aircraft engines

Abstract

The present invention is entitled "Aircraft Engine." The present 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 spindle (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox (30) receiving input from the spindle (26) and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox (30) being an epicyclic gearbox (30) comprising a sun gear (28), a plurality of planetary gears (32), a ring gear (38), and a planet carrier (34) arranged to mount the planetary gears (32); and a gearbox support (40) arranged to at least partially support the gearbox within the engine. The flight cycle ratio: #imgabs0# is less than or equal to 3.20. The present invention also discloses a method (2000) of operating a gas turbine engine.

Description

Technical Field

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

[0002] Gas turbine engines for aircraft propulsion have many design factors that affect overall efficiency and power output or thrust. A general purpose of a gas turbine engine is to provide low specific fuel consumption (SFC). In order to enable higher thrust with high efficiency, a fan with a larger diameter can be used. In order to facilitate the use of larger fan sizes, a gearbox is provided that has an output to a fan shaft via which the fan is driven. The gearbox receives drive from a mandrel of the turbine system connected to the engine core. This gearbox allows the fan to be operated at a reduced rotational speed compared to when a direct drive is used.

[0003] However, when manufacturing engines with larger fan diameters, the inventors have discovered that simply enlarging components of known engine types may not result in an efficient design. For example, there may be issues associated with mounting the gearbox and fan shaft within the engine. Therefore, the properties of the components used to mount the gearbox and the fan shaft need to be considered. Summary of the Invention

[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 turbine, a compressor and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting drive via an output of the gearbox to a fan shaft so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and

[0005] A gearbox support member arranged to at least partially support a gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and the ratio of radial bending stiffness to moment of inertia:

[0006]

[0007] Greater than or equal to 2.5×10 -2 Nkg -1 m -3 .

[0008] The ratio of the radial bending stiffness to the moment of inertia may be greater than or equal to 0.05 Nkg -1 m -3 The ratio of the radial bending stiffness to the moment of inertia can be 2.5×10 -2 Nkg -1 m -3 Up to 6.0Nkg -1 m -3 The ratio of the radial bending stiffness to the moment of inertia can be within the range of 0.05Nkg -1 m -3 Up to 3.0Nkg -1 m -3 The ratio of the radial bending stiffness to the moment of inertia can be within the range of 0.05Nkg -1 m -3 Up to 0.6Nkg -1 m -3 within the range.

[0009] The ratio of the radial bending stiffness of the fan shaft at the output end of the gearbox to the moment of inertia of the fan may be greater than or equal to 2.5×10 -2 Nkg -1 m -3 , greater than or equal to 0.05Nkg -1 m -3 , in 2.5x10 -2 Nkg -1 m -3 Up to 6.0Nkg -1 m -3 Within the range of 0.05Nkg -1 m -3 Up to 3.0Nkg -1 m -3 Within the range of 0.05Nkg -1 m -3 Up to 0.6Nkg -1 m -3 within the range.

[0010] The ratio of the radial bending stiffness of the gearbox support to the moment of inertia of the fan may be greater than or equal to 3.0×10 -2 Nkg -1 m -3 , greater than or equal to 0.06Nkg -1 m -3 , at 3.0×10 -2 Nkg -1 m -3 Up to 4.0Nkg -1 m -3 Within the range of 0.06Nkg -1 m-3 Up to 2.0Nkg -1 m -3 Within the range of 0.06Nkg -1 m -3 Up to 0.48Nkg -1 m -3 within the range.

[0011] The radial bending stiffness of the fan shaft at the output end of the gearbox may be greater than or equal to 4.00×10 6 N / m. The radial bending stiffness of the fan shaft at the output end of the gearbox may be greater than or equal to 3.7×10 7 N / m. The radial bending stiffness of the fan shaft at the output end of the gearbox can be 4.00×10 6 N / m to 1.50×10 9 The radial bending stiffness of the fan shaft at the output end of the gearbox can be within the range of 3.7×10 7 N / m to 1.0×10 9 In the range of N / m.

[0012] The radial bending stiffness of the gearbox support can be greater than or equal to 1.0×10 7 N / m. The radial bending stiffness of the gearbox support can be greater than or equal to 2.0×10 7 N / m. The radial bending stiffness of the gearbox support can be greater than or equal to 3.0×10 7 N / m. The radial bending stiffness of the gearbox support can be 1.0×10 7 N / m to 4.0×10 8 The radial bending stiffness of the gearbox support can be within the range of 2.0×10 7 N / m to 3.0×10 8 The radial bending stiffness of the gearbox support can be within the range of 3.0×10 7 N / m to 2.0×10 8 In the range of N / m.

[0013] The fan may have a diameter in the range of 240 cm to 280 cm. In such embodiments, the ratio of the radial bending stiffness to the moment of inertia may be greater than or equal to 0.05 Nkg -1 m -3 or at 0.05Nkg -1 m -3 Up to 4.0Nkg -1 m -3 within the range.

[0014] Alternatively, the fan may have a diameter in the range of 330 cm to 380 cm. In such embodiments, the ratio of the radial bending stiffness to the moment of inertia may be greater than or equal to 0.025 Nkg -1 m -3 or at 0.025Nkg -1 m -3 Up to 2.0Nkg -1 m -3 within the range.

[0015] The following products (for example, radial bending stiffness moment of inertia product):

[0016] Radial bending stiffness of at least one of the fan shaft and the gearbox support at the output end of the gearbox × moment of inertia of the fan

[0017] Can be greater than or equal to 2.0×10 14 Nkgm, greater than or equal to 4.0×10 14 Nkgm, greater than or equal to 2.0×10 15 Nkgm, at 2.0×10 14 Nkgm to 1.4×10 18 Nkgm range, within 4.0×10 14 Nkgm to 7.0×10 17 Nkgm, or within the range of 2.0×10 15 Nkgm to 7.0×10 17 within the range of Nkgm.

[0018] Ratio of anti-tilting stiffness to moment of inertia:

[0019]

[0020] Can be greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 The ratio of the anti-tilting stiffness to the moment of inertia may be greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m -1 The ratio of the anti-tilting stiffness to the moment of inertia can be 4.0×10 -4 Nrad- 1 kg -1 m -1 to 2.7×10 -1 Nrad -1 kg -1 m -1 The ratio of the anti-tilting stiffness to the moment of inertia can be within the range of 1.0×10 -3 Nrad-1 kg -1 m -1 to 0.1Nrad -1 kg -1 m -1 The ratio of the anti-tilting stiffness to the moment of inertia can be within the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 1.5×10 -2 Nrad -1 kg -1 m -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 comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle 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 spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and the ratio of anti-tilting stiffness to moment of inertia:

[0022]

[0023] Greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 .

[0024] The ratio of the anti-tilting stiffness to the moment of inertia may be greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m -1 The ratio of the anti-tilting stiffness to the moment of inertia can be 4.0×10 -4 Nrad -1 kg -1 m -1 to 2.7×10 -1 Nrad -1 kg -1 m -1The ratio of the anti-tilting stiffness to the moment of inertia can be within the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 0.1Nrad- 1 kg -1 m -1 The ratio of the anti-tilting stiffness to the moment of inertia can be within the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 1.5×10 -2 Nrad -1 kg -1 m -1 within the range.

[0025] The ratio of the anti-tilt stiffness of the fan shaft at the output end of the gearbox to the moment of inertia of the fan may be greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 , greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 , at 4.0×10 -4 Nrad -1 kg -1 m -1 to 2.7×10 -1 Nrad -1 kg -1 m -1 In the range of 1.0×10 -3 Nrad- 1 kg -1 m -1 to 0.1Nrad -1 kg -1 m -1 in the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 1.5×10 -2 Nrad -1 kg -1 m -1 within the range.

[0026] The ratio of the anti-tilting stiffness of the gearbox support to the moment of inertia of the fan can be greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m-1 , greater than or equal to 2.0×10 -3 Nrad -1 kg -1 m -1 , at 1.0×0 -3 Nrad -1 kg -1 m -1 to 7.0×10 -2 Nrad -1 kg -1 m -1 In the range of 2.0×10 -3 Nrad -1 kg -1 m -1 to 3.0×10 -2 Nrad -1 kg -1 m -1 in the range of 2.0×10 - 3 Nrad -1 kg -1 m -1 to 7.0×10 -3 Nrad -1 kg -1 m -1 within the range.

[0027] The anti-tilt stiffness of the fan shaft at the output end of the gearbox may be greater than or equal to 7.00×10 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 9.5×10 5 Nm / rad. The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be 7.00×10 4 Nm / rad to 7.00×10 7 The anti-tilt stiffness of the fan shaft at the output end of the gearbox can be within the range of 9.5×10 5 Nm / rad to 3.5×0 7 In the range of Nm / rad.

[0028] The anti-tilting stiffness of the gearbox support can be greater than or equal to 1.2×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be greater than or equal to 2.4×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be greater than or equal to 3.9×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be 1.2×10 5 Nm / rad to 2.1×107 The anti-tilt stiffness of the gearbox support can be within the range of 2.4×10 5 Nm / rad to 1.6×10 7 The anti-tilt stiffness of the gearbox support can be within the range of 3.9×10 5 Nm / rad to 9.0×10 6 In the range of Nm / rad.

[0029] Ratio of radial bending stiffness to moment of inertia:

[0030]

[0031] Can be greater than or equal to 2.5×10 -2 Nkg -1 m -3 The ratio of the radial bending stiffness to the moment of inertia may be greater than or equal to 0.05 Nkg -1 m -3 The ratio of the radial bending stiffness to the moment of inertia can be 2.5×10 -2 Nkg -1 m -3 Up to 6.0Nkg -1 m -3 The ratio of the radial bending stiffness to the moment of inertia can be within the range of 0.05Nkg -1 m -3 Up to 3.0Nkg -1 m -3 The ratio of the radial bending stiffness to the moment of inertia can be within the range of 0.05Nkg -1 m -3 Up to 0.6Nkg -1 m -3 within the range.

[0032] The fan may have a diameter in the range of 240 cm to 280 cm. In such an example, the ratio of the anti-tilt stiffness to the moment of inertia may be greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m -1 or at 1.0×10 -3 Nrad -1 kg -1 m -1 to 1.45×10 -2 Nrad - 1 kg -1 m -1 within the range.

[0033] Alternatively, the fan may have a diameter in the range of 330 cm to 380 cm. In such an embodiment, the ratio of the anti-tilting stiffness to the moment of inertia may be greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 or at 4.0×10 4 Nrad -1 kg -1 m -1 to 3.0×10 -2 Nrad -1 kg -1 m -1 within the range.

[0034] The following products (for example, the product of the anti-tilting stiffness and the moment of inertia):

[0035] The anti-tilting stiffness of at least one of the fan shaft and the gearbox support at the output end of the gearbox × the moment of inertia of the fan

[0036] Can be greater than or equal to 3.0×10 12 Nm 3 rad -1 kg, greater than or equal to 6.0×10 12 Nm 3 rad -1 kg, greater than or equal to 2.5×10 13 Nm 3 rad -1 kg, at 3.0×10 12 Nm 3 rad -1 kg to 6.0×10 16 Nm 3 rad -1 kg, within the range of 6.0×10 12 Nm 3 rad -1 kg to 3.0×10 16 Nm 3 rad -1 kg, or within 2.5×10 13 Nm 3 rad -1 kg to 3.0×10 16 Nm 3 rad -1 kg range.

[0037] One or more of the following features may be applicable to both or either of the first and second aspects described above.

