High power epicyclic gearbox and operation thereof

Abstract

The invention is entitled High Power Epicyclic Gearbox and Operation Thereof. The disclosure provides a gas turbine engine (10) for an aircraft, the gas turbine engine comprising: an engine core (11) comprising a turbine (19), a compressor (14) and a spool (26) connecting the turbine to the compressor; a fan (23) upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox (30) receiving an input from the spool (26) and outputting a drive to a fan shaft (36) via an output of the gearbox in order to drive the fan at a lower rotational speed than the spool; a fan shaft mounting structure (503) arranged to mount the fan shaft within the engine, the fan shaft mounting structure (503) comprising at least two support bearings (506a, 506b) connected to the fan shaft (36). A fan-gearbox axial distance (110) is defined as the axial distance between the output of the gearbox and the fan axial centreline, the fan-gearbox axial distance being greater than or equal to 0.35m. A system radial bending stiffness is defined as: a fan shaft mounting radial bending stiffness ratio: greater than or equal to 3.9 x 10 6 .

Description

TECHNICAL FIELD

[0001] The present disclosure relates to gas turbine engines, in particular, to gas turbine engines for aircraft. Aspects of the present disclosure relate to aircraft comprising the gas turbine engines and methods of operating the gas turbine engines. BACKGROUND

[0002] Gas turbine engines for aircraft propulsion have a number of design factors that impact overall efficiency and power output or thrust. A general aim of a gas turbine engine is to provide a low specific fuel consumption (SFC). To enable higher thrust with high efficiency, a larger diameter fan can be used. To facilitate use of the 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 be operated at a reduced rotational speed compared to the case of using direct drive.

[0003] However, when manufacturing an engine with a larger fan diameter, simply scaling up components of a known engine type can not lead to an efficient design. For example, there can be issues associated with mounting the fan shaft within the engine. Therefore, the properties of the components used to mount the fan shaft, the properties of the fan shaft itself, and the properties of the gearbox components need to be considered. SUMMARY

[0004] According to a first aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbomachine, a compressor and a core shaft connecting the turbomachine to the compressor; a fan upstream of the engine core, the fan comprising a plurality of fan blades, the fan having an axial centreline of the fan; a gearbox receiving input from the core shaft and outputting 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 fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0005] a fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35m;

[0006] a system radial bending stiffness is defined as:

[0007]

[0008] and a fan shaft mounting radial bending stiffness ratio:

[0009]

[0010] greater than or equal to 1.0 x 10 -3 .

[0011] The fan shaft mounting radial bending stiffness ratio can be greater than or equal to 5.0 x 10 -3 The fan shaft mounting radial bending stiffness ratio can be greater than or equal to 2.0 x 10 -2 The fan shaft mounting radial bending stiffness ratio can be in the range of 1.0 x 10 -3 to 4.0 x 10 -1 The fan shaft mounting radial bending stiffness ratio can be in the range of 5.0 x 10 -3 to 1.5 x 10 -1 The fan shaft mounting radial bending stiffness ratio can be in the range of 5.0 x 10 -3 to 2.0 x 10 -2 The fan shaft mounting radial bending stiffness ratio can be in the range of 2.0 x 10 -2 to 1.5 x 10 -1 .

[0012] The system radial bending stiffness can be greater than or equal to 3.90 x 10 6 N / m. The system radial bending stiffness can be greater than or equal to 3.6 x 10 7 N / m. The system radial bending stiffness can be in the range of 3.90 x 10 6 N / m to 1.40 x 10 9 N / m. The system radial bending stiffness can be in the range of 3.6 x 10 7 N / m to 6.8 x 10 8 N / m.

[0013] The fan shaft mounting structure radial bending stiffness can be greater than or equal to 7.00 x 10 8 N / m. The fan shaft mounting structure radial bending stiffness can be greater than or equal to 1.25 x 10 9 N / m. The fan shaft mounting structure radial bending stiffness can be in the range of 7.00 x 10 8 N / m to 6.00 x 10 11 N / m. The fan shaft mounting structure radial bending stiffness can be in the range of 1.25 x 10 9 N / m to 2.0 x 10 11 N / m.

[0014] The fan shaft radial bending stiffness at the output of the gearbox can be greater than or equal to 4.00 x 10 6 N / m. The fan shaft radial bending stiffness at the output of the gearbox can be greater than or equal to 3.7 x 10 7N / m. The radial bending stiffness of the fan shaft at the output of the gearbox can be in the range of 4.00 x 10 6 N / m to 1.5 x 10 9 N / m. The radial bending stiffness of the fan shaft at the output of the gearbox can be in the range of 4.00 x 10 7 N / m to 1.0 x 10 9 N / m.

[0015] The product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 2.7 x 10 15 (N / m) 2 . The product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 4.0 x 10 15 (N / m) 2 . The product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure can be in the range of 2.7 x 10 15 (N / m) 2 to 9.0 x 10 19 (N / m) 2 . The product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure can be in the range of 4.0 x 10 15 (N / m) 2 to 1.5 x 10 19 (N / m) 2 .

[0016] The system tilt stiffness can be defined as:

[0017]

[0018] and the fan shaft mounting tilt stiffness ratio:

[0019]

[0020] may be greater than or equal to 1.5 x 10 -3 . The fan shaft mounting tilt stiffness ratio can be greater than or equal to 6.0 x 10 -3 . The fan shaft mounting tilt stiffness ratio can be greater than or equal to 2.5 x 10 -2 . The fan shaft mounting tilt stiffness ratio can be in the range of 1.5 x 10 -3 to 5.0 x 10 -1 . The fan shaft mounting tilt stiffness ratio can be in the range of 6.0 x 10 -3 to 2.0 x 10 -1 . The fan shaft mounting tilt stiffness ratio can be in the range of 6.0 x 10 -3 to 2.5 x 10 -2Within the range. The fan shaft mounting anti-tilting stiffness ratio can be 2.5×10. -2 Up to 2.0×10 -1 Within the range.

[0021] According to a second aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades and having a fan axial centerline; a gearbox receiving input from the spindle and outputting drive to the fan shaft via an output end of the gearbox to drive the fan at a lower rotational speed than the spindle; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two support bearings connected to the fan shaft, wherein:

[0022] The fan-gearbox axial distance is defined as the axial distance between the output end of the gearbox and the axial centerline of the fan, and the fan-gearbox axial distance is greater than or equal to 0.35m;

[0023] The system's anti-tilting stiffness is defined as:

[0024]

[0025] Furthermore, the fan shaft mounting anti-tilting stiffness ratio is:

[0026]

[0027] Greater than or equal to 1.5 × 10 -3 .

[0028] The fan shaft mounting anti-tilting stiffness ratio can be greater than or equal to 6.0 × 10⁻⁶. -3 The fan shaft mounting anti-tilting stiffness ratio can be greater than or equal to 2.5 × 10⁻⁶. -2 The fan shaft mounting anti-tilting stiffness ratio can be 1.5 × 10⁻⁶. -3 Up to 5.0×10 -1 Within the range. The fan shaft mounting anti-tilting stiffness ratio can be 6.0 × 10. -3 Up to 2.0×10 -1 Within the range. The fan shaft mounting anti-tilting stiffness ratio can be 6.0 × 10. -3 Up to 2.5×10 -2 Within the range. The fan shaft mounting anti-tilting stiffness ratio can be 2.5×10. -2 Up to 2.0×10 -1 Within the range.

[0029] The system's anti-tilting stiffness can be greater than or equal to 1.10 × 10⁻⁶. 5Nm / rad. The system roll stiffness can be greater than or equal to 8.5 x 10 5 Nm / rad. The system roll stiffness can be in the range of 1.10 x 10 5 Nm / rad to 6.80 x 10 7 Nm / rad. The system roll stiffness can be in the range of 8.5 x 10 5 Nm / rad to 3.4 x 10 7 Nm / rad.

[0030] The fan shaft mounting structure roll stiffness can be greater than or equal to 1.50 x 10 7 Nm / rad. The fan shaft mounting structure roll stiffness can be greater than or equal to 2.1 x 10 7 Nm / rad. The fan shaft mounting structure roll stiffness can be in the range of 1.5 x 10 7 Nm / rad to 2.70 x 10 10 Nm / rad. The fan shaft mounting structure roll stiffness can be in the range of 2.1 x 10 7 Nm / rad to 1 x 10 10 Nm / rad.

[0031] The fan shaft at the output of the gearbox roll stiffness can be greater than or equal to 7.00 x 10 4 Nm / rad. The fan shaft at the output of the gearbox roll stiffness can be greater than or equal to 9.5 x 10 5 Nm / rad. The fan shaft at the output of the gearbox roll stiffness can be in the range of 7.00 x 10 4 Nm / rad to 7.00 x 10 7 Nm / rad. The fan shaft at the output of the gearbox roll stiffness can be in the range of 9.5 x 10 5 Nm / rad to 3.5 x 10 7 Nm / rad.

[0032] The product of the system roll stiffness and the fan shaft mounting structure roll stiffness can be greater than or equal to 1.7 x 10 12 (N / m) 2 . The product of the system roll stiffness and the fan shaft mounting structure roll stiffness can be greater than or equal to 1.6 x 10 13 (N / m) 2 . The product of the system roll stiffness and the fan shaft mounting structure roll stiffness can be in the range of 1.7 x 10 12 (Nm / rad) 2 to 3.0 x 10 17 (Nm / rad) 2the product of the system tilt stiffness and the tilt stiffness of the fan shaft mounting structure can be in the range of 1.6 x 10 13 (Nm / rad) 2 to 3.0 x 10 16 (Nm / rad) 2 .

[0033] One or more of the following features can apply to both the first and second aspects above or any aspect therein.

[0034] The fan shaft is defined as the torque transfer component extending from the output end of the gearbox to the input end of the fan. The fan shaft can include at least a portion of the gearbox output shaft and at least a portion of the fan input shaft.

[0035] The input end of the fan can be a fan input location defined as the point on the fan shaft at the axial midpoint of the interface between the fan and the fan shaft.

[0036] The output end of the gearbox can be defined as the connection point between the fan shaft and the gearbox. The gearbox can be in a planetary configuration and the output end of the gearbox can be a gearbox output location defined as the connection point between the ring gear and the fan shaft. Alternatively, the gearbox can be in a planetary configuration and the output end of the gearbox can be a gearbox output location at the interface between the fan shaft and the planet carrier.

[0037] The at least two support bearings can include a first support bearing and a second support bearing.

[0038] Both the first support bearing and the second support bearing can be located at a position forward of the gearbox. Alternatively, the first support bearing can be located at a position forward of the gearbox and the second support bearing can be located at a position rearward of the gearbox.

[0039] The fan shaft mounting structure can further include a third support bearing. The third support bearing can be located between the fan and the gearbox. The fan shaft can include a gearbox output shaft forming a relatively flexible portion of the fan shaft and the fan shaft mounting structure can include a gearbox output shaft support structure having at least one gearbox output shaft bearing arranged to support the gearbox output shaft. The fan shaft mounting structure can further include one or more soft-mounted non-support bearings. Any one or more of the bearings provided as part of the fan shaft mounting structure can be a dual bearing.

[0040] An axial distance d1 between an input end of the fan and the bearing of the at least two support bearings closest to the fan in a rearward direction can be greater than or equal to 0.12 m. The axial distance d1 can be greater than or equal to 0.13 m. The axial distance d1 can be in a range of 0.12 m to 0.40 m. The axial distance d1 can be in a range of 0.13 m to 0.30 m.

[0041] An axial distance d2 between an output end of the gearbox and the bearing of the at least two support bearings closest to the gearbox in a forward direction can be greater than or equal to 0.15 m. The axial distance d2 can be greater than or equal to 0.16 m. The axial distance d2 can be in a range of 0.15 m to 0.45 m. The axial distance d2 can be in a range of 0.16 m to 0.40 m.

[0042] The fan-gearbox axial distance can be greater than or equal to 0.37 m. The fan-gearbox axial distance can be in a range of 0.35 m to 0.8 m. The fan-gearbox axial distance can be in a range of 0.37 m to 0.75 m.

[0043] The gearbox can be an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier, wherein the planet carrier is arranged to have the plurality of planet gears mounted thereon.

[0044] According to a third aspect, there is provided a propulsor for an aircraft, the propulsor comprising: a fan comprising a plurality of fan blades, the fan having a fan axial centreline; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged for receiving an input from the power unit via a core shaft and outputting a drive to a fan shaft via an output end of the gearbox so as to drive the fan at a lower rotational speed than 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 support bearings connected to the fan shaft, and wherein:

[0045] A fan-gearbox axial distance is defined as an axial distance between an output end of the gearbox and the fan axial centreline, the fan-gearbox axial distance being greater than or equal to 0.35 m;

[0046] A system radial bending stiffness is defined as:

[0047]

[0048] and a fan shaft mounting radial bending stiffness ratio:

[0049]

[0050] is greater than or equal to 1.0 x 10-3 .

[0051] The propulsor of this third aspect can have some or all of the features described above in relation to the gas turbine engine of the first aspect, and in some embodiments can be a gas turbine engine.

[0052] According to a fourth aspect, there is provided a propulsor for an aircraft, the propulsor comprising: a fan comprising a plurality of fan blades, the fan having a fan axial centreline; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged for receiving input from the power unit via the core shaft and outputting drive to a fan shaft via an output end of the gearbox so as to drive the fan at a lower rotational speed than 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 support bearings connected to the fan shaft, and wherein:

[0053] a fan-gearbox axial distance is defined as the axial distance between the output end of the gearbox and the axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35m;

[0054] a system pitch stiffness is defined as

[0055]

[0056] and a fan shaft mounting pitch stiffness ratio:

[0057]

[0058] greater than or equal to 1.5 x 10 -3 .

[0059] The propulsor of this fourth aspect can have some or all of the features described above in relation to the gas turbine engine of the second aspect, and in some embodiments can be a gas turbine engine.

[0060] The third and fourth aspects can be combined. In such aspects, there is provided a propulsor for an aircraft, the propulsor comprising: a fan comprising a plurality of fan blades, the fan having a fan axial centreline; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged for receiving input from the power unit via the core shaft and outputting drive to a fan shaft via an output end of the gearbox so as to drive the fan at a lower rotational speed than 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 support bearings connected to the fan shaft, and wherein:

[0061] The fan-gearbox axial distance is defined as the axial distance between the output end of the gearbox and the axial centerline of the fan, and the fan-gearbox axial distance is greater than or equal to 0.35m;

[0062] The radial bending stiffness of the system is defined as:

[0063]

[0064] Furthermore, the radial bending stiffness ratio of the fan shaft is:

[0065]

[0066] Greater than or equal to 1.0 × 10 -3 ; and / or

[0067] The system's anti-tilting stiffness is defined as

[0068]

[0069] Furthermore, the fan shaft mounting anti-tilting stiffness ratio is:

[0070]

[0071] Greater than or equal to 1.5 × 10 -3 The propulsion unit in this aspect may have some or all of the features described above regarding the gas turbine engine in the first and second aspects, and in some embodiments it may be a gas turbine engine.

[0072] According to a fifth aspect, a method of operating a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades and having a fan axial centerline; a gearbox receiving input from the spindle and outputting drive to the fan shaft via an output end of the gearbox to drive the fan at a lower rotational speed than the spindle; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two support bearings connected to the fan shaft, wherein:

[0073] The fan-gearbox axial distance is defined as the axial distance between the output end of the gearbox and the axial centerline of the fan, and the fan-gearbox axial distance is greater than or equal to 0.35m;

[0074] The radial bending stiffness of the system is defined as:

[0075]

[0076] and a fan shaft mounting radial stiffness ratio:

[0077]

[0078] greater than or equal to 3.9 x 10 6 The method comprises operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0079] The method of the fifth aspect can be a method of operating a gas turbine engine or propulsor of the first aspect or the third aspect respectively. Thus, any of the features, ratios and parameters introduced above in connection with the first aspect or the third aspect can also apply to the fifth aspect.

[0080] 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 upstream of the engine core, the fan comprising a plurality of fan blades, the fan having a fan axial centreline; a gearbox receiving an input from the core shaft and outputting a 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 fan shaft mounting arrangement arranged to mount the fan shaft within the engine, the fan shaft mounting arrangement comprising at least two support bearings connected to the fan shaft, and wherein:

[0081] a fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35m;

[0082] a system pitch stiffness is defined as:

[0083]

[0084] and a fan shaft mounting pitch stiffness ratio:

[0085]

[0086] greater than or equal to 1.5 x 10 -3 The method comprises operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0087] The method of the sixth aspect can be a method of operating a gas turbine engine or propulsor of the second aspect or the fourth aspect respectively. Thus, any of the features, ratios and parameters introduced above in connection with the second aspect or the fourth aspect can also apply to the sixth aspect.

[0088] The fifth and sixth aspects can be combined. Such aspects provide 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 upstream of the engine core, the fan comprising a plurality of fan blades, the fan having a fan axial centerline; a gearbox receiving an input from the core shaft and outputting a 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 fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0089] The fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the axial centerline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35 m;

[0090] The system radial bending stiffness is defined as:

[0091]

[0092] And the fan shaft mounting radial bending stiffness ratio:

[0093]

[0094] is greater than or equal to 3.9 x 10 6 ; and / or

[0095] The system roll stiffness is defined as:

[0096]

[0097] And the fan shaft mounting roll stiffness ratio:

[0098]

[0099] is greater than or equal to 1.5 x 10 -3 . The method includes operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft. This aspect can be a method of operating any of the preceding aspects of a gas turbine or propulsor.

[0100] The present inventors have discovered that designing the fan shaft and the fan shaft mounting structure such that the fan shaft mounting radial bending / tip stiffness ratio is within a specified range allows for improved isolation from loads transmitted from the fan while maintaining adequate positioning of the fan within the engine. The present inventors have discovered that if the stiffness of the fan shaft mounting structure is reduced such that the fan shaft mounting stiffness ratio is outside of the specified range, the fan will not be adequately positioned within the engine, resulting in fan tip clearance control issues. For example, movement of the fan relative to its surrounding structure can result in the need for higher fan tip clearances, which will reduce the overall efficiency of the engine. The present inventors have also discovered that further increasing the stiffness of the fan shaft mounting structure such that the fan shaft mounting stiffness ratio is outside of the specified range provides little or no practical benefit in positioning the fan, but will result in undesirably increased weight of the mounting structure. The present inventors have discovered that increasing the stiffness of the fan shaft such that the fan shaft mounting stiffness ratio is outside of the specified limit will result in excessive loads being transmitted from the fan to the gearbox. While reducing the fan shaft stiffness is therefore advantageous, the present inventors have discovered that further reducing the stiffness of the fan shaft such that the fan shaft mounting stiffness ratio is outside of the above range will result in undesirably low modal lateral vibrations in the fan shaft having excessive amplitudes.

[0101] In other aspects, instead of a value range for the ratio of the components of the fan shaft mounting radial bending / tip stiffness ratio, a value range for the product of the components of the fan shaft mounting radial bending / tip stiffness ratio can be specified, or both a value range for the ratio of the components of the fan shaft mounting radial bending / tip stiffness ratio and a value range for the product of the components of the fan shaft mounting radial bending / tip stiffness ratio can be specified.

[0102] According to one such aspect, the first aspect introduced above can be reduced to providing an aspect of 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 upstream of the engine core, the fan comprising a plurality of fan blades, the fan having a fan axial centerline; a gearbox receiving an input from the core shaft and outputting a drive to the 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 fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0103] a fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the fan axial centerline, the fan-gearbox axial distance being greater than or equal to 0.35 m;

[0104] a system radial bending stiffness is defined as:

[0105]

[0106] And the product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 2.7 x 10 15 (N / m) 2 , greater than or equal to 4.0 x 10 15 (N / m) 2 , in a range of 2.7 x 10 15 (N / m) 2 to 9.0 x 10 19 (N / m) 2 , or in a range of 4.0 x 10 15 (N / m) 2 to 1.5 x 10 19 (N / m) 2 .

[0107] According to another such aspect, the second aspect introduced above can be reduced to providing an aspect of 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 upstream of the engine core, the fan comprising a plurality of fan blades, the fan having a fan axial centerline; a gearbox receiving an input from the core shaft and outputting a 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 fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0108] A fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the axial centerline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35 m;

[0109] A system pitch stiffness is defined as:

[0110]

[0111] And the product of the system pitch stiffness and the pitch stiffness of the fan shaft mounting structure can be greater than or equal to 1.7 x 10 12 (Nm / rad) 2 , greater than or equal to 1.6 x 10 13 (Nm / rad) 2 , in a range of 1.7 x 10 12 (Nm / rad) 2 to 3.0 x 10 17 (Nm / rad) 2 , or in a range of 1.6 x 10 13 (Nm / rad) 2 to 3.0 x 1016 (Nm / rad) 2 in the range of 0.5 to 1.5.

[0112] Those skilled in the art will appreciate that the method and propeller aspects can be formulated accordingly.

[0113] According to a seventh aspect, there is provided 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a 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 via an input of the fan, and wherein:

[0114] a fan-gearbox axial distance is defined as an axial distance between the output of the gearbox and an axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35m, and

[0115] a fan shaft radial bending stiffness ratio:

[0116]

[0117] greater than or equal to 6.0 x 10 -3 .

[0118] The fan shaft radial bending stiffness ratio can be greater than or equal to 0.015. The fan shaft radial bending stiffness ratio can be in the range of 6.0 x 10 -3 to 2.5 x 10 1 . The fan shaft radial bending stiffness ratio can be in the range of 0.015 to 2.5.

[0119] The radial bending stiffness of the fan shaft at the input of the fan can be greater than or equal to 3.00 x 10 6 N / m. The radial bending stiffness of the fan shaft at the input of the fan can be greater than or equal to 6.3 x 10 6 N / m. The radial bending stiffness of the fan shaft at the input of the fan can be in the range of 3.00 x 10 6 N / m to 2.00 x 10 9 N / m. The radial bending stiffness of the fan shaft at the input of the fan can be in the range of 6.3 x 10 6 N / m to 1.0 x 10 9 N / m.

[0120] The radial bending stiffness of the fan shaft at the output of the gearbox can be greater than or equal to 4.00 x 10 6N / m. The radial bending stiffness of the fan shaft at the output end of the gearbox can be greater than or equal to 3.7 x 10 7 N / m. The radial bending stiffness of the fan shaft at the output end of the gearbox can be in the range of 4.00 x 10 6 N / m to 1.5 x 10 9 N / m. The radial bending stiffness of the fan shaft at the output end of the gearbox can be in the range of 3.7 x 10 7 N / m to 1.0 x 10 9 N / m.

[0121] The diameter of the fan can be in the range of 240 cm to 280 cm. In such embodiments, the fan shaft radial bending stiffness ratio can be greater than or equal to 0.03, or in the range of 0.03 to 0.85.

[0122] The diameter of the fan can be in the range of 330 cm to 380 cm. In such embodiments, the fan shaft radial bending stiffness ratio can be greater than or equal to 0.02 or in the range of 0.02 to 1.5.

[0123] The fan shaft anti-tilt stiffness ratio:

[0124]

[0125] may be greater than or equal to 2.5 x 10 -2 .

[0126] The fan shaft anti-tilt stiffness ratio can be greater than or equal to 0.05. The fan shaft anti-tilt stiffness ratio can be in the range of 2.5 x 10 -2 to 3.7 x 10 2 . The fan shaft anti-tilt stiffness ratio can be in the range of 0.05 to 4.0 x 10 1 .

[0127] The following product:

[0128] (radial bending stiffness of the fan shaft at the input end of the fan) x (radial bending stiffness of the fan shaft at the output end of the gearbox)

[0129] may be greater than or equal to 1.2 x 10 13 (N / m) 2 , greater than or equal to 2.4 x 10 14 (N / m) 2 , in the range of 1.2 x 10 13 (N / m) 2 to 3.0 x 10 18 (N / m) 2 , or in the range of 2.4 x 10 14 (N / m) 2 to 3.0 x 1017 (N / m) 2 in the range of 0.5 to 1.5.

[0130] According to an eighth aspect, there is provided 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a drive to a fan shaft via an output of the gearbox so as to drive the fan via an input of the fan at a lower rotational speed than the core shaft, wherein:

[0131] a fan-gearbox axial distance is defined as an axial distance between the output of the gearbox and an axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35, and

[0132] a fan shaft tilt stiffness ratio:

[0133]

[0134] greater than or equal to 2.5 x 10 -2 .

[0135] The fan shaft tilt stiffness ratio can be greater than or equal to 0.05. The fan shaft tilt stiffness ratio can be in the range of 2.5 x 10 -2 to 3.7 x 10 2 . The fan shaft tilt stiffness ratio can be in the range of 0.05 to 4.0 x 10 1 .

[0136] The tilt stiffness of the fan shaft at the input of the fan can be greater than or equal to 5.00 x 10 5 Nm / rad. The tilt stiffness of the fan shaft at the input of the fan can be greater than or equal to 9.0 x 10 5 Nm / rad. The tilt stiffness of the fan shaft at the input of the fan can be in the range of 5.00 x 10 5 Nm / rad to 7.00 x 10 8 Nm / rad. The tilt stiffness of the fan shaft at the input of the fan can be in the range of 9.0 x 10 5 Nm / rad to 3.5 x 10 8 Nm / rad.

