Gas turbine engine airfoil with leading edge trench and impingement cooling

Patent metadata

Field	Value
Publication number	US-10240464-B2
Application number	US-201415035837-A
Country	US
Kind code	B2
Filing date	Nov 5, 2014
Priority date	Nov 25, 2013
Publication date	Mar 26, 2019
Grant date	Mar 26, 2019

Abstract

A gas turbine engine airfoil includes an airfoil structure including an exterior surface that is provided by an exterior wall that has a leading edge. A radially extending interior wall within the airfoil structure separates first and second radial cooling passages. The first cooling passage is arranged near the leading edge. A radially extending trench is in the leading edge. An impingement hole is provided in the interior wall and is configured to direct a cooling fluid from the second cooling passage to the first cooling passage and onto the exterior wall at the leading edge.

First claim

What is claimed is:

1. A gas turbine engine airfoil comprising:
an airfoil structure including an exterior surface provided by an exterior wall having a leading edge, a radially extending interior wall within the airfoil structure separating first and second radial cooling passages, the first cooling passage arranged near the leading edge, a radially extending trench in the leading edge, and an impingement hole provided in the interior wall and configured to direct a cooling fluid from the second cooling passage to the first cooling passage and onto the exterior wall at the leading edge;
wherein the leading edge of the exterior wall includes interior pressure and suction side portions having legs joined to one another;
wherein the interior pressure and suction side portions are generally U-shaped; and
wherein the impingement hole is provided by a plurality of impingement hole rows, the plurality of impingement hole rows with each impingement hole directed along a vector towards a respective section of a curvilinear interior surface in which the vector is normal to that curvilinear interior surface to create a respective acute angle relative to the mechanical chord line of the airfoil shape, said respective acute angle being between about 0.degrees. to about 80.degrees.

2. The gas turbine engine airfoil according to claim 1 , wherein the first cooling passage ejects post-impingement air to the external surface of the airfoil through a series of cooling holes.

3. The gas turbine engine airfoil according to claim 1 , wherein the trench includes radially spaced apart trench cooling holes fluidly connecting the first cooling passage to the exterior surface.

4. The gas turbine engine airfoil according to claim 3 , wherein radially spaced apart bridges interconnect the interior pressure and suction side portions, the trench cooling holes provided by the bridges.

5. The gas turbine engine airfoil according to claim 4 , wherein the trench cooling holes are slots.

6. The gas turbine engine airfoil according to claim 1 , wherein the pressure and suction side portion each include legs that are arranged parallel to one another.

7. The gas turbine engine airfoil according to claim 1 , wherein the trench is provided at an aerodynamic stagnation line.

8. The gas turbine engine airfoil according to claim 1 , wherein the plurality of impingement hole rows includes first and second rows of impingement holes, the first row of impingement holes is aligned with an inflection line of the interior pressure side portion, and the second row of impingement holes is aligned with an inflection line of the interior suction side portion.

9. A gas turbine engine airfoil comprising:
an airfoil structure including an exterior surface provided by an exterior wall having a leading edge, a radially extending interior wall within the airfoil structure separating first and second radial cooling passages, the first cooling passage arranged near the leading edge, at least two sets of impingement holes provided in the interior wall and configured to direct a cooling fluid from the second cooling passage to the first cooling passage and onto the exterior wall at the leading edge, and a separating wall provided between the at least two sets of impingement holes that is joined to the interior and exterior walls and separates the first cooling passage into first and second passageways;
wherein the leading edge of the exterior wall includes interior pressure and suction side portions having legs joined to one another;
wherein the interior pressure and suction side portions are generally U-shaped; and
wherein each of the sets of impingement holes are provided by a plurality of impingement hole rows, the plurality of impingement hole rows with each impingement hole directed along a vector towards a respective section of a curvilinear interior surface in which the vector is normal to that curvilinear interior surface to create a respective acute angle relative to the mechanical chord line of the airfoil shape, said respective acute angle being between about 0.degrees. to about 80.degrees.

10. The gas turbine engine airfoil according to claim 9 , wherein the first cooling passage ejects post-impingement air to the external surface of the airfoil through a series of cooling holes.

11. The gas turbine engine airfoil according to claim 9 , wherein the plurality of impingement hole rows includes first and second rows of impingement holes, the first row of impingement holes is aligned with an inflection line of the interior pressure side portion, and the second row of impingement holes is aligned with an inflection line of the interior suction side portion.

12. A method of manufacturing a gas turbine engine component, comprising the steps of:
(a) forming an airfoil structure including an exterior surface provided by an exterior wall having a leading edge connecting a suction and pressure side;
wherein the leading edge of the exterior wall includes interior pressure and suction side portions having legs joined to one another;
wherein the interior pressure and suction side portions are generally U-shaped;

(b) forming a radially extending interior wall within the airfoil structure separating a first and second radial cooling passages in which the first cooling passage is arranged near the leading edge;
(c) forming a plurality of adjacent impingement hole rows in the interior wall and configured to direct a cooling fluid from the second cooling passage to the first cooling passage;
wherein the impingement hole is provided by a plurality of impingement hole rows, the plurality of impingement hole rows with each impingement hole directed along a vector towards a respective section of a curvilinear interior surface in which the vector is normal to that curvilinear interior surface to create a respective acute angle relative to the mechanical chord line of the airfoil shape, said respective acute angle being between about 0.degrees. to about 80.degrees; and

(d) forming a radially extending trench in the exterior wall of the first cavity.

13. The method according to claim 12 , wherein the providing step includes additively manufacturing the airfoil structure.

14. The method according to claim 12 , wherein the providing step includes additively manufacturing a core having a shape corresponding to the airfoil structure.

15. The method according to claim 14 , wherein the shape is a positive of the airfoil structure.

16. The method according to claim 14 , wherein the shape is a negative of the airfoil structure.

Assignees

• United Technologies Corp

Inventors

• Slavens Thomas N
• Snyder Brooks E

Classifications

• F05D2230/31

  Layer deposition · CPC title

• F01D25/12

  Cooling · CPC title

• F01D9/02

  Nozzles; Nozzle boxes; Stator blades; Guide conduits {, e.g. individual nozzles (nozzle boxes F01D9/047)} · CPC title

• F05D2260/202

  by film cooling · CPC title

• F01D5/18

  Hollow blades, {i.e. blades with cooling or heating channels or cavities (structure of hollow blades in general F01D5/147)}; Heating, heat-insulating or cooling means on blades · CPC title