[0038] The moment of inertia of the fan may be greater than or equal to 8.3×10 7 kgm 2 The moment of inertia of the fan can be 7.40×10 7 kgm 2 to 9.00×10 8 kgm 2 The fan's moment of inertia can be within the range of 8.3×10 7 kgm 2 to 6.5×10 8 kgm 2 within the range.

[0039] The fan blades may be formed of a metallic material. The fan blades may be formed of titanium or an aluminum-lithium alloy with a titanium leading edge.

[0040] The fan blade may be formed at least in part from an organic matrix composite material.The fan blade may have a carbon composite body with a metal leading edge.

[0041] The fan shaft may connect the output of the gearbox to the fan. The gearbox output location may be defined as a connection point between the fan shaft and the gearbox. The fan may have an axial centerline. The fan-gearbox axial distance may be defined as an axial distance between the axial location of the gearbox output location and the axial centerline of the fan. The fan-gearbox axial distance multiplied by the moment of inertia of the fan may be greater than or equal to 1.9×10 7 kgm 3 , greater than or equal to 2.9×10 7 kgm 3 , at 1.9×10 7 kgm 3 to 6.2×10 8 kgm 3 In the range of 2.9×10 7 kgm 3 to 3.9×10 8 kgm 3 within the range.

[0042] The fan shaft is defined as a torque transmitting member extending from the output of the gearbox to the input of the fan. The fan shaft may include at least a portion of the gearbox output shaft and at least a portion of the fan input shaft.

[0043] The gearbox may be in a star configuration, and the output of the gearbox may be a gearbox output location defined as the connection point between the ring gear and the fan shaft. Alternatively, the gearbox may be in a planetary configuration, and the output of the gearbox may be a gearbox output location located at the junction between the fan shaft and the planet carrier.

[0044] According to a third aspect, there is provided a propeller for an aircraft, the propeller comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output 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; a gearbox support arranged to at least partially support the gearbox within the propeller, and wherein: the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and the ratio of radial bending stiffness to moment of inertia:

[0045]

[0046] Greater than or equal to 2.5×10 -2 Nkg -1 m -3 .

[0047] The propulsor of the third aspect may have some or all of the features described above in relation to the gas turbine engine of the first aspect, and in some embodiments may be a gas turbine engine.

[0048] According to a fourth aspect, there is provided a propeller for an aircraft, the propeller comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a spindle and output 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 spindle; and a gearbox support arranged to at least partially support the gearbox within the propeller, wherein: the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and the ratio of anti-tilting stiffness to moment of inertia:

[0049]

[0050] Greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 .

[0051] The propulsor of the fourth aspect may have some or all of the features described above in relation to the gas turbine engine of the second aspect, and in some embodiments may be a gas turbine engine.

[0052] According to a fifth aspect, there is provided a propeller for an aircraft, the propeller comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a spindle and output 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 spindle; and a gearbox support arranged to at least partially support the gearbox within the propeller.

[0053] And the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and wherein:

[0054] a) Ratio of radial bending stiffness to moment of inertia:

[0055]

[0056] Greater than or equal to 2.5×10 -2 Nkg -1 m -3 and / or

[0057] b) Ratio of anti-tilting stiffness to moment of inertia:

[0058]

[0059] Greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 .

[0060] The propulsor of the fifth aspect may have some or all of the features described above in relation to the gas turbine engine of the first or second aspect, and in some embodiments may be a gas turbine engine.

[0061] According to a sixth aspect, a method of operating a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting drive via an output of the gearbox to a fan shaft so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and the ratio of radial bending stiffness to moment of inertia:

[0062]

[0063] Greater than or equal to 2.5×10 -2 Nkg -1 m -3 , the method includes operating a gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0064] The method of the sixth aspect may be a method of operating the gas turbine engine or propeller of the first or third aspect, respectively. Therefore, any of the features, ratios and parameters described above in conjunction with the first or third aspect may also apply to the sixth aspect.

[0065] According to a seventh aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the core shaft and outputting drive to the fan shaft via an output end of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine.

[0066] The moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and the ratio of anti-tilting stiffness to moment of inertia:

[0067]

[0068] Greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 , the method includes operating a gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0069] The method of the seventh aspect may be a method of operating the gas turbine engine or propeller of the second or fourth aspect, respectively. Therefore, any of the features, ratios and parameters described above in conjunction with the second or fourth aspect may also apply to the seventh aspect.

[0070] The sixth and seventh aspects are combinable. According to an eighth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting drive to a fan shaft via an output of the gearbox so as to drive the fan at a rotational speed lower than that of the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine, wherein the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 , and where the ratio of radial bending stiffness to moment of inertia is:

[0071]

[0072] Greater than or equal to 2.5×10 -2 Nkg -1 m -3 and / or

[0073] b) Ratio of anti-tilting stiffness to moment of inertia:

[0074]

[0075] Greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 , the method includes operating a gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0076] The method of the eighth aspect may be a method of operating the gas turbine engine or propeller of the first, second or fifth aspects, respectively. Thus, any of the features, ratios and parameters described above in conjunction with the first, second or fifth aspects may also apply to the eighth aspect.

[0077] The present inventors have discovered that, for a given moment of inertia of the fan, the radial bending stiffness / anti-tilting stiffness of one or both of the fan shaft and the gearbox support at the output end of the gearbox needs to be sufficiently low to isolate the gearbox from load transfer from the fan. The rotation of the fan causes a gyroscopic effect, meaning that the fan shaft tends to maintain a stable orientation of its axis of rotation. However, during maneuvering of an aircraft equipped with a gas turbine engine, the orientation of the fan shaft's axis of rotation changes. The present inventors have discovered that this gyroscopic effect causes loads to be transferred to the gearbox. The present inventors have discovered that by using a fan shaft and / or gearbox support with radial bending stiffness / anti-tilting stiffness such that the ratio of radial bending stiffness / anti-tilting stiffness to the moment of inertia, as defined above, is within a specified range, the problem of load transfer from the fan to the gearbox can be addressed. The present inventors have discovered that if the radial bending stiffness / anti-tilting stiffness of the fan shaft or the gearbox support shaft is increased such that this ratio falls outside the specified range, the gearbox is not adequately isolated. The present inventors have also discovered that if the fan shaft / gearbox support stiffness were to be reduced such that this ratio is outside the range described above, this would result in excessive vibration at low modal frequencies.

[0078] In other aspects, value ranges for the products of the components of the radial bending stiffness to moment of inertia ratio and the tilting stiffness to moment of inertia ratio may be specified instead of, or in addition to, the value ranges for these ratios.

[0079] According to one such aspect, the first aspect described above can be summarized as 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 spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle 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 spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and the product of:

[0080] Radial bending stiffness of at least one of the fan shaft and the gearbox support at the output end of the gearbox × moment of inertia of the fan

[0081] Greater than or equal to 2.0×10 14 Nkgm, greater than or equal to 4.0×10 14 Nkgm, greater than or equal to 2.0×10 15 Nkgm, at 2.0×10 14 Nkgm to 1.4×10 18 Nkgm range, within 4.0×10 14 Nkgm to 7.0×10 17 Nkgm, or within the range of 2.0×10 15 Nkgm to 7.0×10 17 within the range of Nkgm.

[0082] According to another such aspect, the second aspect described above can be summarized as 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 spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle 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 spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40×10 7 kgm 2 ; and the product of:

[0083] The anti-tilting stiffness of at least one of the fan shaft and the gearbox support at the output end of the gearbox × the moment of inertia of the fan

[0084] Greater than or equal to 3.0×10 12 Nm 3 rad -1 kg, greater than or equal to 6.0×10 12 Nm 3 rad -1 kg, greater than or equal to 2.5×10 13 Nm 3 rad -1 kg, at 3.0×10 12 Nm 3 rad -1 kg to 6.0×10 16Nm 3 rad -1 kg, within the range of 6.0×10 12 Nm 3 rad -1 kg to 3.0×10 16 Nm 3 rad -1 kg, or within 2.5×10 13 Nm 3 rad -1 kg to 3.0×10 16 Nm 3 rad -1 kg range.

[0085] Those skilled in the art will understand that method and propulsion aspects may be generalized accordingly.

[0086] According to a ninth aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross-sectional area greater than or equal to 2.4×10 -1 m 2 ; and the first gearbox support strength ratio:

[0087]

[0088] Greater than or equal to 7.0×10 -3 .

[0089] The first gearbox support strength ratio may be greater than or equal to 1.0×10 -2 The first gearbox support strength ratio may be greater than or equal to 7.0×10 -3 The first gearbox support strength ratio may be greater than or equal to 2.0×10 -2 The first gearbox support strength ratio can be 7.0×10 -3 to 2.5×10 -1 The strength ratio of the first gearbox support member may be within the range of 1.0×10 -2 to 1.0×10 -1The first gearbox support member strength ratio may be within the range of 7.0×10 -3 to 2.0×10 -2 The strength ratio of the first gearbox support member can be within the range of 2.0×10 -2 to 2.5×10 -1 within the range.

[0090] The radial bending stiffness of the gearbox support can be greater than or equal to 1.0×10 7 N / m. The radial bending stiffness of the gearbox support can be greater than or equal to 2.0×10 7 N / m. The radial bending stiffness of the gearbox support can be greater than or equal to 3.0×10 7 N / m. The radial bending stiffness of the gearbox support can be 1.0×10 7 N / m to 4.0×10 8 The radial bending stiffness of the gearbox support can be within the range of 2.0×10 7 N / m to 3.0×10 8 The radial bending stiffness of the gearbox support can be within the range of 3.0×10 7 N / m to 2.0×10 8 In the range of N / m.

[0091] The torsional strength of the gearbox support can be greater than or equal to 1.60×10 5 Nm. The torsional strength of the gearbox support can be greater than or equal to 1.8×10 5 Nm. The torsional strength of the gearbox support can be 1.60×10 5 Nm to 2.00×10 7 The torsional strength of the gearbox support can be within the range of 1.8×10 5 Nm to 1.5×10 6 Nm range.

[0092] The planet gear pitch angle in radians can be defined as 2π / N, where N is the number of planet gears.

[0093] Second gearbox support strength ratio:

[0094]

[0095] Can be greater than or equal to 1.0×10 -1 The second gearbox support strength ratio may be greater than or equal to 1.5×10 -1 The second gearbox support strength ratio may be greater than or equal to 1.0×10 -1The second gearbox support strength ratio may be greater than or equal to 2.5×10 -1 The second gearbox support strength ratio can be 1.0×10 -1 to 3.5. The second gearbox support member strength ratio may be 1.5×10 -1 The second gearbox support member strength ratio may be in the range of 1.0×10 -1 to 2.5×10 -1 The second gearbox support member strength ratio can be within the range of 2.5×10 -1 to 3.5.

[0096] According to a tenth aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine, wherein: the planetary gear pitch angle in radians is defined as 2π / N, where N is the number of planetary gears; the gearbox has a cross-sectional area greater than or equal to 2.4×10 -1 m 2 ; and the second gearbox support strength ratio:

[0097]

[0098] Greater than or equal to 1.0×10 -1 .

[0099] The second gearbox support strength ratio may be greater than or equal to 1.5×10 -1 The second gearbox support strength ratio may be greater than or equal to 2.5×10 -1 The second gearbox support strength ratio can be 1.0×10 -1 to 3.5. The second gearbox support member strength ratio may be 1.5×10 -1 The second gearbox support member strength ratio may be in the range of 1.0×10 -1 to 2.5×10 -1 The second gearbox support member strength ratio can be within the range of 2.5×10 -1 to 3.5.

