Gas turbine engine installation

Patent metadata

Field	Value
Publication number	US-2020200094-A1
Application number	US-201916411466-A
Country	US
Kind code	A1
Filing date	May 14, 2019
Priority date	Dec 21, 2018
Publication date	Jun 25, 2020
Grant date	—

Abstract

A gas turbine engine for an aircraft comprises 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 extending from a hub; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine has an engine length and a centre of gravity position measured relative to the fan, and a centre of gravity position ratio of:
the centre of gravity position/the engine length

is in a range from 0.43 to 0.6.

First claim

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 extending from a hub and having a fan tip radius in a range from 110 cm to 150 cm; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein:
the gas turbine engine has an engine length and a centre of gravity position measured relative to the fan, the centre of gravity position in a range between 140 cm and 180 cm, and
wherein a centre of gravity position ratio of:
the centre of gravity position/the engine length

is in a range from 0.43 to 0.6.

2 . The gas turbine engine of claim 1 , wherein the centre of gravity position ratio is in a range from 0.45 to 0.6.

3 . The gas turbine engine of claim 1 , wherein
the centre of gravity position ratio is in a range from 0.47 to 0.49.

4 . The gas turbine engine of claim 1 , wherein the engine length is in the range from 200 cm to 500 cm.

5 . The gas turbine engine of claim 1 , wherein the engine length is in the range from 300 cm to 360 cm.

6 . (canceled)

7 . (canceled)

8 . The gas turbine engine of claim 1 , wherein the engine length is defined as an axial distance between a forward region of the fan and a rearward region of the turbine.

9 . The gas turbine engine of claim 1 , wherein the turbine comprises a lowest pressure turbine stage having row of rotor blades, and the engine length is defined as an axial distance between: an intersection of a leading edge of one of the plurality of fan blades and the hub and a mean radius point of a trailing edge of one of the rotor blades of the lowest pressure turbine stage of the turbine.

10 . The gas turbine engine of claim 9 , wherein the mean radius point is a midpoint between a 0% span position and a 100% span position of the rotor blade.

11 . The gas turbine engine of claim 1 , wherein the turbine is a lowest pressure turbine of a plurality of turbines provided in the core.

12 . The gas turbine engine of claim 1 , wherein the centre of gravity position is defined as an axial distance between an intersection of a leading edge of one of the plurality of fan blades and the hub: and the centre of gravity of the gas turbine engine.

13 . The gas turbine engine of claim 1 , wherein a fan speed to centre of gravity ratio of:
the centre of gravity position ratio×maximum take off rotational fan speed
is in a range from 600 rpm to 1350 rpm.

14 . The gas turbine engine of claim 13 , wherein the fan speed to centre of gravity ratio is in a range from 650 rpm to 1276 rpm.

15 . The gas turbine engine of claim 13 , wherein the fan speed to centre of gravity ratio is in a range from 600 rpm to 1290 rpm.

16 . The gas turbine engine of claim 13 , wherein
the fan speed to centre of gravity ratio is in a range from 925 rpm to 1350 rpm.

17 . The gas turbine engine of claim 1 , wherein the maximum take-off rotational fan speed is in a range between 1450 rpm and 3020 rpm.

18 . The gas turbine engine of claim 1 , wherein
the maximum take-off rotational fan speed is in a range between 1970 rpm and 3020 rpm.

19 . A method of operating an aircraft comprising a gas turbine engine 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 extending from a hub and having a fan tip radius in a range from 110 cm to 150 cm; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein: the gas turbine engine has an engine length and a centre of gravity position defined relative to the fan, the centre of gravity position is in a range between 140 cm and 180 cm, and wherein
the method comprises controlling a pitch of the aircraft such that a centre of gravity position ratio of:
the centre of gravity position/the engine length
is in a range from 0.43 to 0.6, and the method comprises using the engine to provide thrust to the aircraft for take-off, and during take-off of the aircraft a fan speed to centre of gravity ratio of:
the centre of gravity position ratio×maximum take off rotational fan speed

has a maximum value in a range from 600 rpm to 1350 rpm.

20 . The method of claim 19 wherein the method comprises controlling the pitch of the aircraft such that the centre of gravity position ratio is in a range from 650 rpm to 1276 rpm.

21 . The gas turbine engine of claim 1 , wherein the centre of gravity position ratio is in a range from 0.46 to 0.6.

22 . The gas turbine engine of claim 1 , wherein the engine length is in the range from 300 cm to 450 cm.

Assignees

• Rolls Royce Plc

Inventors

• Willmot Michael C
• Stretton Richard G

Classifications

• F04D29/388

  characterised by construction · CPC title

• F02K3/06 Primary

  with front fan · CPC title

• F05D2220/36

  specially adapted for the fan of turbofan engines · CPC title

• F02K1/00

  Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto (rocket nozzles F02K9/97) · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title