[0137] The tilt stiffness of the fan shaft at the output of the gearbox can be greater than or equal to 7.00 x 10 4 Nm / rad. The tilt stiffness of the fan shaft at the output of the gearbox can be greater than or equal to 9.5 x 10 5Nm / rad. The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be in the range of 7.00 x 10 4 Nm / rad to 7.00 x 10 7 Nm / rad. The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be in the range of 9.5 x 10 5 Nm / rad to 3.5 x 10 7 Nm / rad.

[0138] The diameter of the fan can be in the range of 240 cm to 280 cm. In such embodiments, the fan shaft anti-tilt stiffness ratio can be greater than or equal to 0.2, or in the range of 0.2 to 5.0.

[0139] The diameter of the fan can be in the range of 330 cm to 380 cm. In such embodiments, the fan shaft anti-tilt stiffness ratio can be greater than or equal to 0.1, or in the range of 0.1 to 1.0 x 10 1 .

[0140] The fan shaft radial bending stiffness ratio:

[0141]

[0142] may be greater than or equal to 6.0 x 10 -3 . The fan shaft radial bending stiffness ratio can be greater than or equal to 0.015. The fan shaft radial bending stiffness ratio can be in the range of 6.0 x 10 -3 to 2.5 x 10 1 . The fan shaft radial bending stiffness ratio can be in the range of 0.015 to 2.5.

[0143] The following product:

[0144] (Anti-tilt stiffness of the fan shaft at the input end of the fan) x (Anti-tilt stiffness of the fan shaft at the output end of the gearbox)

[0145] may be greater than or equal to 3.5 x 10 10 (Nm / rad) 2 , greater than or equal to 7.2 x 10 11 (Nm / rad) 2 , in the range of 3.5 x 10 10 (Nm / rad) 2 to 5.0 x 10 16 (Nm / rad) 2 , or in the range of 7.2 x 10 11 (Nm / rad) 2 to 5.0 x 10 15 (Nm / rad) 2 .

[0146] One or more of the following features can apply to either or both of the preceding two aspects (e.g., the seventh and eighth aspects):

[0147] The fan-geared box axial distance can be greater than or equal to 0.37 m. The fan-geared box axial distance can be in the range of 0.35 m to 0.8 m. The fan-geared box axial distance can be in the range of 0.37 m to 0.75 m.

[0148] The fan shaft is defined as the torque transfer component that extends from the output end of the gearbox to the input end of the fan. The fan shaft can include at least a portion (or all) of the gearbox output shaft and at least a portion (or all) of the fan input shaft.

[0149] The input end of the fan can be the fan input location, which is defined as the point on the fan shaft at the axial midpoint of the interface between the fan and the fan shaft.

[0150] The output end of the gearbox can be defined as the connection point between the fan shaft and the gearbox. The gearbox can be in a spur configuration, and the output end of the gearbox can be the gearbox output location, which is defined as the connection point between the ring gear and the fan shaft. The gearbox can be in a planetary configuration, and the output end of the gearbox can be the gearbox output location at the interface between the fan shaft and the planet carrier.

[0151] The gas turbine engine can further include a fan shaft mounting structure arranged to mount the fan shaft within the engine. The fan shaft mounting structure can include at least two support bearings connected to the fan shaft.

[0152] The at least two support bearings can include a first support bearing and a second support bearing. Both the first support bearing and the second support bearing can be located at a position forward of the gearbox. Alternatively, the first support bearing can be located at a position forward of the gearbox, and the second support bearing can be located at a position rearward of the gearbox.

[0153] The fan shaft mounting structure can include a third support bearing. The third support bearing can be located between the fan and the gearbox.

[0154] The fan shaft can include a gearbox output shaft that forms a relatively flexible portion of the fan shaft. The fan shaft mounting structure can include a gearbox output shaft support structure having at least one gearbox output shaft bearing arranged to support the gearbox output shaft.

[0155] The fan shaft mounting structure can further include one or more soft-mounted non-support bearings.

[0156] Any one or more of the bearings provided as part of the fan shaft mounting structure can be a double bearing.

[0157] An axial distance dl between an input end of the fan and the bearing of the at least two support bearings that is closest to the fan in a rearward direction can be greater than or equal to 0.12 m, greater than or equal to 0.13 m, in a range of 0.12 m to 0.40 m, or in a range of 0.13 m to 0.30 m.

[0158] An axial distance d2 between an output end of the gearbox and the bearing of the at least two support bearings that is closest to the gearbox in a forward direction can be greater than or equal to 0.15 m, greater than or equal to 0.16 m, in a range of 0.15 m to 0.45 m, or in a range of 0.16 m to 0.40 m.

[0159] The gearbox can be an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier, wherein the planet carrier is arranged to mount the plurality of planet gears thereon.

[0160] According to a ninth aspect, there is provided a propulsor for an aircraft, the propulsor 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 for receiving an input from the power unit via a core shaft and outputting a drive to a fan shaft via an output end of the gearbox so as to drive the fan via an input end of the fan at a lower rotational speed than the core shaft, and wherein

[0161] A fan-gearbox axial distance is defined as an axial distance between the output end of the gearbox and an axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35 m; and wherein:

[0162] A fan shaft radial bending stiffness ratio:

[0163]

[0164] greater than or equal to 6.0 x 10 -3 .

[0165] The propulsor of the ninth aspect can have some or all of the features described above in relation to the gas turbine engine of the seventh aspect, and in some embodiments can be a gas turbine engine.

[0166] According to a tenth aspect, there is provided a propulsor for an aircraft, the propulsor 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 for receiving input from the power unit via a core shaft and outputting drive to a fan shaft via an output end of the gearbox, so as to drive the fan via an input end of the fan at a lower rotational speed than the core shaft, and wherein

[0167] a fan-gearbox axial distance is defined as the axial distance between the output end of the gearbox and an axial centre line of the fan, the fan-gearbox axial distance being greater than or equal to 0.35 m; and wherein:

[0168] a fan shaft radial to torsional stiffness ratio:

[0169]

[0170] greater than or equal to 2.5 x 10 -2 .

[0171] The propulsor of the tenth aspect can have some or all of the features described above in relation to the gas turbine engine of the eighth aspect, and in some embodiments can be a gas turbine engine.

[0172] The ninth and tenth aspects can be combined. According to an eleventh aspect, there is provided a propulsor for an aircraft, the propulsor 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 for receiving input from the power unit via a core shaft and outputting drive to a fan shaft via an output end of the gearbox, so as to drive the fan via an input end of the fan at a lower rotational speed than the core shaft, and wherein

[0173] a fan-gearbox axial distance is defined as the axial distance between the output end of the gearbox and an axial centre line of the fan, the fan-gearbox axial distance being greater than or equal to 0.35 m; and wherein:

[0174] a) a fan shaft radial to torsional stiffness ratio:

[0175]

[0176] greater than or equal to 6.0 x 10 -3 ; and / or

[0177] b) a fan shaft radial to torsional stiffness ratio:

[0178]

[0179] greater than or equal to 2.5 x 10 -2 .

[0180] The propulsor of the eleventh aspect can have some or all of the features described above in relation to the gas turbine engine of the seventh or eighth aspect, and in some embodiments can be a gas turbine engine.

[0181] According to a twelfth 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a drive to a fan shaft via an output of the gearbox so as to drive the fan via an input of the fan at a lower rotational speed than the core shaft, and wherein

[0182] The fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and an axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35m, wherein:

[0183] The fan shaft radial bending stiffness ratio is:

[0184]

[0185] greater than or equal to 6.0 x 10 -3 The method comprises operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0186] The method of the twelfth aspect can be a method of operating the gas turbine engine or propulsor of the seventh or ninth aspect respectively. Accordingly, any of the features, ratios and parameters introduced above in relation to the seventh or ninth aspect can also apply to the twelfth aspect.

[0187] According to a thirteenth 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a drive to a fan shaft via an output of the gearbox so as to drive the fan via an input of the fan at a lower rotational speed than the core shaft, and wherein

[0188] The fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and an axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35m, wherein:

[0189] The fan shaft anti-tilt stiffness ratio is:

[0190]

[0191] greater than or equal to 2.5 x 10 -2 The method comprises operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0192] The method of the thirteenth aspect can be a method of operating the gas turbine engine or propulsor of the eighth aspect or the tenth aspect respectively. Accordingly, any of the features, ratios and parameters introduced above in connection with the eighth aspect or the tenth aspect can also apply to the thirteenth aspect.

[0193] The twelfth and thirteenth aspects can be combined. 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a drive to a fan shaft via an output of the gearbox so as to drive the fan via an input of the fan at a lower rotational speed than the core shaft, and wherein:

[0194] a fan-gearbox axial distance is defined as an axial distance between the output of the gearbox and an axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35m, wherein:

[0195] a) a fan shaft radial stiffness ratio:

[0196]

[0197] greater than or equal to 6.0 x 10 -3 ; and / or

[0198] b) a fan shaft pitch stiffness ratio:

[0199]

[0200] greater than or equal to 2.5 x 10 -2 The method comprises operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0201] The method of the fourteenth aspect can be a method of operating the gas turbine engine or propulsor of the seventh aspect, the eighth aspect or the eleventh aspect. Accordingly, any of the features, ratios and parameters introduced above in connection with the seventh aspect, the eighth aspect or the eleventh aspect can also apply to the fourteenth aspect.

[0202] The present inventors have discovered that by designing the fan shaft to have a comparative stiffness at the fan input end compared to at the gearbox output end, within other constraints of the engine, the fan can be accurately located within the engine while also isolating the gearbox from loads transmitted from the fan. The present inventors have discovered that if the stiffness of the fan shaft at the fan input end is such that the fan shaft radial bending / tip stiffness ratio is below a specified range, then the fan will not be sufficiently located (resulting in fan tip control issues) and there will be undesirable low modal vibrations with high amplitude. The inventors have discovered that if the fan shaft is designed such that the stiffness at the fan input end is further increased such that the fan shaft radial bending / tip stiffness ratio is outside the above range, then there is little or no practical benefit in the improved location of the fan with no undesirable increase in overall weight. The present inventors have also discovered that increasing the stiffness of the fan shaft at the output end of the gearbox such that the fan shaft radial bending / tip stiffness ratio is outside the specified range will result in excessive loads being transmitted from the fan into the gearbox. The inventors have also discovered that decreasing the stiffness of the fan shaft at the gearbox output end such that the fan shaft radial bending / tip stiffness ratio is outside the above range will result in undesirable low modal lateral vibrations with excessive amplitude being generated in the fan shaft.

[0203] In other aspects, instead of a value range for the ratio of the fan shaft radial bending stiffness ratio and the fan shaft tip stiffness ratio, a value range for the product of the components of the fan shaft radial bending stiffness ratio and the fan shaft tip stiffness ratio can be specified, or both a value range for the product and a value range for the ratio can be specified.

[0204] According to one such aspect, the seventh aspect introduced above can be reduced to providing an aspect of a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spool connecting the turbine to the compressor; a fan upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spool and outputting a drive to a fan shaft via an output end of the gearbox so as to drive the fan via an input end of the fan at a lower rotational speed than the spool, and wherein:

[0205] a fan-gearbox axial distance is defined as the axial distance between the output end of the gearbox and an axial centerline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35 m, and the following product:

[0206] (radial bending stiffness of the fan shaft at the input end of the fan) x (radial bending stiffness of the fan shaft at the output end of the gearbox)

[0207] may be greater than or equal to 1.2 x 10 13 (N / m) 2 , greater than or equal to 2.4 x 10 14 (N / m) 2in the range of 1.2 x 10 13 (N / m) 2 to 3.0 x 10 18 (N / m) 2 or in the range of 2.4 x 10 14 (N / m) 2 to 3.0 x 10 17 (N / m) 2 .

[0208] According to another such aspect, the eighth aspect introduced above can be reduced to providing an aspect of 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a drive to a fan shaft via an output of the gearbox so as to drive the fan via an input of the fan at a lower rotational speed than the core shaft, wherein:

[0209] a fan-gearbox axial distance is defined as an axial distance between the output of the gearbox and an axial centerline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35, and a product of:

[0210] (a fan shaft tip torsional stiffness at the input of the fan) x (a fan shaft tip torsional stiffness at the output of the gearbox)

[0211] may be greater than or equal to 3.5 x 10 10 (Nm / rad) 2 , greater than or equal to 7.2 x 10 11 (Nm / rad) 2 , in the range of 3.5 x 10 10 (Nm / rad) 2 to 5.0 x 10 16 (Nm / rad) 2 , or in the range of 7.2 x 10 11 (Nm / rad) 2 to 5.0 x 10 15 (Nm / rad) 2 .

[0212] Those skilled in the art will appreciate that method and propeller aspects can be formulated accordingly.

[0213] According to a fifteenth aspect, there is provided 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a drive to a fan shaft via an output of the gearbox so as to drive the fan via an input of the fan at a lower rotational speed than the core shaft; and

[0214] a fan shaft mounting structure arranged to mount the fan shaft within the engine, wherein the fan shaft mounting structure comprises at least two support bearings connected to the fan shaft, wherein:

[0215] the output of the gearbox is located at a gearbox output location and the input of the fan is located at a fan input location;

[0216] a first bearing spacing distance di is defined as an axial distance between the input of the fan and the bearing of the at least two support bearings that is closest to the fan in a rearward direction; and

[0217] a first bearing spacing ratio:

[0218]

[0219] greater than or equal to 1.6 x 10 -1 and an axial distance d4 between the fan input location and the gearbox output location is greater than or equal to 0.43 m.

[0220] The first bearing spacing ratio can be greater than or equal to 1.8 x 10 -1 . The first bearing spacing ratio can be greater than or equal to 1.6 x 10 -1 . The first bearing spacing ratio can be greater than or equal to 2.2 x 10 -1 . The first bearing spacing ratio can be in the range 1.6 x 10 -1 to 3.3 x 10 -1 . The first bearing spacing ratio can be in the range 1.8 x 10 -1 to 3.0 x 10 -1 . The first bearing spacing ratio can be in the range 1.6 x 10 -1 to 2.2 x 10 -1 . The first bearing spacing ratio can be in the range 2.2 x 10 -1 to 3.3 x 10 -1 .

[0221] The first bearing spacing distance di can be greater than or equal to 0.12 m, greater than or equal to 0.13 m, in the range 0.12 m to 0.40 m, or in the range 0.13 m to 0.30 m.

[0222] The second bearing spacing distance d2 is defined as the axial distance between the output of the gearbox and the bearing of the at least two support bearings that is closest to the gearbox in the forward direction. The second bearing spacing distance d2 can be greater than or equal to 0.15 m, greater than or equal to 0.16 m, in a range of 0.15 m to 0.45 m, or in a range of 0.16 m to 0.40 m.

[0223] The axial distance d4 between the fan input location and the gearbox output location can be greater than or equal to 0.46 m, in a range of 0.43 m to 0.95 m, or in a range of 0.46 m to 0.85 m.

[0224] The first bearing spacing product (defined as:

[0225] The first bearing spacing distance (di) x the axial distance (d4) between the fan input location and the gearbox output location

[0226] may be greater than or equal to 5.2 x 10 -2 m 2 . The first bearing spacing product can be greater than or equal to 5.7 x 10 -2 m 2 . The first bearing spacing product can be in a range of 5.2 x 10 -2 m 2 to 2.6 x 10 -1 m 2 . The first bearing spacing product can be in a range of 5.7 x 10 -2 m 2 to 2.4 x 10 -1 m 2 .

[0227] Both of the at least two support bearings can be located at a location forward of the gearbox. The bearing axial spacing d3 can be defined as the axial distance between the support bearing of the at least two support bearings that is closest to the fan in the rearward direction and the support bearing that is the closest to the gearbox in the forward direction of the at least two support bearings. The second bearing spacing ratio (defined as:

[0228]

[0229] may be greater than or equal to 4.1 x 10 -1 . The second bearing spacing ratio can be greater than or equal to 4.5 x 10 -1 . The second bearing spacing ratio can be greater than or equal to 6.0 x 10 -1 . The second bearing spacing ratio can be in a range of 4.1 x 10 -1 to 8.3 x 10 -1 . The second bearing spacing ratio can be in a range of 4.5 x 10 -1Up to 7.7×10 -1 Within the range. The second bearing spacing ratio can be 4.1 × 10. -1 Up to 6.0×10 -1 Within the range. The second bearing spacing ratio can be 6.0 × 10. -1 Up to 8.3×10 -1 Within the range.

[0230] One of the at least two support bearings may be located at the front of the gearbox, and the other of the at least two support bearings may be located at the rear of the gearbox.

[0231] The at least two support bearings may include a first support bearing and a second support bearing, and the fan shaft mounting structure may include a third support bearing. The third support bearing may be located between the fan and the gearbox.

[0232] The fan shaft may include a gearbox output shaft that forms a relatively flexible portion of the fan shaft. The fan shaft mounting structure may include a gearbox output shaft support structure having at least one gearbox output shaft bearing arranged to support the gearbox output shaft.

[0233] The fan shaft mounting structure may also include one or more flexible, unsupported bearings.

[0234] Any one or more of the bearings provided as part of the fan shaft mounting structure can be a dual bearing.

[0235] Fan shaft radial bending stiffness ratio:

[0236]

[0237] It can be greater than or equal to 6.0 × 10 -3 The radial bending stiffness ratio of this fan shaft can be greater than or equal to 0.015. The radial bending stiffness ratio of this fan shaft can be 6.0 × 10⁻⁶. -3 Up to 2.5×10 1 Within a certain range. The radial bending stiffness ratio of the fan shaft can range from 0.015 to 2.5.

[0238] The radial bending stiffness of the fan shaft at the fan input end can be greater than or equal to 3.00 × 10⁻⁶. 6 N / m. The radial bending stiffness of the fan shaft at the fan input end can be greater than or equal to 6.3 × 10 N / m. 6 N / m. The radial bending stiffness of the fan shaft at the fan input end can reach 3.00 × 10 N / m. 6 N / m to 2.00×10 9The radial bending stiffness of the fan shaft at the fan input end is within the range of N / m. 6 N / m to 1.0×10 9 Within the range of N / m.

[0239] The radial bending stiffness of the fan shaft at the output end of the gearbox can be greater than or equal to 4.00 × 10⁻⁶. 6 N / m. The radial bending stiffness of the fan shaft at the output end of the gearbox can be greater than or equal to 3.7 × 10 N / m. 7 N / m. The radial bending stiffness of the fan shaft at the output end of the gearbox can reach 4.00 × 10 N / m. 6 N / m to 1.5×10 9 The radial bending stiffness of the fan shaft at the output end of the gearbox is within the range of N / m. 7 N / m to 1.0×10 9 Within the range of N / m.

[0240] The diameter of the fan can be in the range of 240cm to 280cm. In such embodiments, the radial bending stiffness ratio of the fan shaft can be greater than or equal to 0.03, or in the range of 0.03 to 0.85.

[0241] Alternatively, the fan diameter can be in the range of 330 cm to 380 cm. In such embodiments, the fan shaft radial bending stiffness ratio can be greater than or equal to 0.02 or in the range of 0.02 to 1.5.

[0242] Fan shaft anti-tilting stiffness ratio:

[0243]

[0244] It can be greater than or equal to 2.5 × 10 -2 The fan shaft tilt stiffness ratio can be greater than or equal to 0.05. The fan shaft tilt stiffness ratio can be 2.5 × 10⁻⁶. -2 Up to 3.7×10 2 Within this range, the fan shaft anti-tilting stiffness ratio can range from 0.05 to 4.0 × 10⁻⁶. 1 Within the range.

[0245] The anti-tilting stiffness of the fan shaft at the fan input end can be greater than or equal to 5.00 × 10⁻⁶. 5 Nm / rad. The anti-tilting stiffness of the fan shaft at the fan input end can be greater than or equal to 9.0 × 10 Nm / rad. 5 Nm / rad. The anti-tilting stiffness of the fan shaft at the fan input end can be 5.00 × 10 Nm / rad. 5 Nm / rad to 7.00×10 8Nm / rad. The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be in the range of 7.00 x 10 5 Nm / rad to 3.5 x 10 8 Nm / rad.

[0246] The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be greater than or equal to 7.00 x 10 4 Nm / rad. The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be greater than or equal to 9.5 x 10 5 Nm / rad. The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be in the range of 7.00 x 10 4 Nm / rad to 7.00 x 10 7 Nm / rad. The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be in the range of 9.5 x 10 5 Nm / rad to 3.5 x 10 7 Nm / rad.

[0247] The diameter of the fan can be in the range of 240 cm to 280 cm. In such embodiments, the fan shaft anti-tilt stiffness ratio can be greater than or equal to 0.2, or in the range of 0.2 to 5.0.

[0248] Alternatively, the diameter of the fan can be in the range of 330 cm to 380 cm. In such embodiments, the fan shaft anti-tilt stiffness ratio can be greater than or equal to 0.1, or in the range of 0.1 to 1.0 x 10 1 .

[0249] The fan shaft is defined as the torque transfer component that extends from the output end of the gearbox to the input end of the fan. The fan shaft can include at least a portion of the gearbox output shaft and at least a portion of the fan input shaft.

[0250] The input end of the fan can be the fan input location, which is defined as the point on the fan shaft at the axial midpoint of the interface between the fan and the fan shaft.

[0251] The output end of the gearbox can be defined as the connection point between the fan shaft and the gearbox. The gearbox can be in a spur configuration, and the output end of the gearbox can be the gearbox output location, which is defined as the connection point between the ring gear and the fan shaft. Alternatively, the gearbox can be in a planetary configuration, and the output end of the gearbox can be the gearbox output location at the interface between the fan shaft and the planet carrier.

[0252] The gearbox can be an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier, wherein the planet carrier is arranged to have the plurality of planet gears mounted thereon.

[0253] According to a sixteenth aspect, there is provided a propulsor for an aircraft, the propulsor 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 for receiving input from the power unit via a core shaft and outputting drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount a fan shaft within the propulsor, wherein the fan shaft mounting structure comprises at least two support bearings connected to the fan shaft, wherein:

[0254] an output end of the gearbox is located at a gearbox output location and an input end of the fan is located at a fan input location;

[0255] a first bearing spacing distance di is defined as an axial distance between the input end of the fan and the bearing of the at least two support bearings that is closest to the fan in a rearward direction; and

[0256] a first bearing spacing ratio:

[0257]

[0258] greater than or equal to 1.6 x 10 -1 and an axial distance d4 between the fan input location and the gearbox output location is greater than or equal to 0.43 m.

[0259] The propulsor can have some or all of the features described above in relation to the fifteenth aspect in respect of a gas turbine engine, and in some embodiments can be a gas turbine engine.

[0260] According to a seventeenth 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the core shaft and outputting drive to a fan shaft via an output end of the gearbox so as to drive the fan via an input end of the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount a fan shaft within the engine, wherein the fan shaft mounting structure comprises at least two support bearings connected to the fan shaft, wherein:

[0261] an output end of the gearbox is located at a gearbox output location and an input end of the fan is located at a fan input location;

[0262] a first bearing spacing distance di is defined as an axial distance between the input end of the fan and the bearing of the at least two support bearings that is closest to the fan in a rearward direction; and

[0263] the first bearing spacing ratio is greater than or equal to 1.6 x 10

[0264]

[0265] greater than or equal to 1.6 x 10 -1 and an axial distance (d4) between the fan input location and the gearbox output location is greater than or equal to 0.43 m. The method includes operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0266] The method of the seventeenth aspect can be a method of operating the gas turbine engine or the propulsor of the fifteenth aspect or the sixteenth aspect, respectively. Thus, any of the features, ratios and parameters introduced above in connection with the fifteenth aspect or the sixteenth aspect also apply to the seventeenth aspect.

[0267] The inventors have found that by arranging the bearings such that the first bearing spacing ratio defined above is within the specified range, the fan can be sufficiently located within the engine and the gearbox is isolated from the load from the fan while still providing a suitable fan shaft geometry to fit within the engine. The inventors have found that if the first bearing spacing distance is to be increased such that the ratio is outside the specified range, the fan will not be sufficiently located. The inventors have also found that if the axial distance between the fan input end and the gearbox output end is to be reduced such that the ratio is outside the specified range, an excessive load will be transferred from the fan into the gearbox.

[0268] In other aspects, instead of a value range for the ratio of the components of the first bearing spacing ratio, a value range for the product of the components of the first bearing spacing ratio can be specified, or both a value range for the ratio of the components of the first bearing spacing ratio and a value range for the product of the components of the first bearing spacing ratio.

