U.S. patent application number 16/820903 was filed with the patent office on 2021-06-10 for geared gas turbine engine.
This patent application is currently assigned to ROLLS-ROYCE plc. The applicant listed for this patent is ROLLS-ROYCE plc. Invention is credited to Mark SPRUCE.
Application Number | 20210172378 16/820903 |
Document ID | / |
Family ID | 1000005608806 |
Filed Date | 2021-06-10 |
United States Patent
Application |
20210172378 |
Kind Code |
A1 |
SPRUCE; Mark |
June 10, 2021 |
GEARED GAS TURBINE ENGINE
Abstract
A gas turbine engine (10) for an aircraft comprising: an engine
core (11) comprising a turbine (19), a compressor (14), and a core
shaft (26) connecting the turbine to the compressor; a fan (23)
located upstream of the engine core, the fan comprising a plurality
of fan blades; a gearbox (30) that receives an input from the core
shaft (26) and outputs drive to the fan so as to drive the fan at a
lower rotational speed than the core shaft, the gearbox (30) being
an epicyclic gearbox (30) comprising a sun gear (28), a plurality
of planet gears (32), a ring gear (38), and a planet carrier (34)
arranged to have the plurality of planet gears (32) mounted
thereon; and a gearbox support (40) arranged to at least partially
support the gearbox within the engine. The gearbox (30) has a cross
sectional area, the cross sectional area being greater than or
equal to 2.4.times.10.sup.-1 m.sup.2; and a first gearbox support
strength ratio of: the torsional strength of the gearbox support (
40 ) the radial bending stiffness of the gearbox support ( 40 )
.times. the cross sectional area of the gearbox ##EQU00001## is
greater than or equal to 7.0.times.10.sup.-3. A planet gear spacing
angle in radians (.beta.) is defined as 2.pi./N, where N is the
number of planet gears (32). A second gearbox support strength
ratio of: the torsional strength of the gearbox support ( 40 ) the
tilt stiffness of the gearbox support ( 40 ) .times. the planet
gear spacing angle ( .beta. ) ##EQU00002## is greater than or equal
to 1.0.times.10.sup.-1.
Inventors: |
SPRUCE; Mark; (Derby,
GB) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
ROLLS-ROYCE plc |
London |
|
GB |
|
|
Assignee: |
ROLLS-ROYCE plc
London
GB
|
Family ID: |
1000005608806 |
Appl. No.: |
16/820903 |
Filed: |
March 17, 2020 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02C 3/107 20130101;
F02K 3/06 20130101; F02C 7/36 20130101 |
International
Class: |
F02C 7/36 20060101
F02C007/36; F02K 3/06 20060101 F02K003/06 |
Foreign Application Data
Date |
Code |
Application Number |
Dec 5, 2019 |
GB |
1917769.0 |
Claims
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; a gearbox that
is configured to: (i) receive an input from the core shaft and (ii)
output drive to the fan so as to drive the fan at a lower
rotational speed than the core shaft, the gearbox being an
epicyclic gearbox comprising a sun gear, a plurality of planet
gears, a ring gear, and a planet carrier arranged to have the
plurality of planet gears mounted thereon; and a gearbox support
arranged to at least partially support the gearbox within the
engine, wherein: the gearbox has a cross sectional area in a range
from 2.4.times.10.sup.-1m.sup.2 to 1.10 m.sup.2; and a first
gearbox support strength ratio of: a torsional strength of the
gearbox support a radial bending stiffness of the gearbox support
.times. the cross sectional area of the gearbox ##EQU00064## is in
a range from 7.0.times.10.sup.-3 to 2.5.times.10.sup.-1.
2. The gas turbine engine according to claim 1, wherein the first
gearbox support strength ratio is in a range from
1.0.times.10.sup.-2 to 2.5.times.10.sup.-1.
3. The gas turbine engine according to claim 1, wherein the radial
bending stiffness of the gearbox support is in a range from greater
than or equal to 1.0.times.10.sup.7 N/m to 4.0.times.10.sup.8
N/m.
4. The gas turbine engine according to claim 1, wherein the
torsional strength of the gearbox support is in a range from
1.60.times.10.sup.5 Nm to 2.00.times.10.sup.7 Nm.
5. The gas turbine engine according to claim 1, wherein: a planet
gear spacing angle in radians (.beta.) is defined as 2.pi./N, where
N is a number of the planet gears; and a second gearbox support
strength ratio of: the torsional strength of the gearbox support a
tilt stiffness of the gearbox support .times. the planet gear
spacing angle ( .beta. ) ##EQU00065## is in a range from
1.0.times.10.sup.-1 to 3.5.
6. 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; a gearbox that
is configured to: (i) receive an input from the core shaft and (ii)
output drive to the fan so as to drive the fan at a lower
rotational speed than the core shaft, the gearbox being an
epicyclic gearbox comprising a sun gear, a plurality of planet
gears, a ring gear, and a planet carrier arranged to have the
plurality of planet gears mounted thereon; and a gearbox support
arranged to at least partially support the gearbox within the
engine, wherein: a planet gear spacing angle in radians (.beta.) is
defined as 2.pi./N, where N is a number of the planet gears; the
gearbox has a cross sectional area in a range from
2.4.times.10.sup.-1m.sup.2 to 1.10 m.sup.2; and a second gearbox
support strength ratio of: a torsional strength of the gearbox
support a tilt stiffness of the gearbox support .times. the planet
gear spacing angle ( .beta. ) ##EQU00066## is in a range from
1.0.times.10.sup.-1 to 3.5.
7. The gas turbine engine according to claim 6, wherein the second
gearbox support strength ratio is in a range from
1.5.times.10.sup.-1 to 1.7.
8. The gas turbine engine according to claim 6, wherein the tilt
stiffness of the gearbox support is in a range from
1.2.times.10.sup.5 Nm/rad to 2.1.times.10.sup.7 Nm/rad.
9. The gas turbine engine according to claim 6, wherein the
torsional strength of the gearbox support is in a range from
1.60.times.10.sup.5 Nm to 2.00.times.10.sup.7 Nm.
10. The gas turbine engine according to claim 6, wherein: a first
gearbox support strength ratio of: the torsional strength of the
gearbox support a radial bending stiffness of the gearbox support
.times. the cross sectional area of the gearbox ##EQU00067## is in
a range from 7.0.times.10.sup.-3 to 2.5.times.10.sup.-1.
11. The gas turbine engine according to claim 1, wherein the
gearbox is in a star configuration.
12. The gas turbine engine according to claim 6, wherein the planet
gear spacing angle (.beta.) is in a range between
9.0.times.10.sup.-1 rad to 2.1 rad.
13. The gas turbine engine according to claim 1, wherein the cross
sectional area of the gearbox is in a range from
2.6.times.10.sup.-1m.sup.2 to 1.10 m.sup.2.
14. The gas turbine engine according to claim 1, wherein: a
torsional shear stress of the gearbox support, at maximum take-off
conditions, is in a range from 1.40.times.10.sup.8 N/m.sup.2 to
4.90.times.10.sup.8 N/m.sup.2.
15. 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, the second compressor, and the second core
shaft are arranged to rotate at a higher rotational speed than the
first core shaft.
16. The gas turbine engine according to claim 1, wherein at least
one of the following is satisfied: i) the gearbox is configured to
have a gear ratio in a range from 3.2 to 4.5; ii) the gas turbine
engine is configured to have a specific thrust in a range from 70
to 90 NKg.sup.-1; and iii) the gas turbine engine is configured to
have a bypass ratio at cruise conditions in a range from 12.5 to
18.
17. The gas turbine engine according to claim 1, wherein: the fan
has a fan diameter greater than 240 cm and less than or equal to
380 cm.
18. A propulsor for an aircraft, comprising: a fan comprising a
plurality of fan blades; a gearbox; a power unit for driving the
fan via the gearbox, the gearbox being arranged to: (i) receive an
input from the power unit via a core shaft and (ii) output drive to
a fan shaft so as to drive the fan at a lower rotational speed than
the core shaft, the gearbox being an epicyclic gearbox comprising a
sun gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the propulsor, wherein: the gearbox has
a cross sectional area in a range from 2.4.times.10.sup.-1m.sup.2
to 1.10 m.sup.2, and at least one of the following is satisfied: a)
a first gearbox support strength ratio of: a torsional strength of
the gearbox support a radial bending stiffness of the gearbox
support .times. the cross sectional area of the gearbox
##EQU00068## is in a range from 7.0.times.10.sup.-3 to
2.5.times.10.sup.-1; and b) a planet gear spacing angle in radians
(.beta.) is defined as 2.pi./N, where N is a number of the planet
gears (32); and a second gearbox support strength ratio of: a
torsional strength of the gearbox support a tilt stiffness of the
gearbox support .times. the planet gear spacing angle ( .beta. )
##EQU00069## is in a range from 1.0.times.10.sup.-1 to 3.5.
19. A method of operating the gas turbine engine according to claim
1, the method comprising operating the gas turbine engine to
provide propulsion for the aircraft under cruise conditions.
20. The method of claim 19, further comprising driving the gearbox
with an input torque in a range from: i) 10,000 to 50,000 Nm at
cruise conditions; and/or ii) 28,000 to 135,000 Nm at max-take off
conditions.
21. A method of operating the gas turbine engine according to claim
6, comprising operating the gas turbine engine to provide
propulsion for the aircraft under cruise conditions.
22. The method according to claim 21, further comprising driving
the gearbox with an input torque in a range from: i) 10,000 to
50,000 Nm at cruise conditions; and/or ii) 28,000 to 135,000 Nm at
max-take off conditions.
23. The gas turbine engine according to claim 1, wherein the first
gearbox support strength ratio is in a range from
7.0.times.10.sup.-3 to 2.0.times.10.sup.-2.
24. The gas turbine engine according to claim 1, wherein the first
gearbox support strength ratio is in a range from
2.0.times.10.sup.-2 to 2.5.times.10.sup.-1.
25. The gas turbine engine according to claim 1, wherein the radial
bending stiffness of the gearbox support is in a range from
3.0.times.10.sup.7 N/m to 2.0.times.10.sup.8 N/m.
26. The gas turbine engine according to claim 1, wherein the
torsional strength of the gearbox support is in a range from
1.8.times.10.sup.5 Nm to 1.5.times.10.sup.6 Nm.
27. The gas turbine engine according to claim 6, wherein the second
gearbox support strength ratio is in a range from
1.0.times.10.sup.-1 to 2.5.times.10.sup.-1.
28. The gas turbine engine according to claim 6, wherein the second
gearbox support strength ratio is in a range from
2.5.times.10.sup.-1 to 3.5.
29. The gas turbine engine according to claim 6, wherein the tilt
stiffness of the gearbox support is in a range from
3.9.times.10.sup.5 Nm/rad to 9.times.10.sup.6 Nm/rad.
30. The gas turbine engine according to claim 6, wherein the
torsional strength of the gearbox support is in a range from
1.8.times.10.sup.5 Nm to 1.5.times.10.sup.6 Nm.
31. The gas turbine engine according to claim 1, wherein: a
torsional shear stress of the gearbox support, at maximum take-off
conditions, is in a range 2.0.times.10.sup.8 N/m.sup.2 to
3.5.times.10.sup.8 N/m.sup.2.
Description
TECHNICAL FIELD
[0001] The present disclosure relates to gas turbine engines,
specifically gas turbine engines for aircraft. Aspects of the
present disclosure also relate to an aircraft comprising the gas
turbine engine, and a method of operating the gas turbine
engine.
BACKGROUND
[0002] Gas turbine engines for aircraft propulsion have many design
factors that affect the overall efficiency and power output or
thrust. A general aim for a gas turbine engine is to provide low
specific fuel consumption (SFC). To enable a higher thrust at a
high efficiency, a larger diameter fan may be used. In order to
facilitate use of a larger fan size, a gearbox is provided having
an output to a fan shaft via which the fan is driven. The gearbox
receives drive from a core shaft connected to a turbine system of
the engine core. The gearbox allows the fan to operate at a reduced
rotational speed compared to if a direct drive were used.
SUMMARY
[0003] When making an engine having a larger fan diameter however,
the inventor has discovered that simply scaling up components of a
known engine type may not lead to an efficient design. For example,
there may be problems associated with mounting the gearbox and the
fan shaft within the engine. Consideration of the properties of
components used to mount the gearbox and the properties of the fan
shaft are therefore required.
[0004] According to a first aspect there is provided 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; a gearbox that receives an
input from the core shaft and outputs 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, the gearbox being an epicyclic gearbox
comprising a sun gear, a plurality of planet gears, a ring gear,
and a planet carrier arranged to have the plurality of planet gears
mounted thereon; and a gearbox support arranged to at least
partially support the gearbox within the engine, wherein: the
moment of inertia of the fan is greater than or equal to
7.40.times.10.sup.7 kgm.sup.2; and a radial bending stiffness to
moment of inertia ratio of:
the radial bending stiffness of at least one of the fan shaft at
the output of the gearbox and the gearbox support the moment of
intertia of the fan ##EQU00003## [0005] is greater than or equal to
2.5.times.10.sup.-3.
[0006] The radial bending stiffness to moment of inertia ratio may
be greater than or equal to 0.05 Nkg.sup.-1m.sup.-3. The radial
bending stiffness to moment of inertia ratio may be in the range
from 2.5.times.10.sup.-2 Nkg.sup.-1m.sup.-3 to 6.0
Nkg.sup.-1m.sup.3. The radial bending stiffness to moment of
inertia ratio may be in the range from 0.05 Nkg.sup.-1m.sup.-3 to
3.0 Nkg.sup.-1m.sup.3. The radial bending stiffness to moment of
inertia ratio may be in the range from 0.05 Nkg.sup.-1m.sup.-3 to
0.6 Nkg.sup.-1m.sup.3
[0007] A ratio of the radial bending stiffness of the fan shaft at
the output of the gearbox to the moment of inertia of the fan may
be greater than or equal to 2.5.times.10.sup.-2 Nkg.sup.-1m.sup.3,
greater than or equal to 0.05 Nkg.sup.-1m.sup.3, in the range from
2.5.times.10.sup.-2 Nkg.sup.-1m.sup.-3 to 6.0 Nkg.sup.-1m.sup.3, in
the range from 0.05 Nkg.sup.-1m.sup.-3 to 3.0 Nkg.sup.-1m.sup.3, or
in the range from 0.05 Nkg.sup.-1m.sup.3 to 0.6
Nkg.sup.-1m.sup.3.
[0008] A ratio of the radial bending stiffness of the gearbox
support to the moment of inertia of the fan may be greater than or
equal to 3.0.times.10.sup.-2 Nkg.sup.-1m.sup.3, greater than or
equal to 0.06 Nkg.sup.-1m.sup.3, in the range from
3.0.times.10.sup.-2 Nkg.sup.-1m.sup.-3 to 4.0 Nkg.sup.-1m.sup.3, in
the range from 0.06 Nkg.sup.-1m.sup.-3 to 2.0 Nkg.sup.-1m.sup.3, or
in the range from 0.06 Nkg.sup.-1M.sup.3 to 0.48
Nkg.sup.-1m.sup.-3.
[0009] The radial bending stiffness of the fan shaft at the output
of the gearbox may be greater than or equal to 4.00.times.10.sup.6
N/m. The radial bending stiffness of the fan shaft at the output of
the gearbox may be greater than or equal to 3.7.times.10.sup.7 N/m.
The radial bending stiffness of the fan shaft at the output of the
gearbox may be in the range from 4.00.times.10.sup.6 N/m to
1.50.times.10.sup.9 N/m. The radial bending stiffness of the fan
shaft at the output of the gearbox may be in the range from
3.7.times.10.sup.7 N/m to 1.0.times.10.sup.9 N/m.
[0010] The radial bending stiffness of the gearbox support may be
greater than or equal to 1.0.times.10.sup.7 N/m. The radial bending
stiffness of the gearbox support may be greater than or equal to
2.0.times.10.sup.7 N/m. The radial bending stiffness of the gearbox
support may be greater than or equal to 3.0.times.10.sup.7 N/m. The
radial bending stiffness of the gearbox support may be in the range
from 1.0.times.10.sup.7 N/m to 4.0.times.10.sup.8 N/m. The radial
bending stiffness of the gearbox support may be in the range from
2.0.times.10.sup.7 N/m to 3.0.times.10.sup.8 N/m. The radial
bending stiffness of the gearbox support may be in the range from
3.0.times.10.sup.7 N/m to 2.0.times.10.sup.8 N/m.
[0011] The diameter of the fan may be in the range from 240 cm to
280 cm. In such an embodiment, the radial bending stiffness to
moment of inertia ratio may be greater than or equal to 0.05
Nkg.sup.-1m.sup.-3 or in the range from 0.05 Nkg.sup.-1m.sup.-3 to
4.0 Nkg.sup.-1m.sup.3.
[0012] Alternatively, the diameter of the fan may be in the range
from 330 cm to 380 cm. In such an embodiment, the radial bending
stiffness to moment of inertia ratio may be greater than or equal
to 0.025 Nkg.sup.-1m.sup.-3 or in the range from 0.025
Nkg.sup.-1m.sup.-3 to 2.0 Nkg.sup.-1m.sup.-3.
[0013] A product (e.g. a radial bending stiffness moment of inertia
product) of:
(the radial bending stiffness of at least one of the fan shaft at
the output of the gearbox and the gearbox support).times.(the
moment of inertia of the fan)
may be greater than or equal to 2.0.times.10.sup.14 Nkgm, greater
than or equal to 4.0.times.10.sup.14 Nkgm, greater than or equal to
2.0.times.10.sup.15 Nkgm, in the range from 2.0.times.10.sup.14
Nkgm to 1.4.times.10.sup.18 Nkgm, in the range from
4.0.times.10.sup.14 Nkgm to 7.0.times.10.sup.17 Nkgm, or in the
range from 2.0.times.10.sup.15 Nkgm to 7.0.times.10.sup.17
Nkgm.
[0014] A tilt stiffness to moment of inertia ratio of:
the tilt stiffness of at least one of the fan shaft at the output
of the gearbox and the gearbox support the moment of intertia of
the fan ##EQU00004##
may be greater than or equal to 4.0.times.10.sup.-4 Nrad.sup.-1
kg.sup.-1m.sup.-1. The tilt stiffness to moment of inertia ratio
may be greater than or equal to 1.0.times.10.sup.-3 Nrad.sup.-1
kg.sup.-1m.sup.-1. The tilt stiffness to moment of inertia ratio
may be in the range from 4.0.times.10.sup.-4 Nrad.sup.-1
kg.sup.-1m.sup.-1 to 2.7.times.10.sup.-1 Nrad.sup.-1
kg.sup.-1m.sup.-1. The tilt stiffness to moment of inertia ratio
may be in the range from 1.0.times.10.sup.-3 Nrad.sup.-1
kg.sup.-1m.sup.-1 to 0.1 Nrad.sup.-1 kg.sup.-1m.sup.-1. The tilt
stiffness to moment of inertia ratio may be in the range from
1.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
1.5.times.10.sup.-2Nrad.sup.-1 kg.sup.-1m.sup.-1.
[0015] According to a second aspect there is provided 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; a gearbox that receives an
input from the core shaft and outputs 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, the gearbox being an epicyclic gearbox
comprising a sun gear, a plurality of planet gears, a ring gear,
and a planet carrier arranged to have the plurality of planet gears
mounted thereon; and a gearbox support arranged to at least
partially support the gearbox within the engine, wherein: the
moment of inertia of the fan is greater than or equal to
7.40.times.10.sup.7 kgm.sup.2; and a tilt stiffness to moment of
inertia ratio of:
the tilt stiffness of at least one of the fan shaft at the output
of the gearbox and the gearbox support the moment of intertia of
the fan ##EQU00005## [0016] is greater than or equal to
4.0.times.10.sup.-4 Nrad.sup.-1kg.sup.-1m.sup.-1.
[0017] The tilt stiffness to moment of inertia ratio may be greater
than or equal to 1.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1.
The tilt stiffness to moment of inertia ratio may be in the range
from 4.0.times.10 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
2.7.times.10.sup.-1 Nrad.sup.-1 kg.sup.-1m.sup.-1. The tilt
stiffness to moment of inertia ratio may be in the range from
1.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1 to 0.1
Nrad.sup.-1 kg.sup.-1m.sup.-1. The tilt stiffness to moment of
inertia ratio may be in the range from 1.0.times.10.sup.-3
Nrad.sup.-1 kg.sup.-1m.sup.-1 to 1.5.times.10.sup.-2 Nrad.sup.-1
kg.sup.-1m.sup.-1.
[0018] A ratio of the tilt stiffness of the fan shaft at the output
of the gearbox to the moment of inertia of the fan may be greater
than or equal to 4.0.times.10.sup.-4 Nrad.sup.-1 kg.sup.-1m.sup.-1,
greater than or equal to 1.0.times.10.sup.-3 Nrad.sup.-1
kg.sup.-1m.sup.-1, in the range from 4.0.times.10.sup.-4
Nrad.sup.-1 kg.sup.-1m.sup.-1 to 2.7.times.10.sup.-1 Nrad.sup.-1
kg.sup.-1m.sup.-1, in the range from 1.0.times.10.sup.-3
Nrad.sup.-1 kg.sup.-1m.sup.-1 to 0.1 Nrad.sup.-1 kg.sup.-1m.sup.-1,
or in the range from 1.0.times.10.sup.-3 Nrad.sup.-1
kg.sup.-1m.sup.-1 to 1.5.times.10.sup.-2Nrad.sup.-1
kg.sup.-1m.sup.-1.
[0019] A ratio of the tilt stiffness of the gearbox support to the
moment of inertia of the fan may be greater than or equal to
1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1, greater than or
equal to 2.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1, in the
range from 1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
7.0.times.10.sup.-2 Nrad.sup.-1 kg.sup.-1m.sup.-1, in the range
from 2.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
3.0.times.10.sup.-2 Nrad.sup.-1 kg.sup.-1m.sup.-1, or in the range
from 2.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
7.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1.
[0020] The tilt stiffness of the fan shaft at the output of the
gearbox may be greater than or equal to 7.00.times.10.sup.4 Nm/rad.
The tilt stiffness of the fan shaft at the output of the gearbox
may be greater than or equal to 9.5.times.10.sup.5 Nm/rad. The tilt
stiffness of the fan shaft at the output of the gearbox may be in
the range from 7.00.times.10.sup.4 Nm/rad to 7.00.times.10.sup.7
Nm/rad. The tilt stiffness of the fan shaft at the output of the
gearbox may be in the range from 9.5.times.10.sup.5 Nm/rad to
3.5.times.10.sup.7 Nm/rad.
[0021] The tilt stiffness of the gearbox support may be greater
than or equal to 1.2.times.10.sup.5 Nm/rad. The tilt stiffness of
the gearbox support may be greater than or equal to
2.4.times.10.sup.5 Nm/rad. The tilt stiffness of the gearbox
support may be greater than or equal to 3.9.times.10.sup.5 Nm/rad.
The tilt stiffness of the gearbox support may be in the range from
1.2.times.10.sup.5 Nm/rad to 2.1.times.10.sup.7 Nm/rad. The tilt
stiffness of the gearbox support may be in the range from
2.4.times.10.sup.5 Nm/rad to 1.6.times.10.sup.7 Nm/rad. The tilt
stiffness of the gearbox support may be in the range from
3.9.times.10.sup.5 Nm/rad to 9.0.times.10.sup.6 Nm/rad.
[0022] A radial bending stiffness to moment of inertia ratio
of:
the radial bending stiffness of at least one of the fan shaft at
the output of the gearbox and the gearbox support the moment of
intertia of the fan ##EQU00006##
may be greater than or equal to 2.5.times.10.sup.-2
Nkg.sup.-1m.sup.3. The radial bending stiffness to moment of
inertia ratio may be greater than or equal to 0.05
Nkg.sup.-1m.sup.3. The radial bending stiffness to moment of
inertia ratio may be in the range from 2.5.times.10.sup.-2
Nkg.sup.-1m.sup.-3 to 6.0 Nkg.sup.-1m.sup.3. The radial bending
stiffness to moment of inertia ratio may be in the range from 0.05
Nkg.sup.-1M.sup.3 to 3.0 Nkg.sup.-1m.sup.3. The radial bending
stiffness to moment of inertia ratio may be in the range from 0.05
Nkg.sup.-1M.sup.3 to 0.6 Nkg.sup.-1m.sup.3.
[0023] The diameter of the fan may be in the range from 240 cm to
28 cm. In such an example, the tilt stiffness to moment of inertia
ratio greater may be than or equal to 1.0.times.10.sup.3
Nrad.sup.-1 kg.sup.-1m.sup.-1 or in the range from
1.0.times.10.sup.3 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
1.45.times.10.sup.-2 Nrad.sup.-1 kg.sup.-1m.sup.-1.
[0024] Alternatively, the diameter of the fan may be in the range
from 330 cm to 380 cm. In such an embodiment the tilt stiffness to
moment of inertia ratio may be greater than or equal to
4.0.times.10.sup.-4 Nrad.sup.-1 kg.sup.-1m.sup.-1 or in the range
from 4.0.times.10.sup.-4 Nrad.sup.-1kg.sup.-1m.sup.-1 to
3.0.times.10.sup.-2 Nrad.sup.-1 kg.sup.-1m.sup.-1.
[0025] A product (e.g. the tilt stiffness moment of inertia
product) of:
(the tilt stiffness of at least one of the fan shaft at the output
of the gearbox and the gearbox support).times.(the moment of
inertia of the fan)
may be greater than or equal to 3.0.times.10.sup.12
Nm.sup.3rad.sup.-1kg, greater than or equal to 6.0.times.10.sup.12
Nm.sup.3rad.sup.-1kg, greater than or equal to 2.5.times.10.sup.13
Nm.sup.3rad.sup.-1kg, in the range from 3.0.times.10.sup.12
Nm.sup.3rad.sup.-1kg to 6.0.times.10.sup.16 Nm.sup.3rad.sup.-1kg,
in the range from 6.0.times.10.sup.12 Nm.sup.3rad.sup.-1kg to
3.0.times.10.sup.16 Nm.sup.3rad.sup.-1kg, or in the range from
2.5.times.10.sup.13 Nm.sup.3rad.sup.-1kg to 3.0.times.10.sup.16
Nm.sup.3rad.sup.-1kg.
[0026] One or more of the following features may apply to either or
both of the first and second aspects above:
[0027] The moment of inertia of the fan may be greater than or
equal to 8.3.times.10.sup.7 kgm.sup.2. The moment of inertia of the
fan may be in the range from 7.40.times.10.sup.7 kgm.sup.2 to
9.00.times.10.sup.8 kgm.sup.2. The moment of inertia of the fan may
be in the range from 8.3.times.10.sup.7 kgm.sup.2 to
6.5.times.10.sup.8 kgm.sup.2.
[0028] The fan blades may be formed from a metallic material. The
fan blades may be formed from titanium or an aluminium-lithium
alloy with a titanium leading edge.
[0029] The fan blades may be formed at least partly from an organic
matrix composite. The fan blades may have a carbon composite body
with a metallic leading edge.
[0030] The fan shaft may connect the output of the gearbox to the
fan. A gearbox output position may be defined as the point of
connection between the fan shaft and the gearbox. The fan may have
an axial centerline. A fan-gearbox axial distance may be defined as
the axial distance between the axial position of the gearbox output
position and the fan axial centerline. The fan-gearbox axial
distance multiplied by the moment of inertia of the fan may be
greater than or equal to 1.9.times.10.sup.7 kgm.sup.3, greater than
or equal to 2.9.times.10.sup.7 kgm.sup.3, in the range from
1.9.times.10.sup.7 kgm.sup.3 to 6.2.times.10.sup.8 kgm.sup.3, or in
the range from 2.9.times.10.sup.7 kgm.sup.3 to 3.9.times.10.sup.8
kgm.sup.3.
[0031] The fan shaft may be defined as the torque transfer
component extending from the output of the gearbox to the input to
the fan. The fan shaft may comprise at least part of a gearbox
output shaft and at least part of a fan input shaft.
[0032] The gearbox may be in a star configuration and the output of
the gearbox may be a gearbox output position defined as the point
of connection between the ring gear and the fan shaft.
