US2022298976A1PendingUtilityA1
Aircraft engine
Est. expiryDec 5, 2039(~13.4 yrs left)· nominal 20-yr term from priority
Inventors:Mark Spruce
F05D 2260/40311F02C 7/36F16H 1/28F05D 2240/60B64D 27/10F02K 3/06B64D 27/12F05D 2220/323
68
PatentIndex Score
0
Cited by
0
References
0
Claims
Abstract
A gas turbine engine including: an engine core including a turbine, a compressor, and a core shaft; a fan located upstream of the engine core; a gearbox; and a fan shaft mounting structure including at least two supporting bearings connected to fan shaft. The supporting bearings are located at positions forward of the gearbox. A first bearing separation distance (d1) is defined as an axial distance between the input to the fan and the closest supporting bearing in a rearward direction from the fan. A bearing axial separation (d3) is defined as an axial distance between the closest supporting bearing in the rearward direction from the fan and the closest supporting bearing in a forward direction from the gearbox. A bearing separation ratio defined as:thefirstbearingseparationdistance(d1)thebearingaxialseparation(d3)is in a range from 4.1×10−1 to 8.3×10−1.
Claims
exact text as granted — not AI-modified1 . 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 receive an input from the core shaft and output drive via an output of the gearbox to a fan shaft so as to drive the fan, via an input to the fan, at a lower rotational speed than a rotational speed of the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the gas turbine engine, the fan shaft mounting structure comprising at least two supporting bearings connected to the fan shaft, wherein: the output of the gearbox is at a gearbox output position and the input to the fan is at a fan input position; a first bearing separation distance (d 1 ) is defined as an axial distance between the input to the fan and a closest bearing of the at least two supporting bearings in a rearward direction from the fan; both of the at least two supporting bearings are located at positions forward of the gearbox; a bearing axial separation (d 3 ) is defined as an axial distance between the closest bearing of the at least two supporting bearings in the rearward direction from the fan and a closest bearing of the at least two supporting bearings in a forward direction from the gearbox; and a bearing separation ratio defined as:
the
first
bearing
separation
distance
(
d
1
)
the
bearing
axial
separation
(
d
3
)
is in a range from 4.1×10 −1 to 8.3×10 −1 .
2 . The gas turbine engine according to claim 1 , wherein the bearing separation ratio is in a range from 4.1×10 −1 to 6.0×10 −1 .
3 . The gas turbine engine according to claim 1 , wherein the bearing separation ratio is in a range from 6.0×10 −1 to 8.3×10 −1 .
4 . The gas turbine engine according to claim 1 , wherein the bearing axial separation (d 3 ) is in a range from 0.18 m to 0.32 m.
5 . The gas turbine engine according to claim 1 , wherein at least one of a)—c) is satisfied:
a) the first bearing separation distance (d 1 ) is in a range from 0.12 m to 0.40 m;
b) a second bearing separation distance (d 2 ) is defined as an axial distance between the output of the gearbox and the closest bearing of the at least two supporting bearings in the forward direction from the gearbox, the second bearing separation distance (d 2 ) being in a range from 0.15 m to 0.45 m; and
c) an axial distance between the fan input position and the gearbox output position (d 4 ) is in a range from 0.43 m to 0.95 m.
6 . The gas turbine engine according to claim 1 , wherein a further bearing separation ratio of:
the
first
bearing
separation
distance
(
d
1
)
an
axial
distance
between
the
fan
input
position
and
the
gearbox
output
position
(
d
4
)
is greater than or equal to 1.6×10 −1 , and the axial distance between the fan input position and the gearbox output position (d 4 ) is greater than or equal to 0.43 m.
7 . The gas turbine engine according to claim 6 , wherein the further bearing separation ratio is in a range from 1.6×10 −1 to 2.2×10 −1 .
8 . The gas turbine engine according to claim 6 , wherein the further bearing separation ratio is in a range from 2.2×10 −1 to 3.3×10 −1 .
9 . The gas turbine engine according to claim 6 , wherein a first bearing separation product defined as:
the first bearing separation distance ( d 1 )×the axial distance between the fan input position and the gearbox output position ( d 4 )
is in a range from 5.2×10 −2 m 2 to 2.6×10 −1 m 2 .
10 . The gas turbine engine according to claim 1 , wherein one or more of a) to d) is satisfied:
a) the at least two supporting bearings include a first supporting bearing and a second supporting bearing, and the fan shaft mounting structure further comprises a third supporting bearing, wherein the third supporting bearing is optionally located at a position between the fan and the gearbox; b) the fan shaft comprises a gearbox output shaft forming a relatively flexible portion of the fan shaft, and the fan shaft mounting structure comprises a gearbox output shaft support structure having at least one gearbox output shaft bearing arranged to support the gearbox output shaft; c) the fan shaft mounting structure further comprises one or more non-supporting softly mounted bearings; and d) one or more bearings provided as part of the fan shaft mounting structure are double bearings.
