Geared turbofan with gearbox seal

Patent metadata

Field	Value
Publication number	US-2016003142-A1
Application number	US-201514709595-A
Country	US
Kind code	A1
Filing date	May 12, 2015
Priority date	Jun 11, 2014
Publication date	Jan 7, 2016
Grant date	—

Abstract

A gas turbine engine comprises a fan, a compressor section, a turbine section, and a gear reduction for driving the fan through the turbine section. A rotating element and at least one bearing compartment includes a bearing for supporting the rotating element, a seal for resisting leakage of lubricant outwardly of the bearing compartment, and for allowing pressurized air to flow from a chamber adjacent the seal into the bearing compartment. The seal has a plurality of sealing members extending radially toward a sealing surface.

First claim

1 . A gas turbine engine comprising:
a fan, a compressor section, a turbine section, and a gear reduction for driving said fan through said turbine section; a rotating element and at least one bearing compartment including a bearing for supporting said rotating element, a seal for resisting leakage of lubricant outwardly of said bearing compartment, and for allowing pressurized air to flow from a chamber adjacent said seal into the bearing compartment; and said seal having a plurality of sealing members extending radially toward a sealing surface.

2 . The gas turbine engine as set forth in claim 1 , wherein said seal is a labyrinth seal having a plurality of knife edges.

3 . The gas turbine engine as set forth in claim 2 , wherein a first radius is defined to a radial extent of said knife edges and a second radius may be defined on a drive shaft associated with said fan drive turbine at a location in a plane defined by a fuel nozzle in a combustor in said gas turbine engine, and a diameter ratio of said first radius to said second radius being less than or equal to about 2.0.

4 . The gas turbine engine as set forth in claim 3 , wherein said diameter radius being less than or equal to about 1.75.

5 . The gas turbine engine as set forth in claim 2 , wherein said bearing compartment is associated with said gear reduction.

6 . The gas turbine engine as set forth in claim 2 , wherein said bearing compartment is associated with said fan.

7 . The gas turbine engine as set forth in claim 2 , wherein said bearing compartment is associated with a compressor rotor.

8 . The gas turbine engine as set forth in claim 2 , wherein said bearing compartment is associated with a turbine rotor in said turbine section.

9 . The gas turbine engine as set forth in claim 1 , wherein said seal is a brush seal.

10 . The gas turbine engine as set forth in claim 9 , wherein said bearing compartment is associated with said gear reduction.

11 . The gas turbine engine as set forth in claim 9 , wherein said bearing compartment is associated with said fan.

12 . The gas turbine engine as set forth in claim 9 , wherein said bearing compartment is associated with a compressor rotor.

13 . The gas turbine engine as set forth in claim 9 , wherein said bearing compartment is associated with a turbine rotor in said turbine section.

14 . The gas turbine engine as set forth in claim 1 , wherein said gear reduction having a gear ratio greater than or equal to about 2.6.

15 . The gas turbine engine as set forth in claim 14 , wherein said fan delivering air into a bypass duct as propulsion air and into said compressor section as core air and a bypass ratio of said bypass air to said core air being greater than or equal to about 6.0.

16 . The gas turbine engine as set forth in claim 5 , wherein said bypass ratio being greater than or equal to about 10.0.

17 . The gas turbine engine as set forth in claim 16 , wherein said bypass air being greater than or equal to about 12.0.

18 . The gas turbine engine as set forth in claim 1 , wherein said fan delivering air into a bypass duct as propulsion air and into said compressor section as core air and a bypass ratio of said bypass air to said core air being greater than or equal to about 6.0.

19 . The gas turbine engine as set forth in claim 18 , wherein said bypass ratio being greater than or equal to about 10.0.

20 . The gas turbine engine as set forth in claim 19 , wherein said bypass air being greater than or equal to about 12.0.

Assignees

• United Technologies Corp

Inventors

• Sheridan William G
• Schwarz Frederick M

Classifications

• F02K3/06

  with front fan · CPC title

• F02C3/107 Primary

  with two or more rotors connected by power transmission · CPC title

• F01D11/02

  by non-contact sealings, e.g. of labyrinth type (for sealing space between rotor blade tips and stator F01D11/08) · CPC title

• F02C7/28

  Arrangement of seals · CPC title

• F05D2240/56

  Brush seals · CPC title