Method and system for improving turbine blade performance

• US 10,934,858 B2
• Filed: 03/26/2019
• Issued: 03/02/2021
• Est. Priority Date: 12/11/2015
• Status: Active Grant

• Current Assignee: General Electric Company
• Original Assignee: General Electric Company
• Inventors: Schroeder, Eric Joseph, Vitt, Paul Hadley, Swenson, Timothy John, Gallier, Kirk Douglas, Wadia, Aspi Rustom
• Primary Examiner(s): Bogue, Jesse S
• Application Number:
  US16/365,105
• Publication Number:
  US 20190218918A1
• Time in Patent Office: 707 Days
• 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 5/145 Means for influencing boundary layers or secondary circulations for compressors F04D29/68
• CPC Class Codes: F01D 5/147 Construction, i.e. structural features, e.g. of weight-saving hollow blades F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F01D 5/20 Specially-shaped blade tips to seal space between tips and stator F01D5/225 takes precedence
• CPC Class Codes: F01D 9/04 forming ring or sector
• CPC Class Codes: F02C 3/06 the compressor comprising only axial stages F02C3/10 takes precedence
• CPC Class Codes: F02C 7/18 the medium being gaseous, e.g. air F02C7/125 takes precedence
• CPC Class Codes: F05D 2220/36 specially adapted for the fan of turbofan engines
• CPC Class Codes: F05D 2230/10 by removing material
• CPC Class Codes: F05D 2240/307 related to the tip of a rotor blade
• CPC Class Codes: F05D 2240/35 Combustors or associated equipment
• CPC Class Codes: F05D 2260/20 Heat transfer, e.g. cooling
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A turbine blade is described herein, the turbine blade including a blade root, a blade tip, and an airfoil extending between the blade root and the blade tip. The airfoil has opposite pressure and suction sides extending between a forward leading edge and an aft trailing edge of the airfoil, and a maximum thickness located between the leading edge and the trailing edge. The blade tip includes a winglet extending laterally outward from at least one of the pressure side and the suction side from a leading point between the leading edge and the maximum thickness aftward to a trailing point between the maximum thickness and the trailing edge.

Claims

1. A turbine blade comprising:

a blade root;

a blade tip; and

an airfoil extending between said blade root and said blade tip, said airfoil having opposite pressure and suction sides extending between a forward leading edge and an aft trailing edge of said airfoil, said airfoil having a maximum thickness located between said leading edge and said trailing edge,

said blade tip comprising a winglet extending laterally outward from said suction side from a leading point between said leading edge and said maximum thickness aftward to a trailing point between said maximum thickness and said trailing edge; and

said winglet comprises a forward first portion extending aft from said leading point between said leading edge and said maximum thickness with said leading point located aft of said leading edge, said first portion comprises a convex forward edge, and an aft second portion extending aft from said first portion, said second portion comprises a concave aft edge.

2. The turbine blade of claim 1, wherein said turbine blade is encased by a stationary shroud, said blade tip spaced apart from the stationary shroud by a tip clearance gap, and wherein said leading point of said winglet is positioned at a location where a vector sum of a pressure gradient between said pressure side and said suction side of said airfoil becomes greater than a relative momentum of air near the stationary shroud.

3. The turbine blade of claim 1, wherein said winglet comprises a side face forming a compound curve over at least a portion of said side face of said winglet.

4. The turbine blade of claim 1, wherein said winglet is, spaced a predetermined radial distance from said blade tip.

5. The turbine blade of claim 1, wherein said blade tip comprises a top face and said winglet comprises a top face continuous with said blade tip top face, said winglet extending laterally outward from said suction side.

6. The turbine blade of claim 1, wherein said blade tip further comprises a cooling aperture arrangement configured to channel cooling air to said winglet.

7. The turbine blade of claim 3, wherein said winglet is, spaced a predetermined radial distance from said blade tip.

8. The turbine blade of claim 7, wherein said blade tip comprises a top face and said winglet comprises a top face continuous with said blade tip top face, said winglet extending laterally outward from said suction side.

9. A method of providing a turbine blade, said method comprising:

providing a turbine blade extending radially outward from a blade root, the turbine blade including a blade tip formed at a distal end of the turbine blade and an airfoil extending between the blade root and the blade tip, wherein the airfoil has opposite pressure and suction sides extending between a forward leading edge and an aft trailing edge of the airfoil and a maximum thickness located between the leading edge and the trailing edge; and

providing a winglet coupled to the blade tip, the winglet extending laterally outward from the suction side from a leading point between the leading edge and the maximum thickness aftward to a trailing point between the maximum thickness and the trailing edge, with said leading point located aft of said leading edge, and said winglet comprises a forward first portion extending aft from said leading point between said leading edge and said maximum thickness, said first portion comprises a convex forward edge, and an aft second portion extending aft from said first portion, said second portion comprises a concave aft edge wherein the winglet is configured to guide a flow of leakage air over the blade tip away from the airfoil, said guiding causing a vortex formed from the leakage air to compact, said guiding further reducing a pressure loss across the blade tip through a blade passage defined therearound.

10. The method of claim 9, further comprising forming a compound curve over at least a portion of a side face of the winglet.

11. The method of claim 9, further comprising spacing the winglet a predetermined radial distance from the blade tip.

12. The method of claim 9, further comprising aligning the winglet with a top face of the blade tip, said aligning positioning the winglet in a continuous configuration with respect to the top face of the blade tip.

13. The method of claim 9, further comprising positioning the leading point of the winglet at a vortex onset point of the blade tip.

14. The method of claim 13, wherein a stationary shroud further defines the blade passage, and wherein positioning the leading point of the winglet at a vortex onset point of the blade tip comprises positioning the leading point of the winglet where a vector sum of a pressure gradient between the pressure side and the suction side of the airfoil becomes greater than a relative momentum of air near the stationary shroud.

15. The method of claim 9, further comprising providing a cooling aperture arrangement defined in the blade tip, the cooling aperture arrangement configured to channel cooling air to the winglet.

16. The method of claim 14, further comprising spacing the winglet a predetermined radial distance from the blade tip.

17. A turbofan engine comprising:

a core engine including a multistage compressor; and

a fan powered by a turbine driven by gas generated in said core engine,

said turbine comprising a turbine blade comprising;

a blade root;

a blade tip; and

an airfoil extending between said blade root and said blade tip, said airfoil having opposite pressure and suction sides extending between a forward leading edge and an aft trailing edge of said airfoil, said airfoil having a maximum thickness located between said leading edge and said trailing edge,

said blade tip comprising a winglet extending laterally outward from said suction side from a leading point between said leading edge and said maximum thickness aftward to a trailing point between said maximum thickness and said trailing edge, with said leading point located aft of said leading edge, and

said winglet comprises a forward first portion extending aft from said leading point between said leading edge and said maximum thickness, said first portion comprises a convex forward edge, and an aft second portion extending aft from said first portion, said second portion comprises a concave aft edge.

18. The turbofan engine of claim 17, wherein said turbine blade is encased by a stationary shroud, said blade tip spaced apart from the stationary shroud by a tip clearance gap, and wherein said leading point of said winglet is positioned at a location where a vector sum of a pressure gradient between said pressure side and said suction side of said airfoil becomes greater than a relative momentum of air near the stationary shroud.