Variable immersion lobe mixer for turbofan jet engine exhaust and method of fabricating the same

Patent metadata

Field	Value
Publication number	US-9995245-B2
Application number	US-201314396800-A
Country	US
Kind code	B2
Filing date	Apr 26, 2013
Priority date	Apr 27, 2012
Publication date	Jun 12, 2018
Grant date	Jun 12, 2018

Abstract

A method of fabricating a mixer for a gas turbine engine is provided. The method includes forming a forward end and an aft end of the mixer, and forming an annularly undulating contour that defines a plurality of core immersion lobes and a plurality of bypass immersion lobes between the forward end and the aft end. The plurality of bypass immersion lobes includes a first bypass immersion lobe and a second bypass immersion lobe. The first bypass immersion lobe has a first crown contour line extending from the forward end to the aft end of the mixer, and the second bypass immersion lobe has a second crown contour line extending from the forward end to the aft end of the mixer. The first crown contour line is different than the second crown contour line.

First claim

What is claimed is:

1. A method of fabricating a mixer for a gas turbine engine having a core system for directing a core flow of gas through the gas turbine engine, and a bypass duct external to the core system for directing a bypass flow of gas through the gas turbine engine, said method comprising:
forming a forward end and an aft end of the mixer; and
forming the aft end of the mixer to have a trailing edge which defines an annularly undulating contour that defines a plurality of core immersion lobes and a plurality of bypass immersion lobes between the forward end and the aft end of the mixer, wherein the plurality of bypass immersion lobes includes a first bypass immersion lobe, a second bypass immersion lobe, and a third bypass immersion lobe, the first bypass immersion lobe having a first crown contour line extending from the forward end to the aft end of the mixer, the second bypass immersion lobe having a second crown contour line extending from the forward end to the aft end of the mixer, wherein the first crown contour line is different than the second crown contour line, and
wherein the first bypass immersion lobe extends radially farther into the bypass duct than the second bypass immersion lobe, and
wherein the third bypass immersion lobe extends into the bypass duct a different radial distance than the respective distances to which the first bypass immersion lobe and the second bypass immersion lobe extend.

2. The method of claim 1 , further comprising forming the first crown contour line to have a different curvature than the second crown contour line.

3. The method of claim 1 , further comprising:
forming the first bypass immersion lobe and the second bypass immersion lobe to be generally semi-elliptically shaped, the first bypass immersion lobe having a first semi-major axis at the trailing edge, the second bypass immersion lobe having a second semi-major axis at the trailing edge, wherein the first semi-major axis is longer than the second semi-major axis.

4. The method of claim 1 , further comprising forming the mixer from a ceramic matrix composite (CMC) material,
wherein the mixer comprises a plurality of first bypass immersion lobes,
wherein the mixer comprises a plurality of second bypass immersion lobes, and
wherein there are more second bypass immersion lobes than first bypass immersion lobes.

5. The method of claim 4 , further comprising forming the trailing edge to be a scalloped trailing edge.

6. A mixer for a gas turbine engine having a core system for directing a core flow of gas through said gas turbine engine, and a bypass duct external to said core system for directing a bypass flow of gas through said gas turbine engine, said mixer comprising:
a forward end;
an aft end defining a trailing edge; and
the trailing edge defining an annularly undulating contour defining a plurality of core immersion lobes and a plurality of bypass immersion lobes between said forward end and said aft end of the mixer, wherein said plurality of bypass immersion lobes comprises a first bypass immersion lobe, a second bypass immersion lobe, and a third bypass immersion lobe, said first bypass immersion lobe having a first crown contour line extending from said forward end to said aft end of said mixer, said second bypass immersion lobe having a second crown contour line extending from said forward end to said aft end of said mixer, wherein said first crown contour line is different than said second crown contour line, and
wherein the first bypass immersion lobe extends radially farther into the bypass duct than the second bypass immersion lobe, and
wherein the third bypass immersion lobe extends into the bypass duct a different radial distance than the respective distances to which the first bypass immersion lobe and the second bypass immersion lobe extend.

7. The mixer of claim 6 , wherein said first crown contour line has a different curvature than said second crown contour line.

8. The mixer of claim 6 , wherein said first bypass immersion lobe and said second bypass immersion lobe are generally semi-elliptically shaped, said first bypass immersion lobe having a first semi-major axis at said trailing edge, said second bypass immersion lobe having a second semi-major axis at said trailing edge, wherein said first semi-major axis is different than said second semi-major axis.

9. The mixer of claim 6 , wherein said mixer comprises a plurality of said first bypass immersion lobes and a plurality of said second bypass immersion lobes.

10. The mixer of claim 9 , wherein each of said first bypass immersion lobes is disposed between a pair of adjacent second bypass immersion lobes.

11. The mixer of claim 6 , wherein said trailing edge is a scalloped trailing edge.

12. A gas turbine engine comprising:
a core system comprising a compressor assembly, a combustion assembly aft of said compressor assembly, and a turbine assembly aft of said combustion assembly, wherein said compressor assembly, said combustion assembly, and said turbine assembly are arranged in flow communication along an axial dimension of said gas turbine engine;
a bypass duct extending along the axial dimension external to said core system;
a fan system forward of said compressor assembly and said bypass duct, said fan system configured to provide a core flow of gas into said core system and a bypass flow of gas into said bypass duct; and
an exhaust system aft of said core system and said bypass duct, said exhaust system comprising:
a tailpipe configured to receive the core flow of gas and the bypass flow of gas; and
a mixer coupled to said core system aft of said turbine assembly within said tailpipe to define a bypass flow region of said tailpipe and a core flow region of said tailpipe, said mixer comprising:
a forward end of said mixer;
an aft end of said mixer defining a trailing edge; and
the trailing edge defining an annularly undulating contour defining a plurality of core immersion lobes and a plurality of bypass immersion lobes between said forward end of said mixer and said aft end of said mixer, said core immersion lobes configured to deliver said bypass flow of gas into said core flow region of said tailpipe, said bypass immersion lobes configured to deliver said core flow of gas into said bypass flow region of said tailpipe,
wherein said plurality of bypass immersion lobes comprises a first bypass immersion lobe, a second bypass immersion lobe, and a third bypass immersion lobe, said first bypass immersion lobe having a first crown contour line extending from said forward end of said mixer to said aft end of said mixer, said second bypass immersion lobe having a second crown contour line extending from said forward end to said aft end of said mixer, wherein said first crown contour line is different than said second crown contour line, and
wherein the first bypass immersion lobe extends farther in a radial direction into the bypass duct than the second bypass immersion lobe, and
wherein the third bypass immersion lobe extends into the bypass duct a different radial distance than the respective distances to which the first bypass immersion lobe and the second bypass immersion lobe extend.

13. The gas turbine engine of claim 12 , wherein said first crown contour line has a different curvature than said second crown contour line.

14. The gas turbine engine of claim 12 , wherein said first bypass immersion lobe and said second bypass immersion lobe are generally semi-elliptically shaped, said first bypass immersion lobe having a first semi-major axis at said trailing edge, said second bypass immersion lobe having a second semi-major

Assignees

• Gen Electric

Inventors

• Dindar Mustafa
• Kunze Vaughn Ray

Classifications

• F05D2300/6033

  Ceramic matrix composites [CMC] · CPC title

• F05D2230/60

  Assembly methods · CPC title

• F02K1/38

  Introducing air inside the jet (F02K1/28 takes precedence) · CPC title

• F02K1/48 Primary

  Corrugated nozzles · CPC title

• F02K1/827

  Sound absorbing structures or liners · CPC title