[0100] The anti-tilting stiffness of the gearbox support can be greater than or equal to 1.2×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be greater than or equal to 2.4×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be greater than or equal to 3.9×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be 1.2×10 5 Nm / rad to 2.1×10 7 The anti-tilt stiffness of the gearbox support can be within the range of 2.4×10 5 Nm / rad to 1.6×10 7 The anti-tilt stiffness of the gearbox support can be within the range of 3.9×10 5 Nm / rad to 9.0×10 6 In the range of Nm / rad.

[0101] The torsional strength of the gearbox support can be greater than or equal to 1.60×10 5 Nm. The torsional strength of the gearbox support can be greater than or equal to 1.8×10 5 Nm. The torsional strength of the gearbox support can be 1.60×10 5 Nm to 2.00×10 7 The torsional strength of the gearbox support can be within the range of 1.8×10 5 Nm to 1.5×10 6 Nm range.

[0102] First gearbox support strength ratio:

[0103]

[0104] Can be greater than or equal to 7.0×10 -3 The first gearbox support strength ratio may be greater than or equal to 1.0×10 -2 The first gearbox support strength ratio may be greater than or equal to 2.0×10 -2 The first gearbox support strength ratio can be 7.0×10 -3 to 2.5×10 -1 The strength ratio of the first gearbox support member may be within the range of 1.0×10 -2 to 1.0×10 -1 The first gearbox support member strength ratio may be within the range of 7.0×10 -3 to 2.0×10 -2 The strength ratio of the first gearbox support member can be within the range of 2.0×10-2 to 2.5×10 -1 within the range.

[0105] One or more of the following features may be applicable to both or either of the ninth and tenth aspects described above.

[0106] The gearbox may be in a star configuration.

[0107] The planetary gear pitch angle can be greater than or equal to 9.0×10 -1 rad. The planetary gear pitch angle can be 9.0×10 - 1 rad to 2.1rad.

[0108] The cross-sectional area of ​​the gearbox can be greater than or equal to 2.6×10 -1 m 2 The cross-sectional area of ​​the gearbox can be 2.4×10 -1 m 2 Up to 1.10m 2 The cross-sectional area of ​​the gearbox can be within the range of 2.6×10 -1 m 2 to 9.0×10 -1 m 2 The cross-sectional area (CSA) of the gearbox can be defined as the area of ​​the pitch circle of the ring gear.

[0109] The torsional shear stress of the gearbox support under the maximum takeoff condition can be less than or equal to 4.90×10 8 N / m 2 , less than or equal to 3.5×10 8 N / m 2 , at 1.40×10 8 N / m 2 to 4.90×10 8 N / m 2 In the range of 2.0×10 8 N / m 2 to 3.5×10 8 N / m 2 within the range.

[0110] First gearbox support shear stress ratio:

[0111]

[0112] Can be less than or equal to 4.9×10 -1 m -1 , less than or equal to 20m -1 , in 3.5×10 -1 m-1 to 4.9×10 1 m -1 within the range of 0.70m -1 Up to 20m -1 within the range.

[0113] Second gearbox support shear stress ratio:

[0114]

[0115] Can be less than or equal to 4.1×10 3 rad / m 3 , less than or equal to 1.4×10 3 rad / m 3 , at 6.6rad / m 3 to 4.1×10 3 rad / m 3 In the range of 1.25×10 1 rad / m 3 to 1.4×10 3 rad / m 3 within the range.

[0116] According to an eleventh aspect, there is provided a propeller for an aircraft, the propeller 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 spindle and output drive to a fan shaft so as to drive the fan at a rotational speed lower than that of the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the propeller, wherein: the gearbox has a cross-sectional area greater than or equal to 2.4×10 -1 m 2 ; First gearbox support strength ratio:

[0117]

[0118] Greater than or equal to 7.0×10 -3 .

[0119] The propulsor of the eleventh aspect may have some or all of the features described above with respect to the gas turbine engine of the ninth aspect, and in some embodiments may be a gas turbine engine.

[0120] According to a twelfth aspect, there is provided a propeller for an aircraft, the propeller 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 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 planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the propeller, wherein: the gearbox has a cross-sectional area greater than or equal to 2.4×10 -1 m 2 ; and the planet gear pitch angle in radians is defined as 2π / N, where N is the number of planet gears. And the second gearbox support strength ratio:

[0121]

[0122] Greater than or equal to 1.0×10 -1 .

[0123] The propulsor of the twelfth aspect may have some or all of the features described above in relation to the gas turbine engine of the tenth aspect, and in some embodiments may be a gas turbine engine.

[0124] The eleventh aspect and the twelfth aspect may be combined. According to the fifteenth aspect, there is provided a propeller for an aircraft, the propeller 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 output drive to a fan shaft so as to drive the fan at a rotational speed lower than that of the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the propeller, wherein: the gearbox has a cross-sectional area that is greater than or equal to 2.4×10 -1 m 2 , and wherein: the first gearbox support strength ratio:

[0125]

[0126] Greater than or equal to 7.0×10 -3 and / or

[0127] b) the planet gear pitch angle in radians is defined as 2π / N, where N is the number of planet gears; and the second gearbox support strength ratio:

[0128]

[0129] Greater than or equal to 1.0×10 -1 .

[0130] The propulsor of the thirteenth aspect may have some or all of the features described above in relation to the gas turbine engines of the ninth and tenth aspects, and in some embodiments may be a gas turbine engine.

[0131] 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 spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross-sectional area greater than or equal to 2.4×10 -1 m 2 ; and the first gearbox support strength ratio:

[0132]

[0133] Greater than or equal to 7.0×10 -3 , the method includes operating a gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0134] The method of the fourteenth aspect may be a method of operating the gas turbine engine or propeller of the ninth or eleventh aspect, respectively. Therefore, any of the features, ratios and parameters described above in conjunction with the ninth or eleventh aspect may also apply to the fourteenth aspect.

[0135] According to a fifteenth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross-sectional area greater than or equal to 2.4×10 -1 m 2 ; and the planet gear pitch angle in radians is defined as 2π / N, where N is the number of planet gears. And the second gearbox support strength ratio:

[0136]

[0137] Greater than or equal to 1.0×10 -1 , the method includes operating a gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0138] The method of the fifteenth aspect may be a method of operating the gas turbine engine or propeller of the tenth or twelfth aspect, respectively. Thus, any of the features, ratios and parameters described above in conjunction with the tenth or twelfth aspect may also apply to the fifteenth aspect.

[0139] The fourteenth aspect and the fifteenth aspect may be combined. According to the sixteenth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein the gearbox has a cross-sectional area greater than or equal to 2.4×10 -1 m 2 ,and

[0140] a) First gearbox support strength ratio:

[0141]

[0142] Greater than or equal to 7.0×10 -3 and / or

[0143] b) the planet gear pitch angle in radians is defined as 2π / N, where N is the number of planet gears; and the second gearbox support strength ratio:

[0144]

[0145] Greater than or equal to 1.0×10 -1 ,

[0146] The method includes operating a gas turbine engine under cruise conditions to provide propulsion for the aircraft.

[0147] The method of the sixteenth aspect may be a method of operating the gas turbine engine or propeller of the ninth, tenth or thirteenth aspects. Thus, any of the features, ratios and parameters described above in conjunction with the ninth, tenth or thirteenth aspects may also apply to the sixteenth aspect.

[0148] The present inventors have discovered that by designing the gearbox support so that the ratio of its torsional strength to its radial bending stiffness and cross-sectional area (i.e., the first gearbox support strength ratio) is within a specified range, sufficient support strength is provided to make the engine reliable while providing sufficient stiffness to minimize misalignment of the gears in the gearbox and avoid vibration. The present inventors have discovered that similar considerations apply to the ratio of torsional strength to anti-tilting stiffness and the planet gear pitch angle (i.e., the second gearbox support strength ratio).

[0149] According to a seventeenth 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 spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine, wherein: a first gearbox support shear stress ratio:

[0150]

[0151] Less than or equal to 4.9×10 1 m -1 .

[0152] The shear stress ratio of the first gearbox support can be less than or equal to 20m -1 The shear stress ratio of the first gearbox support can be 3.5×10 -1 m -1 to 4.9×10 1 m -1 The shear stress ratio of the first gearbox support can be within the range of 0.70m -1 Up to 20m -1 within the range.

[0153] The radial bending stiffness of the gearbox support can be greater than or equal to 1.0×10 7 N / m. The radial bending stiffness of the gearbox support can be greater than or equal to 2.0×10 7 N / m. The radial bending stiffness of the gearbox support can be greater than or equal to 3.0×10 7 N / m. The radial bending stiffness of the gearbox support can be 1.0×10 7 N / m to 4.0×10 8 The radial bending stiffness of the gearbox support can be within the range of 2.0×10 7 N / m to 3×10 8 The radial bending stiffness of the gearbox support can be within the range of 3.0×10 7 N / m to 2.0×10 8 In the range of N / m.

[0154] The torsional shear stress of the gearbox support under the maximum takeoff condition can be less than or equal to 4.90×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be less than or equal to 3.5×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be 1.40×10 8 N / m 2 to 4.90×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be within the range of 2.0×10 8 N / m 2 to 3.5×10 8 N / m 2 within the range.

[0155] The fan diameter may be in the range of 240 cm to 280 cm. In such an embodiment, the first gearbox support shear stress ratio may be less than or equal to 35 m-1 Or at 0.70m -1 Up to 35m -1 within the range.

[0156] Alternatively, the fan diameter may be in the range of 330 cm to 380 cm. In such an embodiment, the first gearbox support shear stress ratio may be less than or equal to 12 m -1 Or at 0.50m -1 Up to 12m -1 within the range.

[0157] Second gearbox support shear stress ratio:

[0158]

[0159] Can be less than or equal to 4.1×10 3 rad / m 3 The shear stress ratio of the second gearbox support member may be less than or equal to 1.4×10 3 rad / m 3 The shear stress ratio of the second gearbox support can be 6.6rad / m 3 to 4.1×10 3 rad / m 3 The shear stress ratio of the second gearbox support member may be within the range of 1.25×10 1 rad / m 3 to 1.4×10 3 rad / m 3 within the range.

[0160] 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 spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine, and wherein: a second gearbox support shear stress ratio:

[0161]

[0162] Less than or equal to 4.1×10 3 rad / m 3 .

[0163] The shear stress ratio of the second gearbox support member may be less than or equal to 1.4×10 3 rad / m 3 The shear stress ratio of the second gearbox support can be 6.6rad / m 3 to 4.1×10 3 rad / m 3 The shear stress ratio of the second gearbox support member may be within the range of 1.25×10 1 rad / m 3 to 1.4×10 3 rad / m 3 within the range.

[0164] The anti-tilting stiffness of the gearbox support can be greater than or equal to 1.2×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be greater than or equal to 2.4×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be greater than or equal to 3.9×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be 1.2×10 5 Nm / rad to 2.1×10 7 The anti-tilt stiffness of the gearbox support can be within the range of 2.4×10 5 Nm / rad to 1.6×10 7 The anti-tilt stiffness of the gearbox support can be within the range of 3.9×10 5 Nm / rad to 9.0×10 6 In the range of Nm / rad.

[0165] The torsional shear stress of the gearbox support under the maximum takeoff condition can be less than or equal to 4.90×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be less than or equal to 3.5×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be 1.40×10 8 N / m 2 to 4.90×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be within the range of 2.0×10 8 N / m 2 to 3.5×10 8 N / m 2 within the range.

[0166] The diameter of the fan may be in the range of 240 cm to 280 cm. In such an embodiment, the second gearbox support shear stress ratio may be less than or equal to 2.9×10 3 rad / m 3 Or at 2.9×10 1 rad / m 3 to 2.9×10 3 rad / m 3 within the range.

