GAS TURBINE ENGINE COMPONENTS WITH BLADE TIP COOLING

• US 20140083116A1
• Filed: 09/27/2012
• Published: 03/27/2014
• 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
• Granted Patent: US 9,546,554 B2
• Time in Patent Office: Days
• US Class Current: 60/806
• 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; and

a flow accelerator positioned within the tip cap passage of the internal cooling circuit.

2. The turbine rotor blade of claim 1, wherein the airfoil includes an interior wall that defines the tip cap passage of the internal cooling circuit with the tip cap.

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

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 4, wherein the flow accelerator is generally tear-drop shaped and generally extends in a chordwise direction.

7. The turbine rotor blade of claim 1, wherein the flow accelerator is arranged to divide the tip cap passage into a first portion defined between the flow accelerator and the suction side wall and a second portion defined between the flow accelerator and the pressure side wall.

8. The turbine rotor blade of claim 1, 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.

9. 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.

10. The turbine rotor blade of claim 9, wherein the additional flow accelerators are airfoil-shaped.

11. The turbine rotor blade of claim 9, wherein the additional flow accelerators are pin-shaped.

12. 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.

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

14. 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.

15. The gas turbine engine of claim 14, wherein the turbine rotor includes an interior wall between the airfoil side walls that defines the tip cap passage with the tip cap.

16. The gas turbine engine of claim 15, wherein the flow accelerator extends in a radial direction from the interior wall to the tip cap.

17. The gas turbine engine of claim 14, 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.

18. The gas turbine engine of claim 14, wherein the flow accelerator is hollow with an outer wall and interior space.

19. The gas turbine engine of claim 18, 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.

20. 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,

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

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

an internal cooling circuit forming tip cap passage with the interior wall, the tip cap passage configured to deliver cooling air to the tip cap;

a flow accelerator positioned within the tip cap passage of the internal cooling circuit and extending in a radial direction from the interior wall to the tip cap,

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.