Internally heated engine inlet screen for aircraft engines

Patent metadata

Field	Value
Publication number	US-10450954-B2
Application number	US-201615072662-A
Country	US
Kind code	B2
Filing date	Mar 17, 2016
Priority date	Mar 17, 2016
Publication date	Oct 22, 2019
Grant date	Oct 22, 2019

Abstract

An apparatus for providing foreign object debris protection and anti-icing capabilities to an air intake of an aircraft engine. The apparatus includes a frame having at least a portion configured to conduct fluids. A tube is positioned such that different portions of the frame are fluidly connected. There is a discharge outlet defined within the frame that is configured to discharge fluids from the frame into the air intake. An inlet is defined within the frame and is fluidly connected to a compressor section of the aircraft engine. In this manner, air from the compressor section of the aircraft engine can flow through the inlet, through the tube, and through the discharge outlet, to return to the compressor section of the aircraft engine.

First claim

What is claimed is:

1. An apparatus for providing foreign object debris protection and anti-icing capabilities to an air intake of an aircraft engine, the apparatus comprising;
a frame having at least a portion configured to conduct fluids;
a plurality of tubes through which different portions of the frame are fluidly connected, wherein the plurality of tubes form a mesh assembly supported by the frame, the mesh assembly having a first tier and a second tier with a first portion of the plurality of tubes being in the first tier and a second portion of the plurality of tubes being in the second tier, the first tier and second tier together defining a mesh of the mesh assembly, the first tier stacked in a direction perpendicular to the flow of fluid traveling through the plurality of tubes on the second tier and in thermal communication with the second tier;
a discharge outlet defined within the frame that is configured to discharge fluid into the air intake;
an inlet defined within the frame; and
wherein the inlet is fluidly connected to a compressor section of the aircraft engine such that hot air from the compressor section of the aircraft engine can flow through the inlet, through the plurality of tubes, and through the discharge outlet, to return to the compressor section of the aircraft engine.

2. The apparatus according to claim 1 , wherein the frame is also configured to support a mesh that is not configured to be heated.

3. The apparatus according to claim 2 , wherein the frame is configured to be heated by hot air from the compressor section around its perimeter.

4. The apparatus according to claim 3 , wherein the first portion of the plurality of tubes are arranged in the first tier generally parallel to each other.

5. The apparatus according to claim 4 , wherein the second portion of the plurality of tubes are arranged in the second tier generally parallel to each other, the first portion of the plurality of tubes crossing the second portion of the plurality of tubes to form a mesh.

6. The apparatus according to claim 5 , wherein the mesh is configured to prevent foreign object debris from entering the air intake of the engine.

7. The apparatus according to claim 1 , further including solid wire links positioned in the second tier.

8. A method for providing anti-icing capabilities to the air intake of an aircraft engine wherein the aircraft intake is covered by a mesh assembly supported by a frame and having a plurality of tubes having at least a portion thereof configured to conduct fluids, the mesh assembly having a first tier and a second tier with a first portion of the plurality of tubes being in the first tier and a second portion of the plurality of tubes being in the second tier in thermal communication with the first portion of the plurality of tubes in the first tier, the first tier and second tier together defining a mesh of the mesh assembly with the first tier stacked in a direction perpendicular to the flow of fluid traveling through the plurality of tubes on the second tier, the method comprising the steps of:
conducting heated air from a compressor section of the aircraft engine;
conducting the heated air through the portion of mesh configured to conduct fluids;
conducting heat through the mesh;
elevating the temperature of the mesh; and
returning the air to the compressor section of the aircraft engine.

9. The apparatus according to claim 8 , wherein the frame is configured to be heated by hot air from the compressor section around its perimeter.

10. The apparatus according to claim 9 , wherein the first portion of the plurality of tubes are arranged in the first tier generally parallel to each other and the second portion of the plurality of tubes are arranged in the second tier generally parallel to each other, the first portion of the plurality of tubes crossing the second portion of the plurality of tubes to form the mesh.

11. The apparatus according to claim 10 , wherein the mesh is configured to prevent foreign object debris from entering the air intake of the engine.

12. The apparatus according to claim 8 , further including solid wire links positioned in the second tier.

13. An engine comprising:
a plurality of mesh screens formed of tubes that form at least a portion of a fluidly conductive circuit that includes a compressor section of the engine, the mesh screens, and an air intake portion of the engine, wherein the mesh screens include a first tier and a second tier in thermal communication with the first tier, the first tier stacked in a direction perpendicular to the flow of fluid traveling through the plurality of tubes on the second tier with at least a portion of the tubes being arranged in the first tier and solid wire links being positioned in the second tier.

14. The engine according to claim 13 wherein the plurality of mesh screens are arranged in an array such that each screen is positioned over an associated air intake.

15. The engine according to claim 14 wherein each of the mesh screens are fluidly connected with the compressor section of the engine via a bleed line.

16. The engine according to claim 15 , wherein each of the mesh screens are fluidly connected to a manifold which is fluidly connected via a single bleed line to the compressor section of the engine.

Assignees

• Gen Electric

Inventors

• Grissino Alan Scott
• Manteiga John Alan
• Kerner Jonathan Harry

Classifications

• F04D29/522

  especially adapted for elastic fluid pumps · CPC title

• F05D2220/32

  in gas turbines · CPC title

• B64D2033/022

  comprising bird or foreign object protections · CPC title

• B64D33/02

  of combustion air intakes · CPC title

• B64D2033/0246

  comprising particle separators · CPC title