Alternatively, the gearbox may be in a planetary configuration and
the output of the gearbox may be a gearbox output position at the
interface between the fan shaft and the planet carrier.
[0033] According to a third aspect there is provided a propulsor
for an aircraft, comprising: a fan comprising a plurality of fan
blades; a gearbox; a power unit for driving the fan via the
gearbox, wherein the gearbox is arranged to receive an input from
the power unit via a 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; a gearbox support arranged to
at least partially support the gearbox within the propulsor, and
wherein: the moment of inertia of the fan is greater than or equal
to 7.40.times.10.sup.7 kgm.sup.2; and a radial bending stiffness to
moment of inertia ratio of:
the radial bending stiffness of at least one of the fan shaft at
the output of the gearbox and the gearbox support the moment of
intertia of the fan ##EQU00007## [0034] is greater than or equal to
2.5.times.10.sup.-2 Nkg.sup.-1m.sup.-3.
[0035] The propulsor of the third aspect may have some or all of
the features described above with respect to the gas turbine engine
of the first aspect, and may be a gas turbine engine in some
embodiments.
[0036] According to a fourth aspect there is provided a propulsor
for an aircraft, comprising: a fan comprising a plurality of fan
blades; a gearbox; a power unit for driving the fan via the
gearbox, wherein the gearbox is arranged to receive an input from
the power unit via a 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 gearbox support
arranged to at least partially support the gearbox within the
propulsor, and wherein: the moment of inertia of the fan is greater
than or equal to 7.40.times.10.sup.7 kgm.sup.2; and a tilt
stiffness to moment of inertia ratio of:
the tilt stiffness of at least one of the fan shaft at the output
of the gearbox and the gearbox support the moment of intertia of
the fan ##EQU00008## [0037] is greater than or equal to
4.0.times.10.sup.-4.
[0038] The propulsor of the fourth aspect may have some or all of
the features described above with respect to the gas turbine engine
of the second aspect, and may be a gas turbine engine in some
embodiments.
[0039] The third and fourth aspects may be combined. According to a
fifth aspect there is provided a propulsor for an aircraft,
comprising: a fan comprising a plurality of fan blades; a gearbox;
a power unit for driving the fan via the gearbox, wherein the
gearbox is arranged to receive an input from the power unit via a
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 gearbox support arranged to at least partially
support the gearbox within the propulsor, and wherein the moment of
inertia of the fan is greater than or equal to 7.40.times.10.sup.7
kgm.sup.2; and wherein:
[0040] a) a radial bending stiffness to moment of inertia ratio
of:
the radial bending stiffness of at least one of the fan shaft at
the output of the gearbox and the gearbox support the moment of
intertia of the fan ##EQU00009##
[0041] is greater than or equal to 2.5.times.10.sup.-2
Nkg.sup.-1m.sup.-3; and/or [0042] b) a tilt stiffness to moment of
inertia ratio of:
[0042] the tilt stiffness of at least one of the fan shaft at the
output of the gearbox and the gearbox support the moment of
intertia of the fan ##EQU00010##
is greater than or equal to 4.0.times.10.sup.-4.
[0043] The propulsor of the fifth aspect may have some or all of
the features described above with respect to the gas turbine engine
of the first or second aspect, and may be a gas turbine engine in
some embodiments.
[0044] According to a sixth aspect there is provided a method of
operating a gas turbine engine for an aircraft, the 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; a gearbox that receives an
input from the core shaft and outputs 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, the gearbox being an epicyclic gearbox
comprising a sun gear, a plurality of planet gears, a ring gear,
and a planet carrier arranged to have the plurality of planet gears
mounted thereon; and a gearbox support arranged to at least
partially support the gearbox within the engine, wherein: the
moment of inertia of the fan is greater than or equal to
7.40.times.10.sup.7 kgm.sup.2; and a radial bending stiffness to
moment of inertia ratio of:
the radial bending stiffness of at least one of the fan shaft at
the output of the gearbox and the gearbox support the moment of
intertia of the fan ##EQU00011## [0045] is greater than or equal to
2.5.times.10.sup.-2 Nkg.sup.-1m.sup.-3, the method comprises
operating the gas turbine engine to provide propulsion for the
aircraft under cruise conditions.
[0046] The method of the sixth aspect may be a method of operating
the gas turbine engine or the propulsor of the first aspect or
third aspect respectively. Any of the features, ratios and
parameters introduced above in connection with the first or third
aspect may also therefore apply to the sixth aspect
[0047] According to a seventh aspect there is provided a method of
operating a gas turbine engine for an aircraft, the 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; a gearbox that receives an
input from the core shaft and outputs 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, the gearbox being an epicyclic gearbox
comprising a sun gear, a plurality of planet gears, a ring gear,
and a planet carrier arranged to have the plurality of planet gears
mounted thereon; and a gearbox support arranged to at least
partially support the gearbox within the engine, [0048] wherein the
moment of inertia of the fan is greater than or equal to
7.40.times.10.sup.7 kgm.sup.2; and a tilt stiffness to moment of
inertia ratio of:
[0048] the tilt stiffness of at least one of the fan shaft at the
output of the gearbox and the gearbox support the moment of
intertia of the fan ##EQU00012## [0049] is greater than or equal to
4.0.times.10.sup.-4 Nrad.sup.-1kg.sup.-1m.sup.-1, the method
comprising operating the gas turbine engine to provide propulsion
for the aircraft under cruise conditions.
[0050] The method of the seventh aspect may be a method of
operating the gas turbine engine or the propulsor of the second
aspect or fourth aspect respectively. Any of the features, ratios
and parameters introduced above in connection with the second
aspect or fourth aspect may also therefore apply to the seventh
aspect.
[0051] The sixth and seventh aspects may be combined. According to
an eighth aspect there is provided a method of operating a gas
turbine engine for an aircraft, the 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; a
gearbox that receives an input from the core shaft and outputs
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, the
gearbox being an epicyclic gearbox comprising a sun gear, a
plurality of planet gears, a ring gear, and a planet carrier
arranged to have the plurality of planet gears mounted thereon; and
a gearbox support arranged to at least partially support the
gearbox within the engine, wherein the moment of inertia of the fan
is greater than or equal to 7.40.times.10.sup.7 kgm.sup.2, and
wherein: a radial bending stiffness to moment of inertia ratio
of:
the radial bending stiffness of at least one of the fan shaft at
the output of the gearbox and the gearbox support the moment of
intertia of the fan ##EQU00013## [0052] is greater than or equal to
2.5.times.10.sup.-2 Nkg.sup.-1m.sup.-3; and/or [0053] b) a tilt
stiffness to moment of inertia ratio of:
[0053] the tilt stiffness of at least one of the fan shaft at the
output of the gearbox and the gearbox support the moment of inertia
of the fan ##EQU00014## [0054] is greater than or equal to
4.0.times.10.sup.-4 Nrad.sup.-1kg.sup.-1m.sup.-1, the method
comprising operating the gas turbine engine to provide propulsion
for the aircraft under cruise conditions.
[0055] The method of the eighth aspect may be a method of operating
the gas turbine engine or the propulsor of the first, second or
fifth aspect. Any of the features, ratios and parameters introduced
above in connection with the first, second or fifth aspect may also
therefore apply to the eighth aspect.
[0056] The inventor has discovered that either the radial
bending/tilt stiffness of one or both of the fan shaft at the
gearbox output and the gearbox support need to be low enough for a
given moment of inertia of the fan to isolate the gearbox from the
transmission of loads from the fan. Rotation of the fan will cause
a gyroscopic effect meaning that the fan shaft will tend to
maintain a steady direction of its axis of rotation. During
maneuvering of the aircraft to which the gas turbine engine is
mounted the orientation of the axis of rotation of the fan shaft
will however change. The inventor has found that gyroscopic effects
will result in loads being transmitted to the gearbox. The inventor
has discovered that by using a fan shaft and/or gearbox support
having a radial bending/tilt stiffness so that the radial
bending/stilt stiffness to moment of inertia ratio defined above is
within the specified range the problem of load from the fan being
transmitted to the gearbox can be addressed. The inventor has found
that if the radial bending/tilt stiffness of the fan shaft or the
gearbox support shaft is increased so that the ratio is outside of
the defined range there is inadequate isolation of the gearbox. The
inventor has also found that if the fan shaft/gearbox support
stiffness were to be decreased so that the ratio is outside of the
range above excessive vibration at low modal frequencies would
result.
[0057] In other aspects, value ranges for the product of the
components of the radial bending stiffness to moment of inertia
ratio and the tilt stiffness to moment of inertia ratio may be
specified instead of, or as well as, value ranges for the
ratios.
[0058] According to one such aspect, the first aspect introduced
above may be formulated as an aspect providing 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; a gearbox that receives an
input from the core shaft and outputs 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, the gearbox being an epicyclic gearbox
comprising a sun gear, a plurality of planet gears, a ring gear,
and a planet carrier arranged to have the plurality of planet gears
mounted thereon; and a gearbox support arranged to at least
partially support the gearbox within the engine, wherein: the
moment of inertia of the fan is greater than or equal to
7.40.times.10.sup.7 kgm.sup.2; and a product of: (the radial
bending stiffness of at least one of the fan shaft at the output of
the gearbox and the gearbox support).times.(the moment of inertia
of the fan) is greater than or equal to 2.0.times.10.sup.14 Nkgm,
greater than or equal to 4.0.times.10.sup.14 Nkgm, greater than or
equal to 2.0.times.10.sup.15 Nkgm, in the range from
2.0.times.10.sup.14 Nkgm to 1.4.times.10.sup.18 Nkgm, in the range
from 4.0.times.10.sup.14 Nkgm to 7.0.times.10.sup.17 Nkgm, or in
the range from 2.0.times.10.sup.15 Nkgm to 7.0.times.10.sup.17
Nkgm.
[0059] According to another such aspect, the second aspect
introduced above may be formulated as an aspect providing 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; a gearbox that
receives an input from the core shaft and outputs 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, the gearbox being an
epicyclic gearbox comprising a sun gear, a plurality of planet
gears, a ring gear, and a planet carrier arranged to have the
plurality of planet gears mounted thereon; and a gearbox support
arranged to at least partially support the gearbox within the
engine, wherein: the moment of inertia of the fan is greater than
or equal to 7.40.times.10.sup.7 kgm.sup.2; and a product of:
(the tilt stiffness of at least one of the fan shaft at the output
of the gearbox and the gearbox support).times.(the moment of
inertia of the fan)
is greater than or equal to 3.0.times.10.sup.12
Nm.sup.3rad.sup.-1kg, greater than or equal to 6.0.times.10.sup.12
Nm.sup.3rad.sup.-1kg, greater than or equal to 2.5.times.10.sup.13
Nm.sup.3rad.sup.-1kg, in the range from 3.0.times.10.sup.12
Nm.sup.3rad.sup.-1kg to 6.0.times.10.sup.16 Nm.sup.3rad.sup.-1kg,
in the range from 6.0.times.10.sup.12 Nm.sup.3rad.sup.-1kg to
3.0.times.10.sup.16 Nm.sup.3rad.sup.-1kg, or in the range from
2.5.times.10.sup.13 Nm.sup.3rad.sup.-1kg to 3.0.times.10.sup.16
Nm.sup.3rad.sup.-1kg.
[0060] The skilled person will appreciate that method and propulsor
aspects may be formulated accordingly.
[0061] According to a ninth aspect there is provided 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; 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
gearbox being an epicyclic gearbox comprising a sun gear, a
plurality of planet gears, a ring gear, and a planet carrier
arranged to have the plurality of planet gears mounted thereon; and
a gearbox support arranged to at least partially support the
gearbox within the engine, wherein: the gearbox has a cross
sectional area, the cross sectional area being greater than or
equal to 2.4.times.10.sup.-1m.sup.2; and a first gearbox support
strength ratio of:
the torsional strength of the gearbox support the radial bending
stiffness of the gearbox support .times. the cross sectional area
of the gearbox ##EQU00015## [0062] is greater than or equal to
7.0.times.10.sup.-3.
[0063] The first gearbox support strength ratio may be greater than
or equal to 1.0.times.10.sup.-2. The first gearbox support strength
ratio may be greater than or equal to 7.0.times.10.sup.-3. The
first gearbox support strength ratio may be greater than or equal
to 2.0.times.10.sup.-2. The first gearbox support strength ratio
may be in the range from 7.0.times.10.sup.-3 to
2.5.times.10.sup.-1. The first gearbox support strength ratio may
be in the range from 1.0.times.10.sup.-2 to 1.0.times.10.sup.-1.
The first gearbox support strength ratio may be in the range from
7.0.times.10.sup.-3 to 2.0.times.10.sup.-2. The first gearbox
support strength ratio may be in the range from 2.0.times.10.sup.-2
to 2.5.times.10.sup.-1.
[0064] The radial bending stiffness of the gearbox support may be
greater than or equal to 1.0.times.10.sup.7 N/m. The radial bending
stiffness of the gearbox support may be greater than or equal to
2.0.times.10.sup.7 N/m. The radial bending stiffness of the gearbox
support may be greater than or equal to 3.0.times.10.sup.7 N/m. The
radial bending stiffness of the gearbox support may be in the range
from 1.0.times.10.sup.7 N/m to 4.0.times.10.sup.8 N/m. The radial
bending stiffness of the gearbox support may be in the range from
2.0.times.10.sup.7 N/m to 3.0.times.10.sup.8 N/m. The radial
bending stiffness of the gearbox support may be in the range from
3.0.times.10.sup.7 N/m to 2.0.times.10.sup.8 N/m.
[0065] The torsional strength of the gearbox support may be greater
than or equal to 1.60.times.10.sup.5 Nm. The torsional strength of
the gearbox support may be greater than or equal to
1.8.times.10.sup.5 Nm. The torsional strength of the gearbox
support may be in the range from 1.60.times.10.sup.5 Nm to
2.00.times.10.sup.7 Nm. The torsional strength of the gearbox
support may be in the range from 1.8.times.10.sup.5 Nm to
1.5.times.10.sup.6 Nm.
[0066] A planet gear spacing angle in radians may be defined as
2.pi./N, where N is the number of planet gears.
[0067] A second gearbox support strength ratio of:
the torsional strength of the gearbox support the tilt stiffness of
the gearbox support .times. the planet gear spacing angle
##EQU00016##
may be greater than or equal to 1.0.times.10.sup.-1. The second
gearbox support strength ratio may be greater than or equal to
1.5.times.10.sup.-1. The second gearbox support strength ratio may
be greater than or equal to 1.0.times.10.sup.-1. The second gearbox
support strength ratio may be greater than or equal to
2.5.times.10.sup.-1. The second gearbox support strength ratio may
be in the range from 1.0.times.10.sup.-1 to 3.5. The second gearbox
support strength ratio may be in the range from 1.5.times.10.sup.-1
to 1.7. The second gearbox support strength ratio may be in the
range from 1.0.times.10.sup.-1 to 2.5.times.10.sup.-1. The second
gearbox support strength ratio may be in the range from
2.5.times.10.sup.-1 to 3.5.
[0068] According to a tenth aspect there is provided 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; 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
gearbox being an epicyclic gearbox comprising a sun gear, a
plurality of planet gears, a ring gear, and a planet carrier
arranged to have the plurality of planet gears mounted thereon; and
a gearbox support arranged to at least partially support the
gearbox within the engine, wherein: a planet gear spacing angle in
radians is defined as 2.pi./N, where Nis the number of planet
gears; the gearbox has a cross sectional area, the cross sectional
area being greater than or equal to 2.4.times.10.sup.0.1m.sup.2;
and a second gearbox support strength ratio of:
the torsional strength of the gearbox support the tilt stiffness of
the gearbox support .times. the planet gear spacing angle
##EQU00017## [0069] is greater than or equal to
1.0.times.10.sup.-1.
[0070] The second gearbox support strength ratio may be greater
than or equal to 1.5.times.10.sup.-1. The second gearbox support
strength ratio may be greater than or equal to 2.5.times.10.sup.-1.
The second gearbox support strength ratio may be in the range from
1.0.times.10.sup.-1 to 3.5. The second gearbox support strength
ratio may be in the range from 1.5.times.10.sup.-1 to 1.7. The
second gearbox support strength ratio may be in the range from
1.0.times.10.sup.-1 to 2.5.times.10.sup.-1. The second gearbox
support strength ratio may be in the range from 2.5.times.10.sup.-1
to 3.5.
[0071] The tilt stiffness of the gearbox support may be greater
than or equal to 1.2.times.10.sup.5 Nm/rad. The tilt stiffness of
the gearbox support may be greater than or equal to
2.4.times.10.sup.5 Nm/rad. The tilt stiffness of the gearbox
support may be greater than or equal to 3.9.times.10.sup.5 Nm/rad.
The tilt stiffness of the gearbox support may be in the range from
1.2.times.10.sup.5 Nm/rad to 2.1.times.10.sup.7 Nm/rad. The tilt
stiffness of the gearbox support may be in the range from
2.4.times.10.sup.5 Nm/rad to 1.6.times.10.sup.7 Nm/rad. The tilt
stiffness of the gearbox support may be in the range from
3.9.times.10.sup.5 Nm/rad to 9.0.times.10.sup.6 Nm/rad.
[0072] The torsional strength of the gearbox support may be greater
than or equal to 1.60.times.10.sup.5 Nm. The torsional strength of
the gearbox support may be greater than or equal to
1.8.times.10.sup.5 Nm. The torsional strength of the gearbox
support may be in the range from 1.60.times.10.sup.5 Nm to
2.00.times.10.sup.7 Nm. The torsional strength of the gearbox
support may be in the range from 1.8.times.10.sup.5 Nm to
1.5.times.10.sup.6 Nm.
[0073] A first gearbox support strength ratio of:
the torsional strength of the gearbox support the radial bending
stiffness of the gearbox support .times. the cross sectional area
of the gearbox ##EQU00018##
may be greater than or equal to 7.0.times.10.sup.-3. The first
gearbox support strength ratio may be greater than or equal to
1.0.times.10.sup.-2. The first gearbox support strength ratio may
be greater than or equal to 2.0.times.10.sup.-2. The first gearbox
support strength ratio may be in the range from 7.0.times.10.sup.-3
to 2.5.times.10.sup.-1. The first gearbox support strength ratio
may be in the range from 1.0.times.10.sup.-2 to
1.0.times.10.sup.-1. The first gearbox support strength ratio may
be in the range from 7.0.times.10.sup.-3 to 2.0.times.10.sup.-2.
The first gearbox support strength ratio may be in the range from
2.0.times.10.sup.-2 to 2.5.times.10.sup.-1.
[0074] One or more of the following features may apply to either or
both of the ninth and tenth aspects above: The gearbox may be in a
star configuration.
[0075] The planet gear spacing angle may be greater than or equal
to 9.0.times.10.sup.-1 rad. The planet gear spacing angle may be in
the range between 9.0.times.10.sup.-1 rad to 2.1 rad.
[0076] The cross sectional area of the gearbox may be greater than
or equal to 2.6.times.10.sup.-1m.sup.2. The cross sectional area of
the gearbox may be in the range from 2.4.times.10.sup.-1m.sup.2 to
1.10 m.sup.2. The cross sectional area of the gearbox may be in the
range from 2.6.times.10.sup.-1m.sup.2 to
9.0.times.10.sup.-1m.sup.2. The cross sectional area (CSA) of the
gearbox may be defined as the area of the pitch circle of the ring
gear.
[0077] A torsional shear stress of the gearbox support, at maximum
take-off conditions, may be less than or equal to
4.90.times.10.sup.8 N/m.sup.2, less than or equal to
3.5.times.10.sup.8 N/m.sup.2, in the range from 1.40.times.10.sup.8
N/m.sup.2 to 4.90.times.10.sup.8 N/m.sup.2, or in the range from
2.0.times.10.sup.8 N/m.sup.2 to 3.5.times.10.sup.8 N/m.sup.2.
[0078] A first gearbox support shear stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions the radial bending stiffness of the gearbox support
##EQU00019##
may be less than or equal to 4.9.times.10.sup.-1m.sup.-1, less than
or equal to 20 m.sup.-1, in the range from
3.5.times.10.sup.-1m.sup.-1 to 4.9.times.10.sup.1 m.sup.-1, or in
the range from 0.70 m.sup.-1 to 20 m.sup.-1.
[0079] A second gearbox support shear stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions the tilt stiffness of the gearbox support
##EQU00020##
may be less than or equal to 4.1.times.10.sup.3rad/m.sup.3, less
than or equal to 1.4.times.10.sup.3 rad/m.sup.3, in the range from
6.6 rad/m.sup.3 to 4.1.times.10.sup.3 rad/m.sup.3, or in the range
from 1.25.times.10.sup.1 rad/m.sup.3 to 1.4.times.10.sup.3
rad/m.sup.3.
[0080] According to an eleventh aspect there is provided a
propulsor for an aircraft, comprising: a fan comprising a plurality
of fan blades; a gearbox; a power unit for driving the fan via the
gearbox, wherein the gearbox is arranged to receive an input from
the power unit via a core shaft and output drive to a fan shaft so
as to drive the fan at a lower rotational speed than the core
shaft, the gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the propulsor, wherein: the gearbox has
a cross sectional area, the cross sectional area being greater than
or equal to 2.4.times.10.sup.-1 m.sup.2; a first gearbox support
strength ratio of:
the torsional strength of the gearbox support the radial bending
stiffness of the gearbox support .times. the cross sectional area
of the gearbox ##EQU00021## [0081] is greater than or equal to
7.0.times.10.sup.-3.
[0082] The propulsor of the eleventh aspect may have some or all of
the features described above with respect to the gas turbine engine
of the ninth aspect, and may be a gas turbine engine in some
embodiments.
[0083] According to an twelfth aspect there is provided a propulsor
for an aircraft, comprising: a fan comprising a plurality of fan
blades; a gearbox; a power unit for driving the fan via the
gearbox, wherein the gearbox is arranged to receive an input from
the power unit via a core shaft and output drive to a fan shaft so
as to drive the fan at a lower rotational speed than the core
shaft, the gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the propulsor, wherein: the gearbox has
a cross sectional area, the cross sectional area being greater than
or equal to 2.4.times.10.sup.-1 m.sup.2; and a planet gear spacing
angle in radians is defined as 2.pi./N, where N is the number of
planet gears; and a second gearbox support strength ratio of:
the torsional strength of the gearbox support the tilt stiffness of
the gearbox support .times. the planet gear spacing angle
##EQU00022## [0084] is greater than or equal to
1.0.times.10.sup.-1.
[0085] The propulsor of the twelfth aspect may have some or all of
the features described above with respect to the gas turbine engine
of the tenth aspect, and may be a gas turbine engine in some
embodiments.
[0086] The eleventh and twelfth aspects may be combined. According
to a thirteenth aspect there is provided a propulsor for an
aircraft, comprising: a fan comprising a plurality of fan blades; a
gearbox; a power unit for driving the fan via the gearbox, wherein
the gearbox is arranged to receive an input from the power unit via
a core shaft and output drive to a fan shaft so as to drive the fan
at a lower rotational speed than the core shaft, the gearbox being
an epicyclic gearbox comprising a sun gear, a plurality of planet
gears, a ring gear, and a planet carrier arranged to have the
plurality of planet gears mounted thereon; and a gearbox support
arranged to at least partially support the gearbox within the
propulsor, wherein the gearbox has a cross sectional area, the
cross sectional area being greater than or equal to
2.4.times.10.sup.-1m.sup.2, and wherein: a first gearbox support
strength ratio of:
the torsional strength of the gearbox support the radial bending
stiffness of the gearbox support .times. the cross sectional area
of the gearbox ##EQU00023## [0087] is greater than or equal to
7.0.times.10.sup.-3; and/or [0088] b) a planet gear spacing angle
in radians is defined as 2.pi./N, where N is the number of planet
gears; and a second gearbox support strength ratio of:
[0088] the torsional strength of the gearbox support the tilt
stiffness of the gearbox support .times. the planet gear spacing
angle ##EQU00024## [0089] is greater than or equal to
1.0.times.10.sup.-1.
[0090] The propulsor of the thirteenth aspect may have some or all
of the features described above with respect to the gas turbine
engine of the ninth aspect and tenth aspects, and may be a gas
turbine engine in some embodiments.
[0091] According to a fourteenth aspect there is provided a method
of operating a gas turbine engine for an aircraft, the 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; 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
gearbox being an epicyclic gearbox comprising a sun gear, a
plurality of planet gears, a ring gear, and a planet carrier
arranged to have the plurality of planet gears mounted thereon; and
a gearbox support arranged to at least partially support the
gearbox within the engine, wherein: the gearbox has a cross
sectional area, the cross sectional area being greater than or
equal to 2.4.times.10.sup.0.1m.sup.2; and a first gearbox support
strength ratio of:
the torsional strength of the gearbox support the radial bending
stiffness of the gearbox support .times. the cross sectional area
of the gearbox ##EQU00025## [0092] is greater than or equal to
7.0.times.10.sup.-3, the method comprising operating the gas
turbine engine to provide propulsion for the aircraft under cruise
conditions.
[0093] The method of the fourteenth aspect may be a method of
operating the gas turbine engine or the propulsor of the ninth
aspect or eleventh aspect respectively. Any of the features, ratios
and parameters introduced above in connection with the ninth aspect
or eleventh aspect may also therefore apply to the fourteenth
aspect.
[0094] According to a fifteenth aspect there is provided a method
of operating a gas turbine engine for an aircraft, the 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; 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
gearbox being an epicyclic gearbox comprising a sun gear, a
plurality of planet gears, a ring gear, and a planet carrier
arranged to have the plurality of planet gears mounted thereon; and
a gearbox support arranged to at least partially support the
gearbox within the engine, wherein: the gearbox has a cross
sectional area, the cross sectional area being greater than or
equal to 2.4.times.10.sup.0.1m.sup.2; and a planet gear spacing
angle in radians is defined as 2.pi./N, where N is the number of
planet gears; and a second gearbox support strength ratio of:
the torsional strength of the gearbox support the tilt stiffness of
the gearbox support .times. the planet gear spacing angle
##EQU00026## [0095] is greater than or equal to
1.0.times.10.sup.-1, the method comprising operating the gas
turbine engine to provide propulsion for the aircraft under cruise
conditions.
[0096] The method of the fifteenth aspect may be a method of
operating the gas turbine engine or the propulsor of the tenth
aspect or twelfth aspect respectively. Any of the features, ratios
and parameters introduced above in connection with the tenth aspect
or twelfth aspect may also therefore apply to the fifteenth
aspect.
[0097] The fourteenth and fifteenth aspects may be combined.
According to a sixteenth aspect there is provided a method of
operating a gas turbine engine for an aircraft, the 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; 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
gearbox being an epicyclic gearbox comprising a sun gear, a
plurality of planet gears, a ring gear, and a planet carrier
arranged to have the plurality of planet gears mounted thereon; and
a gearbox support arranged to at least partially support the
gearbox within the engine, wherein the gearbox has a cross
sectional area, the cross sectional area being greater than or
equal to 2.4.times.10.sup.0.1m.sup.2, and [0098] a) a first gearbox
support strength ratio of:
[0098] the torsional strength of the gearbox support the radial
bending stiffness of the gearbox support .times. the cross
sectional area of the gearbox ##EQU00027## [0099] is greater than
or equal to 7.0.times.10.sup.-3; and/or [0100] b) a planet gear
spacing angle in radians is defined as 2.pi./N, where N is the
number of planet gears; and
[0101] a second gearbox support strength ratio of:
the torsional strength of the gearbox support the tilt stiffness of
the gearbox support .times. the planet gear spacing angle
##EQU00028## [0102] is greater than or equal to
1.0.times.10.sup.-1, the method comprising operating the gas
turbine engine to provide propulsion for the aircraft under cruise
conditions.
[0103] The method of the sixteenth aspect may be a method of
operating the gas turbine engine or the propulsor of the ninth,
tenth or thirteenth aspect. Any of the features, ratios and
parameters introduced above in connection with the ninth, tenth or
thirteenth aspect may also therefore apply to the sixteenth
aspect.
