Gas turbine engine with noise attenuating variable area fan nozzle

Patent metadata

Field	Value
Publication number	US-9745918-B2
Application number	US-201514609800-A
Country	US
Kind code	B2
Filing date	Jan 30, 2015
Priority date	Jun 26, 2008
Publication date	Aug 29, 2017
Grant date	Aug 29, 2017

Abstract

A nacelle assembly for a high-bypass gas turbine engine includes a core nacelle defined about an engine centerline axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path. A variable area fan nozzle is in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section. The second fan nacelle section is axially movable relative to the first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes an acoustic system that has an acoustic impedance located on a radially outer surface.

First claim

What is claimed is:

1. A nacelle assembly for a high-bypass gas turbine engine comprising:
a core nacelle defined about an engine centerline axis;
a fan nacelle mounted at least partially around said core nacelle to define a fan bypass flow path; and
a variable area fan nozzle in communication with said fan bypass flow path, said variable area fan nozzle having a first fan nacelle section and a second fan nacelle section, said second fan nacelle section axially movable relative said first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow, said second fan nacelle section including an acoustic system having an acoustic impedance located on a radially outer surface, wherein the acoustic system is located on a fore portion of the radially outer surface, an aft portion of the radially outer surface lacking the acoustic system.

2. The nacelle assembly as recited in claim 1 , wherein said acoustic system is defined at least in part within a leading edge region of said second fan nacelle section.

3. The nacelle assembly as recited in claim 1 , wherein said acoustic system further comprises a forward acoustic system and an aft acoustic system, said forward acoustic system being different than said aft acoustic system.

4. The nacelle assembly as recited in claim 1 , wherein said acoustic system further comprises a forward acoustic system and an aft acoustic system, said forward acoustic system comprises said leading edge of said second fan nacelle section and said aft acoustic system comprises at least a portion of an upper surface of said second fan nacelle section.

5. The nacelle assembly as recited in claim 1 , wherein said acoustic system comprises a perforated inner face sheet and a perforated outer face sheet supported by a structure.

6. The nacelle assembly as recited in claim 1 , wherein said acoustic system comprises a perforated inner face sheet.

7. The nacelle assembly as recited in claim 1 , wherein said acoustic system comprises an outer face sheet.

8. The nacelle assembly as recited in claim 1 , wherein said acoustic system comprises a bulk absorbing material.

9. A nacelle assembly for a high-bypass gas turbine engine comprising:
a core nacelle defined about an engine centerline axis;
a fan nacelle mounted at least partially around said core nacelle to define a fan bypass flow path; and
a variable area fan nozzle in communication with said fan bypass flow path, said variable area fan nozzle having a first fan nacelle section and a second fan nacelle section, said second fan nacelle section axially movable relative said first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow, said second fan nacelle section including an acoustic system having an acoustic impedance located on a radially outer surface;
wherein said acoustic system comprises a forward acoustic system in fluid communication with an aft acoustic system, said forward acoustic system comprises a bulk absorbing material and said aft acoustic system comprises a perforated outer face sheet along at least a portion of an upper surface of said second fan nacelle section.

10. The nacelle assembly as recited in claim 9 , wherein said second fan nacelle section defines a trailing edge of said variable area fan nozzle.

11. The nacelle assembly as recited in claim 9 , wherein said second fan nacelle section is subdivided into a multiple of independently operable sectors, each of said multiple of independently operable sectors axially movable relative to said first fan nacelle section to define an asymmetric fan nozzle exit area.

12. A high-bypass gas turbine engine comprising:
a core engine defined about an axis;
a gear system driven by said core engine;
a turbofan driven by said gear system about said axis;
a core nacelle defined at least partially about said core engine;
a fan nacelle mounted at least partially around said core nacelle to define a fan bypass flow path; and
a variable area fan nozzle in communication with said fan bypass flow path, said variable area fan nozzle having a first fan nacelle section and a second fan nacelle section, said second fan nacelle section axially movable relative said first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow, said second fan nacelle section including an acoustic system located on a leading edge and radially outer surface, wherein the acoustic system is located on a fore portion of the radially outer surface, an aft portion of the radially outer surface lacking the acoustic system.

13. The high-bypass gas turbine engine as recited in claim 12 , wherein said acoustic system comprises a perforated inner face sheet and a perforated outer face sheet supported by a structure.

14. The high-bypass gas turbine engine as recited in claim 12 , wherein said acoustic system comprises a bulk absorbing material.

15. The high-bypass gas turbine engine as recited in claim 12 , wherein said second fan nacelle section defines a trailing edge of said variable area fan nozzle.

16. A method of reducing a total effective perceived noise level of a gas turbine engine with a variable area fan nozzle comprising:
axially moving a second fan nacelle section between a closed position in which said second fan nacelle section is in sequential alignment with a first fan nacelle section in response to a cruise flight condition and an open position in which said second fan nacelle section is aftward of said first fan nacelle section to define an auxiliary port, said second fan nacelle section having an acoustic system located on a radially outer surface which provides an acoustic impedance when said second fan nacelle section is positioned at a non-closed position, wherein the acoustic system is located on a fore portion of the radially outer surface, an aft portion of the radially outer surface lacking the acoustic system.

17. The method as recited in claim 16 , further comprising:
generating said acoustic impedance with at least a portion of an upper surface of said second fan nacelle section.

18. The method as recited in claim 16 , wherein said open position in which said second fan nacelle section is aftward of said first fan nacelle section to define an auxiliary port is in response to a non-cruise flight condition.

Assignees

• United Technologies Corp

Inventors

• Gilson Jonathan
• Cicon Dennis
• Atassi Oliver V

Classifications

• F05D2260/96

  Preventing, counteracting or reducing vibration or noise · CPC title

• F02K1/827

  Sound absorbing structures or liners · CPC title

• F02K1/09 Primary

  by axially moving an external member, e.g. a shroud (F02K1/12 takes precedence) · CPC title

• F02K1/002

  with means to modify the direction of thrust vector (F02K1/54 takes precedence; thrust vectoring of rockets F02K9/80) · CPC title