Additively manufactured booster splitter with integral heating passageways

Patent metadata

Field	Value
Publication number	US-11624320-B2
Application number	US-202117365855-A
Country	US
Kind code	B2
Filing date	Jul 1, 2021
Priority date	Jan 24, 2018
Publication date	Apr 11, 2023
Grant date	Apr 11, 2023

Abstract

A booster splitter for a gas turbine engine and a method of additively manufacturing the booster splitter are provided. The booster splitter includes an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine, an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path, an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port, and a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.

First claim

What is claimed is:

1. A gas turbine engine defining an axial direction and a radial direction, the gas turbine engine comprising:
an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine;
an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path;
an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port; and
a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.

2. The gas turbine engine of claim 1 , wherein the outer wall defines heat exchange fins within the passageway.

3. The gas turbine engine of claim 1 , further including an inlet conduit providing fluid communication between a fluid supply and the inlet port.

4. The gas turbine engine of claim 3 , wherein the outer wall defines multiple fluid passageways spaced apart along a circumferential direction of the outer wall, each of the multiple fluid passageways being in fluid communication with the fluid supply.

5. The gas turbine engine of claim 1 , wherein the outer wall defines an annular discharge plenum extending circumferentially proximate the forward end of the outer wall, the outer wall further defining a plurality of discharge ports spaced circumferentially proximate the forward end and being in fluid communication with the discharge plenum.

6. The gas turbine engine of claim 5 , wherein discharge ports are positioned within the inlet for discharging fluid flow into the compressor flow path.

7. The gas turbine engine of claim 1 , wherein the gas turbine engine includes a first plurality of airfoils spaced circumferentially within the compressor flow path, each of the first plurality of airfoils defining a heating plenum and bleed air ports.

8. The gas turbine engine of claim 7 , wherein an inlet conduit provides fluid communication between a high pressure compressor and one or more inlet ports defined by the outer wall.

9. The gas turbine engine of claim 1 , wherein the passageway defines a serpentine pattern within the outer wall.

10. The gas turbine engine of claim 1 , wherein the gas turbine engine includes an annular nacelle spaced apart from the outer wall along the radial direction to define a bypass passageway therebetween, and wherein the inner wall and the outer wall split a primary flow of air into a flow of bypass air and a flow of core air.

11. The gas turbine engine of claim 1 , wherein the outer wall, the inner wall, or both are integrally formed as a single monolithic component.

12. A gas turbine engine defining an axial direction and a radial direction, the gas turbine engine comprising:
a splitter positioned at a forward end of a core engine and defining a core inlet, the splitter configured for splitting a primary flow of air between the core inlet and a bypass airflow passageway, the bypass airflow passageway defined between the core engine and an annular nacelle, the splitter including:
an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine;
an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path;
an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port; and
a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.

13. The gas turbine engine of claim 12 , wherein the outer wall defines heat exchange fins within the passageway.

14. The gas turbine engine of claim 12 , further including an inlet conduit providing fluid communication between a fluid supply and the inlet port.

15. The gas turbine engine of claim 12 , wherein the gas turbine engine includes a first plurality of airfoils spaced circumferentially within the compressor section, each of the first plurality of airfoils defining a heating plenum and bleed air ports.

16. A method of manufacturing a gas turbine engine, the method comprising:
depositing a layer of additive material on a bed of an additive manufacturing machine; and
selectively directing energy from an energy source onto the layer of additive material to fuse a portion of the additive material and form at least one of:
an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine;
an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path;
an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port; or
a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.

17. The method of claim 16 , wherein the passageway is surrounded and confined by the outer wall.

18. The method of claim 16 , wherein the annular bulkhead defines a plurality of inlet ports, the inlet ports spaced circumferentially about the bulkhead, each of the inlet ports fluidly coupled to the compressor section through a separate inlet conduit.

19. The method of claim 16 , wherein the method of manufacturing includes binder jetting.

20. The method of claim 16 , wherein the passageway defines a serpentine pattern within the outer wall.

Assignees

• Gen Electric

Inventors

• Kray Nicholas Joseph
• Dorbala Sreekanth Kumar
• Jain Nitesh
• Sharma Prashant

Classifications

• F05D2230/31

  Layer deposition · CPC title

• F01D25/02

  De-icing means for engines having icing phenomena · CPC title

• F05D2230/22

  by sintering · CPC title

• F02C7/047 Primary

  Heating to prevent icing · CPC title

• B64D33/02

  of combustion air intakes · CPC title