Gas turbine engine components with blade tip cooling

• US 9,546,554 B2
• Filed: 09/27/2012
• Issued: 01/17/2017
• Est. Priority Date: 09/27/2012
• Status: Active Grant

• Current Assignee: Honeywell International Inc.
• Original Assignee: Honeywell International Inc.
• Inventors: Crites, Daniel Cregg, Halfmann, Steve, Morris, Mark C., Riahi, Ardeshir
• Primary Examiner(s): Nguyen, Ninh H
• Assistant Examiner(s): Lambert, Wayne A
• Application Number:
  US13/629,284
• Publication Number:
  US 20140083116A1
• Time in Patent Office: 1,573 Days
• Field of Search: 415/115, 416/96.R, 416/97.R
• US Class Current: 1/1
• 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: F05D 2260/201 by impingement of a fluid
• CPC Class Codes: F05D 2260/202 by film cooling
• CPC Class Codes: F05D 2260/221 Improvement of heat transfer
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A turbine rotor blade for a turbine section of an engine is provided. The rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path of the turbine section. The airfoil includes a pressure side wall, a suction side wall joined to the pressure side wall at a leading edge and a trailing edge, and a tip cap extending between the suction side wall and the pressure side wall. The rotor blade further includes an internal cooling circuit having a tip cap passage configured to deliver cooling air to the tip cap and a flow accelerator positioned within the tip cap passage of the internal cooling circuit.

Claims

1. A turbine rotor blade for a turbine section of an engine, comprising:

a platform; and

an airfoil extending from the platform into a mainstream gas path of the turbine section, the airfoil comprising

a pressure side wall,

a suction side wall joined to the pressure side wall at a leading edge and a trailing edge, and

a tip cap extending between the suction side wall and the pressure side wall,

an internal cooling circuit having a tip cap passage configured to deliver cooling air to the tip cap, the airfoil including an interior wall that defines the tip cap passage with the tip cap, the pressure side wall, and the suction side wall, wherein the tip cap passage has a chordwise length between an inlet and an outlet at the trailing edge; and

a flow accelerator positioned within the tip cap passage of the internal cooling circuit, the flow accelerator extending in a radial direction from the tip cap toward the interior wall to define at least a first flow area for the cooling air between the flow accelerator and the pressure side wall and a second flow area for the cooling air between the flow accelerator and the suction side wall,

wherein the tip cap passage has at least a first cross-sectional area defined between the tip cap, the internal wall, the suction side wall, and the pressure side wall, and wherein the flow accelerator, at a position corresponding to the first cross-sectional area, has a second cross-sectional area that is at least 50% of the first cross-sectional area, and

wherein the flow accelerator is generally tear-drop or airfoil shaped and generally extends in a chordwise direction to accelerate a flow of the cooling air through the tip cap passage,

wherein the flow accelerator is hollow with an outer wall and interior space, and wherein the internal cooling circuit includes a central passage that delivers cooling air that impinges on an underside of the outer wall within the interior space.

2. The turbine rotor blade of claim 1, wherein the flow accelerator extends in the radial direction from the tip cap to the interior wall.

3. The turbine rotor blade of claim 2, wherein the pressure side wall defines first squealer tip extensions extending beyond the tip cap and the suction side wall defines second squealer tip extensions extending beyond the tip cap, and wherein the tip cap forms a first heat transfer path between the first squealer tip extensions and the flow accelerator and a second heat transfer path between the second squealer tip extensions and the flow accelerator.

4. The turbine rotor blade of claim 1, wherein the tip cap passage of the internal cooling circuit has a radial height defined between the interior wall and the tip cap, and wherein the flow accelerator extends from the tip cap at a distance less than the radial height.

5. The turbine rotor blade of claim 4, wherein the distance is approximately three-quarters of the radial height.

6. The turbine rotor blade of claim 1, wherein the flow accelerator is a first flow accelerator and the turbine rotor further comprises additional flow accelerators positioned within the tip cap passage of the internal cooling circuit.

