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Turbine blade having angled squealer tip

Imported: 24 Feb '17 | Published: 06 Jan '04

David Glenn Cherry, Ching-Pang Lee, Chander Prakash, Aspi Rustom Wadia, Brian David Keith, Steven Robert Brassfield

USPTO - Utility Patents

Abstract

A turbine blade for a gas turbine engine, including an airfoil and integral dovetail for mounting the airfoil along a radial axis to a rotor disk inboard of a turbine shroud. The airfoil further includes: first and second sidewalls joined together at a leading edge and a trailing edge, where the first and second sidewalls extend from a root disposed adjacent the dovetail to a tip plate for channeling combustion gases thereover; and, at least one tip rib extending outwardly from the tip plate between the leading and trailing edges. The tip rib is oriented so that an axis extending longitudinally therethrough is at an angle with respect to the radial axis for at least a designated portion of an axial length of the turbine blade. Such angle may be substantially the same across the designated portion or may vary thereacross. Accordingly, a recirculation zone of the combustion gases is formed adjacent a distal end of the tip rib which reduces a leakage flow of the combustion gases between the airfoil and the shroud for at least the designated portion of an axial length of the turbine blade.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partly sectional, isometric view of an exemplary gas turbine engine rotor blade mounted in a rotor disk within a surrounding shroud, with the blade having a tip in accordance with an exemplary embodiment of the present invention;

FIG. 2 is an isometric view of the blade tip as illustrated in FIG. 1 having a pair of aerodynamic tip ribs in accordance with an exemplary embodiment;

FIG. 3 is a top view of the blade tip illustrated in FIGS. 1 and 2;

FIG. 4 is an elevational, sectional view through the blade tip illustrated in FIG. 3 within the turbine shroud, taken generally along line

4

4, and depicting a maximum angle between a longitudinal axis through the blade tip ribs and the radial axis;

FIG. 5 is an elevational, sectional view through the blade tip illustrated in FIG. 3 within the turbine shroud, taken generally along line

5

5, and depicting a minimum angle between a longitudinal axis through the blade tip ribs and the radial axis;

FIG. 6 is an elevational, sectional view through an alternative blade tip like that illustrated in FIGS. 4 and 5, where a longitudinal axis through the blade tip rib at the pressure side of the airfoil forms an acute angle with respect to the radial axis and the blade tip rib at the suction side of the airfoil is substantially parallel to the radial axis;

FIG. 7 is an elevational, sectional view through a second alternative blade tip like that illustrated in FIGS. 4 and 5, where a longitudinal axis through the blade tip rib at the suction side of the airfoil forms an acute angle with respect to the radial axis in the upstream direction and the blade tip rib at the pressure side of the airfoil is substantially parallel to the radial axis;

FIG. 8 is an elevational, sectional view through a third alternative blade tip like that illustrated in FIGS. 4 and 5, where a longitudinal axis through the blade tip rib at the suction side of the airfoil forms an acute angle with respect to the radial axis in the downstream direction and the blade tip rib at the pressure side of the airfoil is substantially parallel to the radial axis;

FIG. 9 is an elevational, sectional view through a fourth alternative blade tip like that illustrated in FIGS. 4 and 5, where a third intermediate blade tip rib is positioned between the blade tip ribs located adjacent the pressure and suction sides of the airfoil;

FIG. 10A is an enlarged, partial sectional view through the blade tip illustrated in FIG. 4 within the turbine shroud depicting the flow of combustion gases adjacent the pressure side blade tip rib and through the gap between such rib and the turbine shroud; and,

FIG. 10B is an enlarged, partial sectional view through the blade tip illustrated in FIG. 4 within the turbine shroud depicting the flow of combustion gases adjacent the suction side blade tip rib, the area between the pressure and suction side blade tip ribs, and through the gap between such ribs and the turbine shroud.

Claims

1. A turbine blade for a gas turbine engine including an airfoil and integral dovetail for mounting said airfoil along a radial axis to a rotor disk inboard of a turbine shroud, said airfoil comprising:

wherein said tip rib is oriented so that an axis extending longitudinally therethrough is at an angle with respect to said radial axis for at least a designated portion of an axial length of said turbine blade.

2. The turbine blade of claim 1, wherein said angle between said longitudinal axis and said radial axis is substantially the same across said designated portion.

3. The turbine blade of claim 1, wherein said angle between said longitudinal axis and said radial axis varies across said designated portion.

4. The turbine blade of claim 3, wherein a minimum angle between said longitudinal axis of said tip rib and said radial axis is located adjacent said leading and tailing edge and gradually increases to a maximum angle at a designated point therebetween.

