Sound attenuation apparatus and method

Patent metadata

Field	Value
Publication number	US-9771868-B2
Application number	US-201514804577-A
Country	US
Kind code	B2
Filing date	Jul 21, 2015
Priority date	Jul 21, 2015
Publication date	Sep 26, 2017
Grant date	Sep 26, 2017

Abstract

An aircraft engine sound attenuation apparatus includes a perforated face member, a backing member, a plurality of connecting members coupling the perforated face member to the backing member to form a plurality of channels spanning from the perforated face member to the backing member, and a bulk absorber disposed in each of the plurality of channels, wherein the plurality of channels are connected to an interior portion of an aircraft engine nacelle component so that the plurality of channels are oriented in a direction substantially normal to a direction of fluid flow pressure drop passing through the aircraft engine.

First claim

What is claimed is:

1. An aircraft engine sound attenuation apparatus comprising:
a perforated face member;
a backing member;
a plurality of connecting members coupling the perforated face member to the backing member to form a plurality of channels spanning from the perforated face member to the backing member; and
a bulk absorber disposed in each of the plurality of channels;
wherein the plurality of channels are connected to an interior portion of an aircraft engine nacelle component so that the plurality of channels are oriented in a direction substantially normal to a direction of fluid flow pressure drop passing through the aircraft engine, and each channel has a width, in the direction substantially normal to the direction of fluid flow pressure drop passing through the aircraft engine, the at least one of prevents or mitigates sound propagation inside the channel substantially parallel to a surface of the channel in the direction substantially normal to the direction of fluid flow pressure drop passing through the aircraft engine.

2. The aircraft engine sound attenuation apparatus of claim 1 , wherein a longitudinal length of each channel extends a full length of the aircraft engine nacelle component.

3. The aircraft engine sound attenuation apparatus of claim 1 , further comprising at least one septum layer disposed between the perforated face member and the backing member where the at least one septum layer divides each channel into at least two portions.

4. The aircraft engine sound attenuation apparatus of claim 1 , wherein each of the plurality of connecting members extends radially from the perforated face member so as to form a wall of a respective channel where the wall has a circumferentially extending sinewave configuration that undulates in a circumferential direction around the perforated face member.

5. The aircraft engine sound attenuation apparatus of claim 1 , wherein each of the plurality of connecting members extends linearly along a longitudinal axis of each channel.

6. The aircraft engine sound attenuation apparatus of claim 1 , wherein the plurality of connecting members forms a truss core that extends along a longitudinal axis of each channel.

7. The aircraft engine sound attenuation apparatus of claim 1 , wherein the perforated face member includes perforations that place the bulk absorber in each channel in fluid communication with the fluid flow.

8. An aircraft engine comprising:
a nacelle having a nacelle frame to which a plurality of nacelle components are attached;
at least one sound attenuating member, each sound attenuating member including
a perforated face member,
a backing member,
a plurality of connecting members coupling the perforated face member to the backing member to form a plurality of channels spanning from the perforated face member to the backing member; and
a bulk absorber disposed in each of the plurality of channels;

wherein the plurality of channels are connected to an interior portion of a respective one of the plurality of nacelle components so that the plurality of channels are oriented in a direction substantially normal to a direction of fluid flow pressure drop passing through the aircraft engine, and each channel has a width, in the direction substantially normal to the direction of fluid flow pressure drop passing through the aircraft engine, that at least one of prevents or mitigates sound propagation inside the channel substantially parallel to a surface of the channel in the direction substantially normal to the direction of fluid flow pressure drop passing through the aircraft engine.

9. The aircraft engine of claim 8 , wherein a longitudinal length of each channel extends a full length of the respective one of the plurality of nacelle components.

10. The aircraft engine of claim 8 , wherein the nacelle frame includes frame members having cavities therein and the bulk absorber is disposed within the cavities, the bulk absorber being in fluid communication with the fluid flow.

11. The aircraft engine of claim 8 , wherein the plurality of nacelle components includes at least one thrust reverser panel and one of the at least one sound attenuating member is coupled to each of the at least one thrust reverser panel.

12. The aircraft engine of claim 8 , wherein the at least one sound attenuating member further includes at least one septum layer disposed between the perforated face member and unperforated backing member where the at least one septum layer divides each channel into at least two portions.

13. The aircraft engine of claim 8 , wherein the perforated face member includes perforations that place the bulk absorber in each channel in fluid communication with the fluid flow.

14. The aircraft engine of claim 8 , wherein each channel of one of the at least one sound attenuating member are in fluid communication with a corresponding adjacent channel of a second one of the at least sound attenuating member.

15. The aircraft engine of claim 8 , wherein at least one of the perforated face member and the backing member includes an access panel that is configured for passage of the bulk absorber through the access panel to provide access to the bulk absorber in each channel.

16. A method for aircraft engine sound attenuation, the method comprising:
flowing fluid through an aircraft engine nacelle;
receiving at least a portion of the fluid in perforations of the aircraft engine nacelle so that the portion of the fluid flows through the perforations into at least one channel of the aircraft engine nacelle; and
attenuating acoustic waves with a bulk absorber disposed within each of the at least one channel where sound propagation inside the channel parallel to a surface of the at least one channel, in a direction of fluid flow pressure drop passing through the aircraft engine, is at least one of prevented or mitigated by a width of the at least one channel, the width being in a direction substantially normal to the direction of fluid flow pressure drop passing through the aircraft engine.

17. The method of claim 16 , further comprising heating the bulk absorber with the portion of the fluid to vaporize liquids contained within the bulk absorber.

18. The method of claim 17 , further comprising vaporizing the liquids contained within the bulk absorber with heat, from an engine core member disposed within the aircraft engine nacelle, conducting through a backing member of the aircraft engine nacelle.

19. The method of claim 16 , further comprising providing access to the bulk absorber through a surface of the aircraft engine nacelle.

20. The aircraft engine sound attenuation apparatus of claim 3 , further comprising a liquid drainage path disposed in one of the at least two portions, the liquid drainage with being configured to evacuate liquids from the aircraft engine sound attenuation apparatus.

21. The aircraft engine sound attenuation apparatus of claim 3 , wherein the bulk absorber comprises a first hulk absorber disposed in a first one of the at least two portions and a second bulk absorber disposed in a second one of the at least two portions, the second bulk absorber being different than the first bulk absorber.

Assignees

• Boeing Co

Inventors

• Nesbitt Eric H
• Lan Justin H
• Bhat Thonse R S
• Rust Charles W

Classifications

• F05D2250/283

  honeycomb · CPC title

• F02K1/827

  Sound absorbing structures or liners · CPC title

• B64D29/02

  associated with wings · CPC title

• F05D2260/962

  by means of "anti-noise" · CPC title

• F02C7/24 Primary

  Heat or noise insulation (air intakes having provisions for noise suppression F02C7/045; turbine exhaust heads, chambers, or the like F01D25/30; silencing nozzles of jet-propulsion plants F02K1/00) · CPC title