[0269] According to one such aspect, the fifteenth aspect introduced above can be reduced to an aspect of 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a drive to a fan shaft via an output end of the gearbox so as to drive the fan via an input end of the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, wherein the fan shaft mounting structure comprises at least two support bearings connected to the fan shaft, wherein: the output end of the gearbox is located at a gearbox output location and the input end of the fan is located at a fan input location; a first bearing spacing distance dl is defined as an axial distance between the input end of the fan and the bearing of the at least two support bearings that is closest to the fan in a rearward direction; and a first bearing spacing product (defined as:

[0270] First bearing spacing distance (dl)

[0271] The axial distance (d4) between the fan input location and the gearbox output location is greater than or equal to 5.2 x 10 -2 m 2 , greater than or equal to 5.7 x 10 -2 m 2 , in the range 5.2 x 10 -2 m 2 to 2.6 x 10 -1 m 2 , or in the range 5.7 x 10 -2 m 2 to 2.4 x 10 -1 m 2 , and the axial distance d4 between the fan input location and the gearbox output location is greater than or equal to 0.43 m.

[0272] The skilled person will appreciate that the method and propeller aspects can be formulated accordingly.

[0273] According to an eighteenth aspect, there is provided 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a 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, wherein the planet carrier is arranged to have the plurality of planet gears mounted thereon; and

[0274] a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0275] the torsional stiffness of the planet carrier is greater than or equal to 1.60 x 10 8 Nm / rad; and

[0276] the radial bending stiffness of the fan shaft mounting structure is greater than or equal to 7.00 x 10 8 N / m.

[0277] The torsional stiffness of the planet carrier can be greater than or equal to 2.7 x 10 8 Nm / rad. The torsional stiffness of the planet carrier can be in the range 1.60 x 10 8 Nm / rad to 1.00 x 10 11 Nm / rad. The torsional stiffness of the planet carrier can be in the range 2.7 x 108 Nm / rad to 1 x 10 10 Nm / rad.

[0278] The radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 1.25 x 10 9 N / m. The radial bending stiffness of the fan shaft mounting structure can be in a range of 7.00 x 10 8 N / m to 6.00 x 10 11 N / m. The radial bending stiffness of the fan shaft mounting structure can be in a range of 1.25 x 10 9 N / m to 2.0 x 10 11 N / m.

[0279] The fan can have a fan diameter in a range of 240 cm to 280 cm. In such embodiments, the torsional stiffness of the planetary carrier can be greater than or equal to 1.8 x 10 8 Nm / rad, or in a range of 1.8 x 10 8 Nm / rad to 4.8 x 10 9 Nm / rad. Additionally or alternatively, in such embodiments, the radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 7.0 x 10 8 N / m, or in a range of 7.0 x 10 8 N / m to 5.0 x 10 11 N / m.

[0280] The fan can have a fan diameter in a range of 330 cm to 380 cm. In such embodiments, the torsional stiffness of the planetary carrier can be greater than or equal to 6.0 x 10 8 Nm / rad, or in a range of 6.0 x 10 8 Nm / rad to 2.2 x 10 10 Nm / rad. Additionally or alternatively, in such embodiments, the radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 1.4 x 10 9 N / m, or in a range of 1.4 x 10 9 N / m to 6.0 x 10 11 N / m.

[0281] The fan shaft mounting structure can have a torsional stiffness greater than or equal to 1.50 x 10 7 Nm / rad. The fan shaft mounting structure can have a torsional stiffness greater than or equal to 2.1 x 10 7 Nm / rad. The fan shaft mounting structure can have a torsional stiffness in a range of 1.5 x 10 7 Nm / rad to 2.70 x 10 10The anti-tilting stiffness of the fan shaft mounting structure can be in the range of Nm / rad. 7 Nm / rad to 1×10 10 Within the range of Nm / rad.

[0282] According to a nineteenth aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan shaft to drive the fan at a lower rotational speed than the spindle, the gearbox being a planetary gearbox including a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier, wherein the planet carrier is arranged such that the plurality of planetary gears are mounted thereon; and

[0283] A fan shaft mounting structure is arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two support bearings connected to the fan shaft, and wherein:

[0284] The torsional stiffness of the planetary carrier is greater than or equal to 1.60 × 10⁻⁶. 8 Nm / rad; and

[0285] The overturning stiffness of the fan shaft mounting structure is greater than or equal to 1.50 × 10⁻⁶. 7 Nm / rad.

[0286] The torsional stiffness of the planetary carrier can be greater than or equal to 2.7 × 10⁻⁶. 8 Nm / rad. The torsional stiffness of the planetary carrier can reach 1.60 × 10⁻⁶ Nm / rad. 8 Nm / rad to 1.00×10 11 The torsional stiffness of the planetary carrier can be in the range of Nm / rad. 8 Nm / rad to 1×10 10 Within the range of Nm / rad.

[0287] The overturning stiffness of the fan shaft mounting structure can be greater than or equal to 2.1 × 10⁻⁶. 7 Nm / rad. The anti-tilting stiffness of the fan shaft mounting structure can reach 1.50 × 10⁻⁶. 7 Nm / rad up to 2.70×10 10 The anti-tilting stiffness of the fan shaft mounting structure can be in the range of Nm / rad. 7 Nm / rad to 1×10 10 Within the range of Nm / rad.

[0288] The fan can have a fan diameter in a range of 240 cm to 280 cm. In such embodiments, the torsional stiffness of the planet carrier can be greater than or equal to 1.8 x 10 8 Nm / rad, or in a range of 1.8 x 10 8 Nm / rad to 4.8 x 10 9 Nm / rad. Additionally or alternatively, in such embodiments, the tilt stiffness of the fan shaft mounting structure can be greater than or equal to 2.1 x 10 7 Nm / rad, or in a range of 2.1 x 10 7 Nm / rad to 1.9 x 10 10 Nm / rad.

[0289] Alternatively, the fan can have a fan diameter in a range of 330 cm to 380 cm. In such embodiments, the torsional stiffness of the planet carrier can be greater than or equal to 6.0 x 10 8 Nm / rad, or in a range of 6.0 x 10 8 Nm / rad to 2.2 x 10 10 Nm / rad. Additionally or alternatively, in such embodiments, the tilt stiffness of the fan shaft mounting structure can be greater than or equal to 3.8 x 10 7 Nm / rad, or in a range of 3.8 x 10 7 Nm / rad to 2.7 x 10 10 Nm / rad.

[0290] Any one or more of the following can apply to either or both of the preceding two aspects (e.g., the eighteenth aspect and / or the nineteenth aspect):

[0291] The first planet carrier stiffness ratio:

[0292]

[0293] may be greater than or equal to 7.0 x 10 -3 . The first planet carrier stiffness ratio can be greater than or equal to 7.0 x 10 -2 . The first planet carrier stiffness ratio can be in a range of 7.0 x 10 -3 to 1.9 x 10 3 . The first planet carrier stiffness ratio can be in a range of 7.0 x 10 -2 to 9.0 x 10 1 .

[0294] The first planet carrier stiffness product:

[0295] (effective linear torsional stiffness of the planet carrier) x (radial bending stiffness of the fan shaft mounting structure)

[0296] It can be greater than or equal to 2.9 × 10 18 (N / m) 2 The product of the stiffnesses of the first planetary carrier can be greater than or equal to 5.0 × 10⁻⁶. 18 (N / m) 2 The product of the first planetary carrier stiffness can be 2.9 × 10⁻⁶. 18 (N / m) 2 Up to 8.0×10 22 (N / m) 2 Within the range. The product of the first planetary carrier stiffness can be 5.0 × 10⁻⁶. 18 (N / m) 2 Up to 8.0×10 21 (N / m) 2 Within the range.

[0297] The stiffness ratio of the second planetary carrier (defined as:

[0298]

[0299] It can be greater than or equal to 6.0 × 10 -3 The stiffness ratio of the second planetary carrier can be greater than or equal to 6.0 × 10⁻⁶. -2 The stiffness ratio of the second planetary carrier can be 6.0 × 10⁻⁶. -3 Up to 7.0×10 3 Within the range. The stiffness ratio of the second planetary carrier can be 6.0 × 10. -2 Up to 7.0×10 2 Within the range.

[0300] The product of the stiffness of the second planetary carrier is defined as:

[0301] (Torsional stiffness of the planetary carrier) × (Overturning stiffness of the fan shaft mounting structure)

[0302] And it can be greater than or equal to 2.4 × 10 15 (Nm / rad) 2 The product of the stiffnesses of the second planetary carrier can be greater than or equal to 4.9 × 10⁻⁶. 15 (Nm / rad) 2 The product of the stiffness of the second planetary carrier can be 2.4 × 10⁻⁶. 15 (Nm / rad) 2 Up to 2.7×10 21 (Nm / rad) 2 Within the range. The product of the stiffness of the second planetary carrier can be 4.9 × 10. 15 (Nm / rad) 2 Up to 2.0×10 20 (Nm / rad) 2 Within the range.

[0303] The power transmitted by the gearbox can be greater than or equal to 2.25 x 10 7 W. The power transmitted by the gearbox can be greater than or equal to 2.5 x 10 7 W. The power transmitted by the gearbox can be in the range of 2.25 x 10 7 W to 1.00 x 10 8 W. The power transmitted by the gearbox can be in the range of 2.5 x 10 7 W to 8.0 x 10 7 W.

[0304] The moment of inertia of the fan can be greater than or equal to 7.40 x 10 7 kgm 2 . The moment of inertia of the fan can be greater than or equal to 8.3 x 10 7 kgm 2 . The moment of inertia of the fan can be in the range of 7.40 x 10 7 kgm 2 to 9.00 x 10 8 kgm 2 . The moment of inertia of the fan can be in the range of 8.3 x 10 7 kgm 2 to 6.5 x 10 8 kgm 2 .

[0305] The at least two support bearings can include a first support bearing and a second support bearing. Both the first support bearing and the second support bearing can be located at a position forward of the gearbox. The first support bearing can be located at a position forward of the gearbox and the second support bearing can be located at a position rearward of the gearbox.

[0306] The fan shaft mounting structure can further include a third support bearing. The third support bearing can be located between the fan and the gearbox.

[0307] The fan shaft can include a gearbox output shaft forming a relatively flexible portion of the fan shaft and the fan shaft mounting structure can include a gearbox output shaft support structure having at least one gearbox output shaft bearing arranged to support the gearbox output shaft.

[0308] The fan shaft mounting structure can further include one or more soft mounted non-support bearings.

[0309] Any one or more of the bearings provided as part of the fan shaft mounting structure can be a double bearing.

[0310] According to a twentieth aspect, there is provided a propulsor for an aircraft, the propulsor 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 for receiving input from the power unit via a sun shaft and outputting drive to a fan shaft so as to drive the fan at a lower rotational speed than the sun shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a carrier arranged for mounting the plurality of planet gears thereon; and a fan shaft mounting structure arranged to mount the fan shaft within the propulsor, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0311] the torsional stiffness of the carrier is greater than or equal to 1.6 x 10 8 Nm / rad, and the radial bending stiffness of the fan shaft mounting structure is greater than or equal to 7.00 x 10 8 N / m.

[0312] The propulsor of the twentieth aspect can have some or all of the features described above in relation to the eighteenth aspect of a gas turbine engine, and in some embodiments can be a gas turbine engine.

[0313] According to a twenty-first aspect, there is provided a propulsor for an aircraft, the propulsor 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 for receiving input from the power unit via a sun shaft and outputting drive to a fan shaft so as to drive the fan at a lower rotational speed than the sun shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a carrier arranged for mounting the plurality of planet gears thereon; and a fan shaft mounting structure arranged to mount the fan shaft within the propulsor, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0314] the torsional stiffness of the carrier is greater than or equal to 1.6 x 10 8 Nm / rad, and the torsional stiffness of the fan shaft mounting structure is greater than or equal to 1.50 x 10 7 Nm / rad.

[0315] The propulsor of the twenty-first aspect can have some or all of the features described above in relation to the nineteenth aspect of a gas turbine engine, and in some embodiments can be a gas turbine engine.

[0316] Features of the twentieth and twenty-first aspects can be combined. According to a twenty-second aspect, there is provided a propulsor for an aircraft, the propulsor 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 input from the power unit via a core shaft and to 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 carrier arranged to mount the plurality of planet gears thereon; and a fan shaft mounting structure arranged to mount the fan shaft within the propulsor, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0317] a) the torsional stiffness of the carrier is greater than or equal to 1.6 x 10 8 Nm / rad and the radial bending stiffness of the fan shaft mounting structure is greater than or equal to 7.00 x 10 8 N / m; and / or

[0318] b) the torsional stiffness of the carrier is greater than or equal to 1.6 x 10 8 Nm / rad and the tilt stiffness of the fan shaft mounting structure is greater than or equal to 1.50 x 10 7 Nm / rad.

[0319] The propulsor of the twenty-second aspect can have some or all of the features described above in relation to the eighteenth or nineteenth aspects of gas turbine engine, and in some embodiments can be a gas turbine engine.

[0320] According to a twenty-third 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the core shaft and outputting 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 carrier, wherein the carrier is arranged to have the plurality of planet gears mounted thereon; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0321] the torsional stiffness of the carrier is greater than or equal to 1.6 x 10 8 Nm / rad and the radial bending stiffness of the fan shaft mounting structure is greater than or equal to 7.00 x 108 N / m,

[0322] The method includes operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0323] The method of the twenty-third aspect can be a method of operating the gas turbine engine or propulsor of the eighteenth aspect or the twentieth aspect respectively. Thus, any of the features, ratios and parameters introduced above in connection with the eighteenth aspect or the twentieth aspect also apply to the twenty-third aspect.

[0324] According to a twenty-fourth 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a 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, wherein the planet carrier is arranged to have the plurality of planet gears mounted thereon; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0325] the torsional stiffness of the planet carrier is greater than or equal to 1.6 x 10 8 Nm / rad, and the tilt stiffness of the fan shaft mounting structure is greater than or equal to 1.50 x 10 7 Nm / rad,

[0326] The method includes operating the gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0327] The method of the twenty-fourth aspect can be a method of operating the gas turbine engine or propulsor of the nineteenth aspect or the twenty-first aspect respectively. Thus, any of the features, ratios and parameters introduced above in connection with the nineteenth aspect or the twenty-first aspect also apply to the twenty-fourth aspect.

[0328] The twenty-third and twenty-fourth aspects can be combined. According to a twenty-fifth 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a 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, wherein the planet carrier is arranged to have the plurality of planet gears mounted thereon; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0329] a) the torsional stiffness of the planet carrier is greater than or equal to 1.6 x 10 8 Nm / rad, and the radial bending stiffness of the fan shaft mounting structure is greater than or equal to 7.00 x 10 8 N / m; and / or

[0330] b) the torsional stiffness of the planet carrier is greater than or equal to 1.6 x 10 8 Nm / rad, and the pitch stiffness of the fan shaft mounting structure is greater than or equal to 1.50 x 10 7 Nm / rad,

[0331] The method comprises operating the gas turbine engine under cruise conditions to provide propulsive thrust for the aircraft.

[0332] The method of the twenty-fifth aspect can be a method of operating the gas turbine engine or propulsor of the eighteenth, nineteenth, or twentieth aspects. Thus, any of the features, ratios, and parameters described above in connection with the eighteenth, nineteenth, or twenty-second aspects also apply to the twenty-fourth aspect.

[0333] In the eighteenth and nineteenth aspects described above, the torsional stiffness of the carrier can alternatively be defined as the effective linear torsional stiffness of the carrier, as defined elsewhere herein. The same applies to any of the twentieth to twenty-fifth aspects.

[0334] The inventors have discovered that designing a gas turbine engine such that the torsional stiffness of the mount and the radial bending or tilt stiffness of the fan shaft mounting structure are within specified ranges can achieve high propulsive efficiency. The inventors have discovered that radial stiffness of the fan shaft mounting structure within specified ranges provides improved positioning of the fan so as to reduce any performance losses due to fan tip clearance control issues. The inventors have also discovered that by designing the gearbox such that the torsional stiffness of the mount is within specified ranges, the overall weight of the gearbox can be minimized so as to help maintain a low specific fuel consumption (SFC).

[0335] In other aspects, instead of an absolute value range, the product or ratio of the torsional stiffness of the mount and the radial bending or tilt stiffness of the fan shaft mounting structure can be specified, or both the product or ratio and the absolute value range can be specified.

[0336] According to one such aspect, the eighteenth aspect introduced above can be reduced to an aspect of 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a drive to a fan shaft 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 mount the plurality of planet gears thereon; and

[0337] a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0338] a) a first planet carrier stiffness ratio:

[0339]

[0340] greater than or equal to 7.0 x 10 -3 , greater than or equal to 7.0 x 10 -2 , in a range from 7.0 x 10 -3 to 1.9 x 10 3 , or in a range from 7.0 x 10 -2 to 9.0 x 10 1 ; and / or

[0341] b) a first planet carrier stiffness product:

[0342] (effective linear torsional stiffness of the planet carrier) x (radial bending stiffness of the fan shaft mounting structure)

[0343] greater than or equal to 2.9 x 10 18 (N / m) 2 greater than or equal to 5.0 x 10 18 (N / m) 2 in a range of 2.9 x 10 18 (N / m) 2 to 8.0 x 10 22 (N / m) 2 in a range of 5.0 x 10 18 (N / m) 2 to 8.0 x 10 21 (N / m) 2 in a range of 6.0 x 10 -3 greater than or equal to 6.0 x 10 -2 in a range of 6.0 x 10 -3 to 7.0 x 10 3 in a range of 6.0 x 10 -2 to 7.0 x 10 2 ; and / or

[0344] According to another such aspect, the nineteenth aspect introduced above can be said to provide an aspect of 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 upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting a 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, wherein the planet carrier is arranged to mount the plurality of planet gears thereon; and

[0345] a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two support bearings connected to the fan shaft, and wherein:

[0346] a) a second planet carrier stiffness ratio (defined as:

[0347]

[0348] greater than or equal to 6.0 x 10 -3 greater than or equal to 6.0 x 10 -2 in a range of 6.0 x 10 -3 to 7.0 x 10 3 in a range of 6.0 x 10 -2 to 7.0 x 10 2 ; and / or

[0349] b) a second planet carrier stiffness product is defined as:

[0350] (torsional stiffness of the planet carrier) x (polar stiffness of the fan shaft mounting structure),

[0351] and greater than or equal to 2.4 x 10 15 (Nm / rad)2 greater than or equal to 4.9 x 10 15 (Nm / rad) 2 in a range of 2.4 x 10 15 (Nm / rad) 2 to 2.7 x 10 21 (Nm / rad) 2 or in a range of 4.9 x 10 15 (Nm / rad) 2 to 2.0 x 10 20 (Nm / rad) 2 .

[0352] Those skilled in the art will appreciate that methods and impeller aspects can be formulated accordingly.

[0353] In any of the foregoing aspects, any one or more of the following can apply:

[0354] The turbine can be a first turbine, the compressor can be a first compressor, and the core shaft can be a first core shaft. The engine core can further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft can be arranged to rotate at a higher rotational speed than the first core shaft.

[0355] The gearbox can have a gear ratio within any of the ranges disclosed herein, for example a gear ratio in a range of 3.2 to 4.5, and optionally in a range of 3.2 to 4.0.

[0356] The gas turbine engine can have a specific thrust within any of the ranges disclosed herein, for example a specific thrust in a range of 70 N Kg -1 to 90 N Kg -1 .

[0357] The gas turbine engine can have a bypass ratio under cruise conditions within any of the ranges disclosed herein, for example a bypass ratio in a range of 12.5 to 18, and optionally in a range of 13 to 16.

[0358] The fan can have a fan diameter greater than 240 cm and less than or equal to 380 cm. The fan can have a fan diameter greater than 300 cm and less than or equal to 380 cm. The fan can have a fan diameter in a range of 240 cm to 280 cm. The fan can have a fan diameter in a range of 330 cm to 380 cm.

[0359] The method of any of the above-defined aspects can further include driving the gearbox with an input torque of:

[0360] i) greater than or equal to 10,000 Nm at cruise, and optionally 10,000 Nm to 50,000 Nm; and / or

[0361] ii) greater than or equal to 28,000 Nm at maximum takeoff conditions, and optionally 28,000 Nm to 135,000 Nm.

[0362] For any parameter or ratio of parameters X claimed or disclosed herein, the limitation on the value that X can take that is expressed as “X greater than or equal to Y” can alternatively be expressed as “1 / X less than or equal to 1 / Y”. Thus, any ratio or parameter defined in the above aspects and statements can be expressed as “1 / X less than or equal to 1 / Y” instead of “X greater than or equal to Y”. In such cases, zero can be considered the lower limit.

[0363] Various parameters of the adjustable gearbox and its mounting surfaces and / or more generally of the engine can be adjusted to allow the engine to meet the specifications of the various aspects outlined above. Notes on various such parameters are provided below.

[0364] The inventors have found that reducing the stiffness (radial bending stiffness and / or tilt stiffness) of the fan shaft mounting structure outside the ranges defined herein will result in undesirable vibrations at low modal frequencies (a skilled person will appreciate that lower modal vibrations have a greater amplitude / deflection than higher modes, and are thus more desirably avoided). This can be a function of the size of the gearbox and its configuration.

[0365] The inventors have also found that increasing the radial bending stiffness / tilt stiffness of the fan shaft mounting structure above the ranges defined herein will result in excessive weight increase with little actual performance benefit. The inventors have recognised that the maximum stiffness will be affected by the engineering limitations of the material from which the fan shaft mounting structure is made. The Young’s modulus of the material from which the fan shaft mounting structure is made (typically steel) can for example be in the range 100 GPa to 250 GPa or 105 GPa to 215 GPa, and optionally about 210 GPa. Different grades of steel or other types of metal can be selected to achieve different stiffnesses for the same size and geometry. For example, a steel with a Young’s modulus in the range 190 GPa to 215 GPa, a titanium alloy with a Young’s modulus in the range 105 GPa to 120 GPa or a metal such as titanium with a Young’s modulus of about 110 GPa can be used in various embodiments. The inventors have found that increasing the stiffness above the ranges defined herein using materials such as these will increase excessive weight with little or no actual performance benefit (for example, in terms of fan positioning as described above).

[0366] The inventors have found that reducing the radial bending stiffness and / or the tilt stiffness of the fan shaft (at the input of the fan or the output of the gear box) outside the ranges defined herein will result in undesirable dynamic effects, such as lateral vibrations. In particular, the minimum stiffness defined by the ranges specified herein allows for a reduction or avoidance of vibrations at low modal frequencies (as will be understood by those skilled in the art, vibrations at lower modes have greater amplitude / deflection than higher modes, and thus it is more important to avoid vibrations at lower modes). This can be a function of the size of the gear box and its configuration.

[0367] The inventors have also found that the upper limit of the radial bending and / or tilt stiffness of the fan shaft is influenced by the basic properties of the material or materials from which the fan shaft is made. For example, the maximum stiffness is influenced by the engineering limits of the material from which the fan shaft is made. The Young’s modulus of the material (typically steel) from which the fan shaft is made can be, for example, in the range of 100 GPa to 250 GPa or 105 GPa to 215 GPa, and optionally about 210 GPa. Different grades of steel or other types of metal can be selected to achieve different stiffnesses for the same size and geometry. For example, a steel with a Young’s modulus in the range of 190 GPa to 215 GPa, a titanium alloy with a Young’s modulus in the range of 105 GPa to 120 GPa, or a metal such as titanium with a Young’s modulus of about 110 GPa can be used in various embodiments. The inventors have found that increasing the stiffness beyond the ranges defined herein using materials such as these will result in excessive weight without a practical gain in performance (e.g., no more practical gain in terms of positioning of the fan as described above).

[0368] For both the first bearing spacing distance and the second bearing spacing distance, the inventors have found that increasing the distance (dl) beyond the ranges defined herein will result in undesirable lateral vibrations and insufficient fan tip control from vibrations at low modal frequencies. The inventors have also found that decreasing dl beyond the ranges defined herein will result in design space issues, for example, making it difficult to fit the gear box within the engine architecture. For example, the inventors have considered the need to fit other components within the engine. The inventors have found that the ranges specified herein for the distance provide a balance of these factors while giving the desired benefits of gear box isolation and fan positioning.

[0369] The inventors have found that increasing the distance (d2) between the gear box output position and the nearest support bearing forward of the gear box, beyond the ranges claimed herein, will not provide suitable isolation of the gear box and thus will not avoid the transmission of damaging loads into the gear box. As described elsewhere, this can be a function of the size and configuration of the gear box. The inventors have found that decreasing this distance below the specified ranges will result in issues related to the available design space, making it possible to fit the gear box in the engine.

[0370] With respect to the torsional stiffness of the planet carrier, the inventors have found that the upper limit of the defined range is affected by engineering constraints of the typical material chosen for the planet carrier (typically steel) and the gearbox size. The Young's modulus of the material the carrier is made of (typically steel) can be, for example, in the range of 100 GPa to 250 GPa or 105 GPa to 215 GPa, and optionally about 210 GPa. Different grades of steel or other types of metal can be chosen to achieve different stiffnesses for the same size and geometry. For example, a steel with a Young's modulus in the range of 190 GPa to 215 GPa, a titanium alloy with a Young's modulus in the range of 105 GPa to 120 GPa, or a metal such as titanium with a Young's modulus of about 110 GPa can be used in various embodiments. The inventors have found that increasing the torsional stiffness beyond the range defined herein results in excessive component weight, which minimizes the improvement in operation.