[0104] The inventor has discovered that by designing the gearbox
support so that the ratio of its torsional strength to its radial
bending stiffness and cross sectional area (i.e. the first gearbox
support strength ratio) is within the specified range sufficient
strength of the support is provided so that the engine is reliable
with sufficient stiffness provided to minimise misalignment of the
gears in the gearbox and avoid vibration. The inventor has found
that a similar consideration applies to the ratio of the torsional
strength to the tilt stiffness and planet gear spacing angle (i.e.
the second gearbox support strength ratio).
[0105] According to a seventeenth aspect there is provided 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; 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 gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the engine, wherein: a first gearbox
support shear stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions the radial bending stiffness of the gearbox support
##EQU00029## [0106] is less than or equal to
4.9.times.10.sup.1m.sup.-1.
[0107] The first gearbox support shear stress ratio may be less
than or equal to 20 m.sup.-1. The first gearbox support shear
stress ratio may be in the range from 3.5.times.10.sup.-1m.sup.-1
to 4.9.times.10.sup.1m.sup.-1. The first gearbox support shear
stress ratio may be in the range from 0.70 m.sup.-1 to 20
m.sup.-1.
[0108] The radial bending stiffness of the gearbox support may be
greater than or equal to 1.0.times.10.sup.7 N/m. The radial bending
stiffness of the gearbox support may be greater than or equal to
2.0.times.10.sup.7 N/m. The radial bending stiffness of the gearbox
support may be greater than or equal to 3.0.times.10.sup.7 N/m. The
radial bending stiffness of the gearbox support may be in the range
from 1.0.times.10.sup.7 N/m to 4.0.times.10.sup.8 N/m. The radial
bending stiffness of the gearbox support may be in the range from
2.0.times.10.sup.7 N/m to 3.times.10.sup.8 N/m. The radial bending
stiffness of the gearbox support may be in the range from
3.0.times.10.sup.7 N/m to 2.0.times.10.sup.8 N/m.
[0109] The torsional shear stress of the gearbox support, at
maximum take-off conditions, may be less than or equal to
4.90.times.10.sup.8 N/m.sup.2. The torsional shear stress of the
gearbox support, at maximum take-off conditions, may be less than
or equal to 3.5.times.10.sup.8 N/m.sup.2. The torsional shear
stress of the gearbox support, at maximum take-off conditions, may
be in the range from 1.40.times.10.sup.8 N/m.sup.2 to
4.90.times.10.sup.8 N/m.sup.2. The torsional shear stress of the
gearbox support, at maximum take-off conditions, may be in the
range from 2.0.times.10.sup.8 N/m.sup.2 to 3.5.times.10.sup.8
N/m.sup.2.
[0110] The diameter of the fan may be in the range from 240 cm to
280 cm. In such an embodiment, the first gearbox support shear
stress ratio may be less than or equal to 35 m.sup.-1, or in the
range from 0.70 m.sup.-1 to 35 m.sup.-1.
[0111] Alternatively, the diameter of the fan may be in the range
from 330 cm to 380 cm. In such an embodiment the first gearbox
support shear stress ratio may be less than or equal to 12
m.sup.-1, or in the range from 0.50 m.sup.-1 to 12 m.sup.-1.
[0112] A second gearbox support shear stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions the tilt stiffness of the gearbox support
##EQU00030##
may be less than or equal to 4.1.times.10.sup.3rad/m.sup.3. The
second gearbox support shear stress ratio may be less than or equal
to 1.4.times.10.sup.3 rad/m.sup.3. The second gearbox support shear
stress ratio may be in the range from 6.6 rad/m.sup.3 to
4.1.times.10.sup.3 rad/m.sup.3. The second gearbox support shear
stress ratio may be in the range from 1.25.times.10.sup.1
rad/m.sup.3 to 1.4.times.10.sup.3 rad/m.sup.3.
[0113] According to an eighteenth aspect there is provided 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; 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 gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox arranged to at least partially support the
gearbox within the engine, and: a second gearbox support shear
stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions tilt stiffness of the gearbox support ##EQU00031##
[0114] is less than or equal to 4.1.times.10.sup.3 rad/m.sup.3.
[0115] The second gearbox support shear stress ratio may be less
than or equal to 1.4.times.10.sup.3 rad/m.sup.3. The second gearbox
support shear stress ratio may be in the range from 6.6 rad/m.sup.3
to 4.1.times.10.sup.3rad/m.sup.3. The second gearbox support shear
stress ratio may be in the range from 1.25.times.10.sup.1
rad/m.sup.3 to 1.4.times.10.sup.3 rad/m.sup.3.
[0116] The tilt stiffness of the gearbox support may be greater
than or equal to 1.2.times.10.sup.5 Nm/rad. The tilt stiffness of
the gearbox support may be greater than or equal to
2.4.times.10.sup.5 Nm/rad. The tilt stiffness of the gearbox
support may be greater than or equal to 3.9.times.10.sup.5 Nm/rad.
The tilt stiffness of the gearbox support may be in the range from
1.2.times.10.sup.5 Nm/rad to 2.1.times.10.sup.7 Nm/rad. The tilt
stiffness of the gearbox support may be in the range from
2.4.times.10.sup.5 Nm/rad to 1.6.times.10.sup.7 Nm/rad. The tilt
stiffness of the gearbox support may be in the range from
3.9.times.10.sup.5 Nm/rad to 9.0.times.10.sup.6 Nm/rad.
[0117] The torsional shear stress of the gearbox support, at
maximum take-off conditions, may be less than or equal to
4.90.times.10.sup.8 N/m.sup.2. The torsional shear stress of the
gearbox support, at maximum take-off conditions, may be less than
or equal to 3.5.times.10.sup.8 N/m.sup.2. The torsional shear
stress of the gearbox support, at maximum take-off conditions, may
be in the range from 1.40.times.10.sup.8 N/m.sup.2 to
4.90.times.10.sup.8 N/m.sup.2. The torsional shear stress of the
gearbox support, at maximum take-off conditions, may be in the
range from 2.0.times.10.sup.8 N/m.sup.2 to 3.5.times.10.sup.8
N/m.sup.2.
[0118] The diameter of the fan may be in the range from 240 cm to
280 cm. In such an embodiment, the second gearbox support shear
stress ratio may be less than or equal to 2.9.times.10.sup.3
rad/m.sup.3, or in the range from 2.9.times.10.sup.1 rad/m.sup.3 to
2.9.times.10.sup.3 rad/m.sup.3.
[0119] Alternatively, the diameter of the fan may be in the range
from 330 cm to 380 cm. In such an embodiment, the second gearbox
support shear stress ratio may be less than or equal to
7.0.times.10.sup.2rad/m.sup.3 or in the range from
1.0.times.10.sup.1 rad/m.sup.3 to
7.0.times.10.sup.2rad/m.sup.3.
[0120] A first gearbox support shear stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions the radial bending stiffness of the gearbox support
##EQU00032##
may be less than or equal to 4.9.times.10.sup.1m.sup.-1. The first
gearbox support shear stress ratio may be less than or equal to 20
m.sup.-1. The first gearbox support shear stress ratio may be in
the range from 3.5.times.10.sup.-1m.sup.-1 to
4.9.times.10.sup.1m.sup.-1. The first gearbox support shear stress
ratio may be in the range from 0.70 m.sup.-1 to 20 m.sup.-1.
[0121] One or more of the following features may apply to either or
both of the seventeenth and eighteenth aspects above:
[0122] The torsional strength of the gearbox support may be greater
than or equal to 1.60.times.10.sup.5 Nm. The torsional strength of
the gearbox support may be greater than or equal to
1.8.times.10.sup.5 Nm. The torsional strength of the gearbox
support may be in the range from 1.60.times.10.sup.5 Nm to
2.00.times.10.sup.7 Nm. The torsional strength of the gearbox
support may be in the range from 1.8.times.10.sup.5 Nm to
1.5.times.10.sup.6 Nm.
[0123] The cross sectional area of the gearbox may be greater than
or equal to 2.4.times.10.sup.0.1m.sup.2. The cross sectional area
of the gearbox may be greater than or equal to
2.6.times.10.sup.0.1m.sup.2. The cross sectional area of the
gearbox may be in the range from 2.4.times.10.sup.0.1m.sup.2 to
1.10 m.sup.2. The cross sectional area of the gearbox may be in the
range from 2.6.times.10.sup.-1m.sup.2 to
9.0.times.10.sup.0.1m.sup.2.
[0124] A planet gear spacing angle in radians may be defined as
2.pi./N, where N is the number of planet gears. The planet gear
spacing angle may be greater than or equal to 9.0.times.10.sup.-1
rad. The planet gear spacing angle may be in the range between
9.0.times.10.sup.-1 rad to 2.1 rad.
[0125] A first gearbox support strength ratio of:
the torsional strength of the gearbox support the radial bending
stiffness of the gearbox support .times. the cross sectional area
of the gearbox ##EQU00033##
may be greater than or equal to 7.0.times.10.sup.-3. The first
gearbox support strength ratio may be greater than or equal to
1.0.times.10.sup.-2. The first gearbox support strength ratio may
be greater than or equal to 2.0.times.10.sup.-2. The first gearbox
support strength ratio may be in the range from 7.0.times.10.sup.-3
to 2.5.times.10.sup.-1. The first gearbox support strength ratio
may be in the range from 1.0.times.10.sup.-2 to
1.0.times.10.sup.-1. The first gearbox support strength ratio may
be in the range from 7.0.times.10.sup.-3 to 2.0.times.10.sup.-2.
The first gearbox support strength ratio may be in the range from
2.0.times.10.sup.-2 to 2.5.times.10.sup.-1.
[0126] A second gearbox support strength ratio of:
the torsional strength of the gearbox support the tilt stiffness of
the gearbox support .times. the planet gear spacing angle
##EQU00034##
may be greater than or equal to 1.0.times.10.sup.-1. The second
gearbox support strength ratio may be greater than or equal to
1.5.times.10.sup.-1. The second gearbox support strength ratio may
be greater than or equal to 2.5.times.10.sup.-1. The second gearbox
support strength ratio may be in the range from 1.0.times.10.sup.-1
to 3.5. The second gearbox support strength ratio may be in the
range from 1.5.times.10.sup.-1 to 1.7. The second gearbox support
strength ratio may be in the range from 1.0.times.10.sup.-1 to
2.5.times.10.sup.-1. The second gearbox support strength ratio may
be in the range from 2.5.times.10.sup.-1 to 3.5.
[0127] The torque transmitted through the gearbox support, at
maximum take-off conditions, may be greater than or equal to
6.00.times.10.sup.4 Nm, greater than or equal to 7.2.times.10.sup.4
Nm, in the range from 6.00.times.10.sup.4 Nm to 5.00.times.10.sup.5
Nm, or in the range from 7.2.times.10.sup.4 Nm to
4.2.times.10.sup.5 Nm.
[0128] The gearbox may be in a planetary configuration. In such an
embodiment, the torque transmitted through the gearbox support, at
maximum take-off conditions, may be greater than or equal to
6.00.times.10.sup.4 Nm, greater than or equal to 7.2.times.10.sup.4
Nm, in the range from 6.00.times.10.sup.4 Nm to 3.00.times.10.sup.5
Nm, or in the range from 7.2.times.10.sup.4 Nm to
2.6.times.10.sup.5 Nm.
[0129] Alternatively, the gearbox may be in a star configuration.
In such an embodiment the torque transmitted through the gearbox
support, at maximum take-off conditions, may be greater than or
equal to 1.10.times.10.sup.5 Nm, greater than or equal to
1.3.times.10.sup.5 Nm, in the range from 1.10.times.10.sup.5 Nm to
5.00.times.10.sup.5 Nm, or in the range from 1.3.times.10.sup.5 Nm
to 4.2.times.10.sup.5 Nm.
[0130] The maximum take-off conditions may be as defined anywhere
herein, for example may be defined as operating at maximum take-off
thrust for the engine at ISA sea level pressure and
temperature+15.degree. C. with a fan inlet velocity of between 0.25
and 0.27 Mn, and optionally at 0.25 Mn.
[0131] According to a nineteenth aspect there is provided a
propulsor for an aircraft, comprising: a fan comprising a plurality
of fan blades; a gearbox; a power unit for driving the fan via the
gearbox, wherein the gearbox is arranged to receive an input from
the power unit via a core shaft and output drive to a fan shaft so
as to drive the fan at a lower rotational speed than the core
shaft, the gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the propulsor, wherein: a first gearbox
support shear stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions the radial bending stiffness of the gearbox support
##EQU00035## [0132] is less than or equal to
4.9.times.10.sup.1m.sup.-1.
[0133] The propulsor of the nineteenth aspect may have some or all
of the features described above with respect to the gas turbine
engine of the seventeenth aspect, and may be a gas turbine engine
in some embodiments.
[0134] According to a twentieth aspect there is provided a
propulsor for an aircraft, comprising: a fan comprising a plurality
of fan blades; a gearbox; a power unit for driving the fan via the
gearbox, wherein the gearbox is arranged to receive an input from
the power unit via a core shaft and output drive to a fan shaft so
as to drive the fan at a lower rotational speed than the core
shaft, the gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the propulsor, wherein: a second gearbox
support shear stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions the tilt stiffness of the gearbox support
##EQU00036## [0135] is less than or equal to 4.1.times.10.sup.3
rad/m.sup.3.
[0136] The propulsor of the twentieth aspect may have some or all
of the features described above with respect to the gas turbine
engine of the eighteenth aspect, and may be a gas turbine engine in
some embodiments.
[0137] The nineteenth and twentieth aspects may be combined.
According to a twenty-first aspect there is provided a propulsor
for an aircraft, comprising: a fan comprising a plurality of fan
blades; a gearbox; a power unit for driving the fan via the
gearbox, wherein the gearbox is arranged to receive an input from
the power unit via a core shaft and output drive to a fan shaft so
as to drive the fan at a lower rotational speed than the core
shaft, the gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the propulsor, wherein:
[0138] a) a first gearbox support shear stress ratio of:
the torsional shear stress of the gearbox support at maximum take
off conditions the radial bending stiffness of the gearbox support
##EQU00037## [0139] is less than or equal to
4.9.times.10.sup.1m.sup.-1; and/or [0140] b) a second gearbox
support shear stress ratio of:
[0140] the torsional shear stress of the gearbox support at maximum
take off conditions the tilt stiffness of the gearbox support
##EQU00038## [0141] is less than or equal to 4.1.times.10.sup.3
rad/m.sup.3.
[0142] The propulsor of the twenty-first aspect may have some or
all of the features described above with respect to the gas turbine
engine of the seventeenth aspect and eighteenth aspects, and may be
a gas turbine engine in some embodiments.
[0143] According to a twenty-second aspect there is provided a
method of operating a gas turbine engine, the 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; 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 gearbox being an
epicyclic gearbox comprising a sun gear, a plurality of planet
gears, a ring gear, and a planet carrier arranged to have the
plurality of planet gears mounted thereon; and a gearbox support
arranged to at least partially support the gearbox within the
engine, and wherein the method comprises: operating the gas turbine
engine such that, at maximum take-off conditions, a first gearbox
support shear stress ratio of:
the torsional shear strength of the gearbox support radial bending
stiffness of the gearbox support ##EQU00039## [0144] is less than
or equal to 4.9.times.10.sup.1m.sup.-1.
[0145] The method of the twenty-second aspect may be a method of
operating the gas turbine engine or the propulsor of the
seventeenth aspect or the nineteenth aspect respectively. Any of
the features, ratios and parameters introduced above in connection
with the seventeenth aspect or nineteenth aspect may also therefore
apply to the twenty-second aspect.
[0146] According to a twenty-third aspect there is provided a
method of operating a gas turbine engine, the 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; 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 gearbox being an
epicyclic gearbox comprising a sun gear, a plurality of planet
gears, a ring gear, and a planet carrier arranged to have the
plurality of planet gears mounted thereon; and a gearbox support
arranged to at least partially support the gearbox within the
engine, and wherein the method comprises: operating the gas turbine
engine such that, at maximum take-off conditions, a second gearbox
support shear stress ratio of:
the torsional strength of the gearbox support the tilt stiffness of
the gearbox support ##EQU00040## [0147] is less than or equal to
4.1.times.10.sup.3 rad/m.sup.3.
[0148] The method of the twenty-third aspect may be a method of
operating the gas turbine engine or the propulsor of the eighteenth
aspect or the twentieth aspect respectively. Any of the features,
ratios and parameters introduced above in connection with the
eighteenth aspect or twentieth aspect may also therefore apply to
the twenty-second aspect.
[0149] In another aspect, the twenty-second and twenty-third
aspects may be combined. In such an aspect, there is provided a
method of operating a gas turbine engine, the 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; 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 gearbox being an
epicyclic gearbox comprising a sun gear, a plurality of planet
gears, a ring gear, and a planet carrier arranged to have the
plurality of planet gears mounted thereon; and a gearbox support
arranged to at least partially support the gearbox within the
engine, and wherein the method comprises: operating the gas turbine
engine such that, at maximum take-off conditions:
a) a first gearbox support shear stress ratio of:
the torsional shear strength of the gearbox support radial bending
stiffness of the gearbox support ##EQU00041## [0150] is less than
or equal to 4.9.times.10.sup.1 m'; and/or b) a second gearbox
support shear stress ratio of:
[0150] the torsional shear strength of the gearbox support the tilt
stiffness of the gearbox support ##EQU00042## [0151] is less than
or equal to 4.1.times.10.sup.3 rad/m.sup.3.
[0152] The method of the previous aspect may be a method of
operating the gas turbine engine or the propulsor of the
seventeenth, eighteenth or twenty-first aspect. Any of the
features, ratios and parameters introduced above in connection with
the seventeenth, eighteenth or twenty-first aspect may also
therefore apply to this aspect.
[0153] The inventor has discovered that by designing the gearbox
support so that the ratio of its torsional shear stress at max
take-off to its stiffness (radial bending or tilt stiffness) is
within the specified range sufficient margin of strength of the
support is provided at the highest load point in the engine
operating cycle while the stiffness is suitable to ensure a
suitable load share factor.
[0154] According to a twenty-fourth aspect there is provided 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; 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 gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the engine, and wherein: [0155] a flight
cycle ratio of:
[0155] the torsional shear stress of the gearbox support at maximum
take off ( MTO ) conditions the torsional shear stress of the
gearbox support at cruise conditions ##EQU00043## [0156] is less
than or equal to 3.20.
[0157] The flight cycle ratio may be less than or equal to 2.95.
The flight cycle ratio may be less than or equal to 2.9. The flight
cycle ratio may be less than or equal to 2.90. The flight cycle
ratio may be less than or equal to 2.85. The flight cycle ratio may
be less than or equal to 2.75. The flight cycle ratio may be in the
range from 2.10 to 3.20. The flight cycle ratio may be in the range
from 2.3 to 2.9.
[0158] The torsional shear stress of the gearbox support at maximum
take-off conditions may be less than or equal to
4.90.times.10.sup.8 N/m.sup.2. The torsional shear stress of the
gearbox support at maximum take-off conditions may be less than or
equal to 2.0.times.10.sup.8 N/m.sup.2. The torsional shear stress
of the gearbox support at maximum take-off conditions may be in the
range from 1.40.times.10.sup.8 N/m.sup.2 to 4.9.times.10.sup.8
N/m.sup.2. The torsional shear stress of the gearbox support at
maximum take-off conditions may be in the range from
2.0.times.10.sup.8 N/m.sup.2 to 3.5.times.10.sup.8 N/m.sup.2.
[0159] The torsional shear stress of the gearbox support at cruise
conditions may be greater than or equal to 7.00.times.10.sup.7
N/m.sup.2. The torsional shear stress of the gearbox support at
cruise conditions may be greater than or equal to
8.2.times.10.sup.7 N/m.sup.2. The torsional shear stress of the
gearbox support at cruise conditions may be in the range from
7.00.times.10.sup.7 N/m.sup.2 to 1.90.times.10.sup.8 N/m.sup.2. The
torsional shear stress of the gearbox support at cruise conditions
may be in the range from 8.2.times.10.sup.7 N/m.sup.2 to
1.5.times.10.sup.8 N/m.sup.2.
[0160] The diameter of the fan may be in the range from 240 cm to
280 cm. In such an embodiment, the flight cycle ratio may be less
than or equal to 3.2 or in the range from 2.3 to 3.2.
[0161] Alternatively, the diameter of the fan may be in the range
from 330 cm to 380 cm. In such an embodiment the flight cycle ratio
may be less than or equal to 2.8 or in the range from 2.1 to
2.8.
[0162] A product of:
the torsional shear stress of the gearbox support at maximum take
off conditions.times.the torsional shear stress of the gearbox
support at cruise conditions
may be greater than or equal to 1.00.times.10.sup.16
(N/m.sup.2).sup.2, greater than or equal to 2.05.times.10.sup.16
(N/m.sup.2).sup.2, in the range from 1.00.times.10.sup.16
(N/m.sup.2).sup.2 to 7.50.times.10.sup.16 (N/m.sup.2).sup.2, or in
the range from 2.05.times.10.sup.16 (N/m.sup.2) 2 to
4.9.times.10.sup.16 (N/m.sup.2).sup.2. In such embodiments, the
gearbox may, for example, be in a star configuration.
[0163] A first torque transmission ratio of:
the torque transmitted through the gearbox at maximum take off
conditions the torque transmitted through the gearbox at cruise
conditions ##EQU00044##
may be less than or equal to 3.2. The first torque transmission
ratio may be less than or equal to 2.95. The first torque
transmission ratio may be less than or equal to 2.9. The first
torque transmission ratio may be less than or equal to 2.90. The
first torque transmission ratio may be less than or equal to 2.85.
The first torque transmission ratio may be less than or equal to
2.75. The first torque transmission ratio may be in the range from
2.1 to 3.2. The first torque transmission ratio may be in the range
from 2.3 to 2.9.
[0164] A second torque transmission ratio of:
the torque transmitted through the gearbox support at maximum take
off conditions the torque transmitted through the gearbox support
at cruise conditions ##EQU00045##
may be less than or equal to 3.2. The second torque transmission
ratio may be less than or equal to 2.95. The second torque
transmission ratio may be less than or equal to 2.9. The second
torque transmission ratio may be less than or equal to 2.90. The
second torque transmission ratio may be less than or equal to 2.85.
The second torque transmission ratio may be less than or equal to
2.75. The second torque transmission ratio may be in the range from
2.1 to 3.2. The second torque transmission ratio may be in the
range from 2.3 to 2.9.
[0165] The torsional strength of the gearbox support may be greater
than or equal to 1.60.times.10.sup.5 Nm. The torsional strength of
the gearbox support may be greater than or equal to
1.8.times.10.sup.5 Nm. The torsional strength of the gearbox
support may be in the range from 1.60.times.10.sup.5 Nm to
2.00.times.10.sup.7 Nm. The torsional strength of the gearbox
support may be in the range from 1.8.times.10.sup.5 Nm to
1.5.times.10.sup.6 Nm.
[0166] The gearbox may have a cross sectional area, and a first
gearbox support strength ratio of:
the torsional strength of the gearbox support the radial bending
stiffness of the gearbox support .times. the cross sectional area
of the gearbox ##EQU00046##
may be greater than or equal to 7.0.times.10.sup.-3. The first
gearbox support strength ratio may be greater than or equal to
1.0.times.10.sup.-2. The first gearbox support strength ratio may
be greater than or equal to 2.0.times.10.sup.-2. The first gearbox
support strength ratio may be in the range from 7.0.times.10.sup.-3
to 2.5.times.10.sup.-1. The first gearbox support strength ratio
may be in the range from 1.0.times.10.sup.-2 to
1.0.times.10.sup.-1. The first gearbox support strength ratio may
be in the range from 7.0.times.10.sup.-3 to 2.0.times.10.sup.-2.
The first gearbox support strength ratio may be in the range from
2.0.times.10.sup.-2 to 2.5.times.10.sup.-1.
[0167] A planet gear spacing angle in radians may be defined as
2.pi./N, where N is the number of planet gears. A second gearbox
support strength ratio of:
the torsional strength of the gearbox support the tilt stiffness of
the gearbox support .times. the planet gear spacing angle
##EQU00047##
may be greater than or equal to 1.0.times.10.sup.-1. The second
gearbox support strength ratio may be greater than or equal to
1.5.times.10.sup.-1. The second gearbox support strength ratio may
be greater than or equal to 2.5.times.10.sup.-1. The second gearbox
support strength ratio may be in the range from 1.0.times.10 to
3.5. The second gearbox support strength ratio may be in the range
from 1.5.times.10.sup.-1 to 1.7. The second gearbox support
strength ratio may be in the range from 1.0.times.10.sup.-1 to
2.5.times.10.sup.-1. The second gearbox support strength ratio may
be in the range from 2.5.times.10.sup.-1 to 3.5.
[0168] The cross sectional area of the gearbox may be greater than
or equal to 2.4.times.10.sup.-1m.sup.2. The cross sectional area of
the gearbox may be greater than or equal to
2.6.times.10.sup.-1m.sup.2. The cross sectional area of the gearbox
may be in the range from 2.4.times.10.sup.-1m.sup.2 to 1.10
m.sup.2. The cross sectional area of the gearbox may be in the
range from 2.6.times.10.sup.-1m.sup.2 to
9.0.times.10.sup.-1m.sup.2.
[0169] The planet gear spacing angle may be greater than or equal
to 9.0.times.10.sup.-1 rad. The planet gear spacing angle may be in
the range between 9.0.times.10.sup.-1 rad to 2.1 rad.
[0170] The radial bending stiffness of the gearbox support may be
greater than or equal to 1.0.times.10.sup.7 N/m. The radial bending
stiffness of the gearbox support may be greater than or equal to
2.0.times.10.sup.7 N/m. The radial bending stiffness of the gearbox
support may be greater than or equal to 3.0.times.10.sup.7 N/m. The
radial bending stiffness of the gearbox support may be in the range
from 1.0.times.10.sup.7 N/m to 4.0.times.10.sup.8 N/m. The radial
bending stiffness of the gearbox support may be in the range from
2.0.times.10.sup.7 N/m to 3.times.10.sup.8 N/m. The radial bending
stiffness of the gearbox support may be in the range from
3.0.times.10.sup.7 N/m to 2.0.times.10.sup.8 N/m.
[0171] The tilt stiffness of the gearbox support may be greater
than or equal to 1.2.times.10.sup.5 Nm/rad. The tilt stiffness of
the gearbox support may be greater than or equal to
2.4.times.10.sup.5 Nm/rad. The tilt stiffness of the gearbox
support may be greater than or equal to 3.9.times.10.sup.5 Nm/rad.
The tilt stiffness of the gearbox support may be in the range from
1.2.times.10.sup.5 Nm/rad to 2.1.times.10.sup.7 Nm/rad. The tilt
stiffness of the gearbox support may be in the range from
2.4.times.10.sup.5 Nm/rad to 1.6.times.10.sup.7 Nm/rad. The tilt
stiffness of the gearbox support may be in the range from
3.9.times.10.sup.5 Nm/rad to 9.0.times.10.sup.6 Nm/rad.
[0172] The gearbox may be in a star configuration.
[0173] The maximum take-off conditions may be as defined anywhere
herein, for example as operating at maximum take-off thrust for the
engine at ISA sea level pressure and temperature+15.degree. C. with
a fan inlet velocity of between 0.25 and 0.27 Mn, and optionally at
0.25 Mn.
[0174] The cruise conditions may be as defined anywhere herein. The
cruise conditions may mean the conditions at mid-cruise of an
aircraft to which the engine is attached. The cruise conditions may
be conditions experienced by the aircraft and engine at the
midpoint between top of climb and start of decent.
[0175] The forward speed of the gas turbine engine at the cruise
conditions may be in the range of from Mn 0.75 to Mn 0.85. The
forward speed of the gas turbine engine at the cruise conditions
may be Mn 0.8.
[0176] The cruise conditions may correspond to atmospheric
conditions defined by the International Standard Atmosphere at an
altitude of 11582 m and a forward Mach Number of 0.8.
[0177] The cruise conditions may correspond to atmospheric
conditions defined by the International Standard Atmosphere at an
altitude of 10668 m and a forward Mach Number of 0.85.