[0167] Alternatively, the fan diameter may be in the range of 330 cm to 380 cm. In such an embodiment, the second gearbox support shear stress ratio may be less than or equal to 7.0×10 2 rad / m 3 Or at 1.0×10 1 rad / m 3 to 7.0×10 2 rad / m 3 within the range.

[0168] First gearbox support shear stress ratio:

[0169]

[0170] Can be less than or equal to 4.9×10 1 m -1 The shear stress ratio of the first gearbox support member may be less than or equal to 20m -1 The shear stress ratio of the first gearbox support can be 3.5×10 -1 m -1 to 4.9×10 1 m -1 The shear stress ratio of the first gearbox support can be within the range of 0.70m -1 Up to 20m -1 within the range.

[0171] One or more of the following features may be applicable to both or either of the seventeenth and eighteenth aspects described above.

[0172] The torsional strength of the gearbox support can be greater than or equal to 1.60×10 5 Nm. The torsional strength of the gearbox support can be greater than or equal to 1.8×10 5 Nm. The torsional strength of the gearbox support can be 1.60×10 5 Nm to 2.00×10 7 The torsional strength of the gearbox support can be within the range of 1.8×10 5 Nm to 1.5×10 6Nm range.

[0173] The cross-sectional area of ​​the gearbox can be greater than or equal to 2.4×10 -1 m 2 The cross-sectional area of ​​the gearbox can be greater than or equal to 2.6×10 -1 m 2 The cross-sectional area of ​​the gearbox can be 2.4×10 -1 m 2 Up to 1.10m 2 The cross-sectional area of ​​the gearbox can be within the range of 2.6×10 -1 m 2 to 9.0×10 -1 m 2 within the range.

[0174] The planet gear pitch angle in radians can be defined as 2π / N, where N is the number of planet gears. The planet gear pitch angle can be greater than or equal to 9.0×10 -1 rad. The planetary gear pitch angle can be 9.0×10 -1 rad to 2.1rad.

[0175] First gearbox support strength ratio:

[0176]

[0177] Can be greater than or equal to 7.0×10 -3 The first gearbox support strength ratio may be greater than or equal to 1.0×10 -2 The first gearbox support strength ratio may be greater than or equal to 2.0×10 -2 The first gearbox support strength ratio can be 7.0×10 -3 to 2.5×10 -1 The strength ratio of the first gearbox support member may be within the range of 1.0×10 -2 to 1.0×10 -1 The first gearbox support member strength ratio may be within the range of 7.0×10 -3 to 2.0×10 -2 The strength ratio of the first gearbox support member can be within the range of 2.0×10 -2 to 2.5×10 -1 within the range.

[0178] Second gearbox support strength ratio:

[0179]

[0180] Can be greater than or equal to 1.0×10 -1The second gearbox support strength ratio may be greater than or equal to 1.5×10 -1 The second gearbox support strength ratio may be greater than or equal to 2.5×10 -1 The second gearbox support strength ratio can be 1.0×10 -1 to 3.5. The second gearbox support member strength ratio may be 1.5×10 -1 The second gearbox support member strength ratio may be in the range of 1.0×10 -1 to 2.5×10 -1 The second gearbox support member strength ratio can be within the range of 2.5×10 -1 to 3.5.

[0181] The torque transmitted through the gearbox support under maximum takeoff conditions may be greater than or equal to 6.00 × 10 4 Nm, greater than or equal to 7.2×10 4 Nm, at 6.00×10 4 Nm to 5.00×10 5 Nm, or within the range of 7.2×10 4 Nm to 4.2×10 5 Nm range.

[0182] The gearbox may be of a planetary configuration. In such an embodiment, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 6.00×10 4 Nm, greater than or equal to 7.2×10 4 Nm, at 6.00×10 4 Nm to 3.00×10 5 Nm, or in the range of 7.2×10 4 Nm to 2.6×10 5 Nm range.

[0183] Alternatively, the gearbox may be in a star configuration. In such an embodiment, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 1.10×10 5 Nm, greater than or equal to 1.3×10 5 Nm, at 1.10×10 5 Nm to 5.00×10 5 Nm, or within the range of 1.3×10 5 Nm to 4.2×10 5 Nm range.

[0184] The maximum takeoff condition may be defined as anywhere herein, for example may be defined as operating at the engine's maximum takeoff thrust at ISA sea level pressure and temperature +15°C with a fan inlet speed between 0.25 Mn and 0.27 Mn, and optionally 0.25 Mn.

[0185] According to a nineteenth aspect, there is provided a propeller for an aircraft, the propeller 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 spindle and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the propeller, wherein: a first gearbox support shear stress ratio:

[0186]

[0187] Less than or equal to 4.9×10 1 m -1 .

[0188] The propulsor of the nineteenth aspect may have some or all of the features described above in relation to the gas turbine engine of the seventeenth aspect, and in some embodiments may be a gas turbine engine.

[0189] According to a twentieth aspect, there is provided a propeller for an aircraft, the propeller 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 spindle and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the propeller, wherein: a second gearbox support shear stress ratio:

[0190]

[0191] Less than or equal to 4.1×10 3 rad / m 3 .

[0192] The propulsor of the twentieth aspect may have some or all of the features described above in relation to the gas turbine engine of the eighteenth aspect, and in some embodiments may be a gas turbine engine.

[0193] The nineteenth aspect and the twentieth aspect may be combined. According to the twenty-first aspect, there is provided a propeller for an aircraft, the propeller 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 spindle and output drive to a fan shaft so as to drive the fan at a rotational speed lower than that of the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the propeller, wherein:

[0194] a) Shear stress ratio of the first gearbox support:

[0195]

[0196] Less than or equal to 4.9×10 1 m -1 and / or

[0197] b) Second gearbox support shear stress ratio:

[0198]

[0199] Less than or equal to 4.1×10 3 rad / m 3 .

[0200] The propulsor of the twenty-first aspect may have some or all of the features described above in relation to the gas turbine engines of the seventeenth and eighteenth aspects, and in some embodiments may be a gas turbine engine.

[0201] According to a twenty-second aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine, and wherein the method comprises: operating the gas turbine engine such that, under maximum takeoff conditions, a first gearbox support shear stress ratio is:

[0202]

[0203] Less than or equal to 4.9×10 1 m -1 .

[0204] The method of the twenty-second aspect may be a method of operating the gas turbine engine or propeller of the seventeenth or nineteenth aspect, respectively. Therefore, any of the features, ratios, and parameters described above in conjunction with the seventeenth or nineteenth aspect may also apply to the twenty-second aspect.

[0205] According to a twenty-third aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and

[0206] A gearbox support arranged to at least partially support a gearbox within the engine, and wherein the method comprises operating the gas turbine engine such that, under maximum takeoff conditions, a second gearbox support shear stress ratio:

[0207]

[0208] Less than or equal to 4.1×10 3 rad / m 3 .

[0209] The method of the twenty-third aspect may be a method of operating the gas turbine engine or propeller of the eighteenth or twentieth aspect, respectively. Thus, any of the features, ratios and parameters described above in conjunction with the eighteenth or twentieth aspect may also apply to the twenty-third aspect.

[0210] In another aspect, the twenty-second and twenty-third aspects may be combined. In this aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine, wherein the method comprises: operating the gas turbine engine such that, under maximum takeoff conditions:

[0211] a) Shear stress ratio of the first gearbox support:

[0212]

[0213] Less than or equal to 4.9×10 1 m -1 and / or

[0214] b) Second gearbox support shear stress ratio:

[0215]

[0216] Less than or equal to 4.1×10 3 rad / m 3 .

[0217] The method of the aforementioned aspect may be a method of operating the gas turbine engine or propeller of the seventeenth, eighteenth or twenty-first aspect. Thus, any of the features, ratios and parameters described above in conjunction with the seventeenth, eighteenth or twenty-first aspect may also apply to this aspect.

[0218] The present inventors have found that by designing the gearbox support so that the ratio of its torsional shear stress at maximum takeoff condition to its stiffness (radial bending stiffness or anti-tilting stiffness) is within a specified range, a sufficient strength margin of the support is provided at the highest load point in the engine operating cycle, while the stiffness is suitable to ensure a suitable load sharing factor.

[0219] According to a twenty-fourth 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 spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine, and wherein:

[0220] Flight cycle ratio:

[0221]

[0222] Less than or equal to 3.20.

[0223] The flight cycle ratio may be less than or equal to 2.95. The flight cycle ratio may be less than or equal to 2.9. The flight cycle ratio may be less than or equal to 2.90. The flight cycle ratio may be less than or equal to 2.85. The flight cycle ratio may be less than or equal to 2.75. The flight cycle ratio may be in the range of 2.10 to 3.20. The flight cycle ratio may be in the range of 2.3 to 2.9.

[0224] The torsional shear stress of the gearbox support under the maximum takeoff condition can be less than or equal to 4.90×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be less than or equal to 2.0×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be 1.40×10 8 N / m 2 to 4.9×10 8 N / m 2 The torsional shear stress of the gearbox support under the maximum takeoff condition can be within the range of 2.0×10 8 N / m 2 to 3.5×10 8 N / m 2 within the range.

[0225] The torsional shear stress of the gearbox support under cruising conditions can be less than or equal to 7.00×10 7 N / m 2 The torsional shear stress of the gearbox support under cruising conditions can be less than or equal to 8.2×107 N / m 2 The torsional shear stress of the gearbox support under cruising conditions can be 7.00×10 7 N / m 2 to 1.90×10 8 N / m 2 The torsional shear stress of the gearbox support under cruising conditions can be within the range of 8.2×10 7 N / m 2 to 1.5×10 8 N / m 2 within the range.

[0226] The fan may have a diameter in the range of 240 cm to 280 cm. In such embodiments, the flight cycle ratio may be less than or equal to 3.2 or in the range of 2.3 to 3.2.

[0227] Alternatively, the fan diameter may be in the range of 330 cm to 380 cm. In such embodiments, the flight cycle ratio may be less than or equal to 2.8 or in the range of 2.1 to 2.8.

[0228] The product of:

[0229] The torsional shear stress of the gearbox support under maximum takeoff conditions × the torsional shear stress of the gearbox support under cruise conditions

[0230] Can be greater than or equal to 1.00×10 16 (N / m 2 ) 2 , greater than or equal to 2.05×10 16 (N / m 2 ) 2 , at 1.00×10 16 (N / m 2 ) 2 to 7.50×10 16 (N / m 2 ) 2 In the range of 2.05×10 16 (N / m 2 ) 2 to 4.9×10 16 (N / m 2 ) 2 In such an embodiment, the gearbox may, for example, be in a star-type configuration.

[0231] First torque transfer ratio:

[0232]

[0233] The first torque transmission ratio may be less than or equal to 3.2. The first torque transmission ratio may be less than or equal to 2.95. The first torque transmission ratio may be less than or equal to 2.9. The first torque transmission ratio may be less than or equal to 2.90. The first torque transmission ratio may be less than or equal to 2.85. The first torque transmission ratio may be less than or equal to 2.75. The first torque transmission ratio may be in the range of 2.1 to 3.2. The first torque transmission ratio may be in the range of 2.3 to 2.9.

[0234] Second torque transfer ratio:

[0235]

[0236] The second torque transmission ratio may be less than or equal to 3.2. The second torque transmission ratio may be less than or equal to 2.95. The second torque transmission ratio may be less than or equal to 2.9. The second torque transmission ratio may be less than or equal to 2.90. The second torque transmission ratio may be less than or equal to 2.85. The second torque transmission ratio may be less than or equal to 2.75. The second torque transmission ratio may be in the range of 2.1 to 3.2. The second torque transmission ratio may be in the range of 2.3 to 2.9.