[0371] The inventors have also found that the lower limit of the range defined herein with respect to the torsional stiffness of the planet carrier is affected by the maximum allowed deflection; the inventors have recognized that displacement can create misalignments in the gears and bearings, and a certain misalignment can be tolerated, but larger displacements can detrimentally affect the operation of the engine, so the minimum stiffness can be chosen to keep the displacement within acceptable limits.

[0372] As described elsewhere herein, the present disclosure can relate to a gas turbine engine. Such a gas turbine engine can include an engine core comprising a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor. Such a gas turbine engine can include a fan (with fan blades) upstream of the engine core.

[0373] The gas turbine engine can include a gearbox receiving an input from the spool and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spool. The input to the gearbox can be directly from the spool or indirectly from the spool, such as via a spur gear shaft and / or a gear. The spool can rigidly connect the turbine and the compressor such that the turbine and the compressor rotate at the same speed (with the fan rotating at a lower speed). The output from the gearbox can be directly to the fan shaft, or indirectly to the fan shaft, such as via a spur gear shaft and / or a gear.

[0374] A gas turbine engine as described and / or claimed herein can have any suitable general architecture. For example, the gas turbine engine can have any desired number of shafts connecting turbines and compressors, e.g. one shaft, two shafts or three shafts. By way of example only, the turbine connected to the core shaft can be a first turbine, the compressor connected to the core shaft can be a first compressor, and the core shaft can be a first core shaft. The engine core can further include 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 can be arranged to rotate at a higher rotational speed than the first core shaft.

[0375] In such arrangements, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive a flow from the first compressor (e.g. directly, e.g. via a generally annular conduit).

[0376] The gearbox can be arranged to be driven by the core shaft that is configured (e.g. in use) to rotate at the lowest rotational speed, e.g. the first core shaft in the above example. For example, the gearbox can be arranged to be driven by only the core shaft that is configured (e.g. in use) to rotate at the lowest rotational speed, e.g. in the above example only the first core shaft, and not the second core shaft. Alternatively, the gearbox can be arranged to be driven by any one or more of the shafts, e.g. the first shaft and / or the second shaft in the above example.

[0377] The gearbox can be a reduction gearbox (as the output to the fan is at a lower rotational rate than the input from the core shaft). Any type of gearbox can be used. For example, the gearbox can be a “planetary” or “star” gearbox, as described in more detail elsewhere herein. The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), e.g. greater than 2.5, e.g. in the range 3 to 4.2, or 3.2 to 3.8, e.g. about 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. For example, the gear ratio can be between any two of the values in the preceding sentence. By way of example only, the gearbox can be a “star” gearbox having a gear ratio in the range 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio can be outside these ranges.

[0378] In any gas turbine engine as described and / or claimed herein, a combustor can be disposed axially downstream of the fan and one or more compressors. For example, where a second compressor is provided, the combustor can be located directly downstream of (e.g. at the outlet of) the second compressor. By way of another example, where a second turbine is provided, the flow at the outlet of the combustor can be provided to the inlet of the second turbine. The combustor can be disposed upstream of one or more turbines.

[0379] The compressor or each compressor (e.g. the first and second compressors as described above) can comprise any number of stages, for example a plurality of stages. Each stage can comprise a row of rotor blades and a row of stator vanes, which can be variable stator vanes (as the incidence angle of the row of stator vanes can be variable). The row of rotor blades and the row of stator vanes can be axially offset from one another.

[0380] The turbine or each turbine (e.g. the first and second turbines as described above) can comprise any number of stages, for example a plurality of stages. Each stage can comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes can be axially offset from one another.

[0381] Each fan blade can be defined as having a radial span extending from a root (or hub) at a radially inner gas wash location or 0% span location to a tip at a 100% span location. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip can be less than (or about) 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 can be within a range inclusive of any two of the values in the preceding sentence (i.e. these values can form an upper or lower limit), for example, within a range of 0.28 to 0.32. These ratios can be generally referred to as hub-tip ratios. Both the radius at the hub and the radius at the tip can be measured at the leading edge (or axially forward most) portion of the blade. Of course, the hub-tip ratio refers to the gas washing portion of the fan blade, i.e. the portion radially outside of any platforms.

[0382] The radius of the fan can be measured between the engine centreline and the tip of the leading edge of the fan blade. The fan diameter (which can simply be twice the fan radius) can be greater than (or about) any of: 220cm, 230cm, 240cm, 250cm (about 100 inches), 260cm, 270cm (about 105 inches), 280cm (about 110 inches), 290cm (about 115 inches), 300cm (about 120 inches), 310cm, 320cm (about 125 inches), 330cm (about 130 inches), 340cm (about 135 inches), 350cm, 360cm (about 140 inches), 370cm (about 145 inches), 380cm (about 150 inches), 390cm (about 155 inches), 400cm, 410cm (about 160 inches), or 420cm (about 165 inches). The fan diameter can be within an inclusive range defined by any two of the values in the preceding sentence (i.e. these values can form an upper or lower limit), for example within a range of 240cm to 280cm or 330cm to 380cm.

[0383] The rotational speed of the fan can vary in use. Generally, for fans with larger diameters, the rotational speed is lower. By way of non-limiting example only, the rotational speed of the fan under cruise conditions can be less than 2500rpm, for example less than 2300rpm. By way of further non-limiting example only, for engines with a fan diameter in the range 220cm to 300cm (e.g. 240cm to 280cm or 250cm to 270cm), the rotational speed of the fan under cruise conditions can be in the range 1700rpm to 2500rpm, for example in the range 1800rpm to 2300rpm, for example in the range 1900rpm to 2100rpm. By way of further non-limiting example only, for engines with a fan diameter in the range 330cm to 380cm, the rotational speed of the fan under cruise conditions can be in the range 1200rpm to 2000rpm, for example in the range 1300rpm to 1800rpm, for example in the range 1400rpm to 1800rpm.

[0384] In use of a gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tips of the fan blades to move at a speed Utip= 2πRfω, where Rf is the radius of the fan (measured between the engine centreline and the tip of the leading edge of the fan blade) and ω is the rotational speed of the fan. 尖端 The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. The fan tip loading can be defined as dH / Utip= dH / 2πRfω 尖端 2 where dH is the enthalpy rise (e.g. 1-D average enthalpy rise) across the fan, and Utip= 2πRfω is the speed of the fan tips. 尖端is the (translational) velocity of the fan tip, e.g., at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge times the angular velocity). The fan tip loading at cruise conditions can 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 can be within an inclusive range defined by any two of the values in the preceding sentence (i.e., these values can form an upper or lower limit), e.g., within a range of 0.28 to 0.31 or 0.29 to 0.3.

[0385] A gas turbine engine according to the present disclosure can 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 can be greater than (or on the order of) any of: 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 can be within an inclusive range defined by any two of the values in the preceding sentence (i.e., these values can form an upper or lower limit), e.g., within a range of 12 to 16, or a range of 13 to 15, or a range of 13 to 14. The bypass duct can be substantially annular. The bypass duct can be located radially outward of the core engine. A radially outer surface of the bypass duct can be defined by the nacelle and / or the fan casing.

[0386] The total pressure ratio of a gas turbine engine described and / or claimed herein can 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 entering the combustor). By way of non-limiting example, a gas turbine engine as described and / or claimed herein can have a total pressure ratio at cruise of greater than (or on the order of) any of: 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio can be within an inclusive range defined by any two of the values in the preceding sentence (i.e., these values can form an upper or lower limit), e.g., within a range of 50 to 70.

[0387] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow rate through the engine. At cruise conditions, the specific thrust of an engine described and / or claimed herein can be less than (or on the order of) any of: 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg-1 s or 80 Nkg -1 s. This specific thrust can be within an inclusive range defined by any two of the values in the preceding sentence (i.e., these values can form an upper or lower limit), for example, in a range of 80 Nkg -1 s to 100 Nkg -1 s, or 85 Nkg -1 s to 95 Nkg -1 s. Such engines can be particularly efficient compared to conventional gas turbine engines.

[0388] A gas turbine engine as described and / or claimed herein can have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described and / or claimed herein can produce a maximum thrust of at least (or about) any of: 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 can be within an inclusive range defined by any two of the values in the preceding sentence (i.e., these values can form an upper or lower limit). By way of example only, a gas turbine as described and / or claimed herein can be capable of producing a maximum thrust in a range of 330 kN to 420 kN, for example, 350 kN to 400 kN. The thrust mentioned above can be the maximum net thrust at standard atmospheric conditions, at sea level, plus 15°C (ambient pressure 101.3 kPa, temperature 30°C), with the engine stationary.

[0389] In use, the temperature of the flow at the inlet of the high pressure turbine can be particularly high. This temperature, which can be referred to as TET, can be measured at the outlet of the combustor, for example, just upstream of the first turbine wheel which can itself be referred to as a nozzle guide vane. At cruise, this TET can be at least (or about) any of: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K, or 1650 K. The TET at cruise can be within an inclusive range defined by any two of the values in the preceding sentence (i.e., these values can form an upper or lower limit). The maximum TET of the engine in use can be, for example, at least (or about) any of: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, or 2000 K. The maximum TET can be within an inclusive range defined by any two of the values in the preceding sentence (i.e., these values can form an upper or lower limit), for example, in a range of 1800 K to 1950 K. The maximum TET can occur, for example, in high thrust conditions, for example, in maximum take-off (MTO) conditions.

[0390] The fan blades and / or airfoil portions of the fan blades described and / or claimed herein can be manufactured from any suitable material or combination of materials. For example, at least a portion of the fan blades and / or airfoils can be manufactured at least partially from a composite material, such as a metal matrix composite and / or an organic matrix composite, such as carbon fiber. By way of further example, at least a portion of the fan blades and / or airfoils can be manufactured at least partially from a metal, such as a titanium-based metal or an aluminum-based material, such as an aluminum lithium alloy, or a steel-based material. The fan blades can include at least two regions manufactured using different materials. For example, the fan blades can have a protective leading edge that can be manufactured using a material that is better resistant to impact (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge can be manufactured, for example, using titanium or a titanium-based alloy. Thus, by way of example only, the fan blades can have a carbon fiber or an aluminum-based body with a titanium leading edge, such as an aluminum lithium alloy.

[0391] The fans as described and / or claimed herein can include a central portion from which the fan blades can extend, for example, radially. The fan blades can be attached to the central portion in any desired manner. For example, each fan blade can include a fixture that can engage with a corresponding slot in a hub (or disc). By way of example only, such a fixture can be dovetail in form that can be inserted and / or engaged with a corresponding slot in the hub / disc in order to secure the fan blade to the hub / disc. By way of further example, the fan blades can be integrally formed with the central portion. Such an arrangement can be referred to as a blade disc or a blade ring. Any suitable method can be used to manufacture such a blade disc or blade ring. For example, at least a portion of the fan blades can be machined from a block, and / or at least a portion of the fan blades can be attached to the hub / disc by welding, such as linear friction welding.

[0392] The gas turbine engines described and / or claimed herein can or can not be provided with a variable area nozzle (VAN). Such a variable area nozzle can allow the exit area of the bypass duct to vary in use. The general principles of the present disclosure can be applied to engines with or without a VAN.

[0393] The fans of the gas turbines as described and / or claimed herein can have any desired number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades.

[0394] As used herein, maximum take-off (MTO) conditions have the conventional meaning. Maximum take-off conditions can be defined as operating the engine at maximum take-off thrust at the end of the runway under International Standard Atmosphere (ISA) sea level pressure and temperature conditions + 15°C, which is typically defined as a vehicle speed of about 0.25 Mn, or between about 0.24 Mn and 0.27 Mn. Thus, the maximum take-off conditions for an engine can be defined as operating the engine at maximum take-off thrust (e.g., maximum throttle) of the engine under ISA sea level pressure and temperature + 15°C, with a fan inlet speed of 0.25 Mn.

[0395] As used herein, cruise conditions have the conventional meaning and will be readily understood by the skilled person. Thus, for a given gas turbine engine of an aircraft, the skilled person will immediately recognize that cruise conditions refer to the operating point of the engine to which the gas turbine engine is designed for attachment of the aircraft at mid-cruise of a given mission, which can be referred to in the industry as an “economic mission”. In this regard, mid-cruise is a key point in the flight cycle of the aircraft at which 50% of the total fuel burned between the top of climb and the start of descent has been burned (which can approximate the midpoint in time and / or distance between the top of climb and the start of descent). Thus, cruise conditions define the operating point of the gas turbine engine that provides a thrust that, taking into account the number of engines provided to the aircraft, will ensure steady state operation (i.e., maintaining a constant altitude and constant Mach number) of the aircraft to which the gas turbine engine is designed for attachment at mid-cruise. For example, if the engine is designed to be attached to an aircraft having two identical types of engines, at cruise conditions the engine provides half of the total thrust required for steady state operation of the aircraft at mid-cruise.

[0396] In other words, for a given gas turbine engine of an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust at mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at mid-cruise altitude) (need to provide, at a given mid-cruise Mach number, in combination with any other engine on the aircraft, steady state operation of the aircraft to which the gas turbine engine is designed for attachment). For any given gas turbine engine of an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, so the operating point of the engine at cruise conditions is well defined.

[0397] By way of example only, the forward speed under cruise conditions can be anywhere in the range from 0.7 Mach to 0.9 Mach, 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 about 0.8 Mach, about 0.85 Mach or 0.8 to 0.85. Any single speed within these ranges can be part of the cruise conditions. For some aircraft, the cruise conditions can be outside these ranges, for example below 0.7 Mach or above 0.9 Mach.

[0398] By way of example only, the cruise conditions can correspond to standard atmospheric conditions (according to International Standard Atmosphere, ISA) at an altitude in the range 10000m to 15000m, for example in the range 10000m to 12000m, for example in the range 10400m to 11600m (about 38000ft), for example in the range 10500m to 11500m, for example in the range 10600m to 11400m, for example in the range 10700m (about 35000ft) to 11300m, for example in the range 10800m to 11200m, for example in the range 10900m to 11100m, for example about 11000m. The cruise conditions can correspond to standard atmospheric conditions at any given altitude within these ranges.

[0399] By way of example only, the cruise conditions can correspond to an operating point of the engine providing a known required thrust level (for example, a value in the range 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to International Standard Atmosphere) at an altitude of 38000ft (11582m). By way of another example only, the cruise conditions can correspond to an operating point of the engine providing a known required thrust level (for example, a value in the range 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to International Standard Atmosphere) at an altitude of 35000ft (10668m).

[0400] In use, the gas turbine engine described and / or claimed herein can be operated under cruise conditions as defined elsewhere herein. Such cruise conditions can be determined by cruise conditions (for example, intermediate cruise conditions) of an aircraft on which at least one (for example 2 or 4) gas turbine engine can be installed to provide propulsive thrust.

[0401] 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 an aircraft to which the gas turbine engine has been designed for attachment. Thus, the cruise conditions according to this aspect correspond to the intermediate cruise of the aircraft, as defined elsewhere herein.

[0402] According to an aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation can be at cruise conditions (e.g. in terms of thrust, atmospheric conditions and Mach number) as defined elsewhere herein.

[0403] 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 can comprise (or can be) operation at the intermediate cruise of the aircraft, as defined elsewhere herein.

[0404] Although in the arrangements described herein the drive source for the propulsive fan is provided by a gas turbine engine, the gear box configurations disclosed herein can be applied to other forms of aircraft propulsors comprising alternative drive types, as will be appreciated by the skilled person. For example, the gear box arrangements described above can be used in aircraft propulsors comprising propulsive fans driven by electric motors. In such cases, the electric motors can be configured to operate at higher rotational speeds, and thus can have a smaller rotor diameter, and power can be more dense. The gear box configurations of the foregoing aspects can be used to reduce the rotational input speed of 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, the electric propulsion unit comprising an electric machine configured to drive a propulsive fan via a gear box, the gear box and / or its input / output / support and / or the structure by which the fan shaft drives the fan being supported as described and / or claimed herein.

[0405] The skilled person will appreciate that features or parameters described in relation to any one of the above aspects can be applied to any other aspect, unless mutually exclusive. Furthermore, any feature or parameter described herein can be applied to any aspect and / or in combination with any other feature or parameter described herein, unless mutually exclusive.

[0406] As used herein, a range of "from" a value X to a value Y, or "between" a value X and a value Y, etc., means that the range includes the boundary values; the boundary values of X and Y are included. As used herein, the term "axial plane" means a plane extending along the length of the engine, parallel to and including the axial centerline of the engine, and the term "radial plane" means a plane extending perpendicular to the axial centerline of the engine, thus including all radial lines at the axial location of the radial plane. Axial planes can also be referred to as longitudinal planes, as they extend along the length of the engine. Thus, a radial distance or an axial distance is the distance in a radial or axial plane, respectively. BRIEF DESCRIPTION OF DRAWINGS

[0407] Embodiments will now be described, by way of example only, with reference to the accompanying drawings in which:

[0408] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0409] Figure 2 is a close-up cross-sectional side view of an upstream portion of a gas turbine engine;

[0410] Figure 3 is a partial cutaway view of a gearbox for a gas turbine engine;

[0411] Figure 4 is a schematic diagram showing the radial bending stiffness of a cantilever beam;

[0412] Figure 5 is a schematic diagram showing the roll stiffness of a cantilever beam;

[0413] Figure 6 is a schematic diagram showing the torsional stiffness of a shaft;

[0414] Figure 7 is a schematic close-up cross-sectional view of a gas turbine engine around a portion of its gearbox;

[0415] Figure 8 and Figure 9 is a schematic diagram showing the radial bending stiffness of a fan shaft mounting structure;

[0416] Figure 10 is a close-up view showing only the bearings supporting the fan shaft to show the measurement of radial bending stiffness;

[0417] Figure 11 and Figure 12 is a schematic diagram showing the roll stiffness of a fan shaft mounting structure;

[0418] Figure 13 is a close-up view showing only the bearings supporting the fan shaft to show the measurement of roll stiffness;

[0419] Figures 14 to 19is a schematic diagram showing various embodiments of a fan shaft mounting structure;

[0420] Figure 20 is a schematic diagram showing the gear box output position of a gear box in a star configuration;

[0421] Figure 21 is a schematic diagram showing the gear box output position of a gear box in a planetary configuration;

[0422] Figure 22 and Figure 23 is a schematic diagram showing the fan shaft end radial bending stiffness;

[0423] Figure 24 and Figure 25 is a schematic diagram showing the fan shaft end roll stiffness;

[0424] Figure 26 is a schematic diagram showing an alternative interface between the fan shaft and the fan;

[0425] Figure 27 is a schematic diagram showing a planetary carrier;

[0426] Figure 28 is a schematic diagram showing the torsional stiffness of a carrier in side view;

[0427] Figure 29 is a schematic diagram showing the torsional stiffness of an alternative carrier in front view;

[0428] Figure 30 is a schematic diagram showing the torsional stiffness of a carrier of Figure 29 ;

[0429] Figure 31 is a schematic diagram showing a front view of a carrier including lugs;

[0430] Figure 32 shows an aircraft having a gas turbine engine attached to each wing;

[0431] Figure 33 shows a method of operating a gas turbine engine on an aircraft; and

[0432] Figure 34 shows a plot of applied load versus displacement to show a measure of component stiffness. DETAILED DESCRIPTION

[0433] Figure 1A gas turbine engine 10 is shown having a main axis of rotation 9. The engine 10 includes an air intake 12 and a propulsive fan 23 that produces two air flows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 includes, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, a combustion device 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 epicyclic gearbox 30.

[0434] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and is directed into the high pressure compressor 15 for further compression. Compressed air discharged from the high pressure compressor 15 is directed into the combustion device 16 where it is mixed with fuel and the mixture is burned. The resulting hot combustion products are then expanded through the high and low pressure turbines 17, 19 before being exhausted through the nozzle 20, driving the high and low pressure turbines to provide some of the propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 through a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0435] Figure 2 An exemplary arrangement of a geared fan gas turbine engine 10 is shown in FIG. 1. The low pressure turbine 19 (see Figure 1 ) drives a shaft 26 that is coupled to a sun or sun gear 28 of an epicyclic gear arrangement 30. Radially outward of, and intermeshing with, the sun gear 28 are a plurality of planet gears 32 that are coupled together by a carrier 34. The carrier 34 constrains the planet gears 32 for synchronous revolution about the sun gear 28, while free to rotate about their own axes. The carrier 34 is coupled via a link 36 to the fan 23 so as to drive the fan in rotation about the engine axis 9. Radially outward of, and intermeshing with, the planet gears 32 is a ring or annulus gear 38 that is coupled via a link 40 to the fixed support structure 24.

[0436] The link 36 can be referred to as a fan shaft 36 that optionally includes two or more shaft portions coupled together. For example, the fan shaft 36 can include a gearbox output shaft portion 36a that extends from the gearbox 30 and a fan portion 36b that extends between the gearbox output shaft portion and the fan 23. In Figure 1 and Figure 2In the illustrated embodiment, the gear box 30 is a planetary gear box, and the gear box output shaft portion 36a is connected to the planet carrier 34, and thus it can be referred to as a carrier output shaft 36a. In a spur gear box 30, the gear box output shaft portion 36a can be connected to the ring gear 38, and thus it can be referred to as a ring output shaft 36a. In Figure 1 and 2 In the illustrated embodiment, the fan portion 36b of the fan shaft 36 connects the gear box output shaft portion 36a to the fan 23. Thus, the output of the gear box 30 is transmitted to the fan 23 via the fan shaft 36 to cause the fan to rotate. In alternative embodiments, the fan shaft 36 can comprise a single component or more than two components. Unless otherwise indicated or apparent to those skilled in the art, anything described with respect to the engine 10 having a spur gear box 30 can equally apply to an engine having a planetary gear box 30, and vice versa.

[0437] It is noted that the terms "low pressure turbine" and "low pressure compressor" used herein can refer to the lowest pressure turbine stage and the lowest pressure compressor stage (i.e., excluding the fan 23), respectively, and / or the turbine stage and compressor stage connected together by the interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., excluding the gear box output shaft driving the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein can alternatively be referred to as "intermediate pressure turbine" and "intermediate pressure compressor". Where such alternative nomenclature is used, the fan 23 can be referred to as the first or lowest pressure compressor stage.

[0438] In Figure 3 The epicyclic gear box 30 is shown in greater detail by way of example in Figure 3 Only exemplary portions of the teeth are shown in Four planetary gears 32 are shown, but it will be apparent to those skilled in the art that more or fewer planetary gears 32 can be provided within the scope of the claimed invention. Practical applications of planetary epicyclic gear boxes 30 typically include at least three planetary gears 32.

[0439] In Figure 2 and Figure 3The epicyclic gearbox 30 shown by way of example in Figure 1 is planetary, with the carrier 34 coupled to the output shaft via links 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 can be used. By way of another example, the epicyclic gearbox 30 can be a sun and planet arrangement, with the carrier 34 held fixed, allowing the ring gear (or ring) 38 to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. By way of another alternative example, the gearbox 30 can be a differential gearbox, with both the ring gear 38 and the carrier 34 allowed to rotate. In various other example embodiments, the gearbox can be any other type of gearbox, and thus can not be an epicyclic gearbox.

[0440] It will be appreciated that Figure 2 and Figure 3 the arrangements shown in Figures 1 and 2 are merely exemplary, and various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement can be used to position the gearbox 30 in the engine 10 and / or for connecting the gearbox 30 to the engine 10. By way of another example, the connections (such as the links 36, 40 in the example) between the gearbox 30 and other components of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) can have any desired degree of stiffness or flexibility, as defined or claimed herein. By way of another example, any suitable arrangement of bearings between rotating and fixed components of the engine (e.g. between the input and output shafts from the gearbox and the fixed structure such as the gearbox housing) can be used, and the present disclosure is not limited to Figure 2 the example arrangements of Figures 1 and 2. For example, where the gearbox 30 has a sun and planet arrangement (as described above), the skilled person will readily appreciate that the arrangement of output links and support links and bearing locations is generally different to that shown by way of example in Figures 1 and 2 (e.g. as described in relation to other embodiments disclosed herein having a sun and planet gearbox arrangement). Figure 2 Figure 2 By way of another example, the gearbox 30 can be a differential gearbox, with both the ring gear 38 and the carrier 34 allowed to rotate. In various other example embodiments, the gearbox can be any other type of gearbox, and thus can not be an epicyclic gearbox.

[0441] The present disclosure therefore extends to gas turbine engines having any arrangement of gearbox type (e.g. sun and planet or planetary), support structure, input and output shaft arrangement and bearing location.

[0442] Optionally, the gearbox can drive additional and / or alternative components (e.g. an intermediate pressure compressor and / or a booster compressor).

[0443] Other gas turbine engines to which the present disclosure can apply can have alternative configurations. For example, such engines can have alternative numbers of compressors and / or turbines and / or alternative numbers of interconnecting shafts. By way of further example, Figure 1 ​The gas turbine engine shown has split nozzles 18 and 20, meaning that the flow through the bypass duct 22 has its own nozzle 18, which is separate from and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of this disclosure can also be applied 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 mixing nozzle. One or both nozzles (whether mixing or splitting) may have a fixed or variable area. While the described example relates to a turbofan engine, this disclosure is applicable to, for example, any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or, for example, a turboprop engine.

