HIGH PRESSURE TURBINE VANE COOLING HOLE DISTRUBUTION

• US 20120014809A1
• Filed: 07/19/2010
• Published: 01/19/2012
• Est. Priority Date: 07/19/2010
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Rtx Corporation)
• Original Assignee: Pratt & Whitney Canada Corporation (Rtx Corporation)
• Inventors: DI PAOLA, Franco, Lebel, Larry, Doyon, François
• Granted Patent: US 8,568,085 B2
• Time in Patent Office: Days
• US Class Current: 416/97.R
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• CPC Class Codes: F01D 9/04 forming ring or sector
• 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 vane for a gas turbine engine with an airfoil portion including a perimeter wall having first, second, third and fourth sets of cooling holes defined therethrough, including the holes numbered HA-1 to HA-13, HB-1 to HB-13, PA-1 to PA-9, and SA-1 to SA-3, respectively, and 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 Cartesian coordinate values set forth in Table 3.

Claims

1. A turbine vane for a gas turbine engine comprising an airfoil portion defined by a perimeter wall surrounding an enclosure, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the enclosure and a gaspath of the gas turbine engine, the plurality of cooling holes including first, second, third and fourth rows of holes, the first and second rows of holes extending at least substantially radially adjacent a leading edge of the airfoil portion, the third row of holes extending at least substantially radially on one side of the airfoil in proximity of a trailing edge thereof, and the fourth row of holes extending at least substantially radially on the opposed side of the airfoil in proximity of the leading edge, the first, second, third and fourth rows of holes including the holes numbered HA-1 to HA-13, HB-1 to HB-13, PA-1 to PA-9, and SA-1 to SA-3, respectively, and 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 Cartesian coordinate values set forth in Table 3.

2. The turbine vane as defined in claim 1, 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.

3. The turbine vane as defined in claim 2, wherein for each hole, point 2 corresponds to an intersection of the central axis of the hole with an inner surface of the perimeter wall.

4. The turbine vane as defined in claim 1, wherein each hole has a nominal diameter corresponding to the respective diameter value set forth in Table 3 with a tolerance of ±

0.004 inches.

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

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

6. The turbine vane as defined in claim 1, wherein an outer surface of the perimeter wall is defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 4 to 10 set forth in Table 2.

7. The turbine vane as defined in claim 1, wherein the airfoil portion has a nominal restagger and extends between inner and outer platforms having a plurality of platform cooling holes defined therethrough in fluid communication with the enclosure, the plurality of platform cooling holes including a first set of holes defined through the front end and side of the outer platform, and a second set of holes defined through a gaspath surface of the inner platform, the first and second sets of holes including the holes numbered TA-1 to TA-TA-6 and RA-1 to RA-5, respectively, and 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 Cartesian coordinate values set forth in Table 4.

8. The turbine vane as defined in claim 1, wherein each hole of the first and second sets has a nominal diameter corresponding to the respective diameter value set forth in Table 4 with a tolerance of ±

0.004 inches.

9. The turbine vane as defined in claim 7, wherein the X, Y Cartesian coordinate values have a tolerance of ±

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

10. A turbine stator assembly for a gas turbine engine comprising a plurality of vanes, each vane having an airfoil portion defined by a perimeter wall enclosing a cooling cavity, the perimeter wall having a plurality of cooling holes defined therethrough and providing fluid communication between the cooling cavity and a gaspath of the gas turbine engine, the plurality of cooling holes including first, second, third and fourth sets of holes, the first and second sets of holes extending adjacent a leading edge of the airfoil portion, the third set of holes extending on one side of the airfoil in proximity of a trailing edge thereof, and the fourth set of holes extending on the opposed side of the airfoil in proximity of the leading edge, the first, second, third and fourth sets of holes including the holes numbered HA-1 to HA-13, HB-1 to HB-13, PA-1 to PA-9, and SA-1 to SA-3, respectively, each hole having a central axis extending through point 1 and point 2 located at least substantially in accordance with X, Y Cartesian coordinate values set forth in Table 3.

11. The turbine vane as defined in claim 10, 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.

12. The turbine vane as defined in claim 10, wherein each hole has a nominal diameter corresponding to the respective diameter value set forth in Table 3, with a tolerance of ±

0.004 inches.

13. The turbine vane as defined in claim 10, wherein the X, Y Cartesian coordinate values have a tolerance of ±

0.030 inches.

14. The turbine vane as defined in claim 10, wherein an outer surface of the perimeter wall is defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 4 to 10 set forth in Table 2.

15. The turbine vane as defined in claim 10, wherein the airfoil portion has a nominal restagger and extends between platforms defined generally by at least some of the coordinate values given in Table 1, the platforms having a plurality of platform cooling holes defined therethrough in fluid communication with the cooling cavity, the plurality of platform cooling holes including fifth and sixth sets of holes including the holes numbered TA-1 to TA-6 and RA-1 to RA-5, respectively, each having a central axis extending through point 1 and point 2 located at least substantially in accordance with X, Y Cartesian coordinate values set forth in Table 4.

16. The turbine vane as defined in claim 10, wherein each hole of the fifth and sixth sets has a nominal diameter corresponding to the respective diameter value set forth in Table 4 with a tolerance of ±

0.004 inches.

17. A high pressure turbine vane comprising an airfoil having a perimeter wall surrounding a cooling cavity, the perimeter wall having an outer surface lying substantially on the points of Table 2, the airfoil extending between platforms defined generally by at least some of the coordinate values given in Table 1, wherein a fillet radius is applied around the airfoil between the airfoil and platforms, the perimeter wall having a plurality of cooling holes defined therethrough in fluid communication with the cooling cavity, the plurality of cooling holes including first, second, third and fourth sets of holes including the holes numbered HA-1 to HA-13, HB-1 to HB-13, PA-1 to PA-9, and SA-1 to SA-3, respectively, and 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 Cartesian coordinate values set forth in Table 3.

18. The turbine vane as defined in claim 17, wherein the airfoil portion has a nominal restagger, the platforms including a plurality of platform cooling holes defined therethrough in fluid communication with the cooling cavity, the plurality of platform cooling holes including fifth and sixth sets of holes including the holes numbered TA-1 to TA-6 and RA-1 to RA-5, respectively, and 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 Cartesian coordinate values set forth in Table 4.

19. The turbine vane as defined in claim 18, wherein each hole of the first, second, third and fourth sets has a nominal diameter corresponding to the respective diameter value set forth in Table 3, and each hole of the fifth and sixth sets has a nominal diameter corresponding to the respective diameter value set forth in Table 4.

20. The turbine vane as defined in claim 18, wherein the X, Y Cartesian coordinate values have a tolerance of ±

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