TURBINE BLADES WITH TIP PORTIONS HAVING CONVERGING COOLING HOLES

• US 20150078916A1
• Filed: 09/18/2013
• Published: 03/19/2015
• Est. Priority Date: 09/18/2013
• Status: Active Grant

• Current Assignee: Honeywell International Inc.
• Original Assignee: Honeywell International Inc.
• Inventors: Riahi, Ardeshir, Crites, Daniel C., Bedrosyan, Loris
• Granted Patent: US 9,856,739 B2
• Time in Patent Office: Days
• US Class Current: 416/97.R
• CPC Class Codes: F01D 11/10 using sealing fluid, e.g. steam
• CPC Class Codes: F01D 5/186 Film cooling F01D5/187 takes precedence
• 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 2240/307 related to the tip of a rotor blade
• 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 rotor blade is provided with for a turbine section of an engine that includes a shroud surrounding the rotor blade. The rotor blade includes a platform and an airfoil extending from the platform into a mainstream gas path. 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, a tip cap extending between the suction side wall and the pressure side wall, a first squealer tip extension extending from the pressure side wall at a first angle relative to the pressure side wall, the first squealer tip extension defining a first cooling hole that converges between an inlet and an outlet; an internal cooling circuit configured to deliver cooling air to a gap between the pressure side squealer tip extension and the shroud via the first cooling hole.

Claims

1. A turbine rotor blade for a turbine section of an engine, the turbine section including a shroud surrounding the turbine rotor blade, 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;

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

a first squealer tip extension extending from the pressure side wall at a first angle relative to the pressure side wall, the first squealer tip extension defining a first cooling hole that converges between an inlet and an outlet; and

an internal cooling circuit configured to deliver cooling air to a gap between the pressure side squealer tip extension and the shroud via the first cooling hole.

2. The turbine rotor blade of claim 1, further comprising a second squealer tip extension extending from the suction side wall at a second angle relative to the suction side wall.

3. The turbine rotor blade of claim 2, wherein the first angle is approximately equal to the second angle.

4. The turbine rotor blade of claim 2, wherein the first squealer tip extension is approximately parallel to the second squealer tip extension in a chord-wise direction.

5. The turbine rotor blade of claim 2, wherein the second squealer tip extension is generally solid.

6. The turbine rotor blade of claim 1, wherein the pressure side squealer tip extension terminates at a radial wall, and wherein the outlet of the first cooling hole is formed on the radial wall.

7. The turbine rotor blade of claim 1, wherein the first cooling hole has a longitudinal center line oriented at the first angle relative to the pressure side wall.

8. The turbine rotor blade of claim 1, wherein the first cooling hole is oriented such that the cooling air blocks at least a portion of tip leakage between the first squealer tip extension and the shroud.

9. The turbine rotor blade of claim 1, wherein the first angle is approximately 15°

-60°

10. The turbine rotor blade of claim 1, wherein the airfoil further comprises a step formed in the first squealer tip extension.

11. The turbine rotor blade of claim 10, wherein the first cooling hole at least partially extends through the step.

12. The turbine rotor blade of claim 11, wherein the first cooling hole includes a closed channel section through the step and an open channel section through the first squealer tip extension.

13. A turbine section of a gas turbine engine, comprising:

a shroud; and

a rotor assembly with a turbine rotor blade comprising

a platform defining a mainstream gas path; and

an airfoil extending from the platform into the mainstream gas path;

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;

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

a first squealer tip extension extending from the pressure side wall to define a gap between the rotor assembly and the shroud, the first squealer tip extension extending at a first angle relative to the pressure side wall and defining a first cooling hole that converges between an inlet and an outlet; and

an internal cooling circuit configured to deliver cooling air to the gap between the pressure side squealer tip extension and the shroud via the first cooling hole.

14. The turbine section of claim 13, wherein the airfoil further comprises a second squealer tip extension extending from the suction side wall at a second angle relative to the suction side wall.

15. The turbine section of claim 14, wherein the second squealer tip extension is generally solid.

16. The turbine section of claim 13, wherein the pressure side squealer tip extension terminates at a radial wall, and wherein the outlet of the first cooling hole is formed on the radial wall.

17. The turbine section of claim 13, wherein the first cooling hole has a longitudinal center line oriented at the first angle relative to the pressure side wall.

18. The turbine section of claim 13, wherein the first cooling hole is oriented such that the cooling air blocks at least a portion of tip leakage between the first squealer tip extension and the shroud.

19. The turbine section of claim 13, wherein the first angle is approximately 15°

-60°

20. The turbine section of claim 13, wherein the airfoil further comprises a step formed in the first squealer tip extension on the tip cap, and wherein the first cooling hole at least partially extends through the step.