[0444] The geometry of the gas turbine engine 10 and its components are defined by a conventional shaft system, including the axial direction (aligned with the axis of rotation 9) and the radial direction (in... Figure 1 The direction from bottom to top) and the circumferential direction (perpendicular to) Figure 1 (Page in the view). The axial, radial, and circumferential directions are perpendicular to each other.

[0445] The following general definition of stiffness may be used in this article:

[0446] Radial bending stiffness

[0447] Radial bending stiffness is a measure of deformation under a given force applied in any chosen radial direction (i.e., any direction perpendicular to and through the engine axis). This radial bending stiffness reference... Figure 4 Defined based on the deformation of cantilever beam 401. For example... Figure 4 As shown, a force F applied to the free end of the beam in a direction perpendicular to the longitudinal axis of the beam causes a linear vertical deformation δ. The radial bending stiffness is the force applied for a given linear deformation, i.e., F / δ. In this application, the radial bending stiffness is relative to the engine's axis of rotation 9, and therefore relates to the resistance to linear deformation in the radial direction of the engine caused by the radial force. The beam or equivalent cantilever beam component extends along the engine's axis of rotation, the force F is applied radially perpendicular to the engine's axis of rotation, and the displacement δ is measured along the line of action of the force perpendicular to the axis of rotation. The radial bending stiffness as defined herein has SI units of N / m. In this application, unless otherwise stated, the radial bending stiffness is considered to be free-body stiffness, i.e., stiffness measured for an individual component in the cantilever structure, without other components that might affect its stiffness.

[0448] When a force is applied perpendicular to a cantilever beam at its free end, the resultant curvature is not constant, but increases toward the fixed end of the beam.

[0449] Tilt stiffness

[0450] Reference is made to Figure 5 The anti-roll stiffness is defined as the resulting deformation of the cantilever beam 401 under a moment M applied at its free end. The roll stiffness is a measure of the resistance to rotation at the point on the component where the moment is applied. From Figure 5 It can be seen that a moment applied at the free end of the cantilever beam causes a constant curvature along the length of the beam between its free end and fixed end. The applied moment M causes a rotation angle Θ at the point where the moment is applied. For any component of constant section (such as this beam), the angle Θ is constant along the length of the component. The anti-roll stiffness, as defined herein, therefore has SI units of Nm / rad.

[0451] By expressing the roll stiffness as a pair of equal and opposite forces F acting at either end of a radius, rather than a moment, and the arc displacement at that radius (i.e. the displacement measured along the circumference of a circle having that radius), the roll stiffness can be expressed as an effective linear roll stiffness of the component having the given radius. For the purpose of calculating the effective linear stiffness, an approximate or overall roll angle a can be defined. The arc displacement can be referred to as ra. The effective linear roll stiffness is given by the ratio of the effective force divided by the displacement, F / ra, and has units of N / m.

[0452] Torsional stiffness

[0453] The torsional stiffness is a measure of the deformation for a given torque. Figure 6 The definition of the torsional stiffness of a shaft 401 or other body is shown. A torque τ applied to the free end of the beam causes a rotational deformation θ (e.g. twist) along the length of the beam. The torsional stiffness is the torque applied for a given twist angle, i.e. τ / θ. The SI units of the torsional stiffness are Nm / rad.

[0454] An effective linear torsional stiffness can be determined for a component having a given radius. The effective linear torsional stiffness is defined in terms of an equivalent tangential force (the magnitude of the torque divided by the radius) applied at a point on the radius and the distance δ (the magnitude of the radius multiplied by θ) moved by the point corresponding to the rotational deformation θ of the component.

[0455] The following general definitions of other parameters can also be used herein:

[0456] Torque

[0457] A torque can also be referred to as a moment, is the rotational equivalent of a linear force, and can be thought of as a twist on an object.

[0458] The magnitude of the torque on the body, τ, depends on three quantities: the applied force (F), the lever arm vector connecting the origin to the point of force application (r), and the angle between the force and the lever arm vector (A):

[0459] τ = r x F

[0460] τ = |τ| = |r x F| = |r||F| sin A

[0461] where

[0462] τ is the torque vector and τ is the magnitude of the torque;

[0463] r is the position vector or "lever arm" vector (the vector from a selected point on the body to the point of force application);

[0464] F is the force vector;

[0465] x denotes the cross product; and

[0466] A is the angle between the force vector and the lever arm vector (so that when the force vector is perpendicular to the position vector, sin(a) is one, making τ = rF, i.e., the magnitude of the force times the distance between the selected point on the body and the point of force application).

[0467] The measure of torque is [force] x [distance] and can be expressed in units of Newton-meters (N.m).

[0468] The net torque on a body determines the rate of change of the body's angular momentum.

[0469] Moment of inertia

[0470] The moment of inertia (also known as the angular mass or rotational inertia) is the quantity that determines the torque required to change the rotational rate of a body about an axis of rotation, which is essentially equivalent to how mass determines the force required to change a particular acceleration.

[0471] The moment of inertia depends on the mass distribution of the body and the chosen axis, with larger moments of inertia requiring larger torques to change the rotational rate of the body. The measure of the moment of inertia is [mass] x [distance] 2 , and can be expressed in units of kilograms-meters (kg.m 2 ).

[0472] The moment of inertia, I, is defined as the ratio of the body's net angular momentum, L, to its angular velocity, ω, about the principal axis:

[0473]

[0474] Assuming the shape of the body does not change, its moment of inertia appears in Newton's law of motion as the ratio of the torque τ applied to the body to the angular acceleration α about the principal axis:

[0475] T = Ia

[0476] For a body constrained to rotate in a plane, only the moment of inertia about an axis normal to the plane is important, so I can be represented as a scalar value. The skilled person will appreciate that the fan of a gas turbine engine (and, more generally, the fan rotor of a gas turbine engine comprising a fan disc and blades, and optionally also a fan shaft and / or other associated components) is constrained to rotate in only one plane (a plane normal to the engine axis), and will appreciate that the moment of inertia of the fan can therefore be defined by a single scalar value.

[0477] The moment of inertia of the fan about the engine axis can therefore be measured or defined using any standard method.

[0478] For ease of understanding, more specific definitions of stiffness and other parameters relevant to the embodiments described herein are provided below.

[0479] Fan shaft mounting structure stiffness

[0480] Figure 7 One embodiment of a gas turbine engine having a star configuration gearbox is shown in FIG. 1. Reference numerals corresponding to those used when describing the embodiment of FIG. 1 are used. Figure 1 and Figure 2 Reference numerals corresponding to those used when describing the embodiment of FIG. 1 are used. Figure 7 A close-up of the engine core is shown, showing the mounting of the fan shaft 36. 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 engaged with the fan shaft at points axially spaced along the length of the engine. This fan shaft mounting structure 503 can take a variety of different forms, and can comprise one or more separate support structures provided to support the fan shaft. It can also comprise other structures provided to support the fan shaft, such as an inter-shaft bearing. It therefore includes any support structure extending between the bearings in contact with the fan shaft and stationary structure of the engine (e.g. of the engine core).

[0481] In Figure 7In the arrangement shown, the fan shaft mounting structure 503 includes two bearings, a first support bearing 506a and a second support bearing 506b, via which it is coupled to the fan shaft 36. The support bearings 506a, 506b are spaced apart along the axial length of the fan shaft 36. In the arrangement described, both support bearings 506a, 506b are provided at a location forward of the gearbox 30. In other arrangements, one of the two support bearings 506a, 506b for supporting the fan shaft 36 can be located at a location rearward of the gearbox 30, as will be described later. In other arrangements, more than two support bearings can be provided as part of or to the fan shaft mounting structure.

[0482] In addition to the support bearings 506a, 506b described above, the fan shaft mounting structure can also include additional non-support bearings. These non-support bearings can be additional soft-mounted bearings, or provided as part of the gearbox output shaft support structure, as described in more detail later. The support bearings can be defined as those bearings which have a primary function of locating the fan shaft within the engine rather than having a primary function of aligning other components such as gearbox components.

[0483] The support bearings can be considered to be those bearings which transfer an equal share of the total load, which is an order of magnitude greater than any non-support bearings. More particularly, a support bearing can be defined as any bearing which transfers a load greater than 1 / (10n) of the total load transferred by the mounting structure of which it is a component, where n is the total number of bearings provided in the mounting structure. For example, for a mounting structure having three bearings, any bearing which contributes less than 1 / 30 (i.e. (1 / 3)10) would be considered to be insignificant and therefore not to be considered a support bearing within the meaning of the present application.

[0484] Figure 7 A schematic example of a suitable fan shaft mounting structure 503 is shown. However, other forms of mounting structure can be used to support the fan shaft. The fan shaft mounting structure 503 is coupled to the fixed support structure 24 of the engine in order to provide a fixed mounting for the fan shaft within the engine (with rotation relative to the fixed structure of the engine core being provided by the bearings 506a, 506b). In the arrangement presently described, the fixed support structure 24 is an engine section stator (ESS) which serves both as a structural component to provide a fixed mounting for components such as the fan shaft 36, and as a guide vane provided to direct airflow from the fan 23. In other embodiments, the fixed support structure 24 can comprise struts extending across the core airflow path and separate stator vanes provided to direct airflow, or any other suitable fixed structure relative to which the fan shaft can be mounted.

[0485] The fan shaft mounting structure 503 is considered to include one or more components extending between the contact points between each of the bearings 506a, 506b and the fan shaft 36 and the fixed support structure 24. Any number of individual components can be provided between these points in order to provide the coupling between the fan shaft 36 and the fixed support structure 24. The fan shaft mounting structure 503 is shown schematically in Figure 7 FIG. 6 for illustrative purposes only, and other shapes and arrangements can be provided. For example, as noted above, additional bearings can be provided. These bearings can be soft-mounted non-support bearings, and / or additional bearings providing redundancy in the event of high loads being transferred by the support structure. These additional bearings can be provided forward or aft of each of the first bearing 506a and the second bearing 506b. Providing any additional bearings for redundancy can avoid primary components from entering a failure mode in response to normal operating loads (e.g., fatigue) or failure conditions (e.g., fan blade failure) under higher loads. In some embodiments, any of the bearings (e.g., the first bearing 506a and / or the second bearing 506b) can be a dual bearing formed by providing a pair of bearings in the same bearing housing. Such an arrangement can be used, for example, when the magnitude of normal operating loads would exceed the magnitude at which reliable service can be provided using a single bearing.

[0486] The fan shaft mounting structure 503 has a degree of flexibility, characterized by its radial bending stiffness and its tilt stiffness.

[0487] Fan shaft mounting structure radial bending stiffness

[0488] Referring to Figure 8 and Figure 9 The radial bending stiffness of the fan shaft mounting structure 503 is defined with reference to a radial force F applied to a point on the fan (e.g., at the junction between a fan blade and a fan hub portion) along the fan shaftward centerline Z (i.e., along the axial centerline of the fan blades forming the fan). Specifically, the radial force F is applied along the fan shaftward centerline (where the fan shaftward centerline is defined as the axial midpoint of the fan blades, e.g., the midpoint of the fan blades in the axial plane of the engine).

[0489] The application of this force results in a radial displacement δ of the contact points between the support bearings 506a, 506b and the fan shaft. Figure 9 ​The deformation of the fan shaft mounting structure 503 caused by the applied force is shown in the middle, where the shape without the force applied is shown in dashed lines for comparison. The force F is shown as being radially away from the engine axis 9, but it can equivalently be a force in a radial direction towards the centre line 9.

[0490] The radial bending stiffness of the fan shaft mounting structure 503 is defined as the force F divided by the average displacement at the support bearings provided in front of the gearbox 30 as part of the fan shaft mounting structure. These displacements are shown in the close-up view of Figure 10 , where for ease of explanation only the bearings are shown. Thus, for Figure 9 and Figure 10 the arrangement shown, the radial bending stiffness is given by the following formula:

[0491]

[0492] The unit of the radial bending stiffness of the fan shaft mounting structure is N / m.

[0493] In various embodiments, the radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 7.00 x 10 8 N / m, and optionally greater than or equal to 1.25 x 10 9 N / m.

[0494] In some embodiments, for example in embodiments where the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 7.0 x 10 8 N / m. In some embodiments, for example in embodiments where the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness of the fan shaft mounting structure can be greater than or equal to 1.4 x 10 9 N / m.

[0495] In various embodiments, the radial bending stiffness of the fan shaft mounting structure can be in the range of 7.00 x 10 8 N / m to 6.00 x 10 11 N / m, and optionally in the range of 1.25 x 10 9 N / m to 2.0 x 10 11 N / m.

[0496] In some embodiments, for example in embodiments where the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness of the fan shaft mounting structure can be in the range of 7.0 x 10 8 N / m to 5.0 x 10 11 N / m, and optionally in the range of 7.0 x 10 8 N / m to 2.3 x 10 9Within the range of N / m (and can be equal to 1.5 × 10⁻⁶) 9 N / m).

[0497] In some implementations, such as those with fan diameters in the range of 330 cm to 380 cm, the radial bending stiffness of the fan shaft mounting structure can be 1.4 × 10⁻⁶. 9 N / m to 6.0×10 11 In the range of N / m, and optionally in the range of 1.4 × 10 9 N / m to 3.0×10 9 Within the range of N / m (and can be equal to 2.2 × 10⁻⁶) 9 N / m).

[0498] Fan shaft mounting structure tilt stiffness :

[0499] The tilting stiffness of the fan shaft mounting structure 503 is determined in a manner similar to that of the radial bending stiffness, except that a torque M is applied at a point on the axial centerline of the fan blades instead of a force F. An example of how the tilting stiffness can be determined is shown in... Figure 11 and Figure 12 In Chinese, the anti-tilting stiffness of the fan shaft mounting structure represents its ability to resist applied torque. See also... Figure 11 and 12 The torque M is applied to the axial centerline of the fan blades (i.e., as described above). Figure 8 The anti-tilting stiffness is determined by calculating the linear axial angle between a pair of bearing positions where the fan shaft is supported. The application of torque M results in displacements in opposite directions at the two bearing positions, such as... Figure 13 As shown, this figure presents a close-up view, including only the bearings for ease of explanation. The straight-line angle θ is defined as the change in angle of the axis R extending through the contact point between bearings 506a, 506b and the fan shaft when a torque M is applied. (As shown in...) Figure 13 As can be seen, angle θ extends between the axis R when no torque is applied and the axis R' when a torque M is applied. The anti-tilting stiffness is defined as M / θ, with units of Nm / rad.

[0500] The tilting stiffness of the fan shaft mounting structure is defined by measuring the angle θ between the first two support bearings in the direction of rearward movement from the fan. Figure 11 , Figure 12 and Figure 13 In the arrangement shown, this corresponds to the support bearings 506a and 506b located at the front of the gearbox. In other arrangements, the displacement at the support bearing at the rear of the gearbox can be used to define the anti-tilting stiffness (e.g., if only one support bearing is provided at the front of the gearbox).

[0501] In various embodiments, the fan shaft mounting structure can have a tip-tilt stiffness greater than or equal to 1.50 x 10 7 Nm / rad, and optionally greater than or equal to 2.1 x 10 7 Nm / rad.

[0502] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft mounting structure can have a tip-tilt stiffness greater than or equal to 2.1 x 10 7 Nm / rad or greater than or equal to 2.3 x 10 7 Nm / rad. In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft mounting structure can have a tip-tilt stiffness greater than or equal to 3.8 x 10 7 Nm / rad or greater than or equal to 7.3 x 10 7 Nm / rad.

[0503] In various embodiments, the fan shaft mounting structure can have a tip-tilt stiffness in the range of 1.5 x 10 7 Nm / rad to 2.70 x 10 10 Nm / rad, and optionally in the range of 2.1 x 10 7 Nm / rad to 1 x 10 10 Nm / rad.

[0504] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft mounting structure can have a tip-tilt stiffness in the range of 2.1 x 10 7 Nm / rad to 1.9 x 10 10 Nm / rad, and optionally in the range of 2.3 x 10 7 Nm / rad to 4.3 x 10 7 Nm / rad (and can equal 3.3 x 10 7 Nm / rad).

[0505] In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft mounting structure can have a tip-tilt stiffness in the range of 3.8 x 10 7 Nm / rad to 2.7 x 10 10 Nm / rad, and optionally in the range of 7.1 x 10 7 Nm / rad to 9.1 x 10 7 Nm / rad (and can equal 8.1 x 10 7 Nm / rad).

[0506] For other arrangements of the fan shaft mounting structure, equivalent radial bending stiffness and tilting stiffness can be defined. In Figures 14 to 19 Various other arrangements of the fan shaft mounting structure 503 are shown in Figs. 1-3. These embodiments are provided by way of example. The skilled person will appreciate that other arrangements are possible and are considered to fall within the scope of the present disclosure.

[0507] Figure 14 A schematic view of an arrangement is shown in which the fan shaft 36 is supported by a fan shaft mounting structure having a first support bearing 506a and a second support bearing 506b forward of the gearbox 30. In this arrangement, the fan shaft 36 is supported by the first support bearing 506a and the second support bearing 506b. The first support bearing 506a is located forward of the gearbox 30 and the second support bearing 506b is located rearward of the gearbox 30. In this arrangement, the radial bending stiffness is determined by measuring the displacement at the first support bearing 506a. The tilting stiffness is determined by measuring the angular change in the axis connecting the first support bearing 506a and the second support bearing 506b. Figures 14 to 19 In all of the figures in Figs. 1-3, the mandrel 26 is supported by two mandrel bearings 507a, 507b. These bearings are connected to the fixed structure 24a of the engine via a mandrel support structure. In the arrangements shown in Figs. 1-3, the mandrel support structure is connected to the fixed support structure 24 (i.e. the fan shaft support structure 504). Figure 14 In the arrangement of Fig. 4, the fan shaft mounting structure 503 includes the first bearing 506a and the second bearing 506b and the structure connecting them to the fixed support structure 24 (i.e. the fan shaft support structure 504). Any additional bearing or bearings provided as part of this structure are included in the fan shaft mounting structure 503. Figure 14 In the arrangement of Fig. 5, a non-supported bearing 506a' is also provided which is mounted flexibly. As mentioned above, the radial bending stiffness or the tilting stiffness is determined without taking into account the displacement at this non-supported bearing 506a'.

[0508] Figure 15 A schematic view of an arrangement is shown in which the fan shaft is supported by a first support bearing 506a forward of the gearbox 30 and a second support bearing 506b located at a position rearward of the gearbox 30. In this embodiment, the fan shaft 36 extends through the gearbox 30 and is mounted by support bearings 506a, 506b on either side of the gearbox. Thus, the first support bearing 506a is located forward of the gearbox 30 and the second support bearing 506b is located rearward of the gearbox 30. In this embodiment, the second support bearing 506b is an inter-shaft bearing between the fan shaft 36 and the mandrel 26. For this arrangement, the radial bending stiffness is determined by measuring the displacement at the first support bearing 506a only. However, the tilting stiffness is determined by measuring the angular change in the axis connecting the first support bearing 506a and the second support bearing 506b.

[0509] Figure 16An arrangement is shown in which the fan shaft is supported by first and second support bearings 506a, 506b either side of the gearbox. In this embodiment, the fan shaft mounting structure 503 also includes a third support bearing 506c between the first support bearing 506a and the gearbox 30. In this arrangement, the radial bending stiffness of the fan shaft mounting structure is determined by measuring the average displacement at the first and third support bearings 506a, 506c. Any displacement at the second support bearing 506b is not included as this support bearing is not forward of the gearbox 30. The tilt stiffness is determined by measuring the change in angle of the axis connecting the first and third support bearings 506a, 506c as these are the first two bearings rearward of the fan.

[0510] Figure 17 A modification to the arrangement of Figure 14 is shown in which the fan shaft 36 includes a gearbox output shaft 36a. The gearbox output shaft 36a forms a flexible portion of the fan shaft (i.e. more flexible relative to the intermediate portion of the fan shaft to which it is connected, e.g. the fan portion 36a) at the end of the fan shaft 36 where it is connected to the gearbox 30. The gearbox output shaft 36a is supported by a bearing 508a which forms part of a gearbox output shaft support structure 41. In this embodiment, the fan shaft mounting structure 503 therefore includes the first and second bearings 506a, 506b and the structure connecting them to the fixed support structure 24 (i.e. the fan shaft support structure 504) and the gearbox output shaft support structure 41 (including its bearing 508a). As the bearing 508a of the gearbox output shaft support structure is a non-support bearing, it is not considered when measuring the radial bending stiffness. For this arrangement, the radial bending stiffness is again determined by measuring the average of the displacements at the first and second support bearings 506a, 506b. The tilt stiffness is also measured at the first and second bearings 506a, 506b.

[0511] Figure 18An arrangement is shown in which the gearbox output shaft 36a extends forward and rearward in the axial direction of the gearbox 30 and is supported by first and second gearbox output shaft bearings 508a, 508b on either side of the gearbox 30. Both the first and second gearbox output shaft bearings 508a, 508b form part of the gearbox output shaft support structure 41 and are therefore connected to the fixed structure of the engine. In this arrangement, the fan shaft mounting structure 503 comprises the first and second bearings 506a, 506b and the structure connecting them to the fixed support structure 24 (i.e. the fan shaft support structure 504) and the gearbox output shaft support structure 41 (including its two bearings). Again, the output shaft bearings 508a, 508b are not included in the measurement of the radial bending stiffness of the fan shaft mounting structure as they are non-supporting bearings. The roll stiffness is also measured at the first and second supporting bearings.

[0512] Figure 19 An arrangement is shown with the same gearbox output shaft support as Figure 16 but in which the second bearing 506b arranged to support the fan shaft 36 is arranged rearward of the gearbox 30 as described in relation to Figure 16 In this arrangement, the fan shaft mounting structure comprises: the fan shaft support structure 504; the gearbox support structure 41 and its bearings; and the inter-shaft bearing 506b and the components connecting it to the fixed structure of the engine 24a. The radial bending stiffness is determined by measuring the displacement of the first supporting bearing 506a. Any displacement at the second supporting bearing 506b is not included as this supporting bearing is not forward of the gearbox. Neither are any displacements at the gearbox output shaft bearings 508a, 508b included as they are non-supporting bearings. The roll stiffness is determined by measuring the angle of the axis connecting the first and second supporting bearings 506a, 506b.

[0513] Fan shaft stiffness

[0514] Reference is made to Figures 20 to 26The stiffness of the fan shaft is defined. The fan shaft 36 is defined as the torque transmitting component that extends from the output of the gearbox to the input of the fan. It therefore includes any portion of the gearbox output shaft and the fan input shaft that can be provided between these points. For the purposes of defining the stiffness of the fan shaft 36, it is considered to extend between the fan input location and the gearbox output location and to include all torque transmitting components between these points. It therefore does not include any components of the gearbox itself that transmit discrete forces rather than fan shaft torque (e.g. a planet carrier or a web plate coupled thereto). The gearbox output location can therefore be defined as the point of connection between the fan shaft 36 and the gearbox 30. The fan input location can be defined as the point of connection between the fan shaft 36 and the fan.

[0515] Referring to Figure 20 where the gearbox is in a planetary configuration, the gearbox output location 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 ring of the ring gear 38 (any connecting components extending from the outer surface of the ring are 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 centreline of the interface, as shown in Figure 22

[0516] 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 sections or links 704 that can be provided making up the fan shaft 36, as well as any connections 706 therebetween (e.g. a splined connection).

[0517] Where the gearbox 30 is in a planetary configuration, the gearbox output location is likewise defined as the point of connection between the fan shaft 36 and the gearbox 30. An example of such a configuration is shown in Figure 21 which shows a carrier comprising a front plate 34a and a rear plate 34b, with a plurality of pins 33 extending between the front and rear plates and with planetary gears mounted to the front and rear plates. The fan shaft 36 is connected to the front plate 34a via a connection 708 (e.g. a splined connection). In embodiments such as this, the gearbox output location is considered to be any point on the joint between the fan shaft 36 and the front plate 34a. The front plate 34a is considered to transmit discrete forces rather than a single torque and is therefore considered to be part of the gearbox 30 rather than the fan shaft. Figure 21 ​Only one example of one type of connection between the fan shaft and the planet carrier 34 is shown. In embodiments with different connection arrangements, the gearbox output location is still considered to be at the junction between the component that transmits torque (i.e. a portion of the fan shaft) and the component that transmits discrete forces (e.g. a portion of the gearbox). The key connection 708 is only one example of a connection that can be made between the fan shaft and the gearbox (i.e. between the fan shaft and the front plate 34b in the presently described embodiment). In other embodiments, the junction that forms the gearbox output location can be formed by, for example, a curved connection, a bolted joint, or other toothed or mechanically fixed arrangement.

[0518] Referring to Figure 22 , the fan input location is defined as the point on the fan shaft at the axial midpoint of the junction between the fan and the fan shaft. In the presently described arrangement, the fan 23 includes a support arm 23a arranged to connect the fan 23 to the fan shaft 36. The support arm 23a is connected to the fan shaft by a key coupling 36c (as shown in Figure 22 ) that extends along the length of a portion of the fan shaft 36. The fan input location is defined as the axial midpoint of this key coupling, as indicated by the Y axis in Figure 22 . The key coupling shown in Figure 22 is only one example of a coupling that can form the junction between the fan and the fan shaft. In other embodiments, for example, a curved connection, a bolted joint, or other toothed or mechanically fixed arrangement can be used.

