Gas turbine engine including gearbox structure having stiffness
A gas turbine engine including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft mounting structure that includes supporting bearings connected to the fan shaft; and an epicyclic gearbox that can receive input from the core shaft, and output drive to the fan shaft so as to drive the fan at a lower rotational speed than the core shaft. The gearbox includes a sun gear, planet gears, a ring gear, and a planet carrier for mounting the planet gears. A torsional stiffness of the planet carrier is greater than or equal to 1.6×10 8 Nm/rad. The fan shaft mounting structure has: a radial bending stiffness greater than or equal to 7.00×10 8 N/m, and/or a tilt stiffness greater than or equal to 1.50×10 7 Nm/rad.
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 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 on which the plurality of planet gears is mounted; 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:
a torsional stiffness of the planet carrier is in a range from 1.60×10 8 Nm/rad to 1.00×10 11 Nm/rad; and
a radial bending stiffness of the fan shaft mounting structure is in a range from 7.00×10 8 N/m to 6.00×10 11 N/m.
2. The gas turbine engine according to claim 1 , wherein the torsional stiffness of the planet carrier is in a range from 2.7×10 8 Nm/rad to 1.00×10 11 Nm/rad.
3. The gas turbine engine according to claim 1 , wherein the radial bending stiffness of the fan shaft mounting structure is in a range from 1.25×10 9 N/m to 6.00×10 11 N/m.
4. The gas turbine engine of according to claim 1 , wherein:
the fan has a fan diameter in a range from 240 cm to 280 cm, and
at least one of the following is satisfied:
a) the torsional stiffness of the planet carrier is in a range from 1.8×10 8 Nm/rad to 4.8×10 9 Nm/rad; and
b) the radial bending stiffness of the fan shaft mounting structure is in a range from 7.0×10 8 N/m to 5.0×10 11 N/m.
5. The gas turbine engine of according to claim 1 , wherein:
the fan has a fan diameter in a range from 330 cm to 380 cm, and
at least one of the following is satisfied:
a) the torsional stiffness of the planet carrier is in a range from 6.0×10 8 Nm/rad to 2.2×10 10 Nm/rad; and
b) the radial bending stiffness of the fan shaft mounting structure is in a range from 1.4×10 9 N/m to 6.0×10 11 N/m.
6. The gas turbine engine according to claim 1 , wherein a tilt stiffness of the fan shaft mounting structure is in a range from 1.5×10 7 Nm/rad to 2.70×10 10 Nm/rad.
7. The gas turbine engine according to claim 1 , wherein any one or more of the following is satisfied:
a) a first planet carrier stiffness ratio of:
an
effective
linear
torsional
stiffness
of
the
planet
carrier
the
radial
bending
stiffness
of
the
fan
shaft
mounting
structure
is in a range from 7.0×10 −3 to 1.9×10 3 ;
b) a first planet carrier stiffness product of:
(the effective linear torsional stiffness of the planet carrier)×(the radial bending stiffness of the fan shaft mounting structure)
is in a range from 2.9×10 18 (N/m) 2 to 8.0×10 22 (N/m) 2 ;
c) a second planet carrier stiffness ratio is defined as:
the
torsional
stiffness
of
the
planet
carrier
a
tilt
stiffness
of
the
fan
shaft
mounting
structure
is in a range from 6.0×10 −3 to 7.0×10 3 ; and
d) a second planet carrier stiffness product is defined as:
(the torsional stiffness of the planet carrier)×(the tilt stiffness of the fan shaft mounting structure),
and is in a range from 2.4×10 15 (Nm/rad) 2 to 2.7×10 21 (Nm/rad) 2 .
8. The gas turbine engine according to claim 1 , wherein a power transmitted by the gearbox is in a range from 2.25×10 7 W to 1.00×10 8 W.
9. The gas turbine engine according to claim 1 , wherein:
a) the fan has a fan diameter greater than 240 cm and less than or equal to 380 cm; and/or
b) a moment of inertia of the fan is in a range from 7.40×10 7 kgm 2 to 9.00×10 8 kgm 2 .
10. The gas turbine engine according to claim 1 , wherein the at least two supporting bearings comprise a first supporting bearing and second supporting bearing, and wherein:
a) both of the first and the second supporting bearings are located at positions forward of the gearbox; or
b) the first supporting bearing is located at a position forward of the gearbox and the second supporting bearing is located at a position rearward of the gearbox.
11. The gas turbine engine according to claim 10 , wherein any one of more of the following is satisfied:
a) the fan shaft mounting structure further comprises a third supporting bearing, the third supporting bearing being located between the fan and the gearbox;
b) the fan shaft comprises a gearbox output shaft forming a flexible portion of the fan shaft that is more flexible than an intermediate portion of the fan shaft to which the flexible portion is connected, 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 bearings; and
d) any one or more of the bearings provided as part of the fan shaft mounting structure are double bearings.
12. The gas turbine engine according to claim 1 , wherein:
the gearbox has a gear ratio in a range of 3.2 to 4.5.
13. The gas turbine engine according to claim 1 , wherein the gas turbine engine is configured to have a specific thrust in a range from 70 to 90 NKg −1 .
14. The gas turbine engine according to claim 1 , wherein the gas turbine engine is configured to have a bypass ratio at cruise conditions in a range from 12.5 to 18, the cruise conditions being: (i) conditions at mid-cruise of an aircraft to which the gas turbine engine is attached, or (ii) conditions experienced by the aircraft and the gas turbine engine at a midpoint between a top of a climb and a start of a descent.
15. 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 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 on which the plurality of planet gears is mounted; 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:
a torsional stiffness of the planet carrier is in a range from 1.60×10 8 Nm/rad to 1.00×10 11 Nm/rad; and
a tilt stiffness of the fan shaft mounting structure is in a range from 1.50×10 7 Nm/rad to 2.70×10 10 Nm/rad.
16. The gas turbine engine according to claim 15 , wherein the torsional stiffness of the planet carrier is in a range from 2.7×10 8 Nm/rad to 1×10 10 Nm/rad.
17. The gas turbine engine according to claim 15 , wherein the tilt stiffness of the fan shaft mounting structure is in a range from 2.1×10 7 Nm/rad to 2.70×10 10 Nm/rad.
18. The gas turbine engine according to claim 15 , wherein:
the fan has a fan diameter in a range from 240 cm to 280 cm, and
at least one of the following is satisfied:
a) the torsional stiffness of the planet carrier is in a range from 1.8×10 8 Nm/rad to 4.8×10 9 Nm/rad; and
b) the tilt stiffness of the fan shaft mounting structure is in a range from 2.1×10 7 Nm/rad to 1.9×10 10 Nm/rad.
19. The gas turbine engine according to claim 15 , wherein:
the fan has a fan diameter in a range from 330 cm to 380 cm, and
at least one of the following is satisfied:
a) the torsional stiffness of the planet carrier is in a range from 6.0×10 8 Nm/rad to 2.2×10 10 Nm/rad; and
b) the tilt stiffness of the fan shaft mounting structure is in a range from 3.8×10 7 Nm/rad to 2.7×10 10 Nm/rad.