GAS TURBINE ENGINE AIRFOIL

• US 20140154087A1
• Filed: 05/21/2013
• Published: 06/05/2014
• Est. Priority Date: 12/17/2008
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Kirchner, Jody, Dong, Yuan, Hingorani, Sanjay S.
• Granted Patent: US 8,807,951 B2
• Time in Patent Office: Days
• US Class Current: 416/223.R
• CPC Class Codes: F01D 5/12 Blades specially adapted for radial flow machines or engines F01D5/04; blade roots F01D5/30; rotors with blades adjustable in operation F01D7/00; stator blades F01D9/02
• 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: F04D 29/324 Blades
• CPC Class Codes: F05D 2240/301 Cross-sectional characteristics
• CPC Class Codes: Y10T 29/49321 Assembling individual fluid flow interacting members, e.g., blades, vanes, buckets, on rotary support member
• CPC Class Codes: Y10T 29/49336 Blade making
• CPC Class Codes: Y10T 29/49337 Composite blade

Abstract

A blade for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a platform and an airfoil that extends from the platform. The airfoil extends in span between a root and a tip region and in chord between a leading edge and a trailing edge. A sweep angle is defined at the leading edge and a dihedral angle is defined relative to the chord of the airfoil. The sweep angle and the dihedral angle are localized at the tip region and extend over a distance of the airfoil equivalent to about 10% to about 40% of the span.

Claims

1. A blade for a gas turbine engine, comprising:

a platform;

an airfoil that extends from said platform, said airfoil extending in span between a root and a tip region and in chord between a leading edge and a trailing edge;

a sweep angle defined at said leading edge;

a dihedral angle defined relative to said chord of said airfoil; and

said sweep angle and said dihedral angle localized at said tip region and extending over a distance of said airfoil equivalent to about 10% to about 40% of said span.

2. The blade as recited in claim 1, wherein said sweep angle is a forward sweep angle that extends in an upstream direction relative to the gas turbine engine.

3. The blade as recited in claim 1, wherein said dihedral angle is a positive dihedral angle.

4. The blade as recited in claim 3, wherein said positive dihedral angle extends between a suction surface of said airfoil and a shroud assembly adjacent said tip region.

5. The blade as recited in claim 1, wherein said sweep angle is defined parallel to said chord.

6. The blade as recited in claim 1, wherein said dihedral angle is defined tangentially relative to said chord as measured from a center of gravity of said airfoil.

7. The blade as recited in claim 1, wherein said sweep angle and said dihedral angle extend from an outer edge of said tip region radially inward along a radial axis over a distance equal to about 10% to about 40% of said span.

8. The blade as recited in claim 1, wherein said sweep angle and said dihedral angle extend over a distance equal to about 20% to about 30% of said span.

9. The blade as recited in claim 1, wherein said sweep angle and said dihedral angle extend over a distance equal to about 25% to about 30% of said span.

10. The blade as recited in claim 1, wherein said sweep angle and said dihedral angle extend over a distance equal to about 35% of said span.

11. The blade as recited in claim 1, wherein said blade is a rotor blade.

12. The blade as recited in claim 1, wherein said blade is a stator blade.

13. The blade as recited in claim 1, wherein said blade is a fan blade.

14. A gas turbine engine, comprising:

a shroud assembly;

a blade at least partially surrounded by said shroud assembly;

said blade having an airfoil extending in span between a root and a tip region and in chord between a leading edge and a trailing edge, said airfoil including a sweep angle defined at said leading edge and a dihedral angle defined relative to said chord; and

said sweep angle and said dihedral angle are localized at said tip region of said airfoil.

15. The gas turbine engine as recited in claim 14, wherein said sweep angle is a forward sweep angle that extends in an upstream direction relative to the gas turbine engine.

16. The gas turbine engine as recited in claim 14, wherein said dihedral angle is a positive dihedral angle.

17. The gas turbine engine as recited in claim 14, wherein said dihedral angle extends between a suction surface of said airfoil and said shroud assembly adjacent said tip region.

18. The gas turbine engine as recited in claim 14, wherein said sweep angle and said dihedral angle extend over a distance of said airfoil equivalent to about 10% to about 40% of said span.

19. The gas turbine engine as recited in claim 14, wherein said sweep angle and said dihedral angle extend from an outer edge of said tip region radially inward along a radial axis over a distance equal to about 10% to about 40% of said span.

20. The gas turbine engine as recited in claim 14, wherein said sweep angle and said dihedral angle extend over a distance equal to about 25% to about 30% of said span.