[0519] Figure 26 An arrangement in which an alternative coupling is provided between the fan 23 and the fan shaft 36 is shown. Similar to Figure 22 , the fan 23 is coupled to the fan shaft 36 via a support arm 23a. However, in this arrangement, a flange coupling 36d is provided between the support arm 23a and the fan shaft 36. In this embodiment, the support arm 23a can be connected at the rear of the fan hub. The flange coupling 36d can be a curved coupling. In other embodiments, other forms of flange coupling can be provided. In the embodiment of Figure 26 , the fan input location is the axial midpoint of this flange coupling.

[0520] The fan shaft 36 has a degree of flexibility, characterised by its radial bending stiffness and tilting stiffness.

[0521] Fan shaft end stiffness at fan input and gearbox output :

[0522] Referring to Figures 22 to 25 , the stiffness of the fan shaft at each end where it is coupled to the fan 23 and the gearbox 30 is defined.

[0523] The fan shaft input location is defined above (as indicated by the X axis inFigure 22 The radial bending stiffness of the fan shaft 36 at the input end of the fan 23 is measured by applying a force Fl at the fan shaft 36 at the input end of the fan 23 (see FIG. 6). The fan shaft 36 is considered to be a free body and is held fixed at all bearing locations where it is supported (i.e., at the first support bearing 506a and the second support bearing 506b in the embodiment of FIG. 6). Due to the application of the force Fl, the fan shaft 36 deforms such that the gearbox output location is displaced by a distance dl (see FIG. 6). The radial bending stiffness of the fan shaft 36 at the output end of the fan 23 is then given by Fl / dl. Figure 22 Figure 23 The radial bending stiffness of the fan shaft 36 at the input end of the fan 23 is measured by applying a force Fl at the fan shaft 36 at the input end of the fan 23 (see FIG. 6). The fan shaft 36 is considered to be a free body and is held fixed at all bearing locations where it is supported (i.e., at the first support bearing 506a and the second support bearing 506b in the embodiment of FIG. 6). Due to the application of the force Fl, the fan shaft 36 deforms such that the gearbox output location is displaced by a distance dl (see FIG. 6). The radial bending stiffness of the fan shaft 36 at the output end of the fan 23 is then given by Fl / dl.

[0524] In various embodiments, the radial bending stiffness of the fan shaft at the input end of the fan can be greater than or equal to 3.00 x 10 6 N / m, and optionally greater than or equal to 6.3 x 10 6 N / m.

[0525] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness of the fan shaft at the input end of the fan can be greater than or equal to 6.4 x 10 6 N / m. In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness of the fan shaft at the input end of the fan can be greater than or equal to 6.9 x 10 6 N / m or greater than or equal to 8.9 x 10 6 N / m.

[0526] In various embodiments, the radial bending stiffness of the fan shaft at the input end of the fan can be in the range of 3.00 x 10 6 N / m to 2.00 x 10 9 N / m, and optionally in the range of 6.3 x 10 6 N / m to 1.0 x 10 9 N / m.

[0527] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness of the fan shaft at the input end of the fan can be in the range of 6.4 x 10 6 N / m to 1.0 x 10 9 N / m, and optionally in the range of 6.4 x 10 6 N / m to 7.6 x 10 6 N / m (and can equal 7.0 x 10 6 N / m.

[0528] ​In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness of the fan shaft at the input end of the fan can be in the range of 6.9 x 10 6 N / m to 2.0 x 10 9 N / m, and optionally in the range of 8.9 x 10 6 N / m to 1.1 x 10 7 N / m (and can equal 9.9 x 10 6 N / m).

[0529] The tipping stiffness of the fan shaft 36 at the input end of the fan 23 is measured by applying a moment M1 to the fan shaft at the fan shaft input location (as shown in FIG. 6A) defined above. The fan shaft 36 is again considered to be a free body and is held fixed at the locations of all bearing locations at which it is supported (i.e., the first support bearing 506a and the second support bearing 506b in the arrangement of FIG. 6A). Due to the moment M1, the fan shaft 36 deforms such that the fan shaft input location is displaced by an angular displacement θ1, as shown in FIG. 6B. The tipping stiffness of the fan shaft 36 at the input end of the fan 23 is then given by M1 / θ1. Figure 24 Figure 24 Figure 25

[0530] In various embodiments, the tipping stiffness of the fan shaft at the input end of the fan can be greater than or equal to 5.00 x 10 5 Nm / rad, and optionally greater than or equal to 9.0 x 10 5 Nm / rad.

[0531] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the tipping stiffness of the fan shaft at the input end of the fan can be greater than or equal to 9.5 x 10 5 Nm / rad. In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the tipping stiffness of the fan shaft at the input end of the fan can be greater than or equal to 1.5 x 10 6 Nm / rad or greater than or equal to 2.5 x 10 6 Nm / rad.

[0532] In various embodiments, the tipping stiffness of the fan shaft at the input end of the fan can be in the range of 5.00 x 10 5 Nm / rad to 7.00 x 10 8 Nm / rad, and optionally in the range of 9.0 x 10 5 Nm / rad to 3.5 x 10 8 Nm / rad. ​​​

[0533] In some embodiments, for example in embodiments in which the fan diameter is in the range 330 cm to 380 cm, the anti-tip stiffness of the fan shaft at the input end of the fan can be in the range 1.5 x 10 5 Nm / rad to 7.0 x 10 8 Nm / rad, and optionally in the range 2.5 x 10 5 Nm / rad to 4.5 x 10 6 Nm / rad (and can equal 3.5 x 10 6 Nm / rad).

[0534] In some embodiments, for example in embodiments in which the fan diameter is in the range 330 cm to 380 cm, the anti-tip stiffness of the fan shaft at the input end of the fan can be in the range 1.5 x 10 6 Nm / rad to 7.0 x 10 8 Nm / rad, and optionally in the range 2.5 x 10 6 Nm / rad to 4.5 x 10 6 Nm / rad (and can equal 3.5 x 10 6 Nm / rad).

[0535] The radial bending stiffness of the fan shaft 36 at the output end of the gearbox 30 is measured by applying a force F2 at the gearbox output location (as shown in Figure 22 The fan shaft 36 is considered to be a free body and is held fixed at the locations of all bearing locations at which it is supported (i.e. the first and second support bearings 506a, 506b in the arrangement of Figure 22 Due to the force F2, the fan shaft 36 deforms such that the gearbox output location is displaced by a distance δ2 (as shown in Figure 23 The radial bending stiffness of the fan shaft 36 at the output end of the gearbox is then given by F2 / δ2.

[0536] In various embodiments, the radial bending stiffness of the fan shaft at the output end of the gearbox can be greater than or equal to 4.00 x 10 6 N / m, and optionally greater than or equal to 3.7 x 10 7 N / m.

[0537] In some embodiments, for example in embodiments in which the fan diameter is in the range 330 cm to 380 cm, the radial bending stiffness of the fan shaft at the output end of the gearbox can be greater than or equal to 3.7 x 10 7N / m. In some embodiments, for example in embodiments where the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness of the fan shaft at the output of the gearbox can be greater than or equal to 3.9 x 10 7 N / m or greater than or equal to 5.0 x 10 7 N / m.

[0538] In various embodiments, the radial bending stiffness of the fan shaft at the output of the gearbox can be in the range of 4.00 x 10 6 N / m to 1.5 x 10 9 N / m, and optionally in the range of 3.7 x 10 7 N / m to 1.0 x 10 9 N / m.

[0539] In some embodiments, for example in embodiments where the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness of the fan shaft at the output of the gearbox can be in the range of 3.7 x 10 7 N / m to 5.0 x 10 8 N / m, and optionally in the range of 3.7 x 10 7 N / m to 5.0 x 10 7 N / m (and can equal 4.0 x 10 7 N / m).

[0540] In some embodiments, for example in embodiments where the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness of the fan shaft at the output of the gearbox can be in the range of 3.9 x 10 7 N / m to 1.5 x 10 9 N / m, and optionally can be in the range of 5.0 x 10 7 N / m to 9.0 x 10 7 N / m (and can equal 7.0 x 10 7 N / m).

[0541] The anti-tipping stiffness of the fan shaft 36 at the output of the gearbox 30 is measured by applying a moment M2 to the fan shaft at the gearbox output location defined above (as shown in FIG. 36). The fan shaft 36 is again considered to be a free body and is held fixed at the locations of all bearing locations at which it is supported (i.e., the first and second support bearings 506a and 506b in the embodiment of FIG. 36). Due to the moment M2, the fan shaft 36 deforms such that the gearbox output location is displaced by an angular displacement Θ2, as shown in FIG. 36. The anti-tipping stiffness of the fan shaft 36 at the output of the gearbox is then given by M2 / Θ2. Figure 24 Figure 24 Figure 25 ​​​

[0542] In various embodiments, the anti-tip stiffness of the fan shaft at the output of the gearbox can be greater than or equal to 7.00 x 10 4 Nm / rad, and optionally can be greater than or equal to 9.5 x 10 5 Nm / rad.

[0543] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the anti-tip stiffness of the fan shaft at the output of the gearbox can be greater than or equal to 9.5 x 10 5 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the anti-tip stiffness of the fan shaft at the output of the gearbox can be greater than or equal to 1.1 x 10 6 Nm / rad, and optionally can be greater than or equal to 2.6 x 10 6Nm / rad .

[0544] In various embodiments, the anti-tip stiffness of the fan shaft at the output of the gearbox can be in the range of 7.00 x 10 4 Nm / rad to 7.00 x 10 7 Nm / rad, and optionally in the range of 9.5 x 10 5 Nm / rad to 3.5 x 10 7 Nm / rad.

[0545] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the anti-tip stiffness of the fan shaft at the output of the gearbox can be in the range of 9.5 x 10 5 Nm / rad to 2.0 x 10 7 Nm / rad, and optionally in the range of 9.5 x 10 5 Nm / rad to 2.2 x 10 6 Nm / rad (and can equal 1.2 x 10 6 Nm / rad).

[0546] In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the anti-tip stiffness of the fan shaft at the output of the gearbox can be in the range of 1.1 x 10 6 Nm / rad to 7.0 x 10 7 Nm / rad, and optionally in the range of 2.6 x 10 6 Nm / rad to 4.6 x 10 6 Nm / rad (and can equal 3.6 x 10 6 Nm / rad).

[0547] Relative distance to fan shaft support bearings and between fan shaft support bearings

[0548] Referring to Figure 22 , the fan shaft end stiffness is determined at least in part based on the respective distances between the fan input location, the gearbox output location, and the bearings supporting the fan shaft. For example, the stiffness of the fan shaft at the fan input end can depend on the axial distance dl between the fan input location and the nearest supporting bearing aft of the fan (i.e., the first supporting bearing 506a in the arrangement of Figure 22 ). The axial location of the first supporting bearing 506a is taken as its axial centerline. The stiffness of the fan shaft at the gearbox output end can depend on the axial distance d2 between the gearbox output location and the nearest supporting bearing forward of the gearbox (i.e., the second supporting bearing 506b in the arrangement of Figure 22 ). The axial location of the second supporting bearing 506b is taken as its axial centerline. The nearest supporting bearing forward of the gearbox output location does not include any non-supporting flexible or soft-mounted bearings that can also be coupled to the fan shaft, as described above in connection with the measurement of the stiffness of the fan shaft mounting structure. For example, the nearest supporting bearing does not include the gearbox output shaft bearing 508a, which is included in the arrangement shown in FIGS. 1, Figure 17 , Figure 18 and Figure 19 . A “supporting bearing” thus likewise means a bearing that transmits a significant amount of load transmitted by the fan shaft support structure during normal operation, as defined elsewhere herein.

[0549] The gas turbine engine 10 can be configured such that the relative locations of the bearings (e.g., the axial distances dl and d2) provide a fan shaft end stiffness within a desired range.

[0550] In various embodiments, the distance 1 can be greater than or equal to 0.12 m, and optionally greater than or equal to 0.13 m.

[0551] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the distance dl can be greater than or equal to 0.12 m, or greater than or equal to 0.13 m. In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the distance dl can be greater than or equal to 0.13 m, or greater than or equal to 0.15 m.

[0552] In various embodiments, the distance dl can be in the range of 0.12 m to 0.40 m, and optionally in the range of 0.13 m to 0.30 m.

[0553] In some embodiments, for example in embodiments where the fan diameter is in the range 240 cm to 280 cm, the distance d1 can be in the range 0.12 m to 0.30 m, and optionally in the range 0.13 m to 0.15 m (and can equal 0.14 m).

[0554] In some embodiments, for example in embodiments where the fan diameter is in the range 330 cm to 380 cm, the distance d1 can be in the range 0.13 m to 0.40 m, and optionally in the range 0.15 m to 0.25 m (and can equal 0.20 m).

[0555] In various embodiments, the distance 2 can be greater than or equal to 0.15 m, and optionally greater than or equal to 0.16 m.

[0556] In some embodiments, for example in embodiments where the fan diameter is in the range 240 cm to 280 cm, the distance d2 can be greater than or equal to 0.15 m, or greater than or equal to 0.16 m. In some embodiments, for example in embodiments where the fan diameter is in the range 330 cm to 380 cm, the distance d2 can be greater than or equal to 0.16 m, or greater than or equal to 0.20 m.

[0557] In various embodiments, the distance d2 can be in the range 0.15 m to 0.45 m, and optionally in the range 0.16 m to 0.40 m.

[0558] In some embodiments, for example in embodiments where the fan diameter is in the range 240 cm to 280 cm, the distance d2 can be in the range 0.15 m to 0.35 m, and optionally in the range 0.16 m to 0.18 m (and can equal 0.17 m).

[0559] In some embodiments, for example in embodiments where the fan diameter is in the range 330 cm to 380 cm, the distance d2 can be in the range 0.16 m to 0.45 m, and optionally in the range 0.20 m to 0.28 m (and can equal 0.24 m).

[0560] In Figure 15 In the arrangement of FIG. 6, the distance d1 corresponds to the distance between the first support bearing 506a and the fan input location. The distance d-2 corresponds to the distance between the first support bearing 506a and the gearbox input location.

[0561] The bearing axial spacing d3 is defined as the axial distance between the first support bearing 506a and the second support bearing 506b, as shown in FIG. 6. Figure 22The distance d3 corresponds to the axial spacing between the nearest support bearing aft of the fan (i.e., the first support bearing 506a) and the nearest support bearing forward of the gearbox (i.e., the second support bearing 506b). Again, the axial positions of the first and second support bearings 506a, 506b are taken as their respective axial centerlines.

[0562] In various embodiments, the distance d3 can be in the range of 0.1 m to 0.4 m, and optionally in the range of 0.18 m to 0.32 mm, and further optionally in the range of 0.20 m to 0.30 m.

[0563] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, d3 can be in the range of 0.19 m to 0.23 m (and can equal 0.21 m).

[0564] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, d3 can be in the range of 0.26 m to 0.30 m (and can equal 0.28 m).

[0565] The axial distance d4 is defined as the axial distance between the fan input location and the gearbox output location, as can be seen in Figure 22 (e.g., between Y and X). The axial distance d 4- is equal to the sum of the distances d1, d2 and d 3- defined above.

[0566] The axial distance d3 can only be defined for arrangements in which the first and second support bearings 506a, 506b are located forward of the gearbox. However, the axial distance d 4- is defined for all embodiments. For cases in which only one support bearing is located forward of the gearbox, the distance d3 is zero. Figure 15 An example of how d1, d2 and d4 can be defined for arrangements in which a single support bearing is provided forward of the gearbox is shown in

[0567] In various embodiments, the distance 4 can be greater than or equal to 0.43 m, and optionally greater than or equal to 0.46 m.

[0568] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the distance d4 can be greater than or equal to 0.43 m, or greater than or equal to 0.48 m. In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the distance d4 can be greater than or equal to 0.56 m, or greater than or equal to 0.65 m.

[0569] In various embodiments, the distance d4may be in the range of 0.43m to 0.95m, and optionally in the range of 0.46m to 0.85m.

[0570] In some embodiments, for example in embodiments where the fan diameter is in the range of 240cm to 280cm, the distance d4may be in the range of 0.43m to 0.62m, and optionally in the range of 0.48m to 0.56m (and can equal 0.52m).

[0571] In some embodiments, for example in embodiments where the fan diameter is in the range of 330cm to 380cm, the distance d4may be in the range of 0.56m to 0.95m, and optionally in the range of 0.65m to 0.76m (and can equal 0.71m).

[0572] Relative fan and gearbox position

[0573] Again referring to Figure 7 , the fan-gearbox axial distance 110 is defined as the axial distance between the output end of the gearbox (i.e., the axial position P of the gearbox output location) and the axial centerline Q of the fan. The axial centerline of the fan is defined as the axial midpoint of the fan blades forming the fan (and can correspond to the axis Z). The gearbox output location is defined as the connection point between the fan shaft 36 and the gearbox, as defined elsewhere herein. This gearbox output location can be defined differently for different types of gearboxes, as described elsewhere herein.

[0574] In various embodiments, the fan-gearbox axial distance can be greater than or equal to 0.35m, and optionally greater than or equal to 0.37m.

[0575] In some embodiments, for example in embodiments where the fan diameter is in the range of 240cm to 280cm, the fan-gearbox axial distance can be greater than or equal to 0.38m, or greater than or equal to 0.40m. In some embodiments, for example in embodiments where the fan diameter is in the range of 330cm to 380cm, the fan-gearbox axial distance can be greater than or equal to 0.48m or greater than or equal to 0.50m.

[0576] In various embodiments, the fan-gearbox axial distance can be in the range of 0.35m to 0.8m, and optionally in the range of 0.37m to 0.75m.

[0577] In some embodiments, for example in embodiments in which the fan diameter is in the range 240cm to 280cm, the fan- gearbox axial distance can be in the range 0.38m to 0.65m, and optionally can be in the range 0.40m to 0.44m (and can equal 0.42m).

[0578] In some embodiments, for example in embodiments in which the fan diameter is in the range 330cm to 380cm, the fan- gearbox axial distance can be in the range 0.48m to 0.8m, and optionally can be in the range 0.50m to 0.68m (and can equal 0.58m).

[0579] Fan moment of inertia

[0580] The fan 23 has a moment of inertia I 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 plurality of fan blades, the fan hub, and any support arms or other links provided for connecting the fan to the fan shaft. Thus, the moment of inertia includes all rotating components separate from the fan shaft. The moment of inertia is the mass moment of inertia or rotational inertia of the fan with respect to rotation about the main rotational axis 9 of the engine. Rotation of the fan will induce a gyroscopic effect, meaning that the fan shaft will tend to maintain a stable orientation of its rotational axis. However, during manoeuvring of the aircraft in which the gas turbine engine is installed, the orientation of the rotational axis of the fan shaft will 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 impact on the magnitude of the gyroscopic effect induced, and thus on the design of the fan shaft mounting structure, as discussed elsewhere herein.

[0581] In various embodiments, the moment of inertia of the fan can be greater than or equal to 7.40 x 10 7 kgm 2 , and optionally can be greater than or equal to 8.3 x 10 7 kgm 2 .

[0582] In some embodiments, for example in embodiments in which the fan diameter is in the range 240cm to 280cm, the moment of inertia of the fan can be greater than or equal to 7.4 x 10 7 kgm 2 or 8.6 x 10 7 kgm 2 . In some embodiments, for example in embodiments in which the fan diameter is in the range 330cm to 380cm, the moment of inertia of the fan can be greater than or equal to 3.0 x 10 8 kgm 2or 4.0 x 10 8 kgm 2 .

[0583] In various embodiments, the moment of inertia of the fan can be in the range of 7.40 x 10 7 kgm 2 to 9.00 x 10 8 kgm 2 and optionally in the range of 8.3 x 10 7 kgm 2 to 6.5 x 10 8 kgm 2 .

[0584] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the moment of inertia of the fan can be in the range of 7.4 x 10 7 kgm 2 to 1.5 x 10 8 kgm 2 and optionally can be in the range of 8.6 x 10 7 kgm 2 to 9.6 x 10 7 kgm 2 (and can equal 9.1 x 10 7 kgm 2 ).

[0585] In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the moment of inertia of the fan can be in the range of 3.0 x 10 8 kgm 2 to 9.0 x 10 8 kgm 2 and optionally can be in the range of 4.0 x 10 8 kgm 2 to 5.0 x 10 8 kgm 2 (and can equal 4.5 x 10 8 kgm 2 ).

[0586] Gearbox transmitted power

[0587] During engine operation, power is delivered by the gearbox. The power values for power delivered by the gearbox given herein are the power delivered by the gearbox under maximum takeoff conditions. Maximum takeoff conditions are as defined elsewhere herein. Power delivered by the gearbox is defined as the power at the gearbox output location as defined elsewhere herein.

[0588] In various embodiments, the power delivered by the gearbox can be greater than or equal to 2.25 x 10 7 W, and optionally greater than or equal to 2.5 x 10 7 W.

[0589] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the power delivered by the gearbox can be greater than or equal to 2.25 x 10 7 W or greater than or equal to 2.7 x 10 7 W. In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the power delivered by the gearbox can be greater than or equal to 4.0 x 10 7 W or greater than or equal to 5.0 x 10 7 W.

[0590] In various embodiments, the power delivered by the gearbox can be in the range of 2.25 x 10 7 W to 1.00 x 10 8 W, and optionally in the range of 2.5 x 10 7 W to 8.0 x 10 7 W.

[0591] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the power delivered by the gearbox can be in the range of 2.25 x 10 7 W to 3.6 x 10 7 W, and optionally in the range of 2.7 x 10 7 W to 3.3 x 10 7 W (and can equal 3.0 x 10 7 W).

[0592] In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the power delivered by the gearbox can be in the range of 4.0 x 10 7 W to 1.0 x 10 8 W, and optionally in the range of 5.0 x 10 7 W to 6.0 x 10 7 W (and can equal 5.5 x 10 7 W).

[0593] Rack torsional stiffness

[0594] Referring to Figures 27 to 31 defines the torsional stiffness of the gearbox support. Figure 27A close-up view of a planet carrier in a gearbox arranged in any configuration of the engine described herein is shown. As already described, the planet carrier 34 holds the planetary gears 32 in place. In a planetary gearbox 30, a relatively large centrifugal force may exist that will react to the carrier 34. In a planetary gearbox, the centrifugal force to be reacted is negligible or zero because there is no carrier rotation.

[0595] In the described arrangement, the planetary carrier 34 includes a front plate 34a and a rear plate 34b, and a pin 33 extending between the two plates, as shown. Figure 27 As shown. Pin 33 is arranged parallel to the engine axis 9. In an alternative embodiment, only plate 34b may be provided on one side of the frame 34, and no plate or only a portion of the plate may be provided on the other side.

[0596] exist Figure 31 In the arrangement shown, the frame 34 also includes a lug 34c (also called a wedge or web) located between the front plate 34a and the rear plate 34b. The lug 34c increases the overall rigidity of the frame 34.

[0597] The stiffness of the frame 34 is chosen to be relatively high to resist centrifugal forces and / or maintain gear alignment. Those skilled in the art will understand that stiffness is a measure of the displacement caused by any applied force or torque and may be independent of the strength of the component. Therefore, any stiffness is acceptable to resist high loads, as long as the resulting displacement is permissible. Thus, how high the stiffness is required to keep the displacement within acceptable limits depends on the position and orientation of the gears, which is commonly referred to as gear alignment (or misalignment).

[0598] The torsional stiffness of a frame is a measure of its ability to resist an applied torque τ, such as... Figure 28 (axial cross section) and Figures 29 to 31 As shown in the radial cross-section. The torque axis is parallel to the engine axis 9.

[0599] The slanted line at the rear end of the frame 30 indicates that plate 34b is considered rigid and non-rotating (like a cantilever beam mount). In embodiments with only one plate 34a, the ends of pin 33 (and lugs 34c, if present) away from the single plate 34a may alternatively be held in place.

[0600] A torque τ (at the axial midpoint of the front plate 34a) is applied to the frame 34 and causes rotational deformation θ (e.g., torsion) along the length of the frame 34. This torsion causes the frame 34 to “roll up” when the ends of the pins 33 (and, if present, lugs 34c) are held on the frame plates 34a, 34b with a fixed radius.

[0601] The angle through which a point on the front plate 34a moving on an imaginary circle 902 passing through the rotational axis of each pin 33 is θ, where θ is the angle measured in radians. The imaginary circle 902 can be referred to as the pin circle diameter (PCD). The PCD can be in the range of 0.38m to 0.65m, for example equal to 0.4m or 0.55m. Thus, as noted above, the radius r (for example, as shown) of the imaginary circle 902 can be used to define an effective linear torsional stiffness for the spider 34. Figure 30

[0602] In various embodiments, the torsional stiffness of the spider 34 is greater than or equal to 1.60 x 10 8 Nm / rad, and optionally greater than or equal to 2.7 x 10 8 Nm / rad.

