HIGH PRESSURE TURBINE BLADE COOLING HOLE DISTRIBUTION

• US 20140219816A1
• Filed: 09/28/2012
• Published: 08/07/2014
• Est. Priority Date: 09/28/2012
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: PLANTE, Ghislain, PAQUET, Rene, PELOQUIN, Patrick
• Granted Patent: US 9,121,289 B2
• Time in Patent Office: Days
• US Class Current: 416/97.R
• CPC Class Codes: F01D 5/141 Shape, i.e. outer, aerodynamic form F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F05D 2250/74 given by a set or table of xyz-coordinates
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

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.

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.

2. The turbine blade as defined in claim 1, wherein for each hole, at least one of point 1 and point 2 corresponds to an intersection of the central axis of the hole with a surface of the perimeter wall.

3. The turbine blade as defined in claim 1, wherein each hole has a nominal diameter of 0.014 inches with a tolerance of ±

0.006 inches.

4. The turbine blade as defined in claim 1, wherein the X, Y, Z Cartesian coordinate values have a tolerance of ±

0.030 inches of the nominal location with respect to the X, Y and Z axes.

5. The turbine blade as defined in claim 1, wherein an outer surface of the perimeter wall is defined by a nominal profile substantially in accordance with the X, Y, and Z Cartesian coordinate values of Sections 3 to 14 set forth in Table 2.

6. The turbine blade as defined in claim 1, wherein the airfoil portion includes two rows of holes selected from the group consisting of the first row, the second row, the third row and the fourth row.

7. The turbine blade as defined in claim 1, wherein the airfoil portion includes three rows of holes selected from the group consisting of the first row, the second row, the third row and the fourth row.

8. The turbine blade as defined in claim 1, wherein the airfoil portion includes the first row of holes, the second row of holes, the third row of holes and the fourth row of holes.

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.

10. The rotor as defined in claim 9, wherein each hole has a nominal diameter of 0.014 inches with a tolerance of ±

0.006 inches.

11. The rotor as defined in claim 9, wherein the X, Y, Z Cartesian coordinate values have a tolerance of ±

0.030 inches.

12. The rotor as defined in claim 9, wherein an outer surface of the perimeter wall is defined by a nominal profile substantially in accordance with the X, Y, and Z Cartesian coordinate values of Sections 3 to 14 set forth in Table 2.

13. The rotor as defined in claim 9, wherein the airfoil portion includes two rows of holes selected from the group consisting of the first row, the second row, the third row and the fourth row.

14. The rotor as defined in claim 9, wherein the airfoil portion includes three rows of holes selected from the group consisting of the first row, the second row, the third row and the fourth row.

15. The rotor as defined in claim 9, wherein the airfoil portion includes the first row of holes, the second row of holes, the third row of holes and the fourth row of holes.

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.

17. The turbine blade as defined in claim 16, wherein the airfoil portion includes the first row of holes, the second row of holes, the third row of holes and the fourth row of holes.

18. The turbine blade as defined in claim 16, wherein the airfoil portion includes three rows of holes selected from the group consisting of the first row, the second row, the third row and the fourth row.

19. The turbine blade as defined in claim 16, wherein each hole of the first, second, third and fourth rows has a nominal diameter of 0.014 inches with a tolerance of ±

0.006 inches.

20. The turbine blade as defined in claim 16, wherein the X, Y, Z Cartesian coordinate values of the holes have a tolerance of ±

0.030 inches of the nominal location with respect to the X, Y and Z axes.