Geared turbofan engine

Patent metadata

Field	Value
Publication number	US-11199196-B2
Application number	US-201916398870-A
Country	US
Kind code	B2
Filing date	Apr 30, 2019
Priority date	Nov 29, 2018
Publication date	Dec 14, 2021
Grant date	Dec 14, 2021

Abstract

A gas turbine engine for an aircraft, includes: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan assembly located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan assembly so as to drive the fan assembly at a lower rotational speed than the core shaft, wherein the fan assembly has fan blades mounted around a hub, the fan blades having blade tips defining an outer diameter of the fan assembly of from around 220 cm to around 400 cm, the hub having slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, wherein a ratio of a mass of the hub to a total mass of the fan blades is within the range of around 0.45 to around 0.7.

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 assembly located upstream of the engine core; and
a gearbox that receives an input from the core shaft and outputs drive to the fan assembly so as to drive the fan assembly at a lower rotational speed than the core shaft,
wherein the fan assembly comprises a plurality of fan blades mounted around a hub, the fan blades having blade tips defining an outer diameter of the fan assembly of from around 220 cm to around 400 cm, the hub comprising a plurality of slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, wherein a ratio of a mass of the hub to a total mass of the plurality of fan blades is within a range of around 0.45 to around 0.7.

2. The gas turbine engine of claim 1 wherein a hub-to-tip ratio of the fan assembly is between around 0.2 and 0.4.

3. The gas turbine engine of claim 2 wherein the hub-to-tip ratio of the fan assembly is between around 0.2 and 0.3.

4. The gas turbine engine of claim 1 , wherein the ratio of the mass of the hub to the total mass of the plurality of fan blades is within a range of around 0.5 to around 0.65.

5. The gas turbine engine of claim 4 , wherein the ratio of the mass of the hub to the total mass of the plurality of fan blades is within a range of around 0.5 to around 0.6.

6. The gas turbine engine of claim 1 wherein a minimum radial thickness of the rim is within a range of around 0.5% to around 1.1% of the outer fan diameter.

7. The gas turbine engine of claim 6 wherein the minimum radial thickness is no greater than 35 mm.

8. The gas turbine engine of claim 6 wherein an average of the minimum rim thickness along a rotation axis of the fan assembly is within a range of around 0.5 to around 1.1% of the outer fan diameter.

9. The gas turbine engine of claim 1 wherein the outer diameter of the fan assembly is around 280 cm or greater.

10. The gas turbine engine of claim 1 wherein the outer diameter of the fan assembly is around 330 cm or greater.

11. The gas turbine engine of claim 1 wherein a gear ratio of the gearbox is in a range of from 3.2 to 5.

12. The gas turbine engine of claim 1 wherein a gear ratio of the gearbox is in a range of from 3.2 to 4.2.

13. The gas turbine engine according to claim 1 , wherein:
the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft;
the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and
the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

14. 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 assembly located upstream of the engine core; and
a gearbox that receives an input from the core shaft and outputs drive to the fan assembly so as to drive the fan assembly at a lower rotational speed than the core shaft,
wherein the fan assembly comprises a plurality of fan blades mounted around a hub, a hub-to-tip ratio of the fan assembly being between around 0.2 and 0.4, the hub comprising a plurality of slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, wherein a ratio of a mass of the hub to a total mass of the plurality of fan blades is within a range of around 0.45 to around 0.7.

15. The gas turbine engine of claim 14 wherein the fan blades have blade tips defining an outer diameter of the fan assembly of from around 220 cm to around 400 cm.

Assignees

• Rolls Royce Plc

Inventors

• Breen Clive

Classifications

• F04D29/329 Primary

  Details of the hub · 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

• F05D2260/96

  Preventing, counteracting or reducing vibration or noise · CPC title

• B64C27/463

  Blade tips · CPC title

• F01D5/10

  Anti- vibration means {(specially adapted for radial flow machines or engines F01D5/04)} · CPC title