7. The turbine rotor blade of claim 6, wherein the additional flow accelerators are airfoil-shaped.

8. The turbine rotor blade of claim 6, wherein the additional flow accelerators are pin-shaped.

9. The turbine rotor blade of claim 6, wherein the additional flow accelerators includes a second flow accelerator with a first portion extending in across the tip cap passage between the first flow accelerator and the suction side wall and a second portion extending in across the tip cap passage between the first flow accelerator and the pressure side wall.

10. The turbine rotor blade of claim 1, wherein the flow accelerator includes a radial portion extending from the tip cap to the interior wall and a lateral portion extending from the pressure side wall to the suction side wall.

11. The turbine rotor blade of claim 1, wherein the pressure side wall includes film cooling holes fluidly coupled to the tip cap passage.

12. The turbine rotor blade of claim 1, wherein the flow accelerator is positioned within an interior of the tip cap passage and separated from the trailing edge and the outlet at a distance.

13. A gas turbine engine, comprising:

a compressor section configured to receive and compress air;

a combustion section coupled to the compressor section and configured to receive the compressed air, mix the compressed air with fuel, and ignite the compressed air and fuel mixture to produce combustion gases; and

a turbine section coupled to the combustion section and configured to receive the combustion gases, the turbine section defining a combustion gas path and comprising a turbine rotor positioned within the combustion gas path, the turbine rotor comprising

a suction side wall;

a pressure side wall joined to the suction side wall at a leading edge and a trailing edge;

a tip cap extending between the pressure side wall and the suction side wall;

an interior wall extending between the pressure side wall and the suction side wall;

an internal cooling circuit including a tip cap passage at least partially defined the pressure side wall, the suction side wall, the tip cap, and the interior wall and configured to direct cooling air to the tip cap, wherein the tip cap passage has a chordwise length between an inlet and an outlet at the trailing edge; and

a flow accelerator positioned within the tip cap passage, the flow accelerator extending in a radial direction from the tip cap toward the interior wall to define at least a first flow area for the cooling air between the flow accelerator and the pressure side wall and a second flow area for the cooling air between the flow accelerator and the suction side wall,

wherein the tip cap passage has at least a first cross-sectional area defined between the tip cap, the internal wall, the suction side wall, and the pressure side wall, and wherein the flow accelerator, at a position corresponding to the first cross-sectional area, has a second cross-sectional area that is at least 50% of the first cross-sectional area, and

wherein the flow accelerator is generally tear-drop or airfoil shaped and generally extends in a chordwise direction to accelerate a flow of the cooling air through the tip cap passage,

wherein the flow accelerator is hollow with an outer wall and interior space, and wherein the internal cooling circuit includes a central passage that delivers cooling air that impinges on an underside of the outer wall within the interior space.

14. The gas turbine engine of claim 13, wherein the flow accelerator extends in the radial direction from the tip cap to the interior wall.

15. The gas turbine engine of claim 13, wherein the tip cap passage of the internal cooling circuit has a radial height defined between the interior wall and the tip cap, and wherein the flow accelerator extends from the tip cap at a distance less than the radial height.

16. The gas turbine engine of claim 13, wherein the flow accelerator is hollow with an outer wall and interior space.

17. A gas turbine engine, comprising:

a compressor section configured to receive and compress air;

a combustion section coupled to the compressor section and configured to receive the compressed air, mix the compressed air with fuel, and ignite the compressed air and fuel mixture to produce combustion gases; and

a turbine section coupled to the combustion section and configured to receive the combustion gases, the turbine section defining a combustion gas path and comprising a turbine rotor positioned within the combustion gas path, the turbine rotor comprising

airfoil side walls;

a tip cap extending between the airfoil side walls;

an internal cooling circuit at least partially defined by the airfoil side walls and including a tip cap passage configured to direct cooling air to the tip cap; and

a flow accelerator positioned within the tip cap passage,

wherein the flow accelerator is hollow with an outer wall and interior space, and

wherein the internal cooling circuit includes a central passage that delivers cooling air that impinges on an underside of the outer wall within the interior space.