High pressure turbine blade cooling hole distribution

• US 9,581,029 B2
• Filed: 09/24/2014
• Issued: 02/28/2017
• Est. Priority Date: 09/24/2014
• 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: LeCuyer, Daniel, Papple, Michael
• Primary Examiner(s): Keasel, Eric
• Assistant Examiner(s): Christensen, Danielle M
• Application Number:
  US14/494,837
• Publication Number:
  US 20160084091A1
• Time in Patent Office: 888 Days
• US Class Current: 1/1
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• 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 set of holes selected from the group consisting of a first set, a second set, a third set, a fourth set, a fifth set and a sixth set, wherein the first, second, third, fourth, fifth and sixth sets of holes respectively include the holes numbered A1 to A8, B1 to B10, C1 to C9, D1 to D6, E1 to E7 and F1 to F6 each 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 set of holes selected from the group consisting of a first set, a second set, a third set, a fourth set, a fifth set and a sixth set, wherein the first, second, third, fourth, fifth and sixth sets of holes respectively include the holes numbered A1 to A8, B1 to B10, C1 to C9, D1 to D6, E1 to E7 and F1 to F6 each located such that a central axis of the hole 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 a 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 0.0135 inches with a tolerance of ±

0.004 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 set forth in Table 2.

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

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

8. The turbine blade as defined in claim 1, wherein the airfoil portion includes four sets of holes selected from the group consisting of the first set, the second set, the third set, the fourth set, the fifth set and the sixth set.

9. The turbine blade as defined in claim 1, wherein the airfoil portion includes five sets of holes selected from the group consisting of the first set, the second set, the third set, the fourth set, the fifth set and the sixth set.

10. The turbine blade as defined in claim 1, wherein the airfoil portion includes the first set of holes, the second set of holes, the third set of holes, the fourth set of holes, the fifth set of holes and the sixth set of holes.

11. 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 set of holes selected from the group consisting of a first set, a second set, a third set, a fourth set, a fifth set and a sixth set, wherein the first, second, third, fourth, fifth and sixth sets of holes respectively include the holes numbered A1 to A8, B1 to B10, C1 to C9, D1 to D6, E1 to E7 and F1 to F6 each 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 a tip of the blade.

12. The turbine blade as defined in claim 11, 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.

13. The turbine blade as defined in claim 11, wherein for each hole, point 1 corresponds to an intersection of the central axis of the hole with an outer surface of the perimeter wall and point 2 corresponds to an intersection of the central axis of the hole with an inner surface of the perimeter wall.

14. The turbine blade as defined in claim 11, wherein each hole has a nominal diameter of 0.0135 inches with a tolerance of ±

0.004 inches.

15. The turbine blade as defined in claim 11, 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.

16. The turbine blade as defined in claim 11, wherein the airfoil portion includes two sets of holes selected from the group consisting of the first set, the second set, the third set, the fourth set, the fifth set and the sixth set.

17. The turbine blade as defined in claim 11, wherein the airfoil portion includes three sets of holes selected from the group consisting of the first set, the second set, the third set, the fourth set, the fifth set and the sixth set.

18. The turbine blade as defined in claim 11, wherein the airfoil portion includes four sets of holes selected from the group consisting of the first set, the second set, the third set, the fourth set, the fifth set and the sixth set.

19. The turbine blade as defined in claim 11, wherein the airfoil portion includes five sets of holes selected from the group consisting of the first set, the second set, the third set, the fourth set, the fifth set and the sixth set.

20. The turbine blade as defined in claim 11, wherein the airfoil portion includes the first set of holes, the second set of holes, the third set of holes, the fourth set of holes, the fifth set of holes and the sixth set of holes.