[0178] The cruise conditions may correspond to atmospheric
conditions at an altitude that is in the range of from 10500 m to
11600 m, and optionally at an altitude of 11000 m.
[0179] According to a twenty-fifth aspect there is provided a
propulsor for an aircraft, comprising: a fan comprising a plurality
of fan blades; a gearbox; a power unit for driving the fan via the
gearbox, wherein the gearbox is arranged to receive an input from
the power unit via a core shaft and output drive to a fan shaft so
as to drive the fan at a lower rotational speed than the core
shaft, the gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the propulsor, wherein: [0180] a flight
cycle ratio of:
[0180] the torsional shear stress of the gearbox support at maximum
take off conditions the torsional shear stress of the gearbox
support at cruise conditions ##EQU00048## [0181] is less than or
equal to 3.20.
[0182] The propulsor of the twenty-fifth aspect may have some or
all of the features described above with respect to the gas turbine
engine of the twenty-fourth aspect, and may be a gas turbine engine
in some embodiments.
[0183] According to another aspect, there is provided an aircraft
comprising the gas turbine engine or the propulsor of the
twenty-fourth aspect or the twenty-fifth aspect mounted thereon,
wherein the aircraft has a maximum take-off operating condition and
a cruise condition.
[0184] According to a twenty-sixth aspect there is provided a
method of operating a gas turbine engine, the 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; 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 gearbox being an
epicyclic gearbox comprising a sun gear, a plurality of planet
gears, a ring gear, and a planet carrier arranged to have the
plurality of planet gears mounted thereon; and a gearbox support
arranged to at least partially support the gearbox within the
engine, and wherein the method comprises: [0185] operating the gas
turbine engine such that a flight cycle ratio of:
[0185] the torsional shear stress of the gearbox support at maximum
takeoff conditions the torsional shear stress of the gearbox
support at cruise conditions ##EQU00049## [0186] is less than or
equal to 3.20.
[0187] The method of the twenty-sixth aspect may be a method of
operating the gas turbine engine or the propulsor of the
twenty-fourth aspect or twenty-fifth aspect respectively. Any of
the features, ratios and parameters introduced above in connection
with the twenty-fourth aspect or twenty-fifth aspect may also
therefore apply to the twenty-sixth aspect. The method of the
twenty-fifth aspect may comprise operating the gas turbine engine
to provide propulsion for the aircraft to which it is mounted under
maximum take-off conditions. The method may further comprise
operating the gas turbine engine to provide propulsion during
cruise conditions. Cruise conditions and max-take off conditions
are as defined elsewhere herein.
[0188] Through consideration of a range of gearbox failure modes
including tooth root bending, tooth surface scuffing, tooth surface
macro and micro pitting and gearbox vibration and the extent to
which those failure modes have an impact at maximum take-off (MTO)
and cruise, the inventor has discovered that providing torsional
shear stresses of the gearbox support structure within the
specified range provides a low weight gearbox capable of
successfully transmitting the largest loads at MTO for limited
durations while providing sufficient integrity against other
failure mechanisms at cruise for extended durations.
[0189] In other aspects, value ranges for the product of the
components of the flight cycle ratio may be specified instead of,
or as well as, value ranges for the ratio.
[0190] According to one such aspect, the twenty-fourth aspect
introduced above may be formulated as an aspect providing 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; 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 gearbox being an epicyclic gearbox comprising a sun
gear, a plurality of planet gears, a ring gear, and a planet
carrier arranged to have the plurality of planet gears mounted
thereon; and a gearbox support arranged to at least partially
support the gearbox within the engine, and wherein: a product (e.g.
a flight cycle product) of:
the torsional shear stress of the gearbox support at maximum take
off conditions.times.the torsional shear stress of the gearbox
support at cruise conditions
is greater than or equal to 1.00.times.10.sup.16 (N/m.sup.2).sup.2,
greater than or equal to 2.05.times.10.sup.16 (N/m.sup.2).sup.2, in
the range from 1.00.times.10.sup.16 (N/m.sup.2).sup.2 to
7.50.times.10.sup.16 (N/m.sup.2).sup.2, or in the range from
2.05.times.10.sup.16 (N/m.sup.2).sup.2 to 4.9.times.10.sup.16
(N/m.sup.2).sup.2. In any of these embodiments, the gearbox may be
in a star configuration.
[0191] The skilled person will appreciate that method and propulsor
aspects may be formulated accordingly.
[0192] In any of the preceding aspects, any one or more of the
following may apply as applicable:
[0193] The turbine may be a first turbine, the compressor may be a
first compressor, and the core shaft may be a first core shaft. The
engine core may further comprise a second turbine, a second
compressor, and a second core shaft connecting the second turbine
to the second compressor. The second turbine, second compressor,
and second core shaft may be arranged to rotate at a higher
rotational speed than the first core shaft.
[0194] The gearbox may have a gear ratio in any range disclosed
herein, for example a gear ratio in the range from 3.2 to 4.5, and
optionally from 3.2 to 4.0.
[0195] The gas turbine engine may have a specific thrust in any
range disclosed herein, for example a specific thrust in the range
from 70 to 90 NKg.sup.-1.
[0196] The gas turbine engine may have a bypass ratio at cruise
conditions in any range disclosed herein, for example in the range
from 12.5 to 18, and optionally from 13 to 16.
[0197] The fan may have a fan diameter greater than 240 cm and less
than or equal to 380 cm. The fan may have a fan diameter greater
than 300 cm and less than or equal to 380 cm.
[0198] The diameter of the fan may be in the range from 240 cm to
280 cm. The diameter of the fan may be in the range from 330 cm to
380 cm.
[0199] The cross sectional area (CSA) of the gearbox may be defined
as the area of the pitch circle of the ring gear.
[0200] The method of any of the aspects defined above may further
comprise driving the gearbox with an input torque of: [0201] i)
greater than or equal to 10,000 Nm, and optionally of 10,000 to
50,000 Nm at cruise; and/or [0202] ii) greater than or equal to
28,000 Nm, and optionally of 28,000 to 135,000 Nm at max-take off
conditions.
[0203] For any parameter or ratio of parameters X claimed or
disclosed herein, a limit on the values that X can take that is
expressed as "X is greater than or equal to Y" can alternatively be
expressed as "1/X is less than or equal to 1/Y". Any of the ratios
or parameters defined in the aspects and statements above may
therefore be expressed as "1/X is less than or equal to 1/Y" rather
than "X is greater than or equal to Y". In such cases, zero can be
taken as a lower bound.
[0204] Various parameters of the gearbox and its mounting
structure, and/or of the engine more generally, may be adjusted to
allow the engine to meet the specifications of the various aspects
summarised above. Comments on various such parameters are provided
below.
[0205] Regarding the stiffness (radial bending and/or tilt) of the
gearbox support the inventor has found that a relatively low
stiffness may be provided to isolate the gearbox from damaging
loads being transmitted into it. The inventor has found that if the
stiffness of the gearbox support is decreased outside of the ranges
specified herein there are resulting problems with dynamic effects
such as lateral vibrations in particular, a stiffness less that is
within ranges defined herein has been found to allow vibrations at
low modal frequencies to be reduced or avoided (the skilled person
would appreciate that the lower modal vibrations have larger
amplitudes/deflections than the higher modes, and so are more
important to avoid). This may be a function of the size of the
gearbox and its configuration. The inventor has also found that the
maximum stiffnesses provided by the ranges defined herein allows
the reduction or avoidance of damaging loads being transmitted to
the gearbox from the fan. This may similarly vary according to size
and gearbox configuration.
[0206] The inventor has found that decreasing the radial bending
and/or tilt stiffness of the fan shaft (at the input to the fan or
the output of the gearbox) outside of the ranges defined herein
would lead to undesirable dynamic effects such as lateral
vibration. In particular, the minimum stiffness defined by the
ranged specified herein allows vibrations at low modal frequencies
to be reduced or avoided (the skilled person would appreciate that
the lower modal vibrations have larger amplitudes/deflections than
the higher modes, and so are more important to avoid). This may be
a function of the size of the gearbox and its configuration.
[0207] The inventor has also found that an upper limit of the fan
shaft radial bending and/or tilt stiffness is affected by the
fundamental properties of the material or materials from which it
is made. For example, a maximum stiffness is affected by the
engineering limit of the material from which it is made. The
materials from which the fan shaft is made (often steels) may, for
example, have a Young's modulus in the range from 100 to 250 GPa,
or 105 to 215 GPa, and optionally around 210 GPa different grades
of steel, or other types of metal, may be selected to achieve
different stiffnesses for the same size and geometry. For example,
steels with a Young's modulus in the range 190 to 215 GPa, titanium
alloys with a Young's modulus in the range 105 to 120 GPa, or a
metal such as titanium with a Young's modulus of around 110 GPa may
be used in various embodiments. The inventor has found increasing
the stiffness outside of the ranges defined herein using materials
such as these would lead to excessive weight with no practical gain
in performance.
[0208] The inventor has discovered that the range of the torsional
strength of the gearbox support defined herein provides a desired
level of torque capacity to provide sufficient reliability, while
not being so great as to add excessive weight to the engine with no
practical performance gain.
[0209] Regarding the other properties of the gearbox support such
as its torsional shear stress, the inventor has found that the
ranges specified herein provide the improvements in performance
described above, but without leading to excessive increases in
overall engine weight.
[0210] As noted elsewhere herein, the present disclosure may relate
to a gas turbine engine. Such a gas turbine engine may comprise an
engine core comprising a turbine, a combustor, a compressor, and a
core shaft connecting the turbine to the compressor. Such a gas
turbine engine may comprise a fan (having fan blades) located
upstream of the engine core.
[0211] The gas turbine engine may comprise 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
input to the gearbox may be directly from the core shaft, or
indirectly from the core shaft, for example via a spur shaft and/or
gear. The core shaft may rigidly connect the turbine and the
compressor, such that the turbine and compressor rotate at the same
speed (with the fan rotating at a lower speed). The output from the
gearbox may be directly to a fan shaft, or indirectly to the fan
shaft, for example via a spur shaft and/or gear.
[0212] The gas turbine engine as described and/or claimed herein
may have any suitable general architecture. For example, the gas
turbine engine may have any desired number of shafts that connect
turbines and compressors, for example one, two or three shafts.
Purely by way of example, the turbine connected to the core shaft
may be a first turbine, the compressor connected to the core shaft
may be a first compressor, and the core shaft may be a first core
shaft. The engine core may further comprise a second turbine, a
second compressor, and a second core shaft connecting the second
turbine to the second compressor. The second turbine, second
compressor, and second core shaft may be arranged to rotate at a
higher rotational speed than the first core shaft.
[0213] In such an arrangement, the second compressor may be
positioned axially downstream of the first compressor. The second
compressor may be arranged to receive (for example directly
receive, for example via a generally annular duct) flow from the
first compressor.
[0214] The gearbox may be arranged to be driven by the core shaft
that is configured to rotate (for example in use) at the lowest
rotational speed (for example the first core shaft in the example
above). For example, the gearbox may be arranged to be driven only
by the core shaft that is configured to rotate (for example in use)
at the lowest rotational speed (for example only be the first core
shaft, and not the second core shaft, in the example above).
Alternatively, the gearbox may be arranged to be driven by any one
or more shafts, for example the first and/or second shafts in the
example above.
[0215] The gearbox may be a reduction gearbox (in that the output
to the fan is a lower rotational rate than the input from the core
shaft). Any type of gearbox may be used. For example, the gearbox
may be a "planetary" or "star" gearbox, as described in more detail
elsewhere herein. The gearbox may have any desired reduction ratio
(defined as the rotational speed of the input shaft divided by the
rotational speed of the output shaft), for example greater than
2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for
example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5,
3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for
example, between any two of the values in the previous sentence.
Purely by way of example, the gearbox may be a "star" gearbox
having a ratio in the range of from 3.1 or 3.2 to 3.8. In some
arrangements, the gear ratio may be outside these ranges.
[0216] In any gas turbine engine as described and/or claimed
herein, a combustor may be provided axially downstream of the fan
and compressor(s). For example, the combustor may be directly
downstream of (for example at the exit of) the second compressor,
where a second compressor is provided. By way of further example,
the flow at the exit to the combustor may be provided to the inlet
of the second turbine, where a second turbine is provided. The
combustor may be provided upstream of the turbine(s).
[0217] The or each compressor (for example the first compressor and
second compressor as described above) may comprise any number of
stages, for example multiple stages. Each stage may comprise a row
of rotor blades and a row of stator vanes, which may be variable
stator vanes (in that their angle of incidence may be variable).
The row of rotor blades and the row of stator vanes may be axially
offset from each other.
[0218] The or each turbine (for example the first turbine and
second turbine as described above) may comprise any number of
stages, for example multiple stages. Each stage may comprise a row
of rotor blades and a row of stator vanes. The row of rotor blades
and the row of stator vanes may be axially offset from each other.
Each fan blade may be defined as having a radial span extending
from a root (or hub) at a radially inner gas-washed location, or 0%
span position, to a tip at a 100% span position. The ratio of the
radius of the fan blade at the hub to the radius of the fan blade
at the tip may be less than (or on the order of) any of: 0.4, 0.39,
0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28,
0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at
the hub to the radius of the fan blade at the tip may be in an
inclusive range bounded by any two of the values in the previous
sentence (i.e. the values may form upper or lower bounds), for
example in the range of from 0.28 to 0.32. These ratios may
commonly be referred to as the hub-to-tip ratio. The radius at the
hub and the radius at the tip may both be measured at the leading
edge (or axially forwardmost) part of the blade. The hub-to-tip
ratio refers, of course, to the gas-washed portion of the fan
blade, i.e. the portion radially outside any platform. The radius
of the fan may be measured between the engine centerline and the
tip of a fan blade at its leading edge. The fan diameter (which may
simply be twice the radius of the fan) may be greater than (or on
the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100
inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110
inches), 290 cm (around 115 inches), 300 cm (around 120 inches),
310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340
cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm
(around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155
inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165
inches). The fan diameter may be in an inclusive range bounded by
any two of the values in the previous sentence (i.e. the values may
form upper or lower bounds), for example in the range of from 240
cm to 280 cm or 330 cm to 380 cm.
[0219] The rotational speed of the fan may vary in use. Generally,
the rotational speed is lower for fans with a higher diameter.
Purely by way of non-limitative example, the rotational speed of
the fan at cruise conditions may be less than 2500 rpm, for example
less than 2300 rpm. Purely by way of further non-limitative
example, the rotational speed of the fan at cruise conditions for
an engine having a fan diameter in the range of from 220 cm to 300
cm (for example 240 cm to 280 cm or 250 cm to 270 cm) may be in the
range of from 1700 rpm to 2500 rpm, for example in the range of
from 1800 rpm to 2300 rpm, for example in the range of from 1900
rpm to 2100 rpm. Purely by way of further non-limitative example,
the rotational speed of the fan at cruise conditions for an engine
having a fan diameter in the range of from 330 cm to 380 cm may be
in the range of from 1200 rpm to 2000 rpm, for example in the range
of from 1300 rpm to 1800 rpm, for example in the range of from 1400
rpm to 1800 rpm.
[0220] In use of the gas turbine engine, the fan (with associated
fan blades) rotates about a rotational axis. This rotation results
in the tip of the fan blade moving with a velocity U.sub.tip. The
work done by the fan blades 13 on the flow results in an enthalpy
rise dH of the flow. A fan tip loading may be defined as
dH/U.sub.tip.sup.2, where dH is the enthalpy rise (for example the
1-D average enthalpy rise) across the fan and U.sub.tip is the
(translational) velocity of the fan tip, for example at the leading
edge of the tip (which may be defined as fan tip radius at leading
edge multiplied by angular speed). The fan tip loading at cruise
conditions may be greater than (or on the order of) any of: 0.28,
0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or
0.4. The fan tip loading may be in an inclusive range bounded by
any two of the values in the previous sentence (i.e. the values may
form upper or lower bounds), for example in the range of from 0.28
to 0.31, or 0.29 to 0.3.
[0221] Gas turbine engines in accordance with the present
disclosure may have any desired bypass ratio, where the bypass
ratio is defined as the ratio of the mass flow rate of the flow
through the bypass duct to the mass flow rate of the flow through
the core at cruise conditions. In some arrangements the bypass
ratio may be greater than (or on the order of) any of the
following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15,
15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass
ratio may be in an inclusive range bounded by any two of the values
in the previous sentence (i.e. the values may form upper or lower
bounds), for example in the range of form 12 to 16, 13 to 15, or 13
to 14. The bypass duct may be substantially annular. The bypass
duct may be radially outside the core engine. The radially outer
surface of the bypass duct may be defined by a nacelle and/or a fan
case.
[0222] The overall pressure ratio of a gas turbine engine as
described and/or claimed herein may be defined as the ratio of the
stagnation pressure upstream of the fan to the stagnation pressure
at the exit of the highest pressure compressor (before entry into
the combustor). By way of non-limitative example, the overall
pressure ratio of a gas turbine engine as described and/or claimed
herein at cruise may be greater than (or on the order of) any of
the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall
pressure ratio may be in an inclusive range bounded by any two of
the values in the previous sentence (i.e. the values may form upper
or lower bounds), for example in the range of from 50 to 70.
[0223] Specific thrust of an engine may be defined as the net
thrust of the engine divided by the total mass flow through the
engine. At cruise conditions, the specific thrust of an engine
described and/or claimed herein may be less than (or on the order
of) any of the following: 110 Nkg.sup.-1s, 105 Nkg.sup.-1s, 100
Nkg.sup.-1s, 95 Nkg.sup.-1s, 90 Nkg.sup.-1s, 85 Nkg.sup.-1s or 80
Nkg.sup.-1s. The specific thrust may be in an inclusive range
bounded by any two of the values in the previous sentence (i.e. the
values may form upper or lower bounds), for example in the range of
from 80 Nkg.sup.-1s to 100 Nkg.sup.-1s, or 85 Nkg.sup.-1s to 95
Nkg.sup.-1s. Such engines may be particularly efficient in
comparison with conventional gas turbine engines.
[0224] A gas turbine engine as described and/or claimed herein may
have any desired maximum thrust. Purely by way of non-limitative
example, a gas turbine as described and/or claimed herein may be
capable of producing a maximum thrust of at least (or on the order
of) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN,
250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The
maximum thrust may be in an inclusive range bounded by any two of
the values in the previous sentence (i.e. the values may form upper
or lower bounds). Purely by way of example, a gas turbine as
described and/or claimed herein may be capable of producing a
maximum thrust in the range of from 330 kN to 420 kN, for example
350 kN to 400 kN. The thrust referred to above may be the maximum
net thrust at standard atmospheric conditions at sea level plus 15
degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),
with the engine static.
[0225] In use, the temperature of the flow at the entry to the high
pressure turbine may be particularly high. This temperature, which
may be referred to as TET, may be measured at the exit to the
combustor, for example immediately upstream of the first turbine
vane, which itself may be referred to as a nozzle guide vane. At
cruise, the TET may be at least (or on the order of) any of the
following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at
cruise may be in an inclusive range bounded by any two of the
values in the previous sentence (i.e. the values may form upper or
lower bounds). The maximum TET in use of the engine may be, for
example, at least (or on the order of) any of the following: 1700K,
1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be
in an inclusive range bounded by any two of the values in the
previous sentence (i.e. the values may form upper or lower bounds),
for example in the range of from 1800K to 1950K. The maximum TET
may occur, for example, at a high thrust condition, for example at
a maximum take-off (MTO) condition.
[0226] A fan blade and/or aerofoil portion of a fan blade described
and/or claimed herein may be manufactured from any suitable
material or combination of materials. For example at least a part
of the fan blade and/or aerofoil may be manufactured at least in
part from a composite, for example a metal matrix composite and/or
an organic matrix composite, such as carbon fibre. By way of
further example at least a part of the fan blade and/or aerofoil
may be manufactured at least in part from a metal, such as a
titanium based metal or an aluminium based material (such as an
aluminium-lithium alloy) or a steel based material. The fan blade
may comprise at least two regions manufactured using different
materials. For example, the fan blade may have a protective leading
edge, which may be manufactured using a material that is better
able to resist impact (for example from birds, ice or other
material) than the rest of the blade. Such a leading edge may, for
example, be manufactured using titanium or a titanium-based alloy.
Thus, purely by way of example, the fan blade may have a
carbon-fibre or aluminium based body (such as an aluminium lithium
alloy) with a titanium leading edge.
[0227] A fan as described and/or claimed herein may comprise a
central portion, from which the fan blades may extend, for example
in a radial direction. The fan blades may be attached to the
central portion in any desired manner. For example, each fan blade
may comprise a fixture which may engage a corresponding slot in the
hub (or disc). Purely by way of example, such a fixture may be in
the form of a dovetail that may slot into and/or engage a
corresponding slot in the hub/disc in order to fix the fan blade to
the hub/disc. By way of further example, the fan blades maybe
formed integrally with a central portion. Such an arrangement may
be referred to as a bladed disc or a bladed ring. Any suitable
method may be used to manufacture such a bladed disc or bladed
ring. For example, at least a part of the fan blades may be
machined from a block and/or at least part of the fan blades may be
attached to the hub/disc by welding, such as linear friction
welding.
[0228] The gas turbine engines described and/or claimed herein may
or may not be provided with a variable area nozzle (VAN). Such a
variable area nozzle may allow the exit area of the bypass duct to
be varied in use. The general principles of the present disclosure
may apply to engines with or without a VAN.
[0229] The fan of a gas turbine as described and/or claimed herein
may have any desired number of fan blades, for example 14, 16, 18,
20, 22, 24 or 26 fan blades.
[0230] As used herein, a maximum take-off (MTO) condition has the
conventional meaning. Maximum take-off conditions may be defined as
operating the engine at International Standard Atmosphere (ISA) sea
level pressure and temperature conditions+15.degree. C. at maximum
take-off thrust at end of runway, which is typically defined at an
aircraft speed of around 0.25 Mn, or between around 0.24 and 0.27
Mn. Maximum take-off conditions for the engine may therefore be
defined as operating the engine at a maximum take-off thrust (for
example maximum throttle) for the engine at ISA sea level pressure
and temperature+15.degree. C. with a fan inlet velocity of 0.25
Mn.
[0231] As used herein, cruise conditions have the conventional
meaning and would be readily understood by the skilled person.
Thus, for a given gas turbine engine for an aircraft, the skilled
person would immediately recognise cruise conditions to mean the
operating point of the engine at mid-cruise of a given mission
(which may be referred to in the industry as the "economic
mission") of an aircraft to which the gas turbine engine is
designed to be attached. In this regard, mid-cruise is the point in
an aircraft flight cycle at which 50% of the total fuel that is
burned between top of climb and start of descent has been burned
(which may be approximated by the midpoint--in terms of time and/or
distance--between top of climb and start of descent. Cruise
conditions thus define an operating point of the gas turbine engine
that provides a thrust that would ensure steady state operation
(i.e. maintaining a constant altitude and constant Mach Number) at
mid-cruise of an aircraft to which it is designed to be attached,
taking into account the number of engines provided to that
aircraft. For example where an engine is designed to be attached to
an aircraft that has two engines of the same type, at cruise
conditions the engine provides half of the total thrust that would
be required for steady state operation of that aircraft at
mid-cruise.
[0232] In other words, for a given gas turbine engine for an
aircraft, cruise conditions are defined as the operating point of
the engine that provides a specified thrust (required to provide in
combination with any other engines on the aircraft--steady state
operation of the aircraft to which it is designed to be attached at
a given mid-cruise Mach Number) at the mid-cruise atmospheric
conditions (defined by the International Standard Atmosphere
according to ISO 2533 at the mid-cruise altitude). For any given
gas turbine engine for an aircraft, the mid-cruise thrust,
atmospheric conditions and Mach Number are known, and thus the
operating point of the engine at cruise conditions is clearly
defined.
[0233] Purely by way of example, the forward speed at the cruise
condition may be any point in the range of from Mach 0.7 to 0.9,
for example 0.75 to 0.85, for example 0.76 to 0.84, for example
0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81,
for example on the order of Mach 0.8, on the order of Mach 0.85 or
in the range of from 0.8 to 0.85. Any single speed within these
ranges may be part of the cruise condition. For some aircraft, the
cruise conditions may be outside these ranges, for example below
Mach 0.7 or above Mach 0.9.
[0234] Purely by way of example, the cruise conditions may
correspond to standard atmospheric conditions (according to the
International Standard Atmosphere, ISA) at an altitude that is in
the range of from 10000m to 15000m, for example in the range of
from 10000m to 12000m, for example in the range of from 10400m to
11600m (around 38000 ft), for example in the range of from 10500m
to 11500m, for example in the range of from 10600m to 11400m, for
example in the range of from 10700m (around 35000 ft) to 11300m,
for example in the range of from 10800m to 11200m, for example in
the range of from 10900m to 11100m, for example on the order of
11000m. The cruise conditions may correspond to standard
atmospheric conditions at any given altitude in these ranges.
[0235] Purely by way of example, the cruise conditions may
correspond to an operating point of the engine that provides a
known required thrust level (for example a value in the range of
from 30 kN to 35 kN) at a forward Mach number of 0.8 and standard
atmospheric conditions (according to the International Standard
Atmosphere) at an altitude of 38000 ft (11582m). Purely by way of
further example, the cruise conditions may correspond to an
operating point of the engine that provides a known required thrust
level (for example a value in the range of from 50 kN to 65 kN) at
a forward Mach number of 0.85 and standard atmospheric conditions
(according to the International Standard Atmosphere) at an altitude
of 35000 ft (10668m).
[0236] In use, a gas turbine engine described and/or claimed herein
may operate at the cruise conditions defined elsewhere herein. Such
cruise conditions may be determined by the cruise conditions (for
example the mid-cruise conditions) of an aircraft to which at least
one (for example 2 or 4) gas turbine engine may be mounted in order
to provide propulsive thrust.
[0237] According to an aspect, there is provided an aircraft
comprising a gas turbine engine as described and/or claimed herein.
The aircraft according to this aspect is the aircraft for which the
gas turbine engine has been designed to be attached. Accordingly,
the cruise conditions according to this aspect correspond to the
mid-cruise of the aircraft, as defined elsewhere herein.
[0238] According to an aspect, there is provided a method of
operating a gas turbine engine as described and/or claimed herein.
The operation may be at the cruise conditions as defined elsewhere
herein (for example in terms of the thrust, atmospheric conditions
and Mach Number).
[0239] According to an aspect, there is provided a method of
operating an aircraft comprising a gas turbine engine as described
and/or claimed herein. The operation according to this aspect may
include (or may be) operation at the mid-cruise of the aircraft, as
defined elsewhere herein.
[0240] Whilst in the arrangements described herein the source of
drive for the propulsive fan is provided by a gas turbine engine,
the skilled person will appreciate the applicability of the gearbox
configurations disclosed herein to other forms of aircraft
propulsor comprising alternative drive types. For example, the
above-mentioned gearbox arrangements may be utilised in aircraft
propulsors comprising a propulsive fan driven by an electric motor.
In such circumstances, the electric motor may be configured to
operate at higher rotational speeds and thus may have a lower rotor
diameter and may be more power-dense. The gearbox configurations of
the aforesaid aspects may be employed to reduce the rotational
input speed for the fan or propeller to allow it to operate in a
more favourable efficiency regime. Thus, according to an aspect,
there is provided an electric propulsion unit for an aircraft,
comprising an electric machine configured to drive a propulsive fan
via a gearbox, the gearbox and/or its inputs/outputs/supports
and/or the structure by which the fan shaft driving the fan is
supported being as described and/or claimed herein.
[0241] The skilled person will appreciate that except where
mutually exclusive, a feature or parameter described in relation to
any one of the above aspects may be applied to any other aspect.
Furthermore, except where mutually exclusive, any feature or
parameter described herein may be applied to any aspect and/or
combined with any other feature or parameter described herein.
[0242] As used herein, a range "from value X to value Y" or
"between value X and value Y", or the like, denotes an inclusive
range; including the bounding values of X and Y. As used herein,
the term "axial plane" denotes a plane extending along the length
of an engine, parallel to and containing an axial centerline of the
engine, and the term "radial plane" denotes a plane extending
perpendicular to the axial centerline of the engine, so including
all radial lines at the axial position of the radial plane. Axial
planes may also be referred to as longitudinal planes, as they
extend along the length of the engine. A radial distance or an
axial distance is therefore a distance in a radial or axial plane,
respectively.