[0237] The torsional strength of the gearbox support can be greater than or equal to 1.60×10 5 Nm. The torsional strength of the gearbox support can be greater than or equal to 1.8×10 5 Nm. The torsional strength of the gearbox support can be 1.60×10 5 Nm to 2.00×10 7 The torsional strength of the gearbox support can be within the range of 1.8×10 5 Nm to 1.5×10 6 Nm range.

[0238] The gearbox may have a cross-sectional area, and a first gearbox support member strength ratio:

[0239]

[0240] Can be greater than or equal to 7.0×10 -3 The first gearbox support strength ratio may be greater than or equal to 1.0×10 -2 The first gearbox support strength ratio may be greater than or equal to 2.0×10 -2 The first gearbox support strength ratio can be 7.0×10 -3 to 2.5×10 -1 The strength ratio of the first gearbox support member may be within the range of 1.0×10 -2 to 1.0×10 -1The first gearbox support member strength ratio may be within the range of 7.0×10 -3 to 2.0×10 -2 The strength ratio of the first gearbox support member can be within the range of 2.0×10 -2 to 2.5×10 -1 within the range.

[0241] The planet gear pitch angle in radians can be defined as 2π / N, where N is the number of planet gears. Second gearbox support strength ratio:

[0242]

[0243] Can be greater than or equal to 1.0×10 -1 The second gearbox support strength ratio may be greater than or equal to 1.5×10 -1 The second gearbox support strength ratio may be greater than or equal to 2.5×10 -1 The second gearbox support strength ratio can be 1.0×10 -1 to 3.5. The second gearbox support member strength ratio may be 1.5×10 -1 The second gearbox support member strength ratio may be in the range of 1.0×10 -1 to 2.5×10 -1 The second gearbox support member strength ratio can be within the range of 2.5×10 -1 to 3.5.

[0244] The cross-sectional area of ​​the gearbox can be greater than or equal to 2.4×10 -1 m 2 The cross-sectional area of ​​the gearbox can be greater than or equal to 2.6×10 -1 m 2 The cross-sectional area of ​​the gearbox can be 2.4×10 -1 m 2 Up to 1.10m 2 The cross-sectional area of ​​the gearbox can be within the range of 2.6×10 -1 m 2 to 9.0×10 -1 m 2 within the range.

[0245] The planetary gear pitch angle can be greater than or equal to 9.0x10 -1 rad. The planetary gear pitch angle can be 9.0×10 - 1 rad to 2.1rad.

[0246] The radial bending stiffness of the gearbox support can be greater than or equal to 1.0×107 N / m. The radial bending stiffness of the gearbox support can be greater than or equal to 2.0×10 7 N / m. The radial bending stiffness of the gearbox support can be greater than or equal to 3.0×10 7 N / m. The radial bending stiffness of the gearbox support can be 1.0×10 7 N / m to 4.0×10 8 The radial bending stiffness of the gearbox support can be within the range of 2.0×10 7 N / m to 3×10 8 The radial bending stiffness of the gearbox support can be within the range of 3.0×10 7 N / m to 2.0×10 8 In the range of N / m.

[0247] The anti-tilting stiffness of the gearbox support can be greater than or equal to 1.2×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be greater than or equal to 2.4×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be greater than or equal to 3.9×10 5 Nm / rad. The anti-tilting stiffness of the gearbox support can be 1.2×10 5 Nm / rad to 2.1×10 7 The anti-tilt stiffness of the gearbox support can be within the range of 2.4×10 5 Nm / rad to 1.6×10 7 The anti-tilt stiffness of the gearbox support can be within the range of 3.9×10 5 Nm / rad to 9.0×10 6 In the range of Nm / rad.

[0248] The gearbox may be in a star configuration.

[0249] The maximum takeoff condition may be as defined anywhere herein, for example as operation at the engine's maximum takeoff thrust at ISA sea level pressure and temperature +15°C with a fan inlet velocity between 0.25 Mn and 0.27 Mn, and optionally 0.25 Mn.

[0250] The cruise condition may be as defined anywhere herein. The cruise condition may mean the conditions in mid-cruise of the aircraft to which the engine is attached. The cruise condition may be the conditions experienced by the aircraft and engine at the midpoint between the highest point of a climb and the start of a descent.

[0251] Under cruise conditions, the forward speed of the gas turbine engine may be in the range of Mn 0.75 to Mn 0.85. Under cruise conditions, the forward speed of the gas turbine engine may be Mn 0.8.

[0252] The cruise conditions may correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 11,582 m and a forward Mach number of 0.8.

[0253] The cruise conditions may correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 10,668 m and a forward Mach number of 0.85.

[0254] The cruising conditions may correspond to atmospheric conditions at an altitude between 10,500 m and 11,600 m, and optionally at an altitude of 11,000 m.

[0255] According to a twenty-fifth aspect, there is provided a propeller for an aircraft, the propeller 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 spindle and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged to at least partially support the gearbox within the propeller, wherein:

[0256] Flight cycle ratio:

[0257]

[0258] Less than or equal to 3.20.

[0259] The propulsor of the twenty-fifth aspect may have some or all of the features described above in relation to the gas turbine engine of the twenty-fourth aspect, and in some embodiments may be a gas turbine engine.

[0260] According to another aspect, there is provided an aircraft comprising the gas turbine engine or propeller of the twenty-fourth or twenty-fifth aspect mounted thereon, wherein the aircraft has a maximum takeoff operating condition and a cruise condition.

[0261] According to a twenty-sixth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and

[0262] a gearbox support arranged to at least partially support a gearbox within the engine, and wherein the method comprises:

[0263] The gas turbine engine is operated so that the flight cycle ratio is:

[0264]

[0265] Less than or equal to 3.20.

[0266] The method of the twenty-sixth aspect can be a method of operating the gas turbine engine or propeller of the twenty-fourth or twenty-fifth aspect, respectively. Therefore, any of the features, ratios, and parameters described above in conjunction with the twenty-fourth or twenty-fifth aspects may also be applicable to the twenty-sixth aspect. The method of the twenty-fifth aspect may include operating the gas turbine engine under maximum takeoff conditions to provide propulsion for the aircraft to which it is installed. The method also includes operating the gas turbine engine under cruise conditions to provide propulsion. Cruise conditions and maximum takeoff conditions are as defined elsewhere herein.

[0267] By considering a range of gearbox failure modes, including tooth root bending, tooth surface wear, tooth surface macro- and micro-depression, and gearbox vibration, and the extent to which these failure modes have an impact at maximum takeoff (MTO) and cruise, the present inventors have discovered that providing torsional shear stresses in the gearbox support structure within a specified range provides a low weight gearbox that can successfully transmit maximum loads at MTO for a limited duration, while providing sufficient integrity against other failure mechanisms at cruise for an extended duration.

[0268] In other aspects, a range of values ​​for the product of the components of the flight cycle ratio may be specified instead of or in addition to the range of values ​​for the ratio.

[0269] According to one such aspect, the twenty-fourth aspect introduced above can be summarized as 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 spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planetary gears, a ring gear, and a planet carrier arranged for mounting the plurality of planetary gears thereon; and a gearbox support arranged for at least partially supporting the gearbox within the engine, and wherein: the product (e.g., the flight cycle product):

[0270] The torsional shear stress of the gearbox support under maximum takeoff conditions × the torsional shear stress of the gearbox support under cruise conditions

[0271] Greater than or equal to 1.00×10 16 (N / m 2 ) 2 , greater than or equal to 2.05×10 16 (N / m 2 ) 2 , at 1.00×10 16 (N / m 2 ) 2 to 7.50×10 16 (N / m 2 ) 2 In the range of 2.05×10 16 (N / m 2 ) 2 to 4.9×10 16 (N / m 2 ) 2 In any of these embodiments, the gearbox can be in a star configuration.

[0272] Those skilled in the art will understand that method and propulsion aspects may be generalized accordingly.

[0273] In any of the foregoing, any one or more of the following may apply:

[0274] The turbine may be a first turbine, the compressor may be a first compressor, and the mandrel may be a first mandrel. The engine core may further include a second turbine, a second compressor, and a second mandrel connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second mandrel may be arranged to rotate at a higher rotational speed than the first mandrel.

[0275] The gearbox can have a gear ratio within any range disclosed herein, for example, within the range of 3.2 to 4.5, and optionally within the range of 3.2 to 4.0.

[0276] The gas turbine engine may have a specific thrust within any range disclosed herein, such as 70 NKg -1 Up to 90NKg -1 Specific thrust within the range.

[0277] The gas turbine engine may have a bypass ratio within any range disclosed herein, such as within the range of 12.5 to 18, and optionally within the range of 13 to 16, at cruise conditions.

[0278] The fan diameter of the fan may be greater than 240 cm and less than or equal to 380 cm, and the fan diameter of the fan may be greater than 300 cm and less than or equal to 380 cm.

[0279] The diameter of the fan may be in the range of 240 cm to 280 cm. The diameter of the fan may be in the range of 330 cm to 380 cm.

[0280] The cross-sectional area (CSA) of the gearbox can be defined as the area of ​​the pitch circle of the ring gear.

[0281] The method of any of the aspects defined above may further comprise driving the gearbox with an input torque of:

[0282] i) greater than or equal to 10,000 Nm at cruise, and optionally between 10,000 Nm and 50,000 Nm; and / or

[0283] ii) greater than or equal to 28,000 Nm, and optionally between 28,000 Nm and 135,000 Nm, at maximum take-off conditions.

[0284] For any parameter or ratio of a parameter X claimed or disclosed herein, a limit on the values ​​that X can take expressed as "X is greater than or equal to Y" can alternatively be expressed as "1 / X is less than or equal to 1 / Y." Thus, any ratio or parameter defined in the above aspects and statements can be expressed as "1 / X is less than or equal to 1 / Y" rather than "X is greater than or equal to Y." In such cases, zero can be considered a lower limit.

[0285] Various parameters of the gearbox and its mounting surface and / or more generally the engine may be adjusted to allow the engine to meet the various specifications outlined above. Commentary on various such parameters is provided below.

[0286] With respect to the stiffness of the gearbox support (radial bending and / or tilt resistance), the inventors have found that a relatively low stiffness can be set to isolate the gearbox from the damaging loads transmitted thereto. The inventors have found that if the stiffness of the gearbox support is reduced outside the range specified herein, there will be problems with dynamic effects such as lateral vibrations. Specifically, it has been found that a lower stiffness within the range defined herein allows for vibrations of low modal frequencies to be reduced or avoided (those skilled in the art will understand that vibrations of lower modes have larger amplitudes / deflections than higher modes, so avoiding vibrations of these lower modes is more important). This can be a function of the size of the gearbox and its configuration. The inventors have also found that the maximum stiffness provided by the range defined herein allows for reducing or avoiding damaging loads transmitted from the fan to the gearbox. This can similarly vary depending on the size and configuration of the gearbox.

[0287] The present inventors have discovered that reducing the radial bending stiffness and / or the tilting stiffness of the fan shaft (at the fan input or the gearbox output) outside the ranges defined herein will result in undesirable dynamic effects, such as lateral vibrations. Specifically, the minimum stiffness defined by the ranges specified herein allows for the reduction or avoidance of vibrations at low modal frequencies (as will be appreciated by those skilled in the art, vibrations at lower modes have greater amplitudes / deflections than higher modes, and therefore avoiding vibrations at lower modes is more important). This may be a function of the size of the gearbox and its configuration.