[0603] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240cm to 280cm, the torsional stiffness of the spider 34 can be greater than or equal to 1.8 x 10 8 Nm / rad, and optionally can be greater than or equal to 2.5 x 10 8 Nm / rad (and optionally can be equal to 4.83 x 10 8 Nm / rad). In some embodiments, for example in embodiments in which the fan diameter is in the range of 330cm to 380cm, the torsional stiffness of the spider 34 can be greater than or equal to 6.0 x 10 8 Nm / rad, and optionally can be greater than or equal to 1.1 x 10 9 Nm / rad (and optionally can be equal to 2.17 x 10 9 Nm / rad).

[0604] In various embodiments, the torsional stiffness of the spider 34 is in the range of 1.60 x 10 8 Nm / rad to 1.00 x 10 11 Nm / rad, and optionally in the range of 2.7 x 10 8 Nm / rad to 1 x 10 10 Nm / rad.

[0605] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240cm to 280cm, the torsional stiffness of the spider 34 can be in the range of 1.8 x 10 8 Nm / rad to 4.8 x 10 9 Nm / rad, and optionally can be in the range of 2.5 x 10 8 Nm / rad to 6.5 x 10 8 Nm / rad (and optionally can be equal to 4.83 x 10 8 ​Nm / rad). In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the torsional stiffness of the shelf 34 can be in the range of 6.0 x 10 8 Nm / rad to 2.2 x 10 10 Nm / rad, and optionally can equal 2.17 x 10 9 Nm / rad to 3.0 x 10 9 Nm / rad (and optionally can equal 2.88 x 10 9 Nm / rad).

[0606] In various embodiments, the effective linear torsional stiffness of the shelf 34 can be greater than or equal to 7.00 x 10 9 N / m, and optionally can be greater than or equal to 9.1 x 10 9 N / m.

[0607] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the effective linear torsional stiffness of the shelf 34 can be greater than or equal to 7.70 x 10 9 N / m. In other such embodiments, the effective linear torsional stiffness of the shelf 34 can be greater than or equal to 9.1 x 10 9 N / m, optionally greater than or equal to 1.1 x 10 10 N / m (and optionally can equal 1.26 x 10 10 N / m). In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the effective linear torsional stiffness of the shelf 34 can be greater than or equal to 1.2 x 10 10 N / m, and optionally can be greater than or equal to 2.1 x 10 10 N / m (and optionally can equal 2.88 x 10 10 N / m).

[0608] In various embodiments, the effective linear torsional stiffness of the shelf 34 can be in the range of 7.00 x 10 9 N / m to 1.20 x 10 11 N / m, and optionally in the range of 9.1 x 10 9 N / m to 8.0 x 10 10 N / m.

[0609] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the effective linear torsional stiffness of the shelf 34 can be in the range of 9.1 x 10 9 N / m to 6.0 x 10 10 N / m, and optionally can be in the range of 7 x 109 N / m to 2 x 10 10 N / m or 8.5 x 10 9 N / m to 2 x 10 10 N / m (and optionally can equal 1.26 x 10 10 N / m). In some embodiments, for example in embodiments in which the fan diameter is in the range 330 cm to 380 cm, the effective linear torsional stiffness of the shelf 34 can be in the range 1.2 x 10 10 N / m to 1.2 x 10 11 N / m (and optionally can equal 1.26 x 10 10 N / m to 5.0 x 10 10 N / m (and optionally can equal 2.88 x 10 10 N / m).

[0610] The torsional stiffness of the shelf 34 can be controlled by adjusting one or more parameters, including the shelf material, the shelf geometry, and whether or not lugs are present, to be in the desired range.

[0611] Parameter ratio

[0612] The inventors have found that ratios (and / or products) of some properties have a considerable effect on the operation of the gearbox and its input / output / support structure. Some or all of the following can apply to any embodiment:

[0613] The system radial bending stiffness is defined by combining the radial bending stiffness of the fan shaft mounting structure 503 and the radial bending stiffness of the fan shaft 36 at the output of the gearbox in series. The system radial bending stiffness is defined as:

[0614]

[0615] where K1 is the radial bending stiffness of the fan shaft mounting structure, K2 is the radial bending stiffness of the fan shaft 36 at the output of the gearbox, which is as defined elsewhere herein.

[0616] In various embodiments, the system radial bending stiffness can be greater than or equal to 3.90 x 10 6 N / m, and optionally greater than or equal to 3.6 x 10 7 N / m.

[0617] In some embodiments, for example in embodiments in which the fan diameter is in the range 240 cm to 280 cm, the system radial bending stiffness can be greater than or equal to 3.6 x 10 7N / m. In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the system radial bending stiffness can be greater than or equal to 4.1 x 10 7 N / m or greater than or equal to 5.8 x 10 7 N / m.

[0618] In various embodiments, the system radial bending stiffness can be in the range of 3.90 x 10 6 N / m to 1.40 x 10 9 N / m, and optionally in the range of 3.6 x 10 7 N / m to 6.8 x 10 8 N / m.

[0619] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the system radial bending stiffness can be in the range of 3.6 x 10 7 N / m to 4.0 x 10 8 N / m, and optionally in the range of 3.6 x 10 7 N / m to 4.9 x 10 7 N / m (and can equal 3.9 x 10 7 N / m).

[0620] In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the system radial bending stiffness can be in the range of 4.1 x 10 7 N / m to 1.4 x 10 9 N / m, and optionally in the range of 5.8 x 10 7 N / m to 7.8 x 10 7 N / m (and can equal 6.8 x 10 7 N / m).

[0621] The fan shaft mounting radial bending stiffness ratio can be defined as:

[0622]

[0623] In various embodiments, the fan shaft mounting radial bending stiffness ratio can be greater than or equal to 1.0 x 10 -3 , and optionally greater than or equal to 5.0 x 10 -3 , or further optionally greater than or equal to 2.0 x 10 -2 .

[0624] In various embodiments, the fan shaft mounting radial bending stiffness ratio can be in the range of 1.0 x 10 -3 to 4.0 x 10 -1and optionally greater than or equal to 4.0 x 10 -3 (N / m) -1 and optionally greater than or equal to 4.0 x 10 -3 (N / m) -2 and optionally greater than or equal to 4.0 x 10 -2 (N / m) -1 .

[0625] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft mounting radial bending stiffness ratio can be in the range of 2.2 x 10 -2 to 3.2 x 10 -2 (N / m) -2 .

[0626] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft mounting radial bending stiffness ratio can be in the range of 2.6 x 10 -2 to 3.6 x 10 -2 (N / m) -2 .

[0627] In various embodiments, in addition to or in lieu of the fan shaft mounting radial bending stiffness ratio, a product of parameters that make up the fan shaft mounting radial bending stiffness ratio can be defined. This product, referred to as the fan shaft mounting radial bending stiffness product, can be defined as:

[0628] System radial bending stiffness x radial bending stiffness of the fan shaft mounting structure (503)

[0629] In various embodiments, the fan shaft mounting radial bending stiffness product can be greater than or equal to 2.7 x 10 15 (N / m) 2 and optionally greater than or equal to 4.0 x 10 15 (N / m) 2 .

[0630] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft mounting radial bending stiffness product can be greater than or equal to 4.3 x 10 16 (N / m) 2 . In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft mounting radial bending stiffness product can be greater than or equal to 4.3 x 10 x 10 16 (N / m) 2 .

[0631] In various implementation schemes, the product of the fan shaft mounting radial bending stiffness can be 2.7 × 10⁻⁶. 15 (N / m) 2 Up to 9.0×10 19 (N / m) 2 Within the range, and optionally within 4.0 × 10 15 (N / m) 2 Up to 1.5×10 19 (N / m) 2 Within the range.

[0632] In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the product of the fan shaft mounting radial bending stiffness can be 4.3 × 10⁻⁶. 16 (N / m) 2 Up to 3.0×10 18 (N / m) 2 Within the range.

[0633] In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the product of the fan shaft mounting radial bending stiffness can be 4.3 × 10⁻⁶. 16 (N / m) 2 Up to 9.0×10 19 (N / m) 2 Within the range.

[0634] The system tilt stiffness is defined by combining the tilt stiffness of the combined fan shaft mounting structure 503 and the tilt stiffness of the fan shaft 36 at the output end of the tandem gearbox. The system tilt stiffness is defined as follows:

[0635]

[0636] Where K3 is the anti-tilting stiffness of the fan shaft mounting structure, and K4 is the anti-tilting stiffness of the fan shaft 36 at the output end of the gearbox, which is defined elsewhere in this document.

[0637] In various implementation schemes, the system's anti-tilting stiffness can be greater than or equal to 1.10 × 10⁻⁶. 5 Nm / rad, and optionally greater than or equal to 8.5 × 10 5 Nm / rad.

[0638] In some implementations, such as those where the fan diameter is in the range of 240 cm to 280 cm, the system's tilt stiffness can be greater than or equal to 8.5 × 10⁻⁶. 5Nm / rad. In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the system anti-yaw stiffness can be greater than or equal to 1.5 x 10 6 Nm / rad or greater than or equal to 2.9 x 10 6 Nm / rad.

[0639] In various embodiments, the system anti-yaw stiffness can be in the range of 1.10 x 10 5 Nm / rad to 6.80 x 10 7 Nm / rad, and optionally in the range of 8.5 x 10 5 Nm / rad to 3.4 x 10 7 Nm / rad.

[0640] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the system anti-yaw stiffness can be in the range of 8.5 x 10 5 Nm / rad to 1.2 x 10 7 Nm / rad, and optionally in the range of 8.5 x 10 5 Nm / rad to 1.7 x 10 6 Nm / rad (and can equal 1.2 x 10 6 Nm / rad).

[0641] In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the system anti-yaw stiffness can be in the range of 1.5 x 10 6 Nm / rad to 6.8 x 10 7 Nm / rad, and optionally in the range of 2.9 x 10 6 Nm / rad to 3.9 x 10 6 Nm / rad (and can equal 3.4 x 10 6 Nm / rad).

[0642] In various embodiments, the fan shaft mounting anti-yaw stiffness ratio is defined as:

[0643]

[0644] In various embodiments, the fan shaft mounting anti-yaw stiffness ratio can be greater than or equal to 1.5 x 10 -3 , and optionally greater than or equal to 6.0 x 10 -3 , and further optionally greater than or equal to 2.5 x 10 -2 .

[0645] In various embodiments, the fan shaft mounting anti-yaw stiffness ratio can be in the range of 1.5 x 10 -3 to 5.0 x 10 -1 , and optionally in the range of 6.0 x 10 -3 to 2.0 x 10 -1 , in the range of 6.0 x 10 -3 to 2.5 x 10 -2 , or in the range of 2.5 x 10 -2 to 2.0 x 10 -1 .

[0646] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft mounting anti-yaw stiffness ratio can be in the range of 3.0 x 10 -2 to 4.0 x 10 -2 (and can equal 3.5 x 10 -2 ).

[0647] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft mounting anti-yaw stiffness ratio can be in the range of 3.7 x 10 -2 to 4.7 x 10 -2 (and can equal 4.2 x 10 -2 ).

[0648] In various embodiments, in addition to or instead of the fan shaft mounting anti-yaw stiffness ratio, a product of the parameters making up the ratio can be defined. The product, referred to as the fan shaft mounting anti-yaw stiffness product, can be defined as:

[0649] System anti-yaw stiffness x anti-yaw stiffness of the fan shaft mounting structure (503)

[0650] In various embodiments, the fan shaft mounting anti-yaw stiffness product can be greater than or equal to 1.7 x 10 12 (Nm / rad) 2 , and optionally greater than or equal to 1.6 x 10 13 (Nm / rad) 2 .

[0651] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft mounting anti-yaw stiffness product can be greater than or equal to 1.9 x 10 13 (Nm / rad) 2 . In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft mounting anti-yaw stiffness product can be greater than or equal to 3.0 x 10 13 (Nm / rad)2 .

[0652] In various embodiments, the fan shaft installation anti-tilt stiffness product can be in the range of 1.7 x 10 12 (Nm / rad) 2 to 3.0 x 10 17 (Nm / rad) 2 and optionally in the range of 1.6 x 10 13 (Nm / rad) 2 to 3.0 x 10 16 (Nm / rad) 2 .

[0653] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft installation anti-tilt stiffness product can be in the range of 1.9 x 10 13 (Nm / rad) 2 to 1.5 x 10 16 (Nm / rad) 2 .

[0654] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft installation anti-tilt stiffness product can be in the range of 3.0 x 10 13 (Nm / rad) 2 to 3.0 x 10 17 (Nm / rad) 2 .

[0655] The fan shaft radial bending stiffness ratio is defined as:

[0656]

[0657] In various embodiments, the fan shaft radial bending stiffness ratio can be greater than or equal to 6.0 x 10 -3 and optionally greater than or equal to 0.015.

[0658] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft radial bending stiffness ratio can be greater than or equal to 0.03, or greater than or equal to 0.07. In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft radial bending stiffness ratio can be greater than or equal to 0.02, or greater than or equal to 0.04.

[0659] In various embodiments, the fan shaft radial bending stiffness ratio can be in the range of 6.0 x 10 -3 to 2.5 x 10 1and optionally in the range of 0.015 to 2.5.

[0660] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft radial bending stiffness ratio can be in the range of 0.03 to 0.85, and optionally in the range of 0.07 to 0.27 (and can equal 0.17).

[0661] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft radial bending stiffness ratio can be in the range of 0.02 to 1.5, and optionally in the range of 0.04 to 0.24 (and can equal 0.14).

[0662] In various embodiments, in addition to or in lieu of the fan shaft radial bending stiffness ratio, a product of the parameters making up the ratio can be defined. The product, referred to as the fan shaft radial bending stiffness product, can be defined as:

[0663] Fan shaft (36) radial bending stiffness at the input end of the fan (23) x fan shaft (36) radial bending stiffness at the output end of the gearbox (30)

[0664] In various embodiments, the fan shaft radial bending stiffness product can be greater than or equal to 1.2 x 10 13 (N / m) 2 and optionally greater than or equal to 2.4 x 10 14 (N / m) 2 .

[0665] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft radial bending stiffness product can be greater than or equal to 2.4 x 10 14 (N / m) 2 In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft radial bending stiffness product can be greater than or equal to 5.0 x 10 13 (N / m) 2 .

[0666] In various embodiments, the fan shaft radial bending stiffness product can be in the range of 1.2 x 10 13 (N / m) 2 to 3.0 x 10 18 (N / m) 2 and optionally in the range of 2.4 x 10 14 (N / m) 2 to 3.0 x 10 17 (N / m)2 Within the range.

[0667] In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the product of the fan shaft radial bending stiffness can be 2.4 × 10⁻⁶. 14 (N / m) 2 Up to 2.7×10 15 (N / m) 2 Within the range.

[0668] In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the product of the fan shaft radial bending stiffness can be 5.0 × 10⁻⁶. 13 (N / m) 2 Up to 3.0×10 18 (N / m) 2 Within the range.

[0669] The fan shaft anti-tilting stiffness ratio is defined as:

[0670]

[0671] In various implementation schemes, the fan shaft anti-tilting stiffness ratio can be greater than or equal to 2.5 × 10⁻⁶. -2 And optionally greater than or equal to 0.05.

[0672] In some embodiments, such as those where the fan diameter is in the range of 240cm to 280cm, the fan shaft anti-tilting stiffness ratio may be greater than or equal to 0.2, or greater than or equal to 0.5. In some embodiments, such as those where the fan diameter is in the range of 330cm to 380cm, the fan shaft anti-tilting stiffness ratio may be greater than or equal to 0.1, or greater than or equal to 0.2.

[0673] In various implementation schemes, the fan shaft anti-tilting stiffness ratio can be 2.5 × 10⁻⁶. -2 Up to 3.7×10 2 Within the range, and optionally from 0.05 to 4.0 × 10⁻⁶. 1 Within the range.

[0674] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft anti-tilting stiffness ratio may be in the range of 0.2 to 5.0, and optionally in the range of 0.5 to 1.5 (and may be equal to 1.00).

[0675] In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the fan shaft anti-tilting stiffness ratio can be from 0.1 to 1.0 × 10⁻⁶.1 Within the range, and optionally within the range of 0.2 to 1.4 (and may be equal to 0.98).

[0676] In various implementations, in addition to or instead of the fan shaft anti-tilting stiffness ratio, a product of parameters constituting the ratio can be defined. This product (referred to as the fan shaft anti-tilting stiffness product) can be defined as:

[0677] Anti-tilting stiffness of fan shaft (36) at the input end of fan (23) × Anti-tilting stiffness of fan shaft (36) at the output end of gearbox (30)

[0678] In various implementation schemes, the product of the fan shaft anti-tilting stiffness can be greater than or equal to 3.5 × 10⁻⁶. 10 (Nm / rad) 2 And optionally greater than or equal to 7.2 × 10 11 (Nm / rad) 2 .

[0679] In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the product of the fan shaft anti-tilting stiffnesses may be greater than or equal to 7.2 × 10⁻⁶. 11 (Nm / rad) 2 In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the product of the fan shaft anti-tilting stiffnesses may be greater than or equal to 3.5 × 10⁻⁶. 11 (Nm / rad) 2 .

[0680] In various implementation schemes, the product of the fan shaft anti-tilting stiffness can be 3.5 × 10⁻⁶. 10 (Nm / rad) 2 Up to 5.0×10 16 (Nm / rad) 2 Within the range, and optionally within 7.2 × 10 11 (Nm / rad) 2 Up to 5.0×10 15 (Nm / rad) 2 Within the range.

[0681] In some implementations, such as those where the fan diameter is in the range of 240 cm to 280 cm, the fan shaft anti-tilting stiffness product can be 7.2 × 10⁻⁶. 11 (Nm / rad) 2 Up to 1.5×10 13 (Nm / rad) 2 Within the range.

[0682] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft anti-tip stiffness product can be in the range of 3.5 x 10 11 (Nm / rad) 2 to 5.0 x 10 16 (Nm / rad) 2 .

[0683] In various embodiments, the first bearing spacing ratio is defined as:

[0684]

[0685] In various embodiments, the first bearing spacing ratio can be greater than or equal to 1.6 x 10 -1 , and optionally greater than or equal to 1.8 x 10 -1 , greater than or equal to 1.6 x 10 -1 , or greater than or equal to 2.2 x 10 -1 .

[0686] In various embodiments, the first bearing spacing ratio can be in the range of 1.6 x 10 -1 to 3.3 x 10 -1 , and optionally in the range of 1.8 x 10 -1 to 3.0 x 10 -1 , or in the range of 1.6 x 10 -1 to 2.2 x 10 -1 , or in the range of 2.2 x 10 -1 to 3.3 x 10 -1 . The values in this paragraph and the preceding paragraph can be applicable, for example, to embodiments in which the fan diameter is in the range of 240 cm to 280 cm or in the range of 330 cm to 380 cm.

[0687] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm or in the range of 330 cm to 380 cm, the bearing spacing ratio can be in the range of 2.5 x 10 -1 to 2.9 x 10 -1 (and can be equal to 2.7 x 10 -1 ).

[0688] In various embodiments, in addition to or instead of the first bearing spacing ratio, a product of the parameters making up the ratio can be defined. The product, referred to as the first bearing spacing product, can be defined as:

[0689] the first bearing spacing distance (dl) x the axial distance between the fan input location and the gearbox output location (d4)

[0690] In various embodiments, the first bearing spacing product can be greater than or equal to 5.2 x 10 -2 m 2 , and optionally greater than or equal to 5.7 x 10 -2 m 2 , or optionally greater than or equal to 7.5 x 10 -2 .

[0691] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the first bearing spacing product can be greater than or equal to 5.2 x 10 -2 m 2 . In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the first bearing spacing product can be greater than or equal to 7.5 x 10 -2 m 2 .

[0692] In various embodiments, the first bearing spacing product can be in the range of 5.2 x 10 -2 m 2 to 2.6 x 10 -1 m 2 , and optionally in the range of 5.7 x 10 -2 m 2 to 2.4 x 10 -1 m 2 , and optionally in the range of 7.5 x 10 -2 m 2 to 2.6 x 10 -1 m 2 .

[0693] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the first bearing spacing product can be in the range of 5.2 x 10 -2 m 2 to 1.4 x 10 -1 m 2 .

[0694] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the first bearing spacing product can be in the range of 7.5 x 10 -2 m 2 to 2.6 x 10 -1 m 2 .

[0695] For embodiments in which there are at least a first support bearing and a second support bearing forward of the gearbox (and rearward of the fan input location), the second bearing spacing ratio is defined as:

[0696]

[0697] In various embodiments, the second bearing spacing ratio can be greater than or equal to 4.1 x 10 -1 , and optionally greater than or equal to 4.5 x 10 -1 , and further optionally greater than or equal to 6.0 x 10 -1 .

[0698] In various embodiments, the second bearing spacing ratio can be in the range of 4.1 x 10 -1 to 8.3 x 10 -1 , and optionally in the range of 4.5 x 10 -1 to 7.7 x 10 -1 , or in the range of 4.1 x 10 -1 to 6.0 x 10 -1 , or in the range of 6.0 x 10 -1 to 8.3 x 10 -1 . The values in this paragraph and the preceding paragraph can apply, for example, to embodiments in which the fan diameter is in the range of 240 cm to 280 cm or in the range of 330 cm to 380 cm.

[0699] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm or in the range of 330 cm to 380 cm, or in other various embodiments, the second bearing spacing ratio can be in the range of 6.4 x 10 -1 to 7.0 x 10 -1 (and can equal 6.7 x 10 -1 ).

[0700] In various embodiments, in addition to or instead of this second bearing spacing ratio, a product of the parameters making up the ratio can be defined. This product, referred to as the second bearing spacing product, can be defined as:

[0701] the first bearing spacing distance (dl) x the bearing axial spacing (d3)

[0702] In various embodiments, the second bearing spacing product can be greater than or equal to 2.0 x 10 -2 m 2 , and optionally greater than or equal to 2.3 x 10 -2 m 2 , and further optionally greater than or equal to 3.5 x 10 -2 m 2 .

[0703] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the second bearing spacing product can be greater than or equal to 2.0 x 10 -2 m 2 In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the second bearing spacing product can be greater than or equal to 3.5 x 10 -2 m 2 .

[0704] In various embodiments, the second bearing spacing product can be in the range of 2.0 x 10 -2 m 2 to 1.1 x 10 -1 m 2 , and optionally in the range of 2.3 x 10 -2 m 2 to 8.5 x 10 -2 m 2 , or can be in the range of 3.5 x 10 -2 m 2 to 1.1 x 10 -1 m 2 .

[0705] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the second bearing spacing product can be in the range of 2.0 x 10 -2 m 2 to 5.6 x 10 -2 m 2 .

[0706] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the second bearing spacing product can be in the range of 3.5 x 10 -2 m 2 to 1.1 x 10 -1 m 2 .

[0707] In various embodiments, the first planetary carrier stiffness ratio can be defined as:

[0708]

[0709] In various embodiments, the first planetary carrier stiffness ratio can be greater than or equal to 7.0 x 10 -3 , and optionally greater than or equal to 7.0 x 10 -2 .

[0710] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the first carrier stiffness ratio can be greater than or equal to 6.9, or can be greater than or equal to 7.0 or 4.0. In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the first carrier stiffness ratio can be greater than or equal to 7.0 or greater than or equal to 1.0 x 10 1 or 5.0.

[0711] In various embodiments, the first carrier stiffness ratio can be in the range of 7.0 x 10 -3 to 1.9 x 10 3 and optionally in the range of 7.0 x 10 -2 to 9.0 x 10 1 .

[0712] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the first carrier stiffness ratio can be in the range of 6.9 to 1.2 x 10 2 and optionally in the range of 7.0 to 11.0 (and can equal 8.5). In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the first carrier stiffness ratio can be in the range of 4.0 to 6.0 (and can equal 5.2).

[0713] In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the first carrier stiffness ratio can be in the range of 7.0 to 1.9 x 10 3 and optionally in the range of 1.0 x 10 1 to 2.0 x 10 1 (and can equal 1.33 x 10 1 ). In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the first carrier stiffness ratio can be in the range of 5.0 to 9.0 (and can equal 7.1).

[0714] In various embodiments, the first carrier stiffness product can be defined as:

[0715] (effective linear torsional stiffness of the carrier (34)) x (radial bending stiffness of the fan shaft mounting structure (503))

[0716] In various embodiments, the first carrier stiffness product can be greater than or equal to 2.9 x 10 18 (N / m) 2 and optionally greater than or equal to 5.0 x 10 18(N / m) 2 .

[0717] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the first planetary carrier stiffness product can be greater than or equal to 8.0 x 10 18 (N / m) 2 or greater than or equal to 9.0 x 10 18 (N / m) 2 In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the first planetary carrier stiffness product can be greater than or equal to 1.0 x 10 19 (N / m) 2 or 2.0 x 10 19 (N / m) 2 or 5.0 x 10 19 (N / m) 2 .

[0718] In various embodiments, the first planetary carrier stiffness product can be in the range of 2.9 x 10 18 (N / m) 2 to 8.0 x 10 22 (N / m) 2 and optionally in the range of 5.0 x 10 18 (N / m) 2 to 8.0 x 10 21 (N / m) 2 .