5. The turbine blade of claim 4, wherein said designated point for said maximum angle is located approximately one-fourth to one-half the distance from said leading edge to said trailing edge.

6. The turbine blade of claim 1, wherein said angle between said longitudinal axis and said radial axis is in a range of approximately 0°-70°.

7. The turbine blade of claim 1, wherein said angle between said longitudinal axis and said radial axis is in a range of approximately 20°-65°.

8. The turbine blade of claim 1, wherein said angle between said longitudinal axis and said radial axis is in a range of approximately 40°-60°.

9. The turbine blade of claim 1, further comprising a first tip rib located adjacent to said first sidewall and a second tip rib located adjacent to said second sidewall, wherein said first rib tip is oriented so that an axis extending longitudinally therethrough is at an angle with respect to said radial axis for at least a designated portion of an axial length of said turbine blade.

10. The turbine blade of claim 9, wherein said first tip rib is recessed with respect to said first sidewall to form a tip shelf adjacent said first tip rib.

11. The turbine blade of claim 1, further comprising a first tip rib located adjacent to said first sidewall and a second tip rib located adjacent to said second sidewall, wherein said second rib tip is oriented so that an axis extending longitudinally therethrough is at an angle with respect to said radial axis for at least a designated portion of an axial length of said turbine blade.

12. The turbine blade of claim 11, wherein said second tip rib is recessed with respect to said second sidewall to form a tip shelf adjacent said second tip rib.

13. The turbine blade of claim 11, wherein said angle between said longitudinal axis and said radial axis is in a range of approximately +60° to −60°.

14. The turbine blade of claim 1, further comprising a first tip rib located adjacent to said first sidewall and a second tip rib located adjacent to said second sidewall, wherein said first and second tip ribs are oriented so that an axis extending longitudinally through each respective tip rib is at an angle with respect to said radial axis for at least a designated portion of an axial length of said turbine blade.

15. The turbine blade of claim 14, further comprising a third tip rib extending outwardly from said tip plate between said leading and trailing edges, said third tip rib being spaced laterally between said first and second tip ribs.

16. The turbine blade of claim 1, wherein said angle between said longitudinal axis of said tip rib and said radial axis is more than approximately 5° for said designated portion of said rib.

17. The turbine blade of claim 1, wherein said designated portion extends for approximately 5-95% of a chord through said blade.

18. The turbine blade of claim 1, wherein said designated portion extends for approximately 7-80% of a chord through said blade.

19. The turbine blade of claim 1, wherein said designated portion extends for approximately 10-70% of a chord through said blade.

20. The turbine blade of claim 1, further comprising a plurality of cooling holes located adjacent to said tip rib in communication with a cooling channel disposed in said airfoil for receiving cooling fluid through said dovetail and providing a cooling film along at least one surface of said tip rib.

21. The turbine blade of claim 20, herein a junction between said first tip rib and said tip shelf is radiused so as to form a recirculation zone therein for said combustion gases and thereby maintain aid cooling film.

22. A turbine blade for a gas turbine engine including an airfoil and integral dovetail for mounting said airfoil along a radial axis to a rotor disk inboard of a turbine shroud, said airfoil comprising:

wherein said tip rib is oriented with respect to said radial axis so that a first recirculation zone of said combustion gases is formed adjacent a distal end of said tip rib which reduces a leakage flow of said combustion gases between said airfoil and said shroud for at least a designed portion of an axial length of said turbine blade.

23. The turbine blade of claim 22, said tip rib further being recessed with respect to said first sidewall to form a tip shelf adjacent said tip rib, wherein a junction between said first tip rib and said tip shelf is radiused so that a second recirculation zone of said combustion gases is formed therein which assists in maintaining a cooling film along said tip rib.

24. The turbine blade of claim 22, further comprising a first tip rib located adjacent to said first sidewall and a second tip rib located adjacent to said second sidewall wherein said first and second tip ribs are oriented with respect to said radial axis so that a first recirculation zone of said combustion gases is formed adjacent a distal end of said first tip rib and a second recirculation zone of said combustion gases is formed adjacent a distal end of said second tip rib, said first and second recirculation zones functioning to reduce a leakage flow of said combustion gases between said airfoil and said shroud for at least a designated portion of an axial length of said turbine blade.

25. The turbine blade of claim 24, wherein a first junction between said first tip rib and said tip plate and a second junction between said second tip rib and said tip plate are radiused so that a third recirculation zone of said combustion gases is formed between said first and second tip ribs.