[0288] The inventors have also found that the upper limit of the radial bending and / or anti-tilting stiffness of the fan shaft is affected by the basic properties of one or more materials from which the fan shaft is made. For example, the maximum stiffness is affected 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 may be, for example, in the range of 100 GPa to 250 GPa or 105 GPa to 215 GPa, and is optionally about 210 GPa. Different grades of steel or other types of metal may be selected to achieve different stiffnesses for the same size and geometry. For example, 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 may be used in various embodiments. The inventors have found that increasing the stiffness beyond the range defined herein using materials such as these will result in excessive weight without actual gain in performance.

[0289] The present inventors have discovered that the torsional strength ranges for the gearbox supports defined herein provide the required level of torque capacity to provide adequate reliability, while not being so great as to add excessive weight to the engine without substantial performance gain.

[0290] With respect to other properties of the gearbox support, such as its torsional shear stress, the inventors have found that the ranges specified herein provide the improvements in performance described above without resulting in an excessive increase in the overall weight of the engine.

[0291] As described elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may include an engine core comprising a turbine, a combustor, a compressor, and a spindle connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having fan blades) located upstream of the engine core.

[0292] The gas turbine engine may include a gearbox that receives input from a mandrel and drives an output to a fan so as to drive the fan at a rotational speed lower than the mandrel. The input to the gearbox may be directly from the mandrel or indirectly from the mandrel, such as via a spur gear shaft and / or a gear. The mandrel may rigidly connect the turbine and the compressor so that the turbine and the compressor rotate at the same speed (wherein the fan rotates at a lower speed). The output from the gearbox may directly reach the fan shaft, or indirectly reach the fan shaft, such as via a spur gear shaft and / or a gear.

[0293] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting the turbine and the compressor, such as one, two, or three shafts. By way of example only, the turbine connected to the mandrel may be a first turbine, the compressor connected to the mandrel may be a first compressor, and the mandrel may be a first mandrel. The engine core may also include a second turbine, a second compressor, and a second mandrel connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second mandrel may be arranged to rotate at a higher rotational speed than the first mandrel.

[0294] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor.The second compressor may be arranged to receive flow from the first compressor (eg directly, such as via a generally annular conduit).

[0295] The gearbox may be arranged to be driven by the spindle configured to rotate (e.g., during use) at the lowest rotational speed (e.g., the first spindle in the above example). For example, the gearbox may be arranged to be driven only by the spindle configured to rotate (e.g., during use) at the lowest rotational speed (e.g., only the first spindle, not the second spindle in the above example). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, such as the first and / or second shafts in the above example.

[0296] The gearbox can be a reduction gearbox (because the output to the fan has a lower rotational rate than the input from the spindle). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "stellar" 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), for example greater than 2.5, such as in the range of 3 to 4.2, or 3.2 to 3.8, for example, 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 values ​​in the previous sentence. By way of example only, the gearbox can be a "stellar" gearbox having a gear ratio in the range of 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio can be outside these ranges.

[0297] In any gas turbine engine as described and / or claimed herein, the combustor may be positioned axially downstream of the fan and one or more compressors. For example, where a second compressor is provided, the combustor may be positioned directly downstream of the second compressor (e.g., at its outlet). By way of another example, where a second turbine is provided, the flow at the combustor outlet may be provided to the inlet of the second turbine. The combustor may be positioned upstream of the one or more turbines.

[0298] The or each compressor (e.g., the first and second compressors described above) may include any number of stages, such as a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes, which may be variable stator vanes (because the angle of incidence of the row of stator vanes may be variable). The row of rotor blades and the row of stator vanes may be axially offset from one another.

[0299] The turbine or each turbine (e.g., the first and second turbines described above) may include any number of stages, such as a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

[0300] Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas scrubbing position or 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or approximately) any one of the following: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, within a range of 0.28 to 0.32. These ratios may generally be referred to as hub-tip ratios. Both the radius at the hub and the radius at the tip may be measured at the leading edge (or axially forward-most) portion of the blade. Of course, the hub-to-tip ratio refers to the gas-washed portion of the fan blade, ie, the portion radially outside of any platform.

[0301] The radius of the fan can be measured between the engine centerline and the tips at the leading edges of the fan blades. The fan diameter (which may be just twice the fan radius) can be greater than (or approximately) any of the following: 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches), or 420 cm (about 165 inches). The fan diameter may be within an inclusive range bounded by any two values ​​in the previous sentence (ie, these values ​​may form an upper or lower limit), for example, within the range of 240 cm to 280 cm or 330 cm to 380 cm.

[0302] The rotational speed of the fan can vary during use. Generally speaking, 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 may be less than 2500 rpm, for example, less than 2300 rpm. By way of another non-limiting example only, for an engine with a fan diameter in the range of 220 cm to 300 cm (e.g., 240 cm to 280 cm or 250 cm to 270 cm), the rotational speed of the fan under cruise conditions may be in the range of 1700 rpm to 2500 rpm, for example, in the range of 1800 rpm to 2300 rpm, for example, in the range of 1900 rpm to 2100 rpm. By way of another non-limiting example only, for an engine with a fan diameter in the range of 330 cm to 380 cm, the rotational speed of the fan under cruise conditions may be in the range of 1200 rpm to 2000 rpm, for example, in the range of 1300 rpm to 1800 rpm, for example, in the range of 1400 rpm to 1800 rpm.

[0303] When a gas turbine engine is in use, the fan (with its associated fan blades) rotates about an axis of rotation. This rotation causes the tips of the fan blades to rotate at a speed U 尖端 The work done by the fan blades 13 on the flow causes the enthalpy of the flow to rise dH. The fan tip load can be defined as dH / U 尖端 2 , where dH is the enthalpy rise across the fan (e.g., 1-D average enthalpy rise), and U 尖端 is the (translational) velocity of the fan tip, for example, at the leading edge of the tip (may be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip load at cruise conditions may be greater than (or approximately) any of the following: 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 load may be within an inclusive range bounded by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 0.28 to 0.31 or 0.29 to 0.3.

[0304] Gas turbine engines according to the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of flow through the bypass duct to the mass flow rate of flow through the core under cruise conditions. In some arrangements, the bypass ratio may be greater than (or approximately) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), such as a range of 12 to 16, a range of 13 to 15, or a range of 13 to 14. The bypass duct may be substantially annular. The bypass duct may be located radially outward from the core engine. The radially outer surface of the bypass duct may be defined by the nacelle and / or the fan case.

[0305] The overall pressure ratio of the 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 highest pressure compressor outlet (before entering the combustor). By way of non-limiting example, the overall pressure ratio of the gas turbine engine as described and / or claimed herein at cruise can be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio can be within an inclusive range bounded by any two values ​​in the previous sentence (i.e., these values ​​can form an upper or lower limit), for example, within a range of 50 to 70.

[0306] The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. Under cruise conditions, the specific thrust of the engine described and / or claimed herein may be less than (or approximately) any of the following: 110 Nkg -1 s, 105Nkg -1 s, 100Nkg -1 s, 95Nkg -1 s, 90Nkg -1 s, 85Nkg -1 s or 80Nkg -1 s. The specific thrust may be within an inclusive range defined by any two values ​​in the previous sentence (ie, these values ​​may form an upper or lower limit), for example, within 80 Nkg -1 s to 100Nkg -1 s, or 85Nkg -1 s to 95Nkg -1 Such engines could be particularly efficient compared to conventional gas turbine engines.

[0307] The gas turbine engine as described and / or claimed herein may have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described and / or claimed herein may produce a maximum thrust of at least (or approximately) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit). By way of example only, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 330 kN to 420 kN, such as 350 kN to 400 kN. The thrust mentioned above may be the maximum net thrust under standard atmospheric conditions, at sea level, plus 15°C (ambient pressure 101.3 kPa, temperature 30°C), with the engine at rest.

[0308] During operation, the temperature of the flow at the inlet of the high-pressure turbine can be particularly high. This temperature, which may be referred to as TET, can be measured at the outlet of the combustor, for example, just upstream of the first turbine blades, which may themselves be referred to as nozzle guide vanes. At cruise, the TET may be at least (or approximately) any one of the following: 1400K, 1450K, 1500K, 1550K, 1600K, or 1650K. The TET at cruise may be within an inclusive range bounded by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit). The maximum TET of the engine during operation may be, for example, at least (or approximately) any one of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, or 2000K. The maximum TET may be within an inclusive range bounded by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, within the range of 1800K to 1950K. Maximum TET may occur, for example, during high thrust conditions, such as during maximum takeoff (MTO) conditions.

[0309] The fan blades and / or airfoil portions of the fan blades described and / or claimed herein may be made of any suitable material or combination of materials. For example, at least a portion of the fan blade and / or airfoil may be made at least in part of a composite material, such as a metal matrix composite and / or an organic matrix composite, such as carbon fiber. In another example, at least a portion of the fan blade and / or airfoil may be made at least in part of 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 blade may include at least two regions made of different materials. For example, the fan blade may have a protective leading edge that may be made of a material that is better able to resist impact (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge may be made, for example, of titanium or a titanium-based alloy. Thus, by way of example only, the fan blade may have carbon fiber or an aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0310] A fan as described and / or claimed herein may include a central portion from which fan blades may extend, for example, radially. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fixing that engages with a corresponding slot in the hub (or disc). By way of example only, such fixings may be in the form of a dovetail that can be inserted into and / or engage with corresponding slots in the hub / disc to secure the fan blade to the hub / disc. By way of another example, the fan blades may be formed integrally with the central portion. Such an arrangement may be referred to as a blade disk or blade ring. Any suitable method may be used to manufacture such a blade disk or blade ring. For example, at least a portion of the fan blade may be machined from a block and / or at least a portion of the fan blade may be attached to the hub / disc by welding (such as linear friction welding).

[0311] The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the outlet area of ​​the bypass duct to vary during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0312] The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24, or 26 fan blades.

[0313] As used herein, maximum takeoff (MTO) conditions have conventional meanings. Maximum takeoff conditions can be defined as operating the engine at maximum takeoff thrust at the end of the runway at International Standard Atmosphere (ISA) sea level pressure and temperature conditions of +15°C, which is typically defined as an aircraft speed of approximately 0.25 mn, or between approximately 0.24 mn and 0.27 mn. Thus, the maximum takeoff condition for an engine can be defined as operating the engine at its maximum takeoff thrust (e.g., maximum throttle) at ISA sea level pressure and temperature of +15°C, with a fan inlet speed of 0.25 mn.

[0314] As used herein, the term "cruise condition" has a conventional meaning and will be readily understood by a skilled artisan. Thus, for a given gas turbine engine for an aircraft, a skilled artisan will immediately recognize that cruise condition refers to the operating point of the gas turbine engine during mid-cruise for a given mission (which may be referred to in the industry as an "economy mission") for which the gas turbine engine is designed to be attached to the aircraft. In this regard, mid-cruise is the critical point in the aircraft's flight cycle at which 50% of the total fuel burned between peak ascent and the start of descent has been burned (which may be approximated in terms of time and / or distance to the midpoint between peak ascent and the start of descent). Thus, cruise condition defines the operating point of the gas turbine engine that, taking into account the number of engines provided to the aircraft, provides thrust that will ensure steady-state operation (i.e., maintaining a constant altitude and a constant Mach number) of the aircraft to which the gas turbine engine is designed to be attached during mid-cruise. For example, if the engine is designed to be attached to an aircraft having two engines of the same type, then, during cruise condition, the engines provide half of the total thrust required for steady-state operation of the aircraft during mid-cruise.

[0315] 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) (necessary to provide steady-state operation of the aircraft to which the gas turbine engine is designed, in combination with any other engines on the aircraft, at a given mid-cruise Mach number). 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.

