High power epicyclic gearbox and operation thereof

Patent metadata

Field	Value
Publication number	US-11268452-B2
Application number	US-202017066713-A
Country	US
Kind code	B2
Filing date	Oct 9, 2020
Priority date	Dec 5, 2019
Publication date	Mar 8, 2022
Grant date	Mar 8, 2022

Abstract

A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft. A fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the fan axial centreline, the fan-gearbox axial distance being greater than or equal to 0.35 m.

First claim

The invention claimed is:

1. A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan having a fan axial centreline;
a gearbox that is configured to:
receive an input from the core shaft, and
output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and

a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two supporting bearings connected to the fan shaft, and wherein:
a fan-gearbox axial distance is defined as an axial distance between the output of the gearbox and the fan axial centreline, the fan-gearbox axial distance being greater than or equal to 0.35 m;
a system radial bending stiffness is defined as:

1

(

1

a
⁢

⁢
radial
⁢

⁢
bending
⁢

⁢
stiffness

of
⁢

⁢
the
⁢

⁢
fan

shaft
⁢

⁢
mounting
⁢

⁢
structure

)

+

(

1

a
⁢

⁢
radial
⁢

⁢
bending

stiffness
⁢

⁢
of
⁢

⁢
the
⁢

⁢
fan

shaft
⁢

⁢
at
⁢

⁢
the
⁢

⁢
output

of
⁢

⁢
the
⁢

⁢
gearbox

)

;

and a product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure is greater than or equal to 2.7×10 15 (N/m) 2 .

2. The gas turbine engine according to claim 1 , wherein the product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure is in a range from 4.0×10 15 (N/m) 2 to 1.5×10 19 (N/m) 2 .

3. The gas turbine engine according to claim 1 , wherein at least one of the following is satisfied:
a) the system radial bending stiffness is in a range from 3.90×10 6 N/m to 1.40×10 9 N/m;
b) the radial bending stiffness of the fan shaft mounting structure is in a range from 7.00×10 8 N/m to 6.00×10 11 N/m; and
c) the radial bending stiffness of the fan shaft at the output of the gearbox is in a range from 4.00×10 6 N/m to 1.5×10 9 N/m.

4. The gas turbine engine according to claim 1 , wherein:
a fan diameter is in a range from 240 to 280 cm; and
the product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure is in a range from 4.3×10 16 (N/m) 2 to 3.0×10 18 (N/m) 2 .

5. The gas turbine engine according to claim 1 , wherein:
a fan diameter is in a range from 330 to 380 cm; and
the product of the system radial bending stiffness and the rad

Assignees

• Rolls Royce Plc

Inventors

• Spruce Mark

Classifications

• F05D2200/14

  Division · CPC title

• F02C7/36 Primary

  Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) · CPC title

• F16H1/32 Primary

  in which the central axis of the gearing lies inside the periphery of an orbital gear · CPC title

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title