11 . The gas turbine engine according to claim 1 , wherein a fan shaft radial bending stiffness ratio of:
a
radial
bending
stiffness
of
the
fan
shaft
at
the
input
to
the
fan
a
radial
bending
stiffness
of
the
fan
shaft
at
the
output
to
the
gearbox
is in a range from 6.0×10 −3 to 2.5×10 1 .
12 . The gas turbine engine according to claim 11 , wherein:
a) the radial bending stiffness of the fan shaft at the input to the fan is in a range from 3.00×10 6 N/m to 2.00×10 9 N/m; and/or b) the radial bending stiffness of the fan shaft at the output of the gearbox is in a range from 4.00×10 6 N/m to 1.5×10 9 N/m.
13 . The gas turbine engine according to claim 1 , wherein a fan shaft tilt stiffness ratio of:
a
tilt
stiffness
of
the
fan
shaft
at
the
input
of
the
fan
a
tilt
stiffness
of
the
fan
at
the
output
of
the
gearbox
is in a range from 2.5×10 −2 to 3.7×10 2 .
14 . The gas turbine engine according to claim 13 , wherein:
a) the tilt stiffness of the fan shaft at the input to the fan is in a range from 5.00×10 5 Nm/rad to 7.00×10 8 Nm/rad; and/or b) the tilt stiffness of the fan shaft at the output of the gearbox is in a range from 7.00×10 4 Nm/rad to 7.00×10 7 Nm/rad.
15 . The gas turbine engine according to claim 1 , wherein the output of the gearbox is defined as a point of connection between the fan shaft and the gearbox, and wherein:
a) the gearbox is in a star configuration and the output of the gearbox is a gearbox output position defined as a point of connection between a ring gear and the fan shaft; or b) the gearbox is in a planetary configuration and the output of the gearbox is a gearbox output position at an interface between the fan shaft and a planet carrier.
16 . The gas turbine engine according to claim 1 , wherein at least one of i) to iii) is satisfied:
i) the gearbox has a gear ratio in a range of 3.2 to 4.0; ii) the gas turbine engine is configured to have a specific thrust in a range of 70 to 90 NKg −1 ; and iii) the gas turbine engine in configured to have a bypass ratio at cruise conditions of 13 to 16.
17 . The gas turbine engine according to claim 1 , wherein the fan has a fan diameter in a range from 220 cm to 240 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 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 a rotational speed of the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the propulsor, the fan shaft mounting structure comprising at least two supporting bearings connected to the fan shaft, wherein: an output of the gearbox is at a gearbox output position and an input to the fan is at a fan input position; a first bearing separation distance (d 1 ) is defined as an axial distance between the input to the fan and a closest bearing of the at least two supporting bearings in a rearward direction from the fan; both of the at least two supporting bearings are located at positions forward of the gearbox; a bearing axial separation (d3) is defined as an axial distance between the closest bearing of the at least two supporting bearings in the rearward direction from the fan and a closest bearing of the at least two supporting bearings in a forward direction from the gearbox; and a bearing separation ratio defined as:
the
first
bearing
separation
distance
(
d
1
)
the
bearing
axial
separation
(
d
3
)
is in a range from 4.1×10 −1 to 8.3×10 −1 .
19 . A method of operating 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 receive an input from the core shaft and output drive via an output of the gearbox to a fan shaft so as to drive the fan via an input to the fan at a lower rotational speed than a rotational speed of the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the gas turbine engine, the fan shaft mounting structure comprising at least two supporting bearings connected to the fan shaft, wherein: the output of the gearbox is at a gearbox output position and the input to the fan is at a fan input position; a first bearing separation distance (d 1 ) is defined as an axial distance between the input to the fan and a closest bearing of the at least two supporting bearings in a rearward direction from the fan; both of the at least two supporting bearings are located at positions forward of the gearbox; a bearing axial separation (d3) is defined as an axial distance between the closest bearing of the at least two supporting bearings in the rearward direction from the fan and a closest bearing of the at least two supporting bearings in a forward direction from the gearbox; and a bearing separation ratio defined as:
the
first
bearing
separation
distance
(
d
1
)
the
bearing
axial
separation
(
d
3
)
is in a range from 4.1×10 −1 to 8.3×10 −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 of:
i) greater than or equal to 10,000 Nm at cruise; and/or ii) greater than or equal to 28,000 Nm at max-take off conditions.Join the waitlist — get patent alerts
Track US2022298976A1 — get alerts on status changes and closely related new filings.
We store only your email — no account needed. See our privacy policy.