[0316] By way of example only, the forward speed under cruise conditions may be anywhere within the range of Mach 0.7 to Mach 0.9, such as 0.75 to 0.85, such as 0.76 to 0.84, such as 0.77 to 0.83, such as 0.78 to 0.82, such as 0.79 to 0.81, such as approximately Mach 0.8, approximately Mach 0.85, or 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise condition may be outside these ranges, such as below Mach 0.7 or above Mach 0.9.

[0317] By way of example only, cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere (ISA)) at an altitude within the following range: 10,000 m to 15,000 m, for example, within the range of 10,000 m to 12,000 m, for example, within the range of 10,400 m to 11,600 m (approximately 38,000 ft), for example, within the range of 10,500 m to 11,500 m, for example, within the range of 10,600 m to 11,400 m, for example, within the range of 10,700 m (approximately 35,000 ft) to 11,300 m, for example, within the range of 10,800 m to 11,200 m, for example, within the range of 10,900 m to 11,100 m, for example, approximately 11,000 m. Cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

[0318] By way of example only, a cruise condition may correspond to an operating point of the engine that provides a known desired thrust level (e.g., a value in the range of 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 ft (11,582 m). By way of another example only, a cruise condition may correspond to an operating point of the engine that provides a known desired thrust level (e.g., a value in the range of 50 kN to 65 kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 ft (10,668 m).

[0319] In use, the gas turbine engines described and / or claimed herein may be operated at cruise conditions defined elsewhere herein. Such cruise conditions may be determined by a cruise condition (e.g., an intermediate cruise condition) of an aircraft on which at least one (e.g., two or four) gas turbine engines may be mounted to provide propulsive thrust.

[0320] According to one aspect, an aircraft is provided, 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 is designed to be attached. Thus, the cruise condition according to this aspect corresponds to an intermediate cruise of the aircraft, as defined elsewhere herein.

[0321] According to one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation may be performed at cruise conditions (e.g., in terms of thrust, atmospheric conditions, and Mach number) as defined elsewhere herein.

[0322] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. Operation according to this aspect may comprise (or may be) operation at mid-cruise of the aircraft, as defined elsewhere herein.

[0323] Although in the arrangement described herein the drive source for the pusher fan is provided by a gas turbine engine, it will be understood by those skilled in the art that the gearbox configuration disclosed herein may be applied to other forms of aircraft propulsion including alternative drive types. For example, the above-described gearbox arrangement may be used in an aircraft propulsion including a pusher fan driven by an electric motor. In such cases, the electric motor may be configured to operate at a higher rotational speed and therefore may have a smaller rotor diameter and may be more power intensive. The gearbox configuration of the aforementioned aspects may be used to reduce the rotational input speed of the fan or propeller to allow it to operate at a more favorable efficiency state. Thus, according to one aspect, there is provided an electric propulsion unit for an aircraft, the electric propulsion unit comprising a motor configured to drive a pusher fan via a gearbox, the gearbox and / or its input / output / support members and / or the structure through which the fan shaft drives the fan being supported as described and / or claimed herein.

[0324] Those skilled in the art will understand that, unless mutually exclusive, features or parameters described with respect to any one of the above aspects may be applied to any other aspect. In addition, unless mutually exclusive, any features or parameters described herein may be applied to any aspect and / or combined with any other features or parameters described herein.

[0325] As used herein, a range "from a value X to a value Y" or "between a value X and a value Y" or the like means an inclusive range; the bounding 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 position of the radial plane. Axial planes may also be referred to as longitudinal planes because they extend along the length of the engine. Thus, a radial distance or an axial distance is a distance in a radial or axial plane, respectively. BRIEF DESCRIPTION OF THE DRAWINGS

[0326] Embodiments will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0327] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0328] Figure 2 is a close-up cross-sectional side view of an upstream portion of a gas turbine engine;

[0329] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

[0330] Figure 4 is a schematic diagram showing the radial bending stiffness of a cantilever beam;

[0331] Figure 5 is a schematic diagram showing the anti-tilting stiffness of a cantilever beam;

[0332] Figure 6 is a schematic diagram showing the torsional stiffness of a shaft;

[0333] Figure 7 is a close-up cross-sectional view of the area of ​​a gas turbine engine surrounding its gearbox.

[0334] Figure 8 and Figure 9 is a schematic diagram showing how the radial bending stiffness of a gearbox support may be defined;

[0335] Figure 10 and Figure 11 is a schematic diagram showing how the anti-tilting stiffness of a gearbox support can be defined;

[0336] Figure 12 shows a schematic cross-sectional view of a gearbox support for use with a gearbox in a star configuration;

[0337] Figure 13 shows a schematic cross-sectional view of a gearbox support for use with a planetary configuration gearbox;

[0338] Figure 14 is a schematic diagram illustrating an alternative joint between a fan shaft and a fan;

[0339] Figure 15 and Figure 16 is a schematic diagram showing how the radial bending stiffness of a fan shaft end can be defined;

[0340] Figure 17 and Figure 18 is a schematic diagram showing how the anti-tilt stiffness of a fan shaft end can be defined;

[0341] Figure 19 An aircraft is shown having a gas turbine engine attached to each wing;

[0342] Figure 20 A method of operating a gas turbine engine on an aircraft is shown. And

[0343] Figure 21 A graph of applied load versus displacement is shown to illustrate a measure of component stiffness. DETAILED DESCRIPTION

[0344] Figure 1 A gas turbine engine 10 having a main axis of rotation 9 is shown. The engine 10 includes an air inlet 12 and a pusher fan 23 that generates two airflows: 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 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 in an axial flow series arrangement. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low-pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0345] In use, the core air flow A is accelerated and compressed by the low-pressure compressor 14 and directed to the high-pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is directed to the combustion device 16, in which the compressed air is mixed with fuel and the mixture is burned. The resulting hot combustion products are then expanded through the high-pressure turbine and the low-pressure turbine 17, 19 before being discharged through the nozzle 20, thereby driving the high-pressure turbine and the low-pressure turbine to provide some 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.

[0346] Figure 2 An exemplary arrangement of a geared fan gas turbine engine 10 is shown in FIG. Figure 1 ) drive shaft 26, which is coupled to a sun gear 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, which are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously about the sun gear 28 while causing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the fan 23 via a connecting rod 36 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 gear or annulus 38, which is coupled to the fixed support structure 24 via a connecting rod 40.

[0347] The connecting rod 36 may be referred to as a fan shaft 36, which may optionally include two or more shaft portions coupled together. For example, the fan shaft 36 may include a gearbox output shaft portion 36a extending from the gearbox 30 and a fan portion 36b extending between the gearbox output shaft portion and the fan 23. Figure 1 and Figure 2 In the embodiment shown, the gearbox 30 is a planetary gearbox, and the gearbox output shaft portion 36a is connected to the planet carrier 34, so it can be referred to as the carrier output shaft 36a. In the sun gearbox 30, the gearbox output shaft portion 36a can be connected to the ring gear 38, so it can be referred to as the ring output shaft 36a. Figure 1 and 2 In the illustrated embodiment, the fan portion 36b of the fan shaft 36 connects the gearbox output shaft portion 36a to the fan 23. Thus, the output of the gearbox 30 is transmitted to the fan 23 via the fan shaft 36 to rotate the fan. In alternative embodiments, the fan shaft 36 may include a single component or more than two components. Unless otherwise indicated or apparent to one skilled in the art, anything described with respect to the engine 10 having a sun-type gearbox 30 may also apply to an engine having a planetary gearbox 30, and vice versa.

[0348] It should be noted that the terms "low-pressure turbine" and "low-pressure compressor" used herein may refer to the lowest-pressure turbine stage and lowest-pressure compressor stage, respectively (i.e., excluding the fan 23), 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 gearbox output shaft that drives the fan 23). In some literature, the "low-pressure turbine" and "low-pressure compressor" mentioned herein may be alternatively referred to as the "intermediate-pressure turbine" and "intermediate-pressure compressor." In the case of using such alternative nomenclature, the fan 23 may be referred to as the first or lowest-pressure compression stage.

[0349] exist Figure 3The epicyclic gearbox 30 is shown in greater detail by way of example in FIG. Each of the sun gear 28, the planetary gears 32, and the ring gear 38 includes teeth around its periphery for intermeshing with the other gears. However, for clarity, Figure 3 Only an exemplary portion of the teeth is shown. Four planet gears 32 are shown, but it will be apparent to one skilled in the art that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of the planetary epicyclic gearbox 30 typically include at least three planet gears 32.

[0350] exist Figure 2 and Figure 3 The epicyclic gearbox 30 shown by way of example in FIG is a planetary type, in which the planet carrier 34 is coupled to the output shaft via connecting rods 36, with the ring gear 38 being fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of another example, the epicyclic gearbox 30 may be a star arrangement in which the planet carrier 34 remains fixed, allowing the ring gear (or ring gear) 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 may be a differential gearbox in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0351] It should be understood that Figure 2 and Figure 3 The arrangement shown in is exemplary only, and various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement may be used to position the gearbox 30 within the engine 10 and / or to connect the gearbox 30 to the engine 10. By way of another example, the connections (such as the gearbox 30 and the other components of the engine 10, such as the input shaft 26, the output shaft, and the fixed structure 24) between the gearbox 30 and the engine 10 may be such that the gearbox 30 is positioned within the engine 10 and / or connected to the engine 10. Figure 2 The connecting rods 36, 40 in the example may have any desired degree of stiffness or flexibility, as defined or required elsewhere herein. By way of another example, any suitable arrangement of bearings between rotating and stationary components of the engine (e.g., between the input and output shafts from the gearbox and a stationary structure such as the gearbox housing) may be used, and the present disclosure is not limited to Figure 2 For example, where the gearbox 30 has a star-shaped arrangement (as described above), the skilled person will readily appreciate that the arrangement of the output and support links and bearing locations will typically be different than Figure 2 The arrangement structure shown by way of example in FIG (for example, as described in conjunction with other embodiments disclosed herein having a sun-type gearbox arrangement structure).

[0352] Thus, the present disclosure extends to gas turbine engines having any arrangement of gearbox type (eg, sun or planetary), support structure, input and output shaft arrangement, and bearing locations.

[0353] Optionally, the gearbox may drive additional and / or alternative components (eg, an intermediate pressure compressor and / or a booster compressor).

[0354] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may 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 in has splitter nozzles 18, 20, which means that the flow through the bypass duct 22 has its own nozzle 18, which is separate from the core engine nozzle 20 and radially outside the core engine nozzle. However, this is not restrictive, and any aspect of the present disclosure may 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. Although the examples described relate to turbofan engines, the present disclosure may be applied to, for example, any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or, for example, a turboprop engine.

[0355] The geometry of the gas turbine engine 10 and its components is defined by conventional shafting, including an axial direction (aligned with the axis of rotation 9), a radial direction (in Figure 1 from bottom to top) and circumferential direction (perpendicular to Figure 1 Page in view). The axial, radial, and circumferential directions are perpendicular to each other.

[0356] The following general definition of stiffness may be used herein:

[0357] Radial bending stiffness

[0358] Radial bending stiffness is a measure of the deformation for a given force applied in any chosen radial direction (i.e. any direction perpendicular to and passing through the engine axis). Figure 4 It is defined based on the deformation of the cantilever beam 401. Figure 4As 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 the present application, the radial bending stiffness is relative to the axis of rotation 9 of the engine and therefore relates to the resistance to linear deformation in the radial direction of the engine caused by radial forces. The beam or equivalent cantilever beam component extends along the axis of rotation of the engine, the force F is applied in any radial direction perpendicular to the axis of rotation of the engine, and the displacement δ is measured perpendicular to the axis of rotation along the line of action of the force. The radial bending stiffness as defined herein has International System of Units (SI) units of N / m. In the present application, unless otherwise stated, the radial bending stiffness is considered to be the free body stiffness, i.e. the stiffness measured for a separate component in a cantilever construction without the presence of other components that may affect its stiffness.