BRIEF DESCRIPTION OF THE DRAWINGS
[0243] Embodiments will now be described by way of example only,
with reference to the Figures, in which:
[0244] FIG. 1 is a sectional side view of a gas turbine engine;
[0245] FIG. 2 is a close up sectional side view of an upstream
portion of a gas turbine engine;
[0246] FIG. 3 is a partially cut-away view of a gearbox for a gas
turbine engine;
[0247] FIG. 4 is a schematic diagram illustrating the radial
bending stiffness of a cantilevered beam;
[0248] FIG. 5 is a schematic diagram illustrating the tilt
stiffness of a cantilevered beam;
[0249] FIG. 6 is a schematic diagram illustrating the torsional
stiffness of a shaft;
[0250] FIG. 7 is a close up sectional view of the region of a gas
turbine engine around its gearbox;
[0251] FIGS. 8 and 9 are schematic diagrams illustrating how the
radial bending stiffness of a gearbox support may be defined;
[0252] FIGS. 10 and 11 are schematic diagrams illustrating how the
tilt stiffness of a gearbox support may be defined;
[0253] FIG. 12 shows a schematic sectional view of a gearbox
support used with a gearbox in a star configuration;
[0254] FIG. 13 shows a schematic sectional view of a gearbox
support used with a gearbox in a planetary configuration;
[0255] FIG. 14 is a schematic diagram illustrating an alternative
interface between a fan shaft and fan;
[0256] FIGS. 15 and 16 are schematic diagrams illustrating how fan
shaft end radial bending stiffnesses may be defined;
[0257] FIGS. 17 and 18 are schematic diagrams illustrating how fan
shaft end tilt stiffnesses may be defined;
[0258] FIG. 19 shows an aircraft having a gas turbine engine
attached to each wing;
[0259] FIG. 20 shows a method of operating a gas turbine engine on
an aircraft; and
[0260] FIG. 21 shows a graph of applied load against displacement
to illustrate measurement of the stiffness of a component.
DETAILED DESCRIPTION
[0261] FIG. 1 illustrates a gas turbine engine 10 having a
principal rotational axis 9. The engine 10 comprises an air intake
12 and a propulsive fan 23 that generates two airflows: a core
airflow A and a bypass airflow B. The gas turbine engine 10
comprises a core 11 that receives the core airflow A. The engine
core 11 comprises, in axial flow series, a low pressure compressor
14, a high-pressure compressor 15, combustion equipment 16, a
high-pressure turbine 17, a low pressure turbine 19 and a core
exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10
and defines a bypass duct 22 and a bypass exhaust nozzle 18. The
bypass airflow B flows through the bypass duct 22. The fan 23 is
attached to and driven by the low pressure turbine 19 via a shaft
26 and an epicyclic gearbox 30.
[0262] In use, the core airflow A is accelerated and compressed by
the low pressure compressor 14 and directed into the high pressure
compressor 15 where further compression takes place. The compressed
air exhausted from the high pressure compressor 15 is directed into
the combustion equipment 16 where it is mixed with fuel and the
mixture is combusted. The resultant hot combustion products then
expand through, and thereby drive, the high pressure and low
pressure turbines 17, 19 before being exhausted through the nozzle
20 to provide some propulsive thrust. The high pressure turbine 17
drives the high pressure compressor 15 by a suitable
interconnecting shaft 27. The fan 23 generally provides the
majority of the propulsive thrust. The epicyclic gearbox 30 is a
reduction gearbox.
[0263] An exemplary arrangement for a geared fan gas turbine engine
10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1)
drives the shaft 26, which is coupled to a sun wheel, or sun gear,
28 of the epicyclic gear arrangement 30. Radially outwardly of the
sun gear 28 and intermeshing therewith is a plurality of planet
gears 32 that are coupled together by a planet carrier 34. The
planet carrier 34 constrains the planet gears 32 to process around
the sun gear 28 in synchronicity whilst enabling each planet gear
32 to rotate about its own axis. The planet carrier 34 is coupled
via linkages 36 to the fan 23 in order to drive its rotation about
the engine axis 9. Radially outwardly of the planet gears 32 and
intermeshing therewith is an annulus or ring gear 38 that is
coupled, via linkages 40, to a stationary supporting structure
24.
[0264] The linkages 36 may be referred to as a fan shaft 36, the
fan shaft 36 optionally comprising two or more shaft portions
coupled together. For example, the fan shaft 36 may comprise a
gearbox output shaft portion 36a extending from the gearbox 30 and
a fan portion 36b extending between the gearbox output shaft
portion and the fan 23. In the embodiment shown in FIGS. 1 and 2,
the gearbox 30 is a planetary gearbox and the gearbox output shaft
portion 36a is connected to the planet carrier 34 it may therefore
be referred to as a carrier output shaft 36a. In star gearboxes 30,
the gearbox output shaft portion 36a may be connected to the ring
gear 38 it may therefore be referred to as a ring output shaft 36a.
In the embodiment shown in FIGS. 1 and 2, the fan portion 36b of
the fan shaft 36 connects the gearbox output shaft portion 36a to
the fan 23. The output of the gearbox 30 is therefore transferred
to the fan 23, to rotate the fan, via the fan shaft 36. In
alternative embodiments, the fan shaft 36 may comprise a single
component, or more than two components. Unless otherwise indicated
or apparent to the skilled person, anything described with respect
to an engine 10 with a star gearbox 30 may equally be applied to an
engine with a planetary gearbox 30, and vice versa.
[0265] Note that the terms "low pressure turbine" and "low pressure
compressor" as used herein may be taken to mean the lowest pressure
turbine stages and lowest pressure compressor stages (i.e. not
including the fan 23) respectively and/or the turbine and
compressor stages that are connected together by the
interconnecting shaft 26 with the lowest rotational speed in the
engine (i.e. not including the gearbox output shaft that drives the
fan 23). In some literature, the "low pressure turbine" and "low
pressure compressor" referred to herein may alternatively be known
as the "intermediate pressure turbine" and "intermediate pressure
compressor". Where such alternative nomenclature is used, the fan
23 may be referred to as a first, or lowest pressure, compression
stage.
[0266] The epicyclic gearbox 30 is shown by way of example in
greater detail in FIG. 3. Each of the sun gear 28, planet gears 32
and ring gear 38 comprise teeth about their periphery to intermesh
with the other gears. However, for clarity only exemplary portions
of the teeth are illustrated in FIG. 3. There are four planet gears
32 illustrated, although it will be apparent to the skilled reader
that more or fewer planet gears 32 may be provided within the scope
of the claimed invention. Practical applications of a planetary
epicyclic gearbox 30 generally comprise at least three planet gears
32.
[0267] The epicyclic gearbox 30 illustrated by way of example in
FIGS. 2 and 3 is of the planetary type, in that the planet carrier
34 is coupled to an output shaft via linkages 36, with the ring
gear 38 fixed. However, any other suitable type of epicyclic
gearbox 30 may be used. By way of further example, the epicyclic
gearbox 30 may be a star arrangement, in which the planet carrier
34 is held fixed, with the ring (or annulus) gear 38 allowed to
rotate. In such an arrangement the fan 23 is driven by the ring
gear 38. By way of further alternative example, the gearbox 30 may
be a differential gearbox in which the ring gear 38 and the planet
carrier 34 are both allowed to rotate.
[0268] It will be appreciated that the arrangement shown in FIGS. 2
and 3 is by way of example only, and various alternatives are
within the scope of the present disclosure. Purely by way of
example, any suitable arrangement may be used for locating the
gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to
the engine 10. By way of further example, the connections (such as
the linkages 36, 40 in the FIG. 2 example) between the gearbox 30
and other parts of the engine 10 (such as the input shaft 26, the
output shaft and the fixed structure 24) may have any desired
degree of stiffness or flexibility as defined or claimed elsewhere
herein. By way of further example, any suitable arrangement of the
bearings between rotating and stationary parts of the engine (for
example between the input and output shafts from the gearbox and
the fixed structures, such as the gearbox casing) may be used, and
the disclosure is not limited to the exemplary arrangement of FIG.
2. For example, where the gearbox 30 has a star arrangement
(described above), the skilled person would readily understand that
the arrangement of output and support linkages and bearing
locations would typically be different to that shown by way of
example in FIG. 2 (for example as described in connection with
other embodiments disclosed herein which have a star gearbox
arrangement).
[0269] Accordingly, the present disclosure extends to a gas turbine
engine having any arrangement of gearbox styles (for example star
or planetary), support structures, input and output shaft
arrangement, and bearing locations.
[0270] Optionally, the gearbox may drive additional and/or
alternative components (e.g. the intermediate pressure compressor
and/or a booster compressor).
[0271] Other gas turbine engines to which the present disclosure
may be applied may have alternative configurations. For example,
such engines may have an alternative number of compressors and/or
turbines and/or an alternative number of interconnecting shafts. By
way of further example, the gas turbine engine shown in FIG. 1 has
a split flow nozzle 18, 20 meaning that the flow through the bypass
duct 22 has its own nozzle 18 that is separate to and radially
outside the core engine nozzle 20. However, this is not limiting,
and any aspect of the present disclosure may also apply to engines
in which the flow through the bypass duct 22 and the flow through
the core 11 are mixed, or combined, before (or upstream of) a
single nozzle, which may be referred to as a mixed flow nozzle. One
or both nozzles (whether mixed or split flow) may have a fixed or
variable area. Whilst the described example relates to a turbofan
engine, the disclosure may apply, for example, to any type of gas
turbine engine, such as an open rotor (in which the fan stage is
not surrounded by a nacelle) or turboprop engine, for example.
[0272] The geometry of the gas turbine engine 10, and components
thereof, is defined by a conventional axis system, comprising an
axial direction (which is aligned with the rotational axis 9), a
radial direction (in the bottom-to-top direction in FIG. 1), and a
circumferential direction (perpendicular to the page in the FIG. 1
view). The axial, radial and circumferential directions are
mutually perpendicular.
[0273] The following general definitions of stiffnesses may be used
herein:
Radial Bending Stiffness
[0274] A radial bending stiffness is a measure of deformation for a
given force applied in any one selected radial direction (i.e. any
direction perpendicular to and passing through the engine axis).
The radial bending stiffness is defined with reference to FIG. 4 in
terms of the deformation of a cantilevered beam 401. As illustrated
in FIG. 4, a force, F, applied at the free end of the beam in a
direction perpendicular to the longitudinal axis of the beam causes
a linear perpendicular deformation, .delta.. The radial bending
stiffness is the force applied for a given linear deformation i.e.
F/.delta.. In the present application, the radial bending stiffness
is taken relative to the rotational axis of the engine 9, and so
relates to the resistance to linear deformation in a radial
direction of the engine caused by a radial force. The beam, or
equivalent cantilevered component, extends along the axis of
rotation of the engine, the force, F, is applied perpendicular to
the axis of rotation of the engine, along any radial direction, and
the displacement .delta. is measured perpendicular to the axis of
rotation, along the line of action of the force. The radial bending
stiffness as defined herein has SI units of N/m. In the present
application, unless otherwise stated, the radial bending stiffness
is taken to be a free-body stiffness i.e. stiffness measured for a
component in isolation in a cantilever configuration, without other
components present which may affect its stiffness.
[0275] When the force is applied perpendicular to the cantilevered
beam, and at the free end of the beam, the resultant curvature is
not constant but rather increases towards the fixed end of the
beam.
Tilt Stiffness
[0276] A tilt stiffness is defined with reference to FIG. 5, which
shows the resulting deformation of a cantilevered beam 401 under a
moment M applied at its free end. The tilt stiffness is a measure
of the resistance to rotation of a point on the component at which
a moment is applied. As can be seen in FIG. 5, an applied moment at
the free end of the cantilevered beam induces a constant curvature
along the length of the beam between its free and fixed ends. The
applied moment M causes a rotation .theta. of the point at which it
is applied. For any component of constant section (like the beam),
the angle .theta. is constant along the length of the component.
The tilt stiffness as defined herein therefore has SI units of
Nm/rad.
Torsional Stiffness
[0277] Torsional stiffness is a measure of deformation for a given
torque. FIG. 6 illustrates the definition of the torsional
stiffness of a shaft 401 or other body. A torque, .tau., applied to
the free end of the beam causes a rotational deformation, .theta.
(e.g. twist) along the length of the beam. The torsional stiffness
is the torque applied for a given angle of twist i.e. r/O. The
torsional stiffness has SI units of Nm/rad.
[0278] The following general definitions of other parameters may
also be used herein:
Torque
[0279] Torque, which may also be referred to as moment, is the
rotational equivalent of linear force, and can be thought of as a
twist to an object. The magnitude, .tau., of torque, .tau., of a
body depends on three quantities: the force applied (F), the lever
arm vector connecting the origin to the point of force application
(r), and the angle (A) between the force and lever arm vectors:
.tau.=r.times.F
.tau.=|.tau.|=|r.times.F|=|r.parallel.F|sin A
where [0280] .tau. is the torque vector and .tau. is the magnitude
of the torque; [0281] r is the position vector or "lever arm"
vector (a vector from the selected point on the body to the point
where the force is applied); [0282] F is the force vector; [0283] x
denotes the cross product; and [0284] A is the angle between the
force vector and the lever arm vector (sin(A) is therefore one when
the force vector is perpendicular to the position vector, such that
.tau.=rF, i.e. magnitude of the force multiplied by distance
between the selected point on the body and the point of application
of the force).
[0285] Torque has dimensions of [force].times.[distance] and may be
expressed in units of Newton metres (Nm).
[0286] The net torque on a body determines the rate of change of
the body's angular momentum.
Moment of Inertia
[0287] Moment of inertia, otherwise known as angular mass or
rotational inertia, is a quantity that determines the torque needed
for a desired angular acceleration of a body about a rotational
axis this is substantially equivalent to how mass determines the
force needed for a particular acceleration.
[0288] Moment of inertia depends on the body's mass distribution
and the axis chosen, with larger moments requiring more torque to
change the body's rotation rate. Moment of inertia has dimensions
of [mass].times.[distance].sup.2 and may be expressed in units of
kilogram meter squared (kgm.sup.2). Moment of inertia I is defined
as the ratio of the net angular momentum L of a body to its angular
velocity .omega. around a principal axis:
I = L .omega. ##EQU00050##
[0289] Provided that the shape of the body does not change, its
moment of inertia appears in Newton's law of motion as the ratio of
an applied torque .tau. on a body to the angular acceleration
.alpha. around the principal axis:
.tau.=I.alpha.
[0290] For bodies constrained to rotate in a plane, only the moment
of inertia about an axis perpendicular to the plane matters, and I
can therefore be represented as a scalar value. The skilled person
would appreciate that a fan of a gas turbine engine (and more
generally a fan rotor of the gas turbine engine comprising the fan
disc and blades, and optionally also the fan shaft and/or other
related components) is constrained to rotate in only one plane--a
plane perpendicular to the engine axis--and that the fan's moment
of inertia can therefore be defined by a single, scaler, value.
[0291] The fan's moment of inertia about the engine axis can
therefore be measured or defined using any standard methodology
Torsional Shear Stress
[0292] Shear stress is the component of stress coplanar with a
material cross section; the stress tending to produce shear. Shear
stress arises from a force vector component parallel to the cross
section of the material. Shear stress may be defined as the
external force acting on an object or surface parallel to the slope
or plane in which it lies.
[0293] When a shaft or other body is subjected to a torque, or
twisting, a shearing stress is produced in the body--the shearing
stress may be referred to as a torsional shear stress as it results
from a torque. The shear stress varies from zero along the axis of
the torque rotation to a maximum at the part of the body furthest
from the axis; a radial distance from the axis at which the shear
stress is to be measured is therefore selected. In the arrangements
being described, a mid-height of the component of which the
torsional shear stress is being measured is selected. For example,
if the component has an outer radius of 20 cm, a mid-height
position of 10 cm from the axis 9 is selected.
[0294] Torsional shear stress has dimensions of
[force]/[distance].sup.2 and may be expressed in units of Pascals
(Pa) or Newtons/metre squared (N/m.sup.2).
Shear Strength
[0295] Shear strength is the strength of a body against the body
failing in shear--Shear strength is a material's ability to resist
forces that can cause the internal structure of the material to
slide against itself. A shear load is a force that tends to produce
a sliding failure on a material along a plane that is parallel to
the direction of the force.
[0296] The ultimate shear strength of a material is the maximum
shear stress that may be sustained before the material will
rupture. The proof strength of a material is the stress at which a
particular degree of permanent deformation occurs--for example, a
0.2% proof strength is the stress at which a 0.2% permanent
deformation occurs.
[0297] The strength of the body against shearing when a rotational,
or twisting, shear load is applied may be referred to as torsional
shear strength. The skilled person would appreciate that the shear
strength of a component is important for selecting the dimensions
and materials to be used for the manufacture or construction of the
component. Shear strength has the same units as torsional shear
stress, as it is a maximum supportable shear stress; it may
therefore be expressed in units of Pa or N/m.sup.2.
[0298] As used herein, the listed shear strength of a material is
the 0.2% proof shear strength of the material unless otherwise
specified. The skilled person would appreciate that this is lower
than the ultimate shear strength.
[0299] Shear strength of a material generally varies with
temperature. The shear strength as used herein may be defined at
room temperature.
[0300] For metals (including metal alloys), shear strength
generally decreases gradually with increasing temperature until a
threshold temperature is reached, beyond which the material
strength decreases rapidly. The shear strength may therefore be
approximately constant below the threshold temperature. The
threshold temperature may be around 400-500.degree. C. for various
steel grades. As the temperatures in and around the gearbox 30
generally do not exceed 120.degree. C., well below the threshold
temperature for likely materials for gearbox, gearbox support
structure, and shaft construction, the shear strength may not be
significantly different from the room temperature shear strength
the choice of a specific temperature for strength assessment may
therefore have little effect.
Torsional Strength
[0301] The torsional strength of a specific component is defined as
the ability of a component to withstand an applied torque without
failure (i.e. without yield or structural failure). The torsional
strength is therefore a measure of the torque capacity of a
component. The torsional strength may be measured by applying a
varying torque to a component and determining the torque at which
the component fails. The torsional strength can be considered as
the torque applied to the component at the point where the shear
strength of that component is reached causing it to fail. The
torsional strength as defined herein has dimensions of
[force].times.[distance] and may be expressed in units of Nm.
[0302] More specific definitions of stiffnesses and other
parameters relating to embodiments described herein are provided
below for ease of understanding.
Gearbox Support Stiffness
[0303] FIG. 7 shows a region of the engine core 11 around the
gearbox in close up. The same reference numbers have been used for
components corresponding to those shown in FIGS. 1 to 3. In the
arrangement shown in FIG. 7 the gearbox 30 has a star arrangement,
in which the ring gear 38 is coupled to the fan shaft 36 and the
carrier 34 is held in a fixed position relative to the static
structure of the engine core (e.g. relative to the stationary
supporting structure 24).
[0304] The fan shaft 36 is mounted within the engine by a fan shaft
mounting structure 503. The fan shaft mounting structure 503
comprises at least two bearings connected to or otherwise in
engagement with the fan shaft at points spaced apart axially along
the length of the engine. The fan shaft mounting structure 503 may
take a number of different forms, and may comprise one or more
separate supporting structures provided to support the fan shaft.
It may also include other structures provided to support the fan
shaft such as inter-shaft bearings. It therefore includes any
supporting structure that extends between a bearing in contact with
the fan shaft and a stationary structure of the engine (e.g. of the
engine core).
[0305] In the arrangement shown in FIG. 7, the fan shaft mounting
structure 503 comprises two bearings, a first supporting bearing
506a and a second supporting bearing 506b, via which it is coupled
to the fan shaft 36. The supporting bearings 506a, 506a are spaced
apart along the axial length of the fan shaft 36. In the described
arrangement, both supporting bearings 506a, 506b are provided at
positions that are forward of the gearbox 30. In other
arrangements, one of the two supporting bearing 506a, 506b used to
support the fan shaft 36 may be located at a position rearward of
the gearbox 30. In yet other arrangements, more than two supporting
bearings may be provided as part of the fan shaft mounting
structure or fan shaft mounting structure.
[0306] The engine core 11 comprises a gearbox support 40
(corresponding to the linkage described with reference to FIG. 2)
arranged to support or mount (e.g. to at least partially support or
mount) the gearbox 30 in a fixed position within the engine. The
gearbox support is coupled at a first end to the stationary
supporting structure 24 which extends across the core duct 502
carrying the core airflow A as illustrated in FIG. 7. In the
presently described arrangement, the stationary support structure
24 is an engine section stator (ESS) that acts as both a structural
component to provide a stationary mounting for core components such
as the gearbox support, and as a guide vane provided to direct
airflow from the fan 23. In other embodiments, the stationary
supporting structure 24 may comprise a strut extending across the
core gas flow path and a separate stator vane provided to direct
airflow. In the presently described arrangement, the gearbox
support 40 is coupled at a second end to the planet carrier 34. The
gearbox support 40 therefore acts against rotation of the planet
carrier 34 relative to the static structure of the engine core
(e.g. relative to the stationary supporting structure 24).
[0307] In embodiments where the gearbox 30 is in a planetary
arrangement, the gearbox support 40 is coupled to the ring gear 38
so as to resist its rotation relative to the static structure of
the engine core (e.g. relative to the stationary supporting
structure 24).
[0308] The gearbox support 40 is defined between the point at which
it connects to the gearbox (e.g. to the planet carrier in the
presently described arrangement) and a point at which it connects
to the stationary supporting structure 24. The gearbox support may
be formed by any number of separate components providing a coupling
between those two points.
[0309] The gearbox support 40 has a degree of flexibility
characterized by its radial bending stiffness and its tilt
stiffness.
Gearbox Support Radial Bending Stiffness:
[0310] The radial bending stiffness of the gearbox support 40 is
defined with reference to FIGS. 8 and 9. The radial bending
stiffness can be considered to represent the resistance of the
gearbox support to a force applied to it in a radial direction of
the engine. The radial bending stiffness is determined by treating
the gearbox support 40 as a free body that is fixed at its point of
connection with the stationary supporting structure 24 of the
engine, and which has a radial force F.sub.1 applied at its point
of connection with the gearbox 30 as illustrated in FIG. 8.
Deformation of the gearbox support 40 caused by the applied force
F.sub.1 is illustrated in FIG. 9, with the shape of the support
with no force applied shown in broken lines for comparison. The
radial bending stiffness is defined in terms of the radial
displacement, .delta..sub.1, of the gearbox support at the position
at which the force F.sub.1 is applied. The force, F.sub.1, is shown
radially towards the engine axis 9 in FIG. 8, but could
equivalently be a force radially away from the engine axis 9. The
radial bending stiffness of the gearbox support 40 is therefore
given by F.sub.1/.delta..sub.1. The gearbox 30 is shown for
reference in FIG. 9 remaining in a stationary position despite the
deformation of the gearbox support 40 to which it is connected this
is for illustration purposes only and reflects the gearbox support
40 being treated as a free body in order to determine the
stiffness.
[0311] In various embodiments, the radial bending stiffness of the
gearbox support may be greater than or equal to 1.0.times.10.sup.7
N/m, and optionally greater than or equal to 2.0.times.10.sup.7 N/m
or greater than or equal to 3.0.times.10.sup.7 N/m.
[0312] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the radial bending
stiffness of the gearbox support may be greater than or equal to
1.0.times.10.sup.7 N/m or greater than or equal to
2.7.times.10.sup.7 N/m. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the radial bending stiffness of the gearbox support may be
greater than or equal to 3.2.times.10.sup.7 N/m or greater than or
equal to 4.0.times.10.sup.7 N/m
[0313] In various embodiments, the radial bending stiffness of the
gearbox support may be in the range from 1.0.times.10.sup.7 N/m to
4.0.times.10.sup.8 N/m, and optionally in the range from
2.0.times.10.sup.7 N/m to 3.times.10.sup.8 N/m or further
optionally in the range from 3.0.times.10.sup.7 N/m to
2.0.times.10.sup.8 N/m.
[0314] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the radial bending
stiffness of the gearbox support may be in the range from
1.0.times.10.sup.7 N/m to 3.1.times.10.sup.8 N/m and optionally in
the range from 2.7.times.10.sup.7 N/m to 3.7.times.10.sup.7 N/m
(and may be equal to 3.2.times.10.sup.7 N/m).
[0315] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the radial bending
stiffness of the gearbox support may be in the range from
3.2.times.10.sup.7 N/m to 4.0.times.10.sup.8 N/m, and optionally in
the range from 4.0.times.10.sup.7 N/m to 5.0.times.10.sup.7 N/m
(and may be equal to 4.5.times.10.sup.7 N/m).
Gearbox Support Tilt Stiffness:
[0316] The tilt stiffness of the gearbox support 40 is defined with
reference to FIGS. 10 and 11. The tilt stiffness can be considered
to represent the resistance of the gearbox support to an applied
moment. The tilt stiffness is determined by treating the gearbox
support 40 as a free body that is fixed at its point of connection
with the stationary supporting structure 24. In order to measure
the tilt stiffness, a moment M.sub.1 is applied at the point of
connection between the gearbox support 40 and the gearbox 30 as
illustrated in FIG. 10. Deformation of the gearbox support 40
caused by the applied moment M.sub.1 is illustrated in FIG. 11,
with the shape of the support with no moment applied shown in
broken lines. The tilt stiffness is defined in terms of the angular
displacement, .theta..sub.1, of the gearbox support 40 at the
position at which the moment M.sub.1 is applied. The tilt stiffness
is therefore given by M.sub.1/.theta..sub.1. The gearbox 30 is
again shown for reference in FIG. 11 remaining in a stationary
position despite the deformation of the gearbox support 40 to which
it is connected.
[0317] In various embodiments, the tilt stiffness of the gearbox
support may be greater than or equal to 1.2.times.10.sup.5 Nm/rad,
and optionally greater than or equal to 2.4.times.10.sup.5 Nm/rad
or optionally greater than or equal to 3.9.times.10.sup.5
Nm/rad.
[0318] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the tilt stiffness
of the gearbox support may be greater than or equal to
1.2.times.10.sup.5 Nm/rad or greater than or equal to
4.5.times.10.sup.5 Nm/rad. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the tilt stiffness of the gearbox support may be greater
than or equal to 5.0.times.10.sup.5 Nm/rad or greater than or equal
to 6.0.times.10.sup.5 Nm/rad.
[0319] In various embodiments, the tilt stiffness of the gearbox
support may be in the range from 1.2.times.10.sup.5 Nm/rad to
2.1.times.10.sup.7 Nm/rad, and optionally in the range from
2.4.times.10.sup.5 Nm/rad to 1.6.times.10.sup.7 Nm/rad, and further
optionally in the range from 3.9.times.10.sup.5 Nm/rad to
9.0.times.10.sup.6 Nm/rad.
[0320] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the tilt stiffness
of the gearbox support may be in the range from 1.2.times.10.sup.5
Nm/rad to 7.0.times.10.sup.6 Nm/rad and optionally in the range
from 4.5.times.10.sup.5 Nm/rad to 6.5.times.10.sup.5 Nm/rad (and
may be equal to 5.5.times.10.sup.5 Nm/rad).
[0321] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the tilt stiffness
of the gearbox support may be in the range from 5.0.times.10.sup.5
Nm/rad to 2.1.times.10.sup.7 Nm/rad, and optionally in the range
from 6.0.times.10.sup.5 Nm/rad to 2.6.times.10.sup.6 Nm/rad (and
may be equal to 1.6.times.10.sup.6 Nm/rad).
Gearbox Support Torsional Shear Stress and Torsional Strength
[0322] As described elsewhere herein the gearbox support is
provided to locate the gearbox within the engine. In order to
resist relative rotation of the gearbox compared to the stationary
structure of the engine a torque is transmitted through the gearbox
support. This varies over the operating cycle of the engine as
different levels of torque are transmitted through the gearbox (as
defined elsewhere herein) at different stages of the flight cycle
of the aircraft to which the engine is mounted.