[0719] In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the first planetary carrier stiffness product can be in the range of 8.0 x 10 18 (N / m) 2 to 8.0 x 10 21 (N / m) 2 and optionally in the range of 9.0 x 10 18 (N / m) 2 to 3.0 x 10 19 (N / m) 2 (and can equal 1.1 x 10 19 (N / m) 2 or can equal 1.9 x 10 19 (N / m) 2 .

[0720] In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the first planetary carrier stiffness product can be in the range of 1.0 x 10 19 (N / m)2 to 8.0 x 10 22 (N / m) 2 to 4.0 x 10 19 (N / m) 2 to 4.0 x 10 19 (N / m) 2 to 4.0 x 10 19 (N / m) 2 to 4.0 x 10 19 (N / m) 2 to 8.0 x 10 19 (N / m 2 to 8.0 x 10 19 (N / m) 2 .

[0721] In various embodiments, the second carrier stiffness ratio can be defined as:

[0722]

[0723] In various embodiments, the second carrier stiffness ratio can be greater than or equal to 6.0 x 10 -3 , and optionally greater than or equal to 6.0 x 10 -2 .

[0724] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the second carrier stiffness ratio can be greater than or equal to 1.36 x 10 1 or 7.9. In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the second carrier stiffness ratio can be greater than or equal to 1.5 x 10 1 or greater than or equal to 1.7 x 10 1 or 0.4 x 10 1 .

[0725] In various embodiments, the second carrier stiffness ratio can be in the range of 6.0 x 10 -3 to 7.0 x 10 3 , and optionally in the range of 6.0 x 10 -2 to 7.0 x 10 2 .

[0726] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the second carrier stiffness ratio can be in the range of 1.36 x 10 1 to 7.0 x 10 2 .Within the range, and optionally within 1.36 × 10 1 Up to 2.4×10 1 Within the range (and can be equal to 1.4 × 10⁻⁶) 1 In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the stiffness ratio of the second planetary carrier may be in the range of 7.9 to 9.9 (and may be equal to 8.9).

[0727] In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the stiffness ratio of the second planetary carrier can be 1.5 × 10⁻⁶. 1 Up to 7.0×10 3 Within the range, and optionally within 1.7 × 10 1 Up to 3.7×10 1 Within the range (and can be equal to 2.7 × 10⁻⁶) 1 In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the stiffness ratio of the second planetary carrier can be 0.4 × 10⁻⁶. 1 Up to 2.4×10 1 Within the range (and can be equal to 1.4 × 10⁻⁶) 1 ).

[0728] In various implementation schemes, the product of the second planetary carrier stiffness can be defined as:

[0729] (Torsional stiffness of planetary carrier (34)) × (Overturning stiffness of fan shaft mounting structure (503))

[0730] In various implementation schemes, the product of the second planetary carrier stiffness can be greater than or equal to 2.4 × 10⁻⁶. 15 (Nm / rad) 2 And optionally greater than or equal to 4.9 × 10 15 (Nm / rad) 2 .

[0731] In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the product of the second planetary carrier stiffness may be greater than or equal to 4.9 × 10⁻⁶. 15 (Nm / rad) 2 Or 7.9×10 15 (Nm / rad) 2 Or 1.0×10 16 (Nm / rad) 2 In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the product of the second planetary carrier stiffness may be greater than or equal to 9.0 × 10⁻⁶.15 (Nm / rad) 2 or 7.4 x 10 16 (Nm / rad) 2 or 1.0 x 10 17 (Nm / rad) 2 .

[0732] In various embodiments, the second planetary carrier stiffness product can be in the range of 2.4 x 10 15 (Nm / rad) 2 to 2.7 x 10 21 (Nm / rad) 2 and optionally in the range of 4.9 x 10 15 (Nm / rad) 2 to 2.0 x 10 20 (Nm / rad) 2 .

[0733] In some embodiments, for example in embodiments where the fan diameter is in the range of 240 cm to 280 cm, the second planetary carrier stiffness product can be in the range of 4.9 x 10 15 (Nm / rad) 2 to 9.0 x 10 19 (Nm / rad) 2 and optionally in the range of 7.9 x 10 15 (Nm / rad) 2 to 1.2 x 10 16 (Nm / rad) 2 (and can equal 9.9 x 10 15 (Nm / rad) 2 ), or optionally in the range of 1.0 x 10 16 (Nm / rad) 2 to 1.2 x 10 16 (Nm / rad 2 ) (and can equal 1.6 x 10 16 (Nm / rad) 2 ).

[0734] In some embodiments, for example in embodiments where the fan diameter is in the range of 330 cm to 380 cm, the second planetary carrier stiffness product can be in the range of 9.0 x 10 15 (Nm / rad) 2 to 2.7 x 10 21 (Nm / rad) 2 and optionally in the range of 7.4 x 10 16 (Nm / rad) 2 to 1.1 x 1017 (Nm / rad) 2 in the range of 1.0 x 10 16 (Nm / rad) 2 , or optionally in the range of 1.0 x 10 17 (Nm / rad) 2 to 2.6 x 10 17 (Nm / rad 2 ) (and can equal 1.8 x 10 17 (Nm / rad) 2 ).

[0735] Figure 32 An example aircraft 1000 is shown 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 turbine 10 can be any of the embodiments described herein. Figure 32 The aircraft shown should be understood to be an aircraft to which a gas turbine engine 10 of any of the embodiments or aspects disclosed herein has been designed to be attached. The aircraft 1000 has cruise conditions corresponding to the cruise conditions defined elsewhere herein and MTO conditions corresponding to the maximum takeoff conditions defined elsewhere herein.

[0736] The present disclosure also relates to a method 2000 of operating a gas turbine engine on an aircraft (e.g., an aircraft of Figure 32 ). In Figure 33 , the method 2000 is shown. The method 2000 includes operating 2010 a gas turbine engine 10 described elsewhere herein to provide propulsion for an aircraft to which it is installed under maximum takeoff conditions. The method also includes operating 2020 the gas turbine engine to provide propulsion under cruise conditions. The gas turbine engine is operated such that any of the parameters or ratios defined herein are within the specified ranges. The cruise conditions and maximum takeoff conditions are as defined elsewhere herein.

[0737] The torque on the core shaft 26 can be referred to as an input torque, as this is the torque input to the gearbox 30. The torque provided by the turbine 19 to the core shaft under cruise conditions (i.e., the torque on the core shaft) can be greater than or equal to 10,000 Nm, and optionally greater than or equal to 11,000 Nm. In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the torque on the core shaft 26 under cruise conditions can be greater than or equal to 10,000 Nm or 11,000 Nm (and optionally can be equal to 12,760 Nm). In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the torque on the core shaft 26 under cruise conditions can be greater than or equal to 25,000 Nm, and optionally greater than or equal to 30,000 Nm (and optionally can be equal to 34,000 Nm).

[0738] The torque on the core shaft under cruise conditions can be in the range of 10,000 Nm to 50,000 Nm, and optionally in the range of 11,000 Nm to 45,000 Nm. In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the torque on the core shaft 26 under cruise conditions can be in the range of 10,000 Nm to 15,000 Nm, and optionally in the range of 11,000 Nm to 14,000 Nm (and optionally can be equal to 12,760 Nm). In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the torque on the core shaft 26 under cruise conditions can be in the range of 25,000 Nm to 50,000 Nm, and optionally in the range of 30,000 Nm to 40,000 Nm (and optionally can be equal to 34,000 Nm).

[0739] Under maximum takeoff (MTO) conditions, the torque on the core shaft 26 can be greater than or equal to 28,000 Nm, and optionally greater than or equal to 30,000 Nm. In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the torque on the core shaft 26 under MTO conditions can be greater than or equal to 28,000 Nm, and optionally greater than or equal to 35,000 Nm (and optionally can be equal to 36,300 Nm). In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the torque on the core shaft 26 under MTO conditions can be greater than or equal to 70,000 Nm, and optionally greater than or equal to 80,000 Nm or 82,000 Nm (and optionally can be equal to 87,000 Nm).

[0740] The torque on the mandrel 26 under maximum take-off (MTO) conditions can be in the range of 28,000 Nm to 135,000 Nm, and optionally in the range of 30,000 Nm to 110,000 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the torque on the mandrel 26 under MTO conditions can be in the range of 28,000 Nm to 50,000 Nm, and optionally in the range of 35,000 Nm to 38,000 Nm (and optionally can equal 36,300 Nm). In some embodiments, for example in embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the torque on the mandrel 26 under MTO conditions can be in the range of 70,000 Nm to 135,000 Nm, and optionally in the range of 80,000 Nm to 90,000 Nm or 82,000 Nm to 92,000 Nm (and optionally can equal 87,000 Nm).

[0741] Torque is in units of [force] x [distance], can be expressed in units of Newton meters (N.m), and is defined in the usual manner understood by those skilled in the art.

[0742] Figure 34 It is shown how the stiffness defined herein can be measured. Figure 34 A graph is shown of the displacement δ resulting from the application of a load L (e.g. a force, moment or torque) to a component whose stiffness is to be measured. Under load levels from zero to L G there is a non-linear region in which the displacement is caused by movement of the component (or relative movement of individual parts of the component) as it takes up the load, rather than deformation of the component (e.g. movement within a gap between parts). Above load levels of L H the elastic limit of the component has been exceeded and the applied load no longer causes elastic deformation, but rather plastic deformation or component failure can occur. Between points G and H the applied load and the resulting displacement have a linear relationship. The stiffness defined herein can be determined by measuring the gradient of the linear region between points G and H (the stiffness is the inverse of this gradient). The gradient of as large a region of the linear region as possible can be found by taking measurements with a larger displacement to increase the accuracy of the measurement. For example, the gradient can be found by applying a load equal to or slightly greater than L G and equal to or slightly less than L H Although the displacement is referred to as δ in this specification, those skilled in the art will understand that the principle of equivalence will apply to linear or angular displacement.

[0743] Unless otherwise indicated, stiffness as defined herein is for the corresponding component with the engine off (i.e. at zero speed / not in operation). Stiffness is typically not varying over the operating range of the engine; thus, the stiffness under cruise conditions of an aircraft using the engine (these cruise conditions being as defined elsewhere herein) can be the same as when the engine is not in use. However, in case stiffness varies over the operating range of the engine, the stiffness as defined herein shall be understood as the value with the engine at room temperature and not in motion.

[0744] Unless otherwise stated or where it is clear that a feature is specific to a particular gearbox type, anything described herein with reference to a planetary gearbox can equally apply to a star gearbox.

[0745] It will be appreciated that the application is not limited to the embodiments described above and various modifications and improvements can be made without departing from the concepts described herein. Any feature can be used individually or in combination with any other feature, unless mutually exclusive, and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A gas turbine engine (10) for an aircraft, comprising: an engine core (11) including a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) upstream of the engine core, the fan including a plurality of fan blades, the fan having a fan axial centerline; a gearbox (30) receiving input from the core shaft (26) and outputting drive to a fan shaft (36) via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure (503) arranged to mount the fan shaft within the engine, the fan shaft mounting structure (503) including at least two support bearings (506a, 506b) connected to the fan shaft (36), and wherein: a fan-gearbox axial distance (110) is defined as an axial distance between the output of the gearbox and the fan axial centerline, the fan-gearbox axial distance being greater than or equal to 0.35 m; and (i) a system radial bending stiffness is defined as: and a fan shaft mounting radial bending stiffness ratio: (ii) a system roll stiffness is defined as: greater than or equal to 1.0 x 10 -3 ; and / or and a fan shaft mounting roll stiffness ratio:

8. The gas turbine engine of claim 1 or claim 2, wherein: greater than or equal to 1.5 x 10 -3 .

2. A gas turbine engine as claimed in claim 1, wherein the fan shaft mounting radial bending stiffness ratio is greater than or equal to 5.0 x 10 -3 .

3. The gas turbine engine of claim 1, wherein the fan shaft has a radial bending stiffness greater than or equal to 2.0 x 10 -2 .

4. The gas turbine engine of claim 1, wherein the fan shaft installed radial bending stiffness is in a range of 1.0 x 10 -3 to 4.0 x 10 -1 .

5. The gas turbine engine of claim 1, wherein the fan shaft installed radial bending stiffness is in a range of 5.0 x 10 -3 to 1.5 x 10 -1 .

6. The gas turbine engine of claim 1, wherein the fan shaft installed radial bending stiffness is in a range of 5.0 x 10 -3 to 2.0 x 10 -2 .

7. The gas turbine engine of claim 1, wherein the fan shaft installed radial bending stiffness is in a range of 2.0 x 10 -2 to 1.5 x 10 -1 .

9. The gas turbine engine of claim 1 or claim 2, wherein: a) the system radial bending stiffness is greater than or equal to 3.90 x 10 6 N / m; and / or b) the radial bending stiffness of the fan shaft mounting structure (503) is greater than or equal to 7.00 x 10 8 N / m; and / or c) the radial bending stiffness of the fan shaft (36) at the output of the gearbox (30) is greater than or equal to 4.00 x 10 6 N / m.

10. The gas turbine engine of claim 1 or claim 2, wherein: a) the system has a radial bending stiffness greater than or equal to 3.6 x 10 7 N / m; and / or b) the radial bending stiffness of the fan shaft mounting structure (503) is greater than or equal to 1.25 x 10 9 N / m; and / or c) the radial bending stiffness of the fan shaft (36) at the output of the gearbox (30) is greater than or equal to 3.7 x 10 7 N / m. and / or a) the radial bending stiffness of the system is in the range of 3.90 x 10 6 N / m to 1.40 x 10 9 N / m; 11. The gas turbine engine of claim 1 or claim 2, wherein: b) the radial bending stiffness of the fan shaft mounting structure (503) is in the range of 7.00 x 10 8 N / m to 6.00 x 10 11 N / m; and / or c) the radial bending stiffness of the fan shaft (36) at the output of the gearbox (30) is in the range of 4.00 x 10 6 N / m to 1.5 x 10 9 N / m. and / or a) the system radial bending stiffness is in the range of 3.6 x 10 7 N / m to 6.8 x 10 8 N / m; 22. The gas turbine engine of claim 1 or claim 2, wherein: b) the radial bending stiffness of the fan shaft mounting structure (503) is in the range of 1.25 x 10 9 N / m to 2.0 x 10 11 N / m; and / or c) the radial bending stiffness of the fan shaft (36) at the output of the gearbox (30) is in the range of 3.7 x 10 7 N / m to 1.0 x 10 9 N / m.

12. A gas turbine engine (10) as claimed in claim 1 or claim 2, wherein the product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure is greater than or equal to 2.7 x 10 15 (N / m) 2 .

13. The gas turbine engine (10) of claim 1 or claim 2, wherein the product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure is greater than or equal to 4.0 x 10 15 (N / m) 2 .

14. The gas turbine engine (10) of claim 1 or claim 2, wherein the product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure is in the range of 2.7 x 10 15 (N / m) 2 to 9.0 x 10 19 (N / m) 2 .

15. A gas turbine engine (10) as claimed in claim 1 or claim 2, wherein the product of the system radial bending stiffness and the radial bending stiffness of the fan shaft mounting structure is in the range 4.0 x 10 15 (N / m) 2 to 1.5 x 10 19 (N / m) 2 .

16. A gas turbine engine as claimed in claim 1 or claim 2, wherein the fan shaft mounting tilt stiffness ratio is greater than or equal to 6.0 x 10 -3 .

17. A gas turbine engine as claimed in claim 1 or claim 2, wherein the fan shaft mounting tilt stiffness ratio is greater than or equal to 2.5 x 10 -2 .

18. A gas turbine engine as claimed in claim 1 or claim 2, wherein the fan shaft mounting pitch stiffness ratio is in the range of 1.5 x 10 -3 to 5.0 x 10 -1 .

19. A gas turbine engine as claimed in claim 1 or claim 2, wherein the fan shaft mounting pitch stiffness ratio is in the range from 6.0 x 10 -3 to 2.0 x 10 -1 .

20. A gas turbine engine as claimed in claim 1 or claim 2, wherein the fan shaft mounting pitch stiffness ratio is in the range from 6.0 x 10 -3 to 2.5 x 10 -2 .

21. A gas turbine engine as claimed in claim 1 or claim 2, wherein the fan shaft mounting pitch stiffness ratio is in the range from 2.5 x 10 -2 to 2.0 x 10 -1 .

23. The gas turbine engine of claim 1 or claim 2, wherein: a) the system roll stiffness is greater than or equal to 1.10 x 107Nm / rad 5 Nm / rad; and / or b) the anti-roll stiffness of the fan shaft mounting structure (503) is greater than or equal to 1.50 x 10 7 Nm / rad; and / or c) said anti-roll stiffness of said fan shaft (36) at said output end of said gearbox (30) is greater than or equal to 7.00 x 10 4 Nm / rad.

24. The gas turbine engine of claim 1 or claim 2, wherein: a) the system roll stiffness is greater than or equal to 8.5 x 10 5 Nm / rad; and / or b) the anti-roll stiffness of the fan shaft mounting structure (503) is greater than or equal to 2.1 x 10 7 Nm / rad; and / or c) said anti-roll stiffness of said fan shaft (36) at said output end of said gearbox (30) is greater than or equal to 9.5 x 10 5 Nm / rad.

25. The gas turbine engine of claim 1 or claim 2, wherein: a) the system roll stiffness is in the range of 1.10 x 105 5 Nm / rad to 6.80 x 105 7 Nm / rad; and / or b) the anti-roll stiffness of the fan shaft mounting structure (503) is in the range of 1.5 x 10 7 Nm / rad to 2.70 x 10 10 Nm / rad; and / or c) said anti-roll stiffness of said fan shaft (36) at said output end of said gearbox (30) is in the range of 7.00 x 10 4 Nm / rad to 7.00 x 10 7 Nm / rad. and / or a) the system roll stiffness is in the range of 8.5 x 10 5 Nm / rad to 3.4 x 10 7 Nm / rad; 30. The gas turbine engine of claim 1 or claim 2, wherein the at least two support bearings include a first support bearing (506a) and a second support bearing (506b), and wherein: b) the anti-roll stiffness of the fan shaft mounting structure (503) is in the range of 2.1 x 10 7 Nm / rad to 1 x 10 10 Nm / rad; and / or c) said anti-roll stiffness of said fan shaft (36) at said output end of said gearbox (30) is in the range of 9.5 x 10 5 Nm / rad to 3.5 x 10 7 Nm / rad.

26. A gas turbine engine (10) as claimed in claim 1 or claim 2, wherein the product of the system anti- pitch stiffness and the anti-pitch stiffness of the fan shaft mounting structure (503) is greater than or equal to 1.7 x 10 12 (Nm / rad) 2 .

27. A gas turbine engine (10) as claimed in claim 1 or claim 2, wherein the product of the system anti- pitch stiffness and the anti-pitch stiffness of the fan shaft mounting structure (503) is greater than or equal to 1.6 x 10 13 (Nm / rad) 2 .

28. A gas turbine engine (10) as claimed in claim 1 or claim 2, wherein the product of the system anti-tilt stiffness and the anti-tilt stiffness of the fan shaft mounting structure (503) is in the range 1.7 x 10 12 (Nm / rad) 2 to 3.0 x 10 17 (Nm / rad) 2 .

29. A gas turbine engine (10) as claimed in claim 1 or claim 2, wherein the product of the system anti-tilt stiffness and the anti-tilt stiffness of the fan shaft mounting structure (503) is in the range 1.6 x 10 13 (Nm / rad) 2 to 3.0 x 10 16 (Nm / rad) 2 . a) both the first and second support bearings (506a, 506b) are located at a position forward of the gearbox (30); or b) the first support bearing (506a) is located at a position forward of the gearbox (30) and the second support bearing (506b) is located at a position aft of the gearbox (30).

31. The gas turbine engine of claim 30, wherein any one or more of the following is true: a) the fan shaft mounting structure further includes a third support bearing (506c), wherein the third support bearing (506c) is located between the fan (23) and the gearbox (30); ​ b) the fan shaft (36) includes a gearbox output shaft (30a) forming a relatively flexible portion of the fan shaft, and the fan shaft mounting structure (503) includes a gearbox output shaft support structure (41) having at least one gearbox output shaft bearing (508a, 508b) arranged to support the gearbox output shaft; c) the fan shaft mounting structure (503) further includes one or more soft mounted non-supporting bearings; and / or d) any one or more of the bearings provided as part of the fan shaft mounting structure (503) are dual bearings.

32. A gas turbine engine in accordance with claim 1 or claim 2, wherein: a) an axial distance (dl) between an input end of the fan and a bearing of the at least two supporting bearings (506a, 506b) closest to the fan (23) in a rearward direction is greater than or equal to 0.12 m; and / or b) an axial distance (d2) between the output end of the gearbox (30) and a bearing of the at least two supporting bearings (506a, 506b) closest to the gearbox (30) in a forward direction is greater than or equal to 0.15 m.

33. A gas turbine engine in accordance with claim 1 or claim 2, wherein: a) an axial distance (dl) between an input end of the fan and a bearing of the at least two supporting bearings (506a, 506b) closest to the fan (23) in a rearward direction is greater than or equal to 0.13 m; and / or b) an axial distance (d2) between the output end of the gearbox (30) and a bearing of the at least two supporting bearings (506a, 506b) closest to the gearbox (30) in a forward direction is greater than or equal to 0.16 m.

34. A gas turbine engine in accordance with claim 1 or claim 2, wherein: a) an axial distance (dl) between an input end of the fan and a bearing of the at least two supporting bearings (506a, 506b) closest to the fan (23) in a rearward direction is in the range of 0.12 m to 0.40 m; and / or b) an axial distance (d2) between the output end of the gearbox (30) and a bearing of the at least two supporting bearings (506a, 506b) closest to the gearbox (30) in a forward direction is in the range of 0.15 m to 0.45 m.

35. A gas turbine engine in accordance with claim 1 or claim 2, wherein: a) an axial distance (dl) between an input end of the fan and a bearing of the at least two supporting bearings (506a, 506b) closest to the fan (23) in a rearward direction is in the range of 0.13 m to 0.30 m; and / or b) an axial distance (d2) between the output end of the gearbox (30) and a bearing of the at least two supporting bearings (506a, 506b) closest to the gearbox (30) in a forward direction is in the range of 0.16 m to 0.35 m. ​ b) an axial distance (d2) between the output of the gearbox (30) and the bearing of the at least two support bearings (506a, 506b) closest to the gearbox (30) in a forward direction is in a range of 0.16m to 0.40m.

36. The gas turbine engine of claim 1 or claim 2, wherein the fan-gearbox axial distance (110) is greater than or equal to 0.37m.

37. The gas turbine engine of claim 1 or claim 2, wherein the fan-gearbox axial distance (110) is in a range of 0.35m to 0.8m.

38. The gas turbine engine of claim 1 or claim 2, wherein the fan-gearbox axial distance (110) is in a range of 0.37m to 0.75m.

39. The gas turbine engine (10) of claim 1 or claim 2, wherein: i) the gearbox has a gear ratio of 3.2 to 4.5; and / or ii) the gas turbine engine has a specific thrust of 70 N Kg -1 to 90 N Kg -1 and / or iii) the gas turbine engine has a bypass ratio under cruise conditions of 12.5 to 18; and / or iv) the fan has a fan diameter greater than 240cm and less than or equal to 380cm.

40. The gas turbine engine (10) of claim 1 or claim 2, wherein: i) the gearbox has a gear ratio of 3.2 to 4.0; and / or iii) the gas turbine engine has a bypass ratio under cruise conditions of 13 to 16; and / or iv) the fan has a fan diameter greater than 300cm and less than or equal to 380cm.

41. The gas turbine engine of claim 1 or claim 2, wherein the gearbox (30) is an epicyclic gearbox (30) including 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.

42. A method (2000) of operating a gas turbine engine (10) for an aircraft, the gas turbine engine comprising: an engine core (11) including a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) upstream of the engine core, the fan including a plurality of fan blades, the fan having a fan axial centerline; a gearbox (30) receiving an input from the core shaft (26) and outputting a drive to a fan shaft (36) via an output of the gearbox to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure (503) arranged to mount the fan shaft within the engine, the fan shaft mounting structure (503) including at least two support bearings (506a, 506b) connected to the fan shaft (36), and wherein: a fan-gearbox axial distance (110) is defined as the axial distance between the output of the gearbox and the axial centreline of the fan, the fan-gearbox axial distance being greater than or equal to 0.35m; and (i) the system radial bending stiffness is defined as: and the fan shaft mounting radial bending stiffness ratio is: greater than or equal to 3.9 x 10 6 ; and / or (ii) the system roll stiffness is defined as: and the fan shaft mounting roll stiffness ratio is: greater than or equal to 1.5 x 10 -3 , the method (2000) comprises operating (2020) the gas turbine engine (10) under cruise conditions to provide propulsion for the aircraft.

43. The method of claim 42, further comprising driving the gearbox with an input torque of: i) greater than or equal to 10000 Nm at cruise; and / or ii) greater than or equal to 28000 Nm at maximum take-off conditions.

44. The method of claim 42, further comprising driving the gearbox with an input torque of: i) from 10000 Nm to 50000 Nm at cruise; and / or ii) from 28000 Nm to 135000 Nm at maximum take-off conditions.

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