the blade.
Specification