[0359] When a force is applied perpendicular to a cantilever beam and at the free end of the beam, the resultant curvature is not constant but increases towards the fixed end of the beam.

[0360] Tilt stiffness

[0361] refer to Figure 5 Defining anti-tilting stiffness, the figure shows the resulting deformation of a cantilever beam 401 under a moment M applied at its free end. Tilt stiffness is a measure of the resistance to rotation at the location on the component where the moment is applied. Figure 5 As can be seen in Figure 1, a moment applied at the free end of a cantilever beam induces a constant curvature along the length of the beam, between its free end and its fixed end. The applied moment M induces a rotation angle θ at the point where the moment is applied. For any component with a constant portion (such as this beam), the angle θ is constant along the length of the component. Therefore, the anti-tilting stiffness, as defined herein, has SI units of Nm / rad.

[0362] Torsional stiffness

[0363] Torsional stiffness is a measure of deformation for a given torque. Figure 6 The definition of torsional stiffness of a shaft 401 or other body is shown. A torque τ applied to the free end of a beam causes a rotational deformation θ (e.g., torsion) along the length of the beam. Torsional stiffness is the applied torque for a given torsion angle, i.e., τ / θ. The SI unit of torsional stiffness is Nm / rad.

[0364] The following general definitions of other parameters may also be used in this article:

[0365] Torque

[0366] Torque, also known as moment of force, is the rotational equivalent of a linear force and can be thought of as twisting an object. The magnitude of the torque, τ, on a body depends on three quantities: the applied force (F), the lever arm vector (r) connecting the origin to the point of force application, and the angle (A) between the force and the lever arm vector:

[0367] τ=r×F

[0368] τ=|τ|=|r×F|=|r||F|sinA

[0369] in

[0370] τ is the torque vector and τ is the magnitude of the torque;

[0371] r is the position vector or "lever arm" vector (the vector from the selected point on the body to the point where the force is applied);

[0372] F is the force vector;

[0373] × represents a cross product; and

[0374] A is the angle between the force vector and the lever arm vector (so sin(a) is unity when the force vector is perpendicular to the position vector, so that τ = rF, which is the magnitude of the force multiplied by the distance between the selected point on the body and the point where the force is applied).

[0375] Torque is measured as [force] x [distance] and can be expressed in Newton meters (Nm).

[0376] The net torque on the body determines the rate of change of the body's angular momentum.

[0377] moment of inertia

[0378] Moment of inertia (also called angular mass or rotational inertia) is the amount of torque required to determine a desired angular acceleration of a body about its axis of rotation, which is essentially equivalent to how mass determines the force required for a specific acceleration.

[0379] The moment of inertia depends on the mass distribution of the body and the chosen axis, where larger moments require larger torques to change the body's rate of rotation. The measure of the moment of inertia is [mass] x [distance] 2 and can be kilograms per square meter (kg.m 2 ). The moment of inertia I is defined as the ratio of the net angular momentum L of a body to its angular velocity ω about its major axis:

[0380]

[0381] Assuming the shape of the body does not change, its moment of inertia appears from Newton's laws of motion as the ratio of the torque τ exerted on the body to the angular acceleration α about the main axis:

[0382] τ=Iα

[0383] For a body constrained to rotate within a plane, only the moment of inertia about an axis perpendicular to the plane is important, and so I can be expressed as a scalar value. The skilled person will understand that a fan of a gas turbine engine (and more generally, a fan rotor of a gas turbine engine, including the fan disk and blades, and optionally also the fan shaft and / or other related components) is constrained to rotate in only one plane (the plane perpendicular to the engine axis), and will understand that the moment of inertia of the fan can therefore be defined by a single scalar value.

[0384] Therefore, the moment of inertia of the fan about the engine axis may be measured or defined using any standard method.

[0385] Torsional shear stress

[0386] Shear stress is a stress component coplanar with the material's cross section; this stress tends to produce shear. Shear stress is caused by a force vector component parallel to the material's cross section. Shear stress can be defined as an external force acting on an object or surface parallel to its slope or plane.

[0387] When a shaft or other body is subjected to torque or torsion, shear stress is generated in the body, which can be referred to as torque-induced torsional shear stress. This shear stress varies from zero along the axis of rotation of the torque to a maximum at the part of the body farthest from the axis; therefore, the radial distance from the axis at which the shear stress is to be measured is selected. In the described arrangement, the mid-height of the component where the torsional shear stress is to be measured is selected. For example, if the component has an outer radius of 20 cm, a mid-height position 10 cm from the axis 9 is selected.

[0388] Torsional shear stress has [force] / [distance] 2 The measurement can be Pascal (Pa) or Newton / square meter (N / m 2 ) units.

[0389] Shear strength

[0390] Shear strength is the resistance of a body to failure in shear - Shear strength is the ability of a material to resist forces that would cause the material's internal structure to slide against itself. A shear load is a force that tends to produce sliding failure in a material along a plane parallel to the direction of the force.

[0391] The ultimate shear strength of a material is the maximum shear stress that can be sustained before the material breaks. The proof strength of a material is the stress at which a specific degree of permanent deformation occurs - for example, 0.2% proof strength is the stress at which 0.2% permanent deformation occurs.

[0392] The strength of a body to resist shear when a rotational or torsional shear load is applied may be referred to as torsional shear strength. A skilled artisan will appreciate that the shear strength of a component is important in selecting the dimensions and materials used to manufacture or construct the component. Shear strength has the same units as torsional shear stress because it is the maximum supportable shear stress; therefore, it may be expressed in Pa or N / m 2 Expressed as a unit.

[0393] As used herein, unless otherwise specified, the shear strength of a material listed is the 0.2% proof shear strength of the material. The skilled person will appreciate that this is lower than the ultimate shear strength.

[0394] The shear strength of a material generally varies with temperature. Shear strength as used herein can be defined at room temperature.

[0395] For metals (including metal alloys), shear strength typically decreases gradually with increasing temperature until a threshold temperature is reached, above which the material strength decreases rapidly. Thus, the shear strength may be approximately constant below the threshold temperature. For various steel grades, the threshold temperature may be approximately 400°C-500°C. Because the temperature in and around the gearbox 30 typically does not exceed 120°C, well below the threshold temperature of the likely materials of the gearbox, gearbox support structure, and shaft construction, this shear strength may not differ significantly from the room temperature shear strength, and thus the choice of a specific temperature for strength evaluation may have little impact.

[0396] Torsional strength

[0397] The torsional strength of a particular component is defined as the ability of the component to withstand an applied torque without failing (i.e., without yielding or structural failure). Torsional strength is therefore a measure of the torque capacity of a component. Torsional strength can be measured by applying varying torques to the component and determining the torque at which the component fails. Torsional strength can be considered to be the torque applied to the component at the point at which the shear strength of the component is reached, resulting in component failure. Torsional strength, as defined herein, has a measure of [force] x [distance] and can be expressed in units of Nm.

[0398] To facilitate understanding, more specific definitions of stiffness and other parameters relevant to the embodiments described herein are provided below.

[0399] Gearbox support stiffness

[0400] Figure 7 The area surrounding the gearbox of the engine core 11 is shown in close-up. The same reference numerals are used for Figures 1 to 3 The components shown correspond to the components in Figure 7In the illustrated arrangement, gearbox 30 has a star-type arrangement, with ring gear 38 coupled to fan shaft 36 and carrier 34 held in a fixed position relative to the static structure of the engine core (eg, relative to stationary support structure 24 ).

[0401] The fan shaft 36 is mounted within the engine via a fan shaft mounting structure 503. The fan shaft mounting structure 503 includes at least two bearings connected to or otherwise engaged with the fan shaft at points axially spaced apart along the length of the engine. The fan shaft mounting structure 503 can take a variety of different forms and can include one or more separate support structures configured to support the fan shaft. It can also include other structures configured to support the fan shaft, such as interaxial bearings. Thus, it includes any support structure extending between a bearing in contact with the fan shaft and a stationary structure of the engine (e.g., the engine core).

[0402] exist Figure 7 In the illustrated arrangement, the fan shaft mounting structure 503 includes two bearings, namely a first support bearing 506a and a second support bearing 506b, via which the fan shaft mounting structure 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 depicted arrangement, both support bearings 506a, 506b are positioned forward of the gearbox 30. In other arrangements, one of the two support bearings 506a, 506b supporting the fan shaft 36 may be positioned rearward of the gearbox 30. In other arrangements, more than two support bearings may be provided as part of or as part of the fan shaft mounting structure.

[0403] The engine core 11 includes a gearbox support 40 (corresponding to the reference Figure 2 The gearbox support is arranged to support or mount the gearbox 30 (at least partially) in a fixed position within the engine. The gearbox support is coupled at a first end to a fixed support structure 24 that extends across the core duct 502 carrying the core airflow A, as shown. Figure 7As shown. In the presently described arrangement, the fixed support structure 24 is an engine section stator (ESS) that serves both as a structural component to provide a fixed mounting for core components such as the gearbox support and as guide vanes provided to direct airflow from the fan 23. In other embodiments, the fixed support structure 24 may include struts extending across the core airflow path and separate stator vanes provided to direct airflow. In the presently described arrangement, the gearbox support 40 is coupled to the planet carrier 34 at a second end. Thus, the gearbox support 40 resists rotation of the planet carrier 34 relative to the static structure of the engine core (e.g., relative to the static support structure 24).

[0404] In embodiments where the gearbox 30 is in a planetary arrangement, the gearbox support 40 is coupled to the ring gear 38 so as to resist rotation thereof relative to the static structure of the engine core (eg, relative to the static support structure 24 ).

[0405] The gearbox support 40 is defined between its point of connection to the gearbox (e.g., to the planet carrier in the presently described arrangement) and its point of connection to the fixed support structure 24. The gearbox support may be formed from any number of separate components providing a coupling between these two points.

[0406] The gearbox support 40 has a certain degree of flexibility, which is characterized by its radial bending stiffness and its tilting stiffness.

[0407] Radial bending stiffness of gearbox support :

[0408] refer to Figure 8 and Figure 9 Define the radial bending stiffness of the gearbox support 40. The radial bending stiffness can be considered to represent the resistance of the gearbox support to forces applied to it in the radial direction of the engine. The radial bending stiffness is determined by considering the gearbox support 40 as a free body, which is fixed at its connection point with the stationary support structure 24 of the engine and has a radial force F1 applied at its connection point with the gearbox 30, as shown in FIG. Figure 8 shown. Figure 9 The deformation of the gearbox support 40 caused by the applied force F1 is shown in FIG, wherein the shape of the support without the applied force is shown in dashed lines for comparison. The radial bending stiffness is defined as the radial displacement δ1 of the gearbox support at the location where the force F1 is applied. Figure 8 In FIG, force F1 is shown as being radially towards the engine axis 9, but it can be equivalently a force radially away from the engine axis 9. Therefore, the radial bending stiffness of the gearbox support 40 is given by F1 / δ1. Even if the gearbox support 40 to which the gearbox 30 is connected is deformed, Figure 9The gearbox 30 is shown for reference in FIG. 3 held in a stationary position for illustrative purposes only and reflects that the gearbox support 40 is considered a free body for purposes of determining stiffness.

[0409] In various embodiments, the radial bending stiffness of the gearbox support may be greater than or equal to