[0323] The torque transmitted through the gearbox support is
defined as the torque at the point of connection between the
gearbox support 40 and the gearbox 30.
[0324] The torsional strength of the gearbox support 40 is defined
as the level of torque applied at the point of connection between
the gearbox support 40 and the gearbox 20 that would result in
failure of the gearbox support.
[0325] When measuring the torsional strength or torque transmitted
through the gearbox support 40 the gearbox support is considered to
be a free body that is fixed at its point of connection with the
stationary supporting structure 24 of the engine with a torque
applied at the point of connection to the gearbox (i.e. in a
similar manner as the application of the force or moment used to
determine the radial bending stiffness and tilt stiffness of the
gearbox support).
[0326] In the arrangement illustrated in FIG. 7, the point of
connection between the gearbox support 40 and the gearbox 30 is the
point of connection to the planet carrier 34 of the gearbox. Where
the gearbox is in a planetary configuration, this point of
connection would be between the gearbox support 40 and the ring
gear 38.
[0327] In various embodiments, the torsional strength of the
gearbox support may be greater than or equal to 1.60.times.10.sup.5
Nm, and optionally greater than or equal to 1.8.times.10.sup.5
Nm.
[0328] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torsional
strength of the gearbox support may be greater than or equal to
1.8.times.10.sup.5 Nm. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the torsional strength of the gearbox support may be
greater than or equal to 4.0.times.10.sup.5 Nm or greater than or
equal to 5.5.times.10.sup.5 Nm.
[0329] In various embodiments, the torsional strength of the
gearbox support may be in the range from 1.60.times.10.sup.5 Nm to
2.00.times.10.sup.7 Nm, and optionally in the range from
1.8.times.10.sup.5 Nm to 1.5.times.10.sup.6 Nm.
[0330] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torsional
strength of the gearbox support may be in the range from
1.8.times.10.sup.5 Nm to 7.0.times.10.sup.5 Nm and optionally in
the range from 1.8.times.10.sup.5 Nm to 2.6.times.10.sup.5 Nm (and
may be equal to 2.2.times.10.sup.5 Nm).
[0331] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the torsional
strength of the gearbox support may be in the range from
4.0.times.10.sup.5 Nm to 2.0.times.10.sup.7 Nm and optionally in
the range from 5.5.times.10.sup.5 Nm to 7.5.times.10.sup.5 Nm (and
may be equal to 6.5.times.10.sup.5 Nm).
Torque Transmitted Through the Gearbox Support at MOT
Conditions:
[0332] In various embodiments, the torque transmitted through the
gearbox support at maximum take-off conditions may be greater than
or equal to 6.00.times.10.sup.4 Nm, and optionally greater than or
equal to 7.2.times.10.sup.4 Nm.
[0333] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torque
transmitted through the gearbox support at maximum take-off
conditions may be greater than or equal to 7.0.times.10.sup.4 Nm.
In some embodiments, for example in embodiments in which the fan
diameter is in the range from 330 to 380 cm, the torque transmitted
through the gearbox support at maximum take-off conditions may be
greater than or equal to 1.8.times.10.sup.5 Nm.
[0334] In various embodiments, the torque transmitted through the
gearbox support at maximum take-off conditions may be in the range
from 6.00.times.10.sup.4 Nm to 5.00.times.10.sup.5 Nm, and
optionally in the range from 7.2.times.10.sup.4 Nm to
4.2.times.10.sup.5 Nm.
[0335] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torque
transmitted through the gearbox support at maximum take-off
conditions may be in the range from 7.0.times.10.sup.4 Nm to
1.9.times.10.sup.5 Nm.
[0336] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, torque transmitted
through the gearbox support at maximum take-off conditions may be
in the range from 1.8.times.10.sup.5 Nm to 4.5.times.10.sup.5
Nm.
[0337] The values in the paragraphs above may apply to any gearbox
configuration (i.e. star or planetary, or other gearbox
arrangements).
[0338] In various embodiments, for example in which the gearbox is
in a star configuration, the torque transmitted through the gearbox
support at maximum take-off conditions may be greater than or equal
to 1.10.times.10.sup.5 Nm, and optionally greater than or equal to
1.3.times.10.sup.5 Nm.
[0339] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm and/or in which the
gearbox is in a star configuration, the torque transmitted through
the gearbox support at maximum take-off conditions may be greater
than or equal to 1.2.times.10.sup.5 Nm or greater than or equal to
1.4.times.10.sup.5 Nm. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm and/or in which the gearbox is in a star configuration, the
torque transmitted through the gearbox support at maximum take-off
conditions may be greater than or equal to 3.0.times.10.sup.5 Nm or
is greater than or equal to 3.4.times.10.sup.5 Nm.
[0340] In various embodiments, for example in which the gearbox is
in a star configuration, the torque transmitted through the gearbox
support at maximum take-off conditions may be in the range from
1.10.times.10.sup.5 Nm to 5.00.times.10.sup.5 Nm, and optionally in
the range from 1.3.times.10.sup.5 Nm to 4.2.times.10.sup.5 Nm.
[0341] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm and/or in which the
gearbox is in a star configuration, the torque transmitted through
the gearbox support at maximum take-off conditions may be in the
range from 1.2.times.10.sup.5 Nm to 1.9.times.10.sup.5 Nm and
optionally in the range from 1.4.times.10.sup.5 Nm to
1.8.times.10.sup.5 Nm (and may be equal to 1.6.times.10.sup.5
Nm).
[0342] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm and/or the gearbox
is in a star configuration, torque transmitted through the gearbox
support at maximum take-off conditions may be in the range from
3.0.times.10.sup.5 Nm to 4.5.times.10.sup.5 Nm and optionally in
the range from 3.4.times.10.sup.5 Nm to 4.2.times.10.sup.5 Nm (and
may be equal to 3.8.times.10.sup.5 Nm).
[0343] In various embodiments, for example in which the gearbox is
in a planetary configuration, the torque transmitted through the
gearbox support at maximum take-off conditions may be greater than
or equal to 6.00.times.10.sup.4 Nm, and optionally greater than or
equal to 7.2.times.10.sup.4 Nm.
[0344] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm and/or in which the
gearbox is in a planetary configuration, the torque transmitted
through the gearbox support at maximum take-off conditions may be
greater than or equal to 7.0.times.10.sup.4 Nm. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm and/or in which the gearbox is
in a planetary configuration, the torque transmitted through the
gearbox support at maximum take-off conditions may be greater than
or equal to 1.8.times.10.sup.5 Nm.
[0345] In various embodiments, for example in which the gearbox is
in a planetary configuration, the torque transmitted through the
gearbox support at maximum take-off conditions may be in the range
from 6.00.times.10.sup.4 Nm to 3.00.times.10.sup.5 Nm, and
optionally in the range from 7.2.times.10.sup.4 Nm to
2.6.times.10.sup.5 Nm.
[0346] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm and/or in which the
gearbox is in a planetary configuration, the torque transmitted
through the gearbox support at maximum take-off conditions may be
in the range from 7.0.times.10.sup.4 Nm to 1.1.times.10.sup.5
Nm.
[0347] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm and/or in which the
gearbox is in a planetary configuration, torque transmitted through
the gearbox support at maximum take-off conditions may be in the
range from 1.8.times.10.sup.5 Nm to 2.5.times.10.sup.5 Nm.
Torque Transmitted Through the Gearbox Support at Cruise
Conditions:
[0348] In various embodiments, the torque transmitted through the
gearbox support at cruise conditions may be greater than or equal
to 2.00.times.10.sup.4 Nm, and optionally greater than or equal to
2.2.times.10.sup.4 Nm.
[0349] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torque
transmitted through the gearbox support at cruise conditions may be
greater than or equal to 2.2.times.10.sup.4 Nm. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the torque transmitted through
the gearbox support at cruise conditions may be greater than or
equal to 6.8.times.10.sup.4 Nm.
[0350] In various embodiments, the torque transmitted through the
gearbox support at cruise conditions may be in the range from
2.00.times.10.sup.4 Nm to 2.00.times.10.sup.5 Nm, and optionally in
the range from 2.2.times.10.sup.4 Nm to 1.7.times.10.sup.5 Nm.
[0351] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torque
transmitted through the gearbox support at cruise conditions may be
in the range from 2.2.times.10.sup.4 Nm to 6.6.times.10.sup.4
Nm.
[0352] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the torque
transmitted through the gearbox support at cruise conditions may be
in the range from 6.8.times.10.sup.4 Nm to 1.8.times.10.sup.5
Nm.
[0353] The values in the paragraphs above may apply to any gearbox
configuration (i.e. star or planetary, or other gearbox
arrangements).
[0354] In various embodiments, for example in which the gearbox is
in a star configuration, the torque transmitted through the gearbox
support at cruise conditions may be greater than or equal to
4.00.times.10.sup.4 Nm, and optionally greater than or equal to
4.8.times.10.sup.4 Nm.
[0355] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm and/or in which the
gearbox is in a star configuration, the torque transmitted through
the gearbox support at cruise conditions may be greater than or
equal to 4.5.times.10.sup.4 Nm or greater than or equal to
5.0.times.10.sup.4 Nm. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm and/or the gearbox is in a star configuration, the torque
transmitted through the gearbox support at cruise conditions may be
greater than or equal to 1.2.times.10.sup.5 Nm.
[0356] In various embodiments, for example in which the gearbox is
in a star configuration, the torque transmitted through the gearbox
support at cruise conditions may be in the range from
4.00.times.10.sup.4 Nm to 2.00.times.10.sup.5 Nm, and optionally in
the range from 4.8.times.10.sup.4 Nm to 1.7.times.10.sup.5 Nm.
[0357] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm and/or the gearbox
is in a star configuration, the torque transmitted through the
gearbox support at cruise conditions may be in the range from
4.5.times.10.sup.4 Nm to 6.6.times.10.sup.4 Nm and optionally in
the range from 5.0.times.10.sup.4 Nm to 6.0.times.10.sup.4 Nm (and
may be equal to 5.5.times.10.sup.4 Nm).
[0358] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm and/or the gearbox
is in a star configuration, torque transmitted through the gearbox
support at cruise conditions may be in the range from
1.2.times.10.sup.5 Nm to 1.8.times.10.sup.5 Nm and optionally in
the range from 1.2.times.10.sup.5 Nm to 1.8.times.10.sup.5 Nm (and
may be equal to 1.5.times.10.sup.5 Nm).
[0359] In various embodiments, for example in which the gearbox is
in a planetary configuration, the torque transmitted through the
gearbox support at cruise conditions may be greater than or equal
to 2.00.times.10.sup.4 Nm, and optionally greater than or equal to
2.2.times.10.sup.4 Nm.
[0360] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm and/or in which the
gearbox is in a planetary configuration, the torque transmitted
through the gearbox support at cruise conditions may be greater
than or equal to 2.2.times.10.sup.4 Nm. In some embodiments, for
example in embodiments in which the fan diameter is in the range
from 330 to 380 cm and/or the gearbox is in a planetary
configuration, the torque transmitted through the gearbox support
at cruise conditions may be greater than or equal to
6.8.times.10.sup.4 Nm.
[0361] In various embodiments, for example in which the gearbox is
in a planetary configuration, the torque transmitted through the
gearbox support at cruise conditions may be in the range from
2.00.times.10.sup.4 Nm to 1.30.times.10.sup.5 Nm, and optionally in
the range from 2.2.times.10.sup.4 Nm to 1.1.times.10.sup.5 Nm.
[0362] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm and/or the gearbox
is in a planetary configuration, the torque transmitted through the
gearbox support at cruise conditions may be in the range from
2.2.times.10.sup.4 Nm to 3.3.times.10.sup.4 Nm.
[0363] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm and/or the gearbox
is in a planetary configuration, torque transmitted through the
gearbox support at cruise conditions may be in the range from
6.8.times.10.sup.4 Nm to 1.1.times.10.sup.5 Nm.
Gearbox Support Torsional Shear Stress
[0364] The gearbox support 40 has a torsional shear stress that
also represents the resistance of the gearbox support 40 to a
torque applied by the gearbox 30. The torsional shear stress is as
defined elsewhere herein.
[0365] In various embodiments, the torsional shear stress of the
gearbox support at maximum take-off conditions may be less than or
equal to 4.90.times.10.sup.8 N/m.sup.2, and optionally less than or
equal to 3.5.times.10.sup.8 N/m.sup.2.
[0366] In various embodiments, the torsional shear stress of the
gearbox support at maximum take-off conditions may be in the range
from 1.40.times.10.sup.8 N/m.sup.2 to 4.90.times.10.sup.8
N/m.sup.2, and optionally in the range from 2.0.times.10.sup.8
N/m.sup.2 to 3.5.times.10.sup.8 N/m.sup.2.
[0367] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm or in the range
from 330 to 380 cm, the torsional shear stress of the gearbox
support at maximum take-off conditions may be in the range from
2.3.times.10.sup.8 N/m.sup.2 to 3.7.times.10.sup.8 N/m.sup.2 (and
may be equal to 2.5.times.10.sup.8 N/m.sup.2).
[0368] In various embodiments, the torsional shear stress of the
gearbox support at cruise conditions may be greater than or equal
to 7.00.times.10.sup.7 N/m.sup.2, and optionally greater than or
equal to 8.2.times.10.sup.7 N/m.sup.2.
[0369] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torsional
shear stress of the gearbox support at cruise conditions may be
greater than or equal to 8.0.times.10.sup.7 N/m.sup.2. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the torsional shear stress of
the gearbox support at cruise conditions may be greater than or
equal to 8.5.times.10.sup.7 N/m.sup.2 or greater than or equal to
9.0.times.10.sup.7 N/m.sup.2.
[0370] In various embodiments, the torsional shear stress of the
gearbox support at cruise conditions may be in the range from
7.00.times.10.sup.7 N/m.sup.2 to 1.90.times.10.sup.8 N/m.sup.2, and
optionally in the range from 8.2.times.10.sup.7 N/m.sup.2 to
1.5.times.10.sup.8 N/m.sup.2.
[0371] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torsional
shear stress of the gearbox support at cruise conditions may be in
the range from 8.0.times.10.sup.7 N/m.sup.2 to 1.5.times.10.sup.8
N/m.sup.2 and optionally in the range from 8.0.times.10.sup.7
N/m.sup.2 to 9.2.times.10.sup.7 N/m.sup.2 (and may be equal to
8.6.times.10.sup.7 N/m.sup.2).
[0372] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the torsional
shear stress of the gearbox support at cruise conditions may be in
the range from 8.5.times.10.sup.7 N/m.sup.2 to 1.9.times.10.sup.8
N/m.sup.2 and optionally in the range from 9.0.times.10.sup.7
Nm.sup.2 to 1.2.times.10.sup.8 N/m.sup.2 (and may be equal to
9.6.times.10.sup.7 N/m.sup.2).
[0373] The maximum take-off conditions and cruise conditions
referred to in this section may be as defined elsewhere herein.
[0374] The shear stress may be measured through the highest loaded
plane section of the component (i.e. the gearbox support). It
excludes the effects of stress concentrations at small radii and
holes. The skilled person will understand that for the shear stress
of the gearbox support defined herein since the predominant load is
a torque, the effective radius also affects the measurement of the
shear as well as the area.
Fan Moment of Inertia
[0375] The fan 23 has a moment of inertial I.sub.F. The moment of
inertia of the fan is measured based on the total mass of the rotor
forming the fan, i.e. including the total mass of the plurality of
fan blades, the fan hub and any support arm or other linkages
provided to connect the fan to the fan shaft. The moment of
inertial therefore includes all rotating components apart from the
fan shaft. The moment of inertia is the mass moment of inertia or
rotational inertia of the fan with respect to rotation around the
principal rotational axis 9 of the engine. Rotation of the fan will
cause a gyroscopic effect meaning that the fan shaft will tend to
maintain a steady direction of its axis of rotation. During
maneuvering of the aircraft to which the gas turbine engine is
mounted the orientation of the axis of rotation of the fan shaft
will however change. The gyroscopic effect will result in a
reaction force at the fan shaft mounting structure to resist the
tendency of the fan shaft to maintain its orientation. The moment
of inertia of the fan will have an effect on the magnitude of the
gyroscopic effect produced, and so has an impact on the design of
the fan shaft and the fan shaft mounting structure as is discussed
elsewhere herein.
[0376] In various embodiments, the moment of inertia of the fan may
be greater than or equal to 7.40.times.10.sup.7 kgm.sup.2, and
optionally greater than or equal to 8.3.times.10.sup.7
kgm.sup.2.
[0377] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the moment of
inertia of the fan may be greater than or equal to
7.4.times.10.sup.7 kgm.sup.2 or 8.6.times.10.sup.7 kgm.sup.2. In
some embodiments, for example in embodiments in which the fan
diameter is in the range from 330 to 380 cm, the moment of inertia
of the fan may be greater than or equal to 3.0.times.10.sup.8
kgm.sup.2 or 4.0.times.10.sup.8 kgm.sup.2.
[0378] In various embodiments, the moment of inertia of the fan may
be in the range from 7.40.times.10.sup.7 kgm.sup.2 to
9.00.times.10.sup.8 kgm.sup.2, and optionally in the range from
8.3.times.10.sup.7 kgm.sup.2 to 6.5.times.10.sup.8 kgm.sup.2.
[0379] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the moment of
inertia of the fan may be in the range from 7.4.times.10.sup.7
kgm.sup.2 to 1.5.times.10.sup.8 kgm.sup.2 and optionally in the
range from 8.6.times.10.sup.7 kgm.sup.2 to 9.6.times.10.sup.7
kgm.sup.2 (and may be equal to 9.1.times.10.sup.7 kgm.sup.2).
[0380] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the moment of
inertia of the fan may be in the range from 3.0.times.10.sup.8
kgm.sup.2 to 9.0.times.10.sup.8 kgm.sup.2 and optionally in the
range from 4.0.times.10.sup.8 kgm.sup.2 to 5.0.times.10.sup.8
kgm.sup.2 (and may be equal to 4.5.times.10.sup.8 kgm.sup.2).
Relative Fan and Gearbox Positions
[0381] Referring again to FIG. 7, a fan-gearbox axial distance 110
is defined as the axial distance between the axial position P of
the gearbox output position (also labelled X in FIGS. 15 to 18) and
the fan axial centerline Q. The fan axial centerline is defined as
the axial midpoint of the fan blades forming the fan. The gearbox
output position is defined as the point of connection between the
fan shaft 36 and the gearbox. This may be defined differently for
different types of gearbox as is described below.
[0382] In various embodiments, the fan-gearbox axial distance may
be greater than or equal to 0.35 m, and optionally greater than or
equal to 0.37 m.
[0383] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan-gearbox
axial distance may be greater than or equal to 0.38 m or greater
than or equal to 0.40 m. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the fan-gearbox axial distance may be greater than or equal
to 0.48 m or greater than or equal to 0.50 m.
[0384] In various embodiments, the fan-gearbox axial distance may
be in the range from 0.35 m to 0.8 m, and optionally in the range
from 0.37 m to 0.75 m.
[0385] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan-gearbox
axial distance may be in the range from 0.38 m to 0.65 m and
optionally in the range from 0.40 m to 0.44 m (and may be equal to
0.42 m).
[0386] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the fan-gearbox
axial distance may be in the range from 0.48 m to 0.8 m and
optionally in the range from 0.50 m to 0.66 m (and may be equal to
0.58 m).
[0387] The fan shaft 36 is defined as the torque transfer component
that extends from the output of the gearbox to the fan input. It
therefore includes part or all of any gearbox output shaft and fan
input shaft that may be provided between those points. For the
purposes of defining the stiffness of the fan shaft 36 it is
considered to extend between a fan input position and a gearbox
output position, and includes all of the torque transfer components
between those points. It does not therefore include any components
of the gearbox (e.g. the planet carrier or connecting plate coupled
to it) which transmit discrete forces, rather than the fan shaft
torque. The gearbox output position therefore may be defined as the
point of connection between the fan shaft 36 and the gearbox 30.
The fan input position may be defined as the point of connection
between the fan shaft 36 and the fan.
[0388] Referring to FIG. 12, where the gearbox is in a star
configuration, the gearbox output position is defined as the point
of connection 702 between the ring gear 38 and the fan shaft 36.
More specifically, it is the point of connection to the annulus of
the ring gear 38 (with any connection component extending from the
outer surface of the annulus being considered to be part of the
ring gear). Where the point of connection is formed by an interface
extending in a direction having an axial component, the point of
connection is considered to be the axial centerline of that
interface as illustrated in FIG. 7.
[0389] The fan shaft 36 includes all torque transmitting components
up to the point of connection 702 with the ring gear 38. It
therefore includes any flexible portions or linkages 704 of the fan
shaft 36 that may be provided, and any connection(s) 706 (e.g.
spline connections) between them.
[0390] Where the gearbox 30 is in a planetary configuration, the
gearbox output position is again defined as the point of connection
between the fan shaft 36 and the gearbox 30. An example of this is
illustrated in FIG. 13, which shows a carrier comprising a forward
plate 34a and rearward plate 34b, with a plurality of pins 33
extending between them and on which the planet gears are mounted.
The fan shaft 36 is connected to the forward plate 34a via a spline
connection 708. In an arrangement such as this, the gearbox output
position is taken as any point on the interface between the fan
shaft 36 and the forward plate 34a. The forward plate 34a is
considered to transmit discrete forces, rather than a single
torque, and so is taken to be part of the gearbox 30 rather than
the fan shaft.
[0391] FIG. 13 shows only one example of a type of connection
between the fan shaft and planet carrier 34. In embodiments having
different connection arrangements, the gearbox output position is
still taken to be at the interface between components transmitting
a torque (i.e. that are part of the fan shaft) and those
transmitting discrete forces (e.g. that are part of the gearbox).
The spline connection 708 is only one example of a connection that
may be formed between the fan shaft and gearbox (i.e. between the
fan shaft and the forward plate 34b in the presently described
arrangement). In other embodiments, the interface which forms the
gearbox output position may be formed by, for example, a curvic
connection, a bolted joint or other toothed or mechanically dogged
arrangement.
[0392] The fan input position is defined as a point on the fan
shaft at the axial midpoint of the interface between the fan and
the fan shaft. In the presently described arrangement, the fan 23
comprises a support arm 23a (as can be seen, for example, in FIG.
7) arranged to connect the fan 23 to the fan shaft 36. The support
arm 23a is connected to the fan shaft by a spline coupling 36c that
extends along the length of a portion of the fan shaft 36. The fan
input position is defined as the axial midpoint of the spline
coupling. The spline coupling shown is only one example of a
coupling that may form the interface between the fan and fan shaft.
In other embodiments, for example, a curvic connection, a bolted
joint or other toothed or mechanically dogged arrangement may be
used. For example, a flange coupling may be provided between the
support arm 23a and the fan shaft 36. In such an embodiment, the
support arm can be connected at the rear of the fan hub. In this
embodiment the fan input position is the axial midpoint of the
flange coupling. FIG. 14 illustrates an arrangement in which an
alternative coupling is provided between the fan 23 and the fan
shaft 36. Similarly to FIG. 7, the fan 23 is coupled to the fan
shaft 36 via a support arm 23a. In this arrangement however, a
flange coupling 36d is provided between the support arm 23a and the
fan shaft 36. In this embodiment, the support arm 23a is connected
at the rear of the fan hub. The flange coupling 36d may be a curvic
coupling. In other embodiments, other forms of flange coupling may
be provided. In the embodiment of FIG. 14, the fan input position
is the axial midpoint (labelled Y) of the flange coupling.
[0393] The fan shaft 36 has a degree of flexibility characterized
by its radial bending stiffness and tilt stiffness.
Fan Shaft End Stiffness at Gearbox Output:
[0394] The stiffness of the fan shaft where it couples to the
gearbox 30 is defined with reference to FIGS. 15 to 18.
[0395] The radial bending stiffness of the fan shaft 36 at the
output of the gearbox 30 is measured by applying a force F.sub.2 to
the fan shaft at the gearbox output position defined above
(illustrated in FIG. 15). The fan shaft 36 is treated as a free
body, and is held fixed at the position of all of bearing positions
at which it is supported i.e. the first and second bearings 506a,
506b in the arrangement of FIG. 15. As a result of force F.sub.2
the fan shaft 36 deforms so that the gearbox output position is
displaced by a distance of .delta..sub.2 (as illustrated in FIG.
16). The radial bending stiffness of the fan shaft 36 at the output
of the gearbox is then given by F.sub.2/.delta..sub.2.
[0396] In various embodiments, the radial bending stiffness of the
fan shaft at the output of the gearbox may be greater than or equal
to 4.00.times.10.sup.6 N/m, and optionally greater than or equal to
3.7.times.10.sup.7 N/m.
[0397] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the radial bending
stiffness of the fan shaft at the output of the gearbox may be
greater than or equal to 3.7.times.10.sup.7 N/m. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the radial bending stiffness of
the fan shaft at the output of the gearbox may be greater than or
equal to 3.9.times.10.sup.7 N/m or greater than or equal to
5.0.times.10.sup.7 N/m.
[0398] In various embodiments, the radial bending stiffness of the
fan shaft at the output of the gearbox may be in the range from
4.00.times.10.sup.6 N/m to 1.5.times.10.sup.9 N/m, and optionally
in the range from 3.7.times.10.sup.7 N/m to 1.0.times.10.sup.9
N/m.
[0399] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the radial bending
stiffness of the fan shaft at the output of the gearbox may be in
the range from 3.7.times.10.sup.7 N/m to 5.0.times.10.sup.8 N/m and
optionally in the range from 3.7.times.10.sup.7 N/m to
4.3.times.10.sup.7 N/m (and may be equal to 4.0.times.10.sup.7
N/m).
[0400] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the radial bending
stiffness of the fan shaft at the output of the gearbox may be in
the range from 3.9.times.10.sup.7 N/m to 1.5.times.10.sup.9 N/m,
and optionally in the range from 5.0.times.10.sup.7 N/m to
9.0.times.10.sup.7 N/m (and may be equal to 7.0.times.10.sup.7
N/m).
[0401] The tilt stiffness of the fan shaft 36 at the output of the
gearbox 30 is measured by applying a moment M.sub.2 to the fan
shaft at the gearbox output position defined above (illustrated in
FIG. 17). The fan shaft 36 is again treated as a free body, and is
held fixed at the position of all of the bearing positions at which
it is supported i.e. the first and second bearings 506a, 506b in
the arrangement of FIG. 15. As a result of moment M.sub.2 the fan
shaft 36 deforms so that the gearbox output position is displaced
by an angular displacement of .theta..sub.2 as illustrated in FIG.
17. The tilt stiffness of the fan shaft 36 at the output of the
gearbox is then given by M.sub.2/.theta..sub.2.
[0402] In various embodiments, the tilt stiffness of the fan shaft
at the output of the gearbox may be greater than or equal to
7.00.times.10.sup.4 Nm/rad, and optionally greater than or equal to
9.5.times.10.sup.5 Nm/rad.
[0403] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the tilt stiffness
of the fan shaft at the output of the gearbox may be greater than
or equal to 9.5.times.10.sup.5, and optionally may be greater than
or equal to 9.5.times.10.sup.5 Nm/rad. In some embodiments, for
example in embodiments in which the fan diameter is in the range
from 330 to 380 cm, the tilt stiffness of the fan shaft output of
the gearbox may be greater than or equal to 1.1.times.10.sup.6
Nm/rad and optionally may be greater than or equal to
2.0.times.10.sup.6 Nm/rad.
[0404] In various embodiments, the tilt stiffness of the fan shaft
at the output of the gearbox may be in the range from
7.00.times.10.sup.4 Nm/rad to 7.00.times.10.sup.7 Nm/rad, and
optionally in the range from 9.5.times.10.sup.5 Nm/rad to
3.5.times.10.sup.7 Nm/rad.
[0405] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the tilt stiffness
of the output of the gearbox may be in the range from
9.5.times.10.sup.5 Nm/rad to 2.0.times.10.sup.7 Nm/rad and
optionally in the range from 9.5.times.10.sup.5 Nm/rad to
2.4.times.10.sup.6 Nm/rad (and may be equal to 1.2.times.10.sup.6
Nm/rad).
[0406] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the tilt stiffness
of the fan shaft at the output of the gearbox may be in the range
from 1.1.times.10.sup.6 Nm/rad to 7.0.times.10.sup.7 Nm/rad, and
optionally in the range from 2.0.times.10.sup.6 Nm/rad to
5.2.times.10.sup.6 Nm/rad (and may be equal to 3.6.times.10.sup.6
Nm/rad).
Torque Transmission by the Gearbox
[0407] The gearbox provides a torque conversion between the toque
at its input (i.e. the core shaft) and its output (i.e. the fan
shaft). The torque transmitted through the gearbox is defined as
the torque at the output position of the gearbox (the output
position being as defined elsewhere herein). The torque transmitted
through the gearbox varies over the operating cycle of the
engine.
[0408] In various embodiments, the torque transmitted through the
gearbox at maximum take-off conditions may be greater than or equal
to 7.00.times.10.sup.4 Nm, and optionally greater than or equal to
1.0.times.10.sup.5 Nm.
[0409] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torque
transmitted through the gearbox at maximum take-off conditions may
be greater than or equal to 1.1.times.10.sup.5 Nm. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the torque transmitted through
the gearbox at maximum take-off conditions may be greater than or
equal to 1.5.times.10.sup.5 Nm or greater than or equal to
2.0.times.10.sup.5 Nm.
[0410] In various embodiments, the torque transmitted through the
gearbox at maximum take-off conditions may be in the range from
7.00.times.10.sup.4 Nm to 5.00.times.10.sup.5 Nm, and optionally in
the range from 1.0.times.10.sup.5 Nm to 3.5.times.10.sup.5 Nm.
[0411] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torque
transmitted through the gearbox at maximum take-off conditions may
be in the range from 1.1.times.10.sup.5 Nm to 1.5.times.10.sup.5 Nm
and optionally in the range from 1.1.times.10.sup.5 Nm to
1.3.times.10.sup.5 Nm (and may be equal to 1.2.times.10.sup.5
Nm).
[0412] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the torque
transmitted through the gearbox at maximum take-off conditions may
be in the range from 1.5.times.10.sup.5 Nm to 5.0.times.10.sup.5 Nm
and optionally in the range from 2.0.times.10.sup.5 Nm to
3.8.times.10.sup.5 Nm (and may be equal to 2.9.times.10.sup.5
Nm).
[0413] In various embodiments, the torque transmitted through the
gearbox at cruise conditions may be greater than or equal to
2.30.times.10.sup.4 Nm, and optionally greater than or equal to
3.1.times.10.sup.4 Nm.
[0414] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torque
transmitted through the gearbox at cruise conditions may be greater
than or equal to 3.2.times.10.sup.4 Nm or greater than or equal to
3.8.times.10.sup.4 Nm. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the torque transmitted through the gearbox at cruise
conditions may be greater than or equal to 7.3.times.10.sup.4 Nm or
greater than or equal to 9.8.times.10.sup.4 Nm.
[0415] In various embodiments, the torque transmitted through the
gearbox at cruise conditions may be in the range from
2.30.times.10.sup.4 Nm to 1.80.times.10.sup.5 Nm, and optionally in
the range from 3.1.times.10.sup.4 Nm to 1.5.times.10.sup.5 Nm.
[0416] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the torque
transmitted through the gearbox at cruise conditions may be in the
range from 3.2.times.10.sup.4 Nm to 7.2.times.10.sup.4 Nm and
optionally in the range from 3.8.times.10.sup.4 Nm to
4.6.times.10.sup.4 Nm (and may be equal to 4.2.times.10.sup.4
Nm).
[0417] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, torque transmitted
through the gearbox at cruise conditions may be in the range from
7.3.times.10.sup.4 Nm to 1.8.times.10.sup.5 Nm and optionally in
the range from 9.8.times.10.sup.4 Nm to 1.4.times.10.sup.5 Nm (and
may be equal to 1.1.times.10.sup.5 Nm).
Gearbox CSA
[0418] The cross sectional area (CSA) of the gearbox is defined as
the area of the pitch circle of the ring gear. The pitch circle of
a gear is an imaginary circle that rolls without slipping with the
pitch circle of any other gear with which the first gear is meshed.
The pitch circle passes through the points where the teeth of two
gears meet as the meshed gears rotate the pitch circle of a gear
generally passes through a mid-point of the length of the teeth of
the gear. The CSA of the gearbox can be found by measuring the
pitch circle diameter (PCD) of the gear. The PCD can be roughly
estimated by taking the average of the diameter between tips of the
gear teeth and the diameter between bases of the gear teeth.
[0419] In various embodiments, the CSA of the gearbox may be
greater than or equal to 2.4.times.10.sup.0.1m.sup.2 and optionally
greater than or equal to 2.6.times.10.sup.0.1m.sup.2.
[0420] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the CSA of the
gearbox may be greater than or equal to 2.4.times.10.sup.0.1m.sup.2
or greater than or equal to 2.5.times.10.sup.-1m.sup.2.
[0421] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the CSA of the
gearbox may be greater than or equal to 4.5.times.10.sup.0.1m.sup.2
or greater than or equal to 5.5.times.10.sup.-1m.sup.2.
[0422] In various embodiments, the CSA of the gearbox may be in the
range from 2.4.times.10.sup.-1m.sup.2 to 1.10 m.sup.2, and
optionally in the range from 2.6.times.10.sup.0.1m.sup.2 to
9.0.times.10.sup.0.1m.sup.2.
[0423] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the CSA of the
gearbox may be in the range from 2.4.times.10.sup.0.1m.sup.2 to
5.0.times.10.sup.0.1m.sup.2 and optionally in the range from
2.5.times.10.sup.-1m.sup.2 to 3.4.times.10.sup.-1m.sup.2 (and may
be equal to 2.9.times.10.sup.-1m.sup.2).
[0424] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the CSA of the
gearbox may be in the range from 4.5.times.10.sup.-1m.sup.2 to
1.1m.sup.2, and optionally in the range from
5.5.times.10.sup.-1m.sup.2 to 6.4.times.10.sup.-1m.sup.2 (and may
be equal to 5.9.times.10.sup.-1m.sup.2).
Angle Between Adjacent Planet Gears
[0425] A planet gear spacing angle in radians (.beta.) is defined
as 2.pi./N, where N is the number of planet gears 32 provided in
the gearbox. The planet gear spacing angle is illustrated in FIG.
3. The planet gearbox spacing angle corresponds to the average
angle (in rad) between all the adjacent pairs of planet gears.
[0426] In various embodiments, the planet gear spacing angle ((3)
may be greater than or equal to 9.0.times.10.sup.1 rad, and
optionally in the range between 9.0.times.10.sup.-1 rad to 2.1 rad,
and further optionally in the range between 1.1 rad and 1.3 rad
(and may be equal to 1.26 rad)
Parameter Ratios
[0427] The inventor has discovered that the ratios (and/or
products) of some properties have a considerable impact on the
operation of the gearbox and its inputs/outputs/support structure.
Some or all of the below may apply to any embodiment:
[0428] In various embodiments, a radial bending stiffness to moment
of inertia ratio may be defined as:
the radial bending stiffness of at least one of the fan shaft ( 36
) at the output of the gearbox and the gearbox support ( 40 ) the
moment of inertia of the fan ( 23 ) ##EQU00051##
[0429] In various embodiments, the radial bending stiffness to
moment of inertia ratio may be greater than or equal to
2.5.times.10.sup.-2 Nkg.sup.-1m.sup.-3 (i.e. (N/m)/(kgm.sup.2)) and
optionally greater than or equal to 0.05 Nkg.sup.-1m.sup.-3.
[0430] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the radial bending
stiffness to moment of inertia ratio may be greater than or equal
to 0.05 Nkg.sup.-1m.sup.3. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the radial bending stiffness to moment of inertia ratio may
be greater than or equal to 0.025 Nkg.sup.-1m.sup.-3.
[0431] In various embodiments, the radial bending stiffness to
moment of inertia ratio may be in the range from
2.5.times.10.sup.-2 Nkg.sup.-1 m.sup.-3 to 6.0 Nkg.sup.-1m.sup.3,
and optionally in the range from 0.05 Nkg.sup.-1 m.sup.-3 to 3.0
Nkg.sup.-1m.sup.-3, and further optionally in the range from 0.05
Nkg.sup.-1m.sup.3 to 0.6 Nkg.sup.-1m.sup.3.
[0432] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the radial bending
stiffness to moment of inertia ratio may be in the range from 0.05
Nkg.sup.-31m.sup.3 to 4.0 Nkg.sup.-1m.sup.-3.
[0433] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the radial bending
stiffness to moment of inertia ratio may be in the range from 0.025
Nkg.sup.-1M.sup.3 to 2.0 Nkg.sup.-1m.sup.3.
[0434] In various embodiments, in addition to or alternatively to
the radial bending stiffness to moment of inertia ratio, a product
of the parameters making up that ratio may be defined. This product
(referred to as the radial bending stiffness to moment of inertia
product) may be defined as:
(the radial bending stiffness of at least one of the fan shaft (36)
at the output of the gearbox and the gearbox support
(40)).times.(the moment of inertia of the fan (23))
[0435] In various embodiments, the radial bending stiffness to
moment of inertia product may be greater than or equal to
2.0.times.10.sup.14 Nkgm (i.e. (N/m)(kgm.sup.2)), and optionally
greater than or equal to 4.0.times.10.sup.14 Nkgm, and further
optionally greater than or equal to 2.0.times.10.sup.15 Nkgm.
[0436] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the radial bending
stiffness to moment of inertia product may be greater than or equal
to 1.5.times.10.sup.15 Nkgm. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the radial bending stiffness to moment of inertia product
may be greater than or equal to 2.0.times.10.sup.15 Nkgm.
[0437] In various embodiments, the radial bending stiffness to
moment of inertia product may be in the range from
2.0.times.10.sup.14 Nkgm to 1.4.times.10.sup.18 Nkgm, and
optionally in the range from 4.0.times.10.sup.14 Nkgm to
7.0.times.10.sup.17 Nkgm, and further optionally in the range from
2.0.times.10.sup.15 Nkgm to 7.0.times.10.sup.17 Nkgm.
[0438] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the radial bending
stiffness to moment of inertia product may be in the range from
1.5.times.10.sup.15 Nkgm to 1.3.times.10.sup.17 Nkgm.
[0439] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the radial bending
stiffness to moment of inertia product may be in the range from
2.0.times.10.sup.15 Nkgm to 1.4.times.10.sup.18 Nkgm.
[0440] In various embodiments, a fan shaft radial bending stiffness
to moment of inertia ratio may be defined as:
the radial bending stiffness of the fan shaft ( 36 ) at the output
of the gearbox the moment of inertia of the fan ( 23 )
##EQU00052##
[0441] In various embodiments, the fan shaft radial bending
stiffness to moment of inertia ratio may be greater than or equal
to 2.5.times.10.sup.-2 Nkg.sup.-1M.sup.3 (i.e. (N/m)/(kgm.sup.2)),
and optionally greater than or equal to 0.05 (N/m)/(kgm.sup.2).
[0442] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan shaft
radial bending stiffness to moment of inertia ratio may be greater
than or equal to 0.05 Nkg.sup.-1m.sup.-3 or greater than or equal
to 0.4 Nkg.sup.-1m.sup.3. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the fan shaft radial bending stiffness to moment of inertia
ratio may be greater than or equal to 0.025 Nkg.sup.-1m.sup.-3 or
optionally greater than or equal to 0.06 Nkg.sup.-1m.sup.3.
[0443] In various embodiments, the fan shaft radial bending
stiffness to moment of inertia ratio may be in the range from
2.5.times.10.sup.-2 Nkg.sup.-1M.sup.-3 to 6.0 Nkg.sup.-1m.sup.3,
and optionally in the range from 0.05 Nkg.sup.-1M.sup.3 to 3.0
Nkg.sup.-1m.sup.3, and further optionally in the range from 0.05
Nkg.sup.-1M.sup.3 to 0.6 Nkg.sup.-1m.sup.3.
[0444] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan shaft
radial bending stiffness to moment of inertia ratio may be in the
range from 0.05 Nkg.sup.-1m.sup.3 to 0.6 Nkg.sup.-1m.sup.-3 and
optionally in the range from 0.4 Nkg.sup.-1m.sup.3 to 0.5
Nkg.sup.-1m.sup.3 (and may be equal to 0.44
Nkg.sup.-1m.sup.-3).
[0445] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the fan shaft
radial bending stiffness to moment of inertia ratio may be in the
range from 0.025 Nkg.sup.-1m.sup.3 to 0.6 Nkg.sup.-1m.sup.3, and
optionally in the range from 0.06 Nkg.sup.-1m.sup.3 to 0.26
Nkg.sup.-1m.sup.3 (and may be equal to 0.16 Nkg.sup.-1m.sup.3).
[0446] In various embodiments, in addition to or alternatively to
the fan shaft radial bending stiffness to moment of inertia ratio,
a product of the parameters making up that ratio may be defined.
This product (referred to as the fan shaft radial bending stiffness
to moment of inertia product) may be defined as:
[0447] the radial bending stiffness of the fan shaft (36) at the
output of the gearbox.times.the moment of inertia of the fan
(23)
[0448] In various embodiments, the fan shaft radial bending
stiffness to moment of inertia product may be greater than or equal
to 3.0.times.10.sup.14 Nkgm (i.e. (N/m)(kgm.sup.2)), and optionally
greater than or equal to 6.0.times.10.sup.14 Nkgm, and further
optionally greater than or equal to 2.0.times.10.sup.15 Nkgm.
[0449] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan shaft
radial bending stiffness to moment of inertia product may be
greater than or equal to 2.0.times.10.sup.15 Nkgm. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the fan shaft radial bending
stiffness to moment of inertia product may be greater than or equal
to 2.3.times.10.sup.15 Nkgm.
[0450] In various embodiments, the fan shaft radial bending
stiffness to moment of inertia product may be in the range from
3.0.times.10.sup.14 Nkgm to 1.4.times.10.sup.18 Nkgm, and
optionally in the range from 6.0.times.10.sup.14 Nkgm to
7.0.times.10.sup.17 Nkgm, and further optionally in the range from
2.0.times.10.sup.15 Nkgm to 7.0.times.10.sup.17 Nkgm.
[0451] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan shaft
radial bending stiffness to moment of inertia product may be in the
range from 2.0.times.10.sup.15 Nkgm to 7.5.times.10.sup.16
Nkgm.
[0452] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the fan shaft
radial bending stiffness to moment of inertia product may be in the
range from 2.3.times.10.sup.15 Nkgm to 1.4.times.10.sup.18
Nkgm.
[0453] In various embodiments, a gearbox support radial bending
stiffness to moment of inertia ratio may be defined as
the radial bending stiffness of the gearbox support ( 40 ) the
moment of inertia of the fan ( 23 ) ##EQU00053##
[0454] In various embodiments, the gearbox support radial bending
stiffness to moment of inertia ratio may be greater than or equal
to 3.0.times.10.sup.-2 Nkg.sup.-1m.sup.3 (i.e. (N/m)/(kgm.sup.2)),
and optionally greater than or equal to 0.06
Nkg.sup.-1m.sup.-3.
[0455] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the gearbox
support radial bending stiffness to moment of inertia ratio may be
greater than or equal to 0.06 Nkg.sup.-1M.sup.3 or greater than or
equal to 0.25 Nkg.sup.-1m.sup.3. In some embodiments, for example
in embodiments in which the fan diameter is in the range from 330
to 380 cm, the gearbox support radial bending stiffness to moment
of inertia ratio may be greater than or equal to 0.03
Nkg.sup.-1m.sup.3 or greater than or equal to 0.06
Nkg.sup.-1m.sup.3.
[0456] In various embodiments, the gearbox support radial bending
stiffness to moment of inertia ratio may be in the range from
3.0.times.10.sup.-2 Nkg.sup.-1m.sup.-3 to 4.0 Nkg.sup.-1m.sup.-3,
and optionally in the range from 0.06 Nkg.sup.-1m.sup.-3 to 2.0
Nkg.sup.-1m.sup.3, and further optionally in the range from 0.06
Nkg.sup.-1m.sup.-3 to 0.48 Nkg.sup.-1m.sup.3.
[0457] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the gearbox
support radial bending stiffness to moment of inertia ratio may be
in the range from 0.06 Nkg.sup.-1m.sup.3 to 4.0 Nkg.sup.-1m.sup.3
and optionally in the range from 0.25 Nkg.sup.-1m.sup.-3 to 0.45
Nkg.sup.-1m.sup.-3 (and may be equal to 0.35
Nkg.sup.-1m.sup.-3).
[0458] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the gearbox
support radial bending stiffness to moment of inertia ratio may be
in the range from 0.03 Nkg.sup.-1m.sup.3 to 2.0 Nkg.sup.-1m.sup.-3,
and optionally in the range from 0.06 Nkg.sup.-1m.sup.-3 to 0.6
Nkg.sup.-1m.sup.-3 (and may be equal to 0.1)
Nkg.sup.-1m.sup.-3.
[0459] In various embodiments, in addition to or alternatively to
the gearbox support radial bending stiffness to moment of inertia
ratio, a product of the parameters making up that ratio may be
defined. This product (referred to as the gearbox support radial
bending stiffness to moment of inertia product) may be defined
as:
the radial bending stiffness of the gearbox support (40).times.the
moment of inertia of the fan (23)
[0460] In various embodiments, the gearbox support radial bending
stiffness to moment of inertia product may be greater than or equal
to 2.0.times.10.sup.14 Nkgm (i.e. (N/m)(kgm.sup.2)), and optionally
greater than or equal to 4.0.times.10.sup.14 Nkgm or further
optionally greater than or equal to 2.0.times.10.sup.15 Nkgm.
[0461] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the gearbox
support radial bending stiffness to moment of inertia product may
be greater than or equal to 1.5.times.10.sup.15 Nkgm. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the gearbox support radial
bending stiffness to moment of inertia product may be greater than
or equal to 2.0.times.10.sup.15 Nkgm.
[0462] In various embodiments, the gearbox support radial bending
stiffness to moment of inertia product may be in the range from
2.0.times.10.sup.14 Nkgm to 3.0.times.10.sup.17 Nkgm, and
optionally in the range from 4.0.times.10.sup.14 Nkgm to
1.3.times.10.sup.17 Nkgm, and yet further optionally in the range
from 2.0.times.10.sup.15 Nkgm to 1.3.times.10.sup.17 Nkgm.
[0463] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the gearbox
support radial bending stiffness to moment of inertia product may
be in the range from 1.5.times.10.sup.15 Nkgm to
1.3.times.10.sup.17 Nkgm.
[0464] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the gearbox
support radial bending stiffness to moment of inertia product may
be in the range from 2.0.times.10.sup.15 Nkgm to
3.0.times.10.sup.17 Nkgm.
[0465] In various embodiments, a tilt stiffness to moment of
inertia ratio may be defined as:
the tilt stiffness of at least one of the fan shaft ( 36 ) at the
output of the gearbox and the gearbox support ( 40 ) the moment of
inertia of the fan ( 23 ) ##EQU00054##
[0466] In various embodiments, the tilt stiffness to moment of
inertia ratio may be greater than or equal to 4.0.times.10.sup.-4
Nrad.sup.-1kg.sup.-1m.sup.-1 (i.e. (Nm/rad)/(kgm.sup.2)), and
optionally greater than or equal to 1.0.times.10.sup.-3
Nrad.sup.-1kg.sup.-1m.sup.-1.
[0467] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the tilt stiffness
to moment of inertia ratio may be greater than or equal to
1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the tilt stiffness to moment of
inertia ratio may be greater than or equal to 4.0.times.10.sup.-4
Nrad.sup.-1kg.sup.-1m.sup.-1.
[0468] In various embodiments, the tilt stiffness to moment of
inertia ratio may be in the range from 4.0.times.10.sup.-4
Nrad.sup.-1kg.sup.-1m.sup.-1 to 2.7.times.10.sup.-1
Nrad.sup.-1kg.sup.-1m.sup.-1, and optionally in the range from
1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to 0.1
Nrad.sup.-1kg.sup.-1m.sup.-1, and further optionally in the range
from 1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
1.5.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1.
[0469] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the tilt stiffness
to moment of inertia ratio may be in the range from
1.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
1.45.times.10.sup.-2 Nrad.sup.-1 kg.sup.-1m.sup.-1.
[0470] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the tilt stiffness
to moment of inertia ratio may be in the range from
4.0.times.10.sup.-4 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
3.0.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1.
[0471] In various embodiments, in addition to or alternatively to
the tilt stiffness to moment of inertia ratio, a product of the
parameters making up that ratio may be defined. This product
(referred to as the tilt stiffness to moment of inertia product)
may be defined as:
(the tilt stiffness of at least one of the fan shaft (36) at the
output of the gearbox and the gearbox support (40)).times.(the
moment of inertia of the fan (23))
[0472] In various embodiments, the tilt stiffness to moment of
inertia product may be greater than or equal to 3.0.times.10.sup.12
Nm.sup.3rad.sup.-1 kg (i.e. (Nm/rad)(kgm.sup.2)), and optionally
greater than or equal to 6.0.times.10.sup.12 Nm.sup.3rad.sup.-1 kg,
and further optionally greater than or equal to 2.5.times.10.sup.13
Nm.sup.3rad.sup.-1 kg.
[0473] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the tilt stiffness
to moment of inertia product may be greater than or equal to
2.0.times.10.sup.13 Nm.sup.3rad.sup.-1 kg. In some embodiments, for
example in embodiments in which the fan diameter is in the range
from 330 to 380 cm, the tilt stiffness to moment of inertia product
may be greater than or equal to 2.5.times.10.sup.13
Nm.sup.3rad.sup.-1 kg.
[0474] In various embodiments, the tilt stiffness to moment of
inertia product may be in the range from 3.0.times.10.sup.12
Nm.sup.3rad.sup.-1 kg to 6.0.times.10.sup.16 Nm.sup.3rad.sup.-1 kg,
and optionally in the range from 6.0.times.10.sup.12
Nm.sup.3rad.sup.-1 kg to 3.0.times.10.sup.16 Nm.sup.3rad.sup.-1kg,
and further optionally in the range from 2.5.times.10.sup.13
Nm.sup.3rad.sup.-1 kg to 3.0.times.10.sup.16 Nm.sup.3rad.sup.-1
kg.
[0475] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the tilt stiffness
to moment of inertia product may be in the range from
2.0.times.10.sup.13 Nm.sup.3rad.sup.-1 kg to 4.0.times.10.sup.15
Nm.sup.3rad.sup.-1 kg.
[0476] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the tilt stiffness
to moment of inertia product may be in the range from
2.5.times.10.sup.13 Nm.sup.3rad.sup.-1 kg to 6.0.times.10.sup.16
Nm.sup.3rad.sup.-1 kg.
[0477] In various embodiments, a fan shaft tilt stiffness to moment
of inertia ratio may be defined as:
the tilt stiffness of the fan shaft ( 36 ) at the output of the
gearbox the moment of inertia of the fan ( 23 ) ##EQU00055##
[0478] In various embodiments, the fan shaft tilt stiffness to
moment of inertia ratio may be greater than or equal to
4.0.times.10.sup.-4 Nrad.sup.-1 kg.sup.-1m.sup.-1 (i.e.
(Nm/rad)/(kgm.sup.2)), and optionally greater than or equal to
1.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1.
[0479] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan shaft tilt
stiffness to moment of inertia ratio may be greater than or equal
to 1.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1 or greater
than or equal to 3.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1.
In some embodiments, for example in embodiments in which the fan
diameter is in the range from 330 to 380 cm, the fan shaft tilt
stiffness to moment of inertia ratio may be greater than or equal
to 4.0.times.10 Nrad.sup.-1 kg.sup.-1m.sup.-1 or greater than or
equal to 0.7.times.10.sup.-3Nrad.sup.-1kg.sup.-1m.sup.-1.
[0480] In various embodiments, the fan shaft tilt stiffness to
moment of inertia ratio may be in the range from 4.0.times.10
Nrad.sup.-1kg.sup.-1m.sup.-1 to 0.27 Nrad.sup.-1kg.sup.-1m.sup.-1,
and optionally in the range from 1.0.times.10.sup.-3
Nrad.sup.-1kg.sup.-1m.sup.-1 to 0.1 Nrad.sup.-1kg.sup.-1m.sup.-1,
and further optionally in the range from 1.0.times.10.sup.-3
Nrad.sup.-1kg.sup.-1m.sup.-1 to
1.5.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1.
[0481] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan shaft tilt
stiffness to moment of inertia ratio may be in the range from
1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
1.45.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1 and optionally in
the range from 3.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
2.3.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1 (and may be equal
to 1.3.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1).
[0482] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the fan shaft tilt
stiffness to moment of inertia ratio may be in the range from
4.0.times.10.sup.-4 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
1.4.times.10.sup.-2Nrad.sup.-1 kg.sup.-1m.sup.-1, and optionally in
the range from 7.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1 to
9.0.times.10.sup.-3 Nrad.sup.-1 kg.sup.-1m.sup.-1 (and may be equal
to 8.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1).
[0483] In various embodiments, in addition to or alternatively to
the fan shaft tilt stiffness to moment of inertia ratio, a product
of the parameters making up that ratio may be defined. This product
(referred to as the fan shaft tilt stiffness to moment of inertia
product) may be defined as:
the tilt stiffness of the fan shaft (36) at the output of the
gearbox.times.the moment of inertia of the fan (23)
[0484] In various embodiments, the fan shaft tilt stiffness to
moment of inertia product may be greater than or equal to
5.0.times.10.sup.12 Nm.sup.3rad.sup.-1kg (i.e.
(Nm/rad)(kgm.sup.2)), and optionally greater than or equal to
1.0.times.10.sup.13 Nm.sup.3rad.sup.-1kg or greater than or equal
to 6.0.times.10.sup.13 Nm.sup.3rad.sup.-1kg.
[0485] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan shaft tilt
stiffness to moment of inertia product may be greater than or equal
to 6.0.times.10.sup.13 Nm.sup.3rad.sup.-1kg. In some embodiments,
for example in embodiments in which the fan diameter is in the
range from 330 to 380 cm, the fan shaft tilt stiffness to moment of
inertia product may be greater than or equal to 1.0.times.10.sup.14
Nm.sup.3rad.sup.-1kg.
[0486] In various embodiments, the fan shaft tilt stiffness to
moment of inertia product may be in the range from
5.0.times.10.sup.12 Nm.sup.3rad.sup.-1 kg to 6.0.times.10.sup.16
Nm.sup.3rad.sup.-1 kg, and optionally in the range from
1.0.times.10.sup.13 Nm.sup.3rad.sup.-1 kg to 3.0.times.10.sup.16
Nm.sup.3rad.sup.-1 kg, and further optionally in the range from
6.0.times.10.sup.13 Nm.sup.3rad.sup.-1 kg to 3.0.times.10.sup.16
Nm.sup.3rad.sup.-1 kg.
[0487] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan shaft tilt
stiffness to moment of inertia product may be in the range from
6.0.times.10.sup.13 Nm.sup.3rad.sup.-1kg to 3.0.times.10.sup.15
Nm.sup.3rad.sup.-1kg.
[0488] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the fan shaft tilt
stiffness to moment of inertia product may be in the range from
1.0.times.10.sup.14 Nm.sup.3rad.sup.-1kg to 6.0.times.10.sup.16
Nm.sup.3rad.sup.-1kg.
[0489] In various embodiments, a gearbox support tilt stiffness to
moment of inertia ratio may be defined as:
the tilt stiffness of the gearbox support ( 40 ) the moment of
inertia of the fan ( 23 ) ##EQU00056##
[0490] In various embodiments, the gearbox support tilt stiffness
to moment of inertia ratio may be greater than or equal to
1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 (i.e.
(Nm/rad)/(kgm.sup.2)), and optionally greater than or equal to
2.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1.
[0491] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the gearbox
support tilt stiffness to moment of inertia ratio may be greater
than or equal to 2.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1
or greater than or equal to 5.0.times.10.sup.-3
Nrad.sup.-1kg.sup.-1m.sup.-1. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the gearbox support tilt stiffness to moment of inertia
ratio may be greater than or equal to 1.0.times.10.sup.-3
Nrad.sup.-1kg.sup.-1m.sup.-1 or greater than or equal to
2.6.times.10.sup.-3 m.
[0492] In various embodiments, the gearbox support tilt stiffness
to moment of inertia ratio may be in the range from
1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
7.0.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1, and optionally in
the range from 2.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
3.0.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1, and further
optionally in the range from 2.0.times.10.sup.-3
Nrad.sup.-1kg.sup.-1m.sup.-1 to 7.0.times.10.sup.-3
Nrad.sup.-1kg.sup.-1m.sup.-1.
[0493] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the gearbox
support tilt stiffness to moment of inertia ratio may be in the
range from 2.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
7.2.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 and optionally in
the range from 5.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
7.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 (and may be equal
to 6.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1).
[0494] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the gearbox
support tilt stiffness to moment of inertia ratio may be in the
range from 1.0.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
3.0.times.10.sup.-2Nrad.sup.-1kg.sup.-1m.sup.-1, and optionally in
the range from 2.6.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 to
4.6.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1 (and may be equal
to 3.6.times.10.sup.-3 Nrad.sup.-1kg.sup.-1m.sup.-1).
[0495] In various embodiments, in addition to or alternatively to
the gearbox support tilt stiffness to moment of inertia ratio, a
product of the parameters making up that ratio may be defined. This
product (referred to as the gearbox support tilt stiffness to
moment of inertia product) may be defined as:
the tilt stiffness of gearbox support (40).times.the moment of
inertia of the fan (23)
[0496] In various embodiments, the gearbox support tilt stiffness
to moment of inertia product may be greater than or equal to
3.0.times.10.sup.12 Nm.sup.3rad.sup.-1kg (i.e.
(Nm/rad)(kgm.sup.2)), and optionally greater than or equal to
6.0.times.10.sup.12 Nm.sup.3rad.sup.-1 kg, and further optionally
greater than or equal to 2.5.times.10.sup.13 Nm.sup.3rad.sup.-1
kg.
[0497] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the gearbox
support tilt stiffness to moment of inertia product may be greater
than or equal to 2.0.times.10.sup.13 Nm.sup.3rad.sup.-1kg. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the gearbox support tilt
stiffness to moment of inertia product may be greater than or equal
to 2.5.times.10.sup.13 Nm.sup.3rad.sup.-1kg.
[0498] In various embodiments, the gearbox support tilt stiffness
to moment of inertia product may be in the range from
3.0.times.10.sup.12 Nm.sup.3rad.sup.-1kg to 9.0.times.10.sup.15
Nm.sup.3rad.sup.-1kg, and optionally in the range from
6.0.times.10.sup.12 Nm.sup.3rad.sup.-1kg to 4.0.times.10.sup.15
Nm.sup.3rad.sup.-1kg, and further optionally in the range from
2.5.times.10.sup.13 Nm.sup.3rad.sup.-1kg to 4.0.times.10.sup.15
Nm.sup.3rad.sup.-1kg.
[0499] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the gearbox
support tilt stiffness to moment of inertia product may be in the
range from 2.0.times.10.sup.13 Nm.sup.3rad.sup.-1kg to
4.0.times.10.sup.15 Nm.sup.3rad.sup.-1kg.
[0500] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the gearbox
support tilt stiffness to moment of inertia product may be in the
range from 2.5.times.10.sup.13 Nm.sup.3rad.sup.-1kg to
9.0.times.10.sup.15 Nm.sup.3rad.sup.-1kg.
[0501] In various embodiments, a product of the fan-gearbox axial
distance multiplied by the moment of inertia of the fan may be
defined.
[0502] In various embodiments, the product of the fan-gearbox axial
distance and the moment of inertia of the fan may be greater than
or equal to 1.9.times.10.sup.7 kgm.sup.3, and optionally greater
than or equal to 2.9.times.10.sup.7 kgm.sup.3.
[0503] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the product of the
fan-gearbox axial distance and the moment of inertia of the fan may
be greater than or equal to 2.0.times.10.sup.7 kgm.sup.3. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 330 to 380 cm, the product of the fan-gearbox
axial distance and the moment of inertia of the fan may be greater
than or equal to 1.2.times.10.sup.8 kgm.sup.3.
[0504] In various embodiments, the product of the fan-gearbox axial
distance and the moment of inertia of the fan may be in the range
from 1.9.times.10.sup.7 kgm.sup.3 to 6.2.times.10.sup.8 kgm.sup.3,
and optionally in the range from 2.9.times.10.sup.7 kgm.sup.3 to
3.9.times.10.sup.8 kgm.sup.3.
[0505] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the product of the
fan-gearbox axial distance and the moment of inertia of the fan may
be in the range from 2.0.times.10.sup.7 kgm.sup.3 to
8.0.times.10.sup.7 kgm.sup.3.
[0506] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the product of the
fan-gearbox axial distance and the moment of inertia of the fan may
be in the range from 1.2.times.10.sup.8 kgm.sup.3 to
6.2.times.10.sup.8 kgm.sup.3.
[0507] In various embodiments, a ratio given by the fan-gearbox
axial distance divided by the moment of inertia of the fan may also
be defined.
[0508] In various embodiments, the fan-gearbox axial distance
divided by the moment of inertia of the fan may be less than or
equal to 8.8.times.10.sup.-9 m/kgm.sup.2 (i.e. kg.sup.-1m.sup.-1),
and optionally less than or equal to 6.2.times.10.sup.-9
m/kgm.sup.2.
[0509] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan-gearbox
axial distance divided by the moment of inertia of the fan may be
less than or equal to 6.5.times.10.sup.-9 m/kgm.sup.2 or less than
or equal to 5.2.times.10.sup.-9 m/kgm.sup.2. In some embodiments,
for example in embodiments in which the fan diameter is in the
range from 330 to 380 cm, the fan-gearbox axial distance divided by
the moment of inertia of the fan may be less than or equal to
2.8.times.10.sup.-9 m/kgm.sup.2 or less than or equal to
1.8.times.10.sup.-9 m/kgm.sup.2.
[0510] In various embodiments, the fan-gearbox axial distance
divided by the moment of inertia of the fan may be in the range
from 5.3.times.10.sup.-10 m/kgm.sup.2 to 8.8.times.10.sup.-9
m/kgm.sup.2, and optionally in the range from
8.8.times.10.sup.-1.degree. m/kgm.sup.2 to 6.2.times.10.sup.-9
m/kgm.sup.2.
[0511] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the fan-gearbox
axial distance divided by the moment of inertia of the fan may be
in the range from 2.1.times.10.sup.-9 m/kgm.sup.2 to
6.5.times.10.sup.-9 m/kgm.sup.2 and optionally in the range from
4.0.times.10.sup.-9 m/kgm.sup.2 to 5.2.times.10.sup.-9 m/kgm.sup.2
(and may be equal to 4.6.times.10.sup.-9 m/kgm.sup.2).
[0512] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the fan-gearbox
axial distance divided by the moment of inertia of the fan may be
in the range from 5.3.times.10.sup.-10 m/kgm.sup.2 to
2.8.times.10.sup.-9 m/kgm.sup.2, and optionally in the range from
8.0.times.10.sup.-1.degree. m/kgm.sup.2 to 1.8.times.10.sup.-9
m/kgm.sup.2 (and may be equal to 1.3.times.10.sup.-9
m/kgm.sup.2).
[0513] In various embodiments, a first gearbox support strength
ratio may be defined as
the torsional strength of the gearbox ( 40 ) radial bending
stiffness of the gearbox support ( 40 ) .times. the cross sectional
area of the gearbox ##EQU00057##
[0514] In various embodiments, the first gearbox support strength
ratio may be greater than or equal to 7.0.times.10.sup.-3, and
optionally greater than or equal to 1.0.times.10.sup.-2 or greater
than or equal to 2.0.times.10.sup.-2.
[0515] In various embodiments, the first gearbox support strength
ratio may be in the range from 7.0.times.10.sup.-3 to
2.5.times.10.sup.-1, and optionally in the range from
1.0.times.10.sup.-2 to 1.0.times.10.sup.-1, in the range from
7.0.times.10.sup.-3 to 2.0.times.10.sup.-2 or in the range from
2.0.times.10.sup.-2 to 2.5.times.10.sup.-1.
[0516] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm or in the range
from 330 to 380 cm, the first gearbox support strength ratio may be
in the range from 1.9.times.10.sup.-2 to 2.9.times.10.sup.-2 (and
may be equal to 2.4.times.10.sup.-2).
[0517] In various embodiments, a second gearbox support strength
may be defined as:
the torsional strength of the gearbox ( 40 ) tilt stiffness of the
gearbox support ( 40 ) .times. the planet gear spacing angle (
.beta. ) ##EQU00058##
[0518] In various embodiments, the second gearbox support strength
ratio may be greater than or equal to 1.0.times.10.sup.-1, and
optionally greater than or equal to 1.5.times.10.sup.-1, greater
than or equal to 1.0.times.10.sup.-1 or greater than or equal to
2.5.times.10.sup.-1.
[0519] In various embodiments, the second gearbox support strength
ratio may be in the range from 1.0.times.10.sup.-1 to 3.5, and
optionally in the range from 1.5.times.10.sup.-1 to 1.7, in the
range from 1.0.times.10.sup.-1 to 2.5.times.10.sup.-1 or in the
range from 2.5.times.10.sup.-1 to 3.5.
[0520] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm or in the range
from 330 to 380 cm, the second gearbox support strength ratio may
be in the range from 2.8.times.10.sup.-1 to 3.8.times.10.sup.-1
(and may be equal to 3.3.times.10.sup.-1).
[0521] In various embodiments, a first gearbox support shear stress
ratio, may be defined as:
the torsional shear stress of the gearbox support ( 40 ) at MTO
radial bending stiffness of the gearbox support ( 40 )
##EQU00059##
[0522] In various embodiments, the first gearbox support shear
stress ratio may be less than or equal to
4.9.times.10.sup.1m.sup.-1, and optionally less than or equal to 20
m.sup.-1.
[0523] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the first gearbox
support shear stress ratio may be less than or equal to 35 m.sup.-1
or less than or equal to 10.0 m.sup.-1. In some embodiments, for
example in embodiments in which the fan diameter is in the range
from 330 to 380 cm, the first gearbox support shear stress ratio
may be less than or equal to 12 m.sup.-1 or less than or equal to
6.0 m.sup.-1.
[0524] In various embodiments, the first gearbox support shear
stress ratio may be in the range from 0.35 m.sup.-1 to
4.9.times.10.sup.1m.sup.-1, and optionally in the range from 0.70
m.sup.-1 to 20 m.sup.-1.
[0525] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the first gearbox
support shear stress ratio may be in the range from 0.70 m.sup.-1
to 35 m.sup.-1 and optionally in the range from 6.0 m.sup.-1 to
10.0 m.sup.-1 (and may be equal to 7.8 m.sup.-1).
[0526] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the first gearbox
support shear stress ratio may be in the range from 0.50 m.sup.-1
to 12 m.sup.-1, and optionally in the range from 3.0 m.sup.-1 to
6.0 m.sup.-1 (and may be equal to 5.4 m.sup.-1).
[0527] In various embodiments a second gearbox support shear stress
ratio may be defined as:
the torsional shear stress of the gearbox support ( 40 ) at maximum
take off conditions tilt stiffness of the gearbox support ( 40 )
##EQU00060##
[0528] In various embodiments, the second gearbox support shear
stress ratio may be less than or equal to 4.1.times.10.sup.3
rad/m.sup.3, and optionally less than or equal to
1.4.times.10.sup.3 rad/m.sup.3.
[0529] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the second gearbox
support shear stress ratio may be less than or equal to
2.9.times.10.sup.3 rad/m.sup.3 or less than or equal to
6.5.times.10.sup.2 rad/m.sup.3. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the second gearbox support shear stress ratio may be less
than or equal to 7.0.times.10.sup.2 rad/m.sup.3 or less than or
equal to 2.5.times.10.sup.2rad/m.sup.3.
[0530] In various embodiments, the second gearbox support shear
stress ratio may be in the range from 6.6 rad/m.sup.3 to
4.1.times.10.sup.3 rad/m.sup.3, and optionally in the range from
1.25.times.10.sup.1 rad/m.sup.3 to 1.4.times.10.sup.3
rad/m.sup.3.
[0531] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the second gearbox
support shear stress ratio may be in the range from
2.9.times.10.sup.1 rad/m.sup.3 to 2.9.times.10.sup.3 rad/m.sup.3
and optionally in the range from 2.5.times.10.sup.2 rad/m.sup.3 to
6.5.times.10.sup.2 rad/m.sup.3 (and may be equal to
4.5.times.10.sup.2 rad/m.sup.3).
[0532] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the second gearbox
support shear stress ratio may be in the range from
1.0.times.10.sup.1 rad/m.sup.3 to 7.0.times.10.sup.2 rad/m.sup.3,
and optionally in the range from 50 rad/m.sup.3 to
2.5.times.10.sup.2 rad/m.sup.3 (and may be equal to
1.5.times.10.sup.2 rad/m.sup.3).
[0533] In various embodiments, a flight cycle ratio may be defined
as:
the torsional shear stress of the gearbox support ( 40 ) at maximum
take off conditions the torsional shear stress of the gearbox
support ( 40 ) at cruise conditions ##EQU00061##
[0534] In various embodiments, the flight cycle ratio may be less
than or equal to 3.20, and optionally less than or equal to 2.95,
optionally less than or equal to 2.9 (or 2.90), optionally less
than or equal to 2.85, or optionally less than or equal to
2.75.
[0535] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the flight cycle
ratio may be less than or equal to 3.2 or less than or equal to
3.0. In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the flight cycle
ratio may be less than or equal to 2.8 or less than or equal to
2.7.
[0536] In various embodiments, the flight cycle ratio may be in the
range from 2.10 to 3.20, and optionally in the range from 2.3 to
2.9.
[0537] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the flight cycle
ratio may be in the range from 2.3 to 3.2 and optionally in the
range from 2.8 to 3.0 (and may be equal to 2.9).
[0538] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the flight cycle
ratio may be in the range from 2.1 to 2.8, and optionally in the
range from 2.3 to 2.7 (and may be equal to 2.6).
[0539] In various embodiments, in addition to or alternatively to
the flight cycle ratio, a product of the parameters making up that
ratio may be defined. This product (referred to as the flight cycle
product) may be defined as:
the torsional shear stress of the gearbox support (40) at maximum
take off conditions.times.the torsional shear stress of the gearbox
support (40) at cruise conditions
[0540] In various embodiments, the flight cycle product may be
greater than or equal to 1.00.times.10.sup.16 (N/m.sup.2).sup.2,
and optionally greater than or equal to 2.05.times.10.sup.16
(N/m.sup.2).sup.2.
[0541] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the flight cycle
product may be greater than or equal to 2.0.times.10.sup.16
(N/m.sup.2).sup.2. In some embodiments, for example in embodiments
in which the fan diameter is in the range from 330 to 380 cm, the
flight cycle product may be greater than or equal to
2.2.times.10.sup.16 (N/m.sup.2).sup.2.
[0542] In various embodiments, the flight cycle product may be in
the range from 1.00.times.10.sup.16 (N/m.sup.2).sup.2 to
7.50.times.10.sup.16 (N/m.sup.2).sup.2, and optionally in the range
from 2.05.times.10.sup.16 (N/m.sup.2).sup.2 to 4.8.times.10.sup.16
(N/m.sup.2).sup.2.
[0543] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the flight cycle
product may be in the range from 2.0.times.10.sup.16
(N/m.sup.2).sup.2 to 4.9.times.10.sup.16 (N/m.sup.2).sup.2.
[0544] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the flight cycle
product may be in the range from 2.2.times.10.sup.16
(N/m.sup.2).sup.2 to 6.2.times.10.sup.16 (N/m.sup.2).sup.2.
[0545] In various embodiments, a first torque transmission ratio
may be defined as:
the torque transmitted through the gearbox ( 30 ) at maximum
takeoff conditions the torque transmitted through the gearbox ( 30
) at cruise conditions ##EQU00062##
[0546] In various embodiments, the first torque transmission ratio
may be less than or equal to 3.2, and optionally less than or equal
to 2.95, optionally less than or equal to 2.9 (or 2.90), optionally
less than or equal to 2.85 or optionally less than or equal to
2.75.
[0547] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the first torque
transmission ratio may be less than or equal to 3.2 or less than or
equal to 3.0. In some embodiments, for example in embodiments in
which the fan diameter is in the range from 330 to 380 cm, the
first torque transmission ratio may be less than or equal to 2.8 or
less than or equal to 2.7.
[0548] In various embodiments, the first torque transmission ratio
may be in the range from 2.1 to 3.2, and optionally in the range
from 2.3 to 2.9.
[0549] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the first torque
transmission ratio may be in the range from 2.3 to 3.2 and
optionally in the range from 2.8 to 3.0 (and may be equal to
2.9).
[0550] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the first torque
transmission ratio may be in the range from 2.1 to 2.8, and
optionally in the range from 2.5 to 2.7 (and may be equal to
2.6).
[0551] In various embodiments, in addition to or alternatively to
the first torque transmission ratio, a product of the parameters
making up that ratio may be defined. This product (referred to as
the first torque transmission product) may be defined as:
the torque transmitted through the gearbox (30) at maximum take off
conditions.times.the torque transmitted through the gearbox (30) at
cruise conditions
[0552] In various embodiments, the first torque transmission
product may be greater than or equal to 2.1.times.10.sup.9
(Nm).sup.2, and optionally greater than or equal to
3.5.times.10.sup.9 (Nm).sup.2.
[0553] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the first torque
transmission product may be greater than or equal to
4.0.times.10.sup.9 (Nm).sup.2. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the first torque transmission product may be greater than
or equal to 9.0.times.10.sup.9 (Nm).sup.2.
[0554] In various embodiments, the first torque transmission
product may be in the range from 2.1.times.10.sup.9 (Nm).sup.2 to
9.0.times.10.sup.10 (Nm).sup.2, and optionally in the range from
3.5.times.10.sup.9 (Nm).sup.2 to 5.2.times.10.sup.10
(Nm).sup.2.
[0555] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the first torque
transmission product may be in the range from 4.0.times.10.sup.9
(Nm).sup.2 to 9.0.times.10.sup.9 (Nm).sup.2.
[0556] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the first torque
transmission product may be in the range from 9.0.times.10.sup.9
(Nm).sup.2 to 9.0.times.10.sup.10 (Nm).sup.2.
[0557] In various embodiments, a second torque transmission ratio
may be defined as:
the torque transmitted through the gearbox support ( 40 ) at
maximum take off conditions the torque transmitted through the
gearbox ( 40 ) at cruise conditions ##EQU00063##
[0558] In various embodiments, the second torque transmission ratio
may be less than or equal to 3.2, and optionally less than or equal
to 2.95, or optionally less than or equal to 2.9 (or 2.90),
optionally less than or equal to 2.85, or optionally less than or
equal to 2.75.
[0559] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the second torque
transmission ratio may be less than or equal to 3.2 or less than or
equal to 3.0. In some embodiments, for example in embodiments in
which the fan diameter is in the range from 330 to 380 cm, the
second torque transmission ratio may be greater than or equal to
2.8 or less than or equal to 2.7.
[0560] In various embodiments, the second torque transmission ratio
may be in the range from 2.1 to 3.2, and optionally in the range
from 2.3 to 2.9.
[0561] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the second torque
transmission ratio may be in the range from 2.3 to 3.2 and
optionally in the range from 2.8 to 3.0 (and may be equal to
2.9).
[0562] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the second torque
transmission ratio may be in the range from 2.1 to 2.8, and
optionally in the range from 2.3 to 2.7 (and may be equal to
2.6).
[0563] In various embodiments, in addition to or alternatively to
the second torque transmission ratio, a product of the parameters
making up that ratio may be defined. This product (referred to as
the second torque transmission product) may be defined as:
the torque transmitted through the gearbox support (40) at maximum
take off conditions.times.the torque transmitted through the
gearbox support (40) at cruise conditions
[0564] In various embodiments, the second torque transmission
product may be greater than or equal to 4.1.times.10.sup.9
(Nm).sup.2, and optionally greater than or equal to
6.1.times.10.sup.9 (Nm).sup.2.
[0565] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the second torque
transmission product may be greater than or equal to
4.1.times.10.sup.9 (Nm).sup.2. In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the second torque transmission product may be greater than
or equal to 1.1.times.10.sup.10 (Nm).sup.2.
[0566] In various embodiments, the second torque transmission
product may be in the range from 4.1.times.10.sup.9 (Nm).sup.2 to
9.0.times.10.sup.10 (Nm).sup.2, and optionally in the range from
6.1.times.10.sup.9 (Nm).sup.2 to 8.0.times.10.sup.10
(Nm).sup.2.
[0567] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 240 to 280 cm, the second torque
transmission product may be in the range from 4.1.times.10.sup.9
(Nm).sup.2 to 3.9.times.10.sup.10.
[0568] In some embodiments, for example in embodiments in which the
fan diameter is in the range from 330 to 380 cm, the second torque
transmission product may be in the range from 1.1.times.10.sup.10
(Nm).sup.2 to 9.0.times.10.sup.10 (Nm).sup.2.
[0569] In the ratios defined above the maximum take-off and cruise
conditions are as defined anywhere herein.
[0570] FIG. 19 illustrates an example aircraft 1000 having a gas
turbine engine 10 attached to each wing 1002a, 1002b thereof. Each
gas turbine engine 10 is attached via a respective pylon 1004a,
1004b. The gas turbines 10 may be that of any embodiment described
herein. The aircraft shown in FIG. 19 is to be understood as the
aircraft for which the gas turbine engine 10 of any embodiment or
aspect disclosed herein has been designed to be attached. The
aircraft 1000 has a cruise condition corresponding to the cruise
conditions defined elsewhere herein and a max take-off condition
corresponding to the MTO conditions defined elsewhere herein.
[0571] The present disclosure also relates to a method 2000 of
operating a gas turbine engine on an aircraft (e.g. the aircraft of
FIG. 19). The method 2000 is illustrated in FIG. 20. The method
2000 comprises operating 2010 the gas turbine engine 10 described
elsewhere herein to provide propulsion for the aircraft to which it
is mounted under maximum take-off conditions. The method further
comprises operating 2020 the gas turbine engine to provide
propulsion during cruise conditions. The gas turbine engine is
operated such that any of the parameters or ratios defined herein
are within the specified ranges. Cruise conditions and max-take off
conditions are as defined elsewhere herein.
[0572] The torque on the core shaft 26 may be referred to as the
input torque, as this is the torque which is input to the gearbox
30. The torque supplied by the turbine 19 to the core shaft (i.e.
the torque on the core shaft) at cruise conditions may be greater
than or equal to 10,000 Nm, and optionally greater than or equal to
11,000 Nm. In some embodiments, for example in embodiments in which
the fan diameter is in the range from 240 to 280 cm, the torque on
the core shaft 26 at cruise conditions may be greater than or equal
to 10,000 or 11,000 Nm (and optionally may be equal to 12,760 Nm).
In some embodiments, for example in embodiments in which the fan
diameter is in the range from 330 to 380 cm, the torque on the core
shaft 26 at cruise conditions may be greater than or equal to
25,000 Nm, and optionally greater than or equal to 30,000 Nm (and
optionally may be equal to 34,000 Nm).
[0573] The torque on the core shaft at cruise conditions may be in
the range from 10,000 to 50,000 Nm, and optionally from 11,000 to
45,000 Nm. In some embodiments, for example in embodiments in which
the fan diameter is in the range from 240 to 280 cm, the torque on
the core shaft 26 at cruise conditions may be in the range from
10,000 to 15,000 Nm, and optionally from 11,000 to 14,000 Nm (and
optionally may be equal to 12,760 Nm). In some embodiments, for
example in embodiments in which the fan diameter is in the range
from 330 to 380 cm, the torque on the core shaft 26 at cruise
conditions may be in the range from 25,000 Nm to 50,000 Nm, and
optionally from 30,000 to 40,000 Nm (and optionally may be equal to
34,000 Nm).
[0574] Under maximum take-off (MTO) conditions, the torque on the
core shaft 26 may be greater than or equal to 28,000 Nm, and
optionally greater than or equal to 30,000 Nm. In some embodiments,
for example in embodiments in which the fan diameter is in the
range from 240 to 280 cm, the torque on the core shaft 26 under MTO
conditions may be greater than or equal to 28,000, and optionally
greater than or equal to 35,000 Nm (and optionally may be equal to
36,300 Nm). In some embodiments, for example in embodiments in
which the fan diameter is in the range from 330 to 380 cm, the
torque on the core shaft 26 under MTO conditions may greater than
or equal to 70,000 Nm, and optionally greater than or equal to
80,000 or 82,000 Nm (and optionally may be equal to 87,000 Nm).
[0575] Under maximum take-off (MTO) conditions, the torque on the
core shaft 26 may be in the range from 28,000 Nm to 135,000 Nm, and
optionally in the range from 30,000 to 1:10,000 Nm. In some
embodiments, for example in embodiments in which the fan diameter
is in the range from 240 to 280 cm, the torque on the core shaft 26
under MTO conditions may be in the range from 28,000 to 50,000 Nm,
and optionally from 35,000 to 38,000 Nm (and optionally may be
equal to 36,300 Nm). In some embodiments, for example in
embodiments in which the fan diameter is in the range from 330 to
380 cm, the torque on the core shaft 26 under MTO conditions may be
in the range from 70,000 Nm to 135,000 Nm, and optionally from
80,000 to 90,000 Nm or 82,000 to 92,000 Nm (and optionally may be
equal to 87,000 Nm).
[0576] Torque has units of [force].times.[distance] and may be
expressed in units of Newton metres (Nm), and is defined in the
usual way as would be understood by the skilled person.
[0577] FIG. 21 illustrates how the stiffnesses defined herein may
be measured. FIG. 21 shows a plot of the displacement 6 resulting
from the application of a load L (e.g. a force, moment or torque)
applied to a component for which the stiffness is being measured.
At levels of load from zero to L.sub.R there is a non-linear region
in which displacement is caused by motion of the component (or
relative motion of separate parts of the component) as it is
loaded, rather than deformation of the component; for example
moving within clearance between parts. At levels of load above
L.sub.S the elastic limit of the component has been exceeded and
the applied load no longer causes elastic deformation--plastic
deformation or failure of the component may occur instead. Between
points R and S the applied load and resulting displacement have a
linear relationship. The stiffnesses defined herein may be
determined by measuring the gradient of the linear region between
points R and S (with the stiffness being the inverse of that
gradient). The gradient may be found for as large a region of the
linear region as possible to increase the accuracy of the
measurement by providing a larger displacement to measure. For
example, the gradient may be found by applying a load equal to or
just greater than L.sub.R and equal to or just less than L.sub.S.
Although the displacement is referred to as .delta. in this
description, the skilled person would appreciate that equivalent
principles would apply to a linear or angular displacement.
[0578] The stiffnesses defined herein, unless otherwise stated, are
for the corresponding component(s) when the engine is off (i.e. at
zero speed/on the bench). The stiffness generally does not vary
over the operating range of the engine; the stiffness at cruise
conditions of the aircraft to which the engine is used (those
cruise conditions being as defined elsewhere herein) may therefore
be the same as for when the engine is not in use. However, where
the stiffness varies over the operating range of the engine, the
stiffnesses defined herein are to be understood as being values for
when the engine is at room temperature and unmoving. Values for
component strength (e.g. the torsional strength of the gearbox
support) as given herein are also at room temperature unless
otherwise stated.
[0579] Anything described herein with reference to a planetary type
gearbox can apply equally to a star type gearbox unless otherwise
stated or where it is apparent that a feature is specific to a
particular gearbox type.
[0580] It will be understood that the invention is not limited to
the embodiments above-described and various modifications and
improvements can be made without departing from the concepts
described herein. Except where mutually exclusive, any of the
features may be employed separately or in combination with any
other features and the disclosure extends to and includes all
combinations and sub-combinations of one or more features described
herein.
* * * * *