Gas turbine engine airfoil leading edge cooling

Patent metadata

Field	Value
Publication number	US-10329923-B2
Application number	US-201514633677-A
Country	US
Kind code	B2
Filing date	Feb 27, 2015
Priority date	Mar 10, 2014
Publication date	Jun 25, 2019
Grant date	Jun 25, 2019

Abstract

An example gas turbine engine component includes an airfoil having a leading edge area, a first circuit to cool a first section of the leading edge area, and a second circuit to cool a second section of the leading edge area. The first circuit separate and distinct from the second circuit within the airfoil.

First claim

We claim:

1. A gas turbine engine component, comprising:
an airfoil having a leading edge area that corresponds to a stagnation traverse region of the airfoil;
a first circuit to cool a first section of the leading edge area; and
a second circuit to cool a second section of the leading edge area, the first circuit separate and distinct from the second circuit within the airfoil.

2. The component of claim 1 , wherein the first circuit and the second circuit are both contained within the airfoil.

3. The component of claim 1 , wherein flow across the airfoil has an incidence angle relative to the airfoil, and the leading edge area represents an area of the airfoil traversed by a maximum stagnation point of flow across the airfoil through one-hundred-eighty degrees of incident angle shift.

4. The component of claim 1 , wherein the airfoil is a turbine airfoil.

5. The component of claim 1 , wherein the airfoil has a forward apex relative to a direction of flow across the airfoil, the forward apex extending radially along the airfoil, and the first circuit and the second circuit each cooling at least some of the airfoil at the forward apex.

6. The component of claim 5 , wherein the first circuit is configured to communicate air to a radially inner portion of the forward apex, and the second circuit is configured to communicate air to a radially outer portion of the forward apex.

7. The component of claim 1 , wherein the first circuit and the second circuit are configured to receive flow from a radially inner end of the airfoil.

8. The component of claim 1 , wherein the second circuit comprises a serpentine cooling circuit.

9. The component of claim 1 , wherein the airfoil includes at least two other cooling circuits.

10. A gas turbine engine, comprising:
an airfoil;
a variable vane array upstream from the airfoil;
a first circuit within the airfoil to cool a leading edge area of the airfoil; and
a second circuit within the airfoil to cool a leading edge area of the airfoil, wherein the leading edge area corresponds to a stagnation traverse region of the airfoil.

11. The gas turbine engine of claim 10 , wherein the airfoil is a turbine airfoil.

12. The gas turbine engine of claim 11 , wherein the turbine airfoil is in a first stage of a turbine section of the gas turbine engine.

13. The gas turbine engine of claim 10 , wherein flow across the airfoil has an incidence angle relative to the airfoil, and the leading edge area represents an area of the airfoil traversed by a maximum stagnation point of flow across the airfoil through one-hundred-eighty degrees of incident angle shift.

14. The gas turbine engine of claim 10 , wherein the airfoil has a forward apex relative to a direction of flow across the airfoil, the forward apex extending radially along the airfoil, and the first circuit and the second circuit each cooling at least some of the airfoil at the forward apex.

15. The gas turbine engine of claim 10 , wherein the first circuit is separate and distinct from the second circuit.

16. A method of cooling an airfoil, comprising:
communicating a cooling flow to a leading edge area of an airfoil through both a first cooling circuit and a second cooling circuit that is separate and distinct from the first cooling circuit within the airfoil, the leading edge area corresponding to a stagnation traverse region of the airfoil.

17. The method of claim 16 , wherein flow across the airfoil has an incidence angle relative to the airfoil, and the leading edge area represents an area of the airfoil traversed by a maximum stagnation point of flow across the airfoil through one-hundred-eighty degrees of incident angle shift.

18. The method of claim 16 , further comprising communicating the cooling flow from both the first circuit and the second circuit to a forward apex of the airfoil, the forward apex extending radially along the airfoil.

19. The method of claim 16 , wherein the first circuit is separate and distinct from the second circuit.

Assignees

• United Technologies Corp

Inventors

• Slavens Thomas N
• Mongillo Jr Dominic J
• Devore Matthew A

Classifications

• F01D5/08

  Heating, heat-insulating or cooling means {(specially adapted for radial flow machines or engines F01D5/04)} · CPC title

• F01D5/186

  Film cooling (F01D5/187 takes precedence) · CPC title

• F01D9/041

  using blades (F01D5/148 takes precedence) · CPC title

• F05D2260/202

  by film cooling · CPC title

• F05D2240/12

  Fluid guiding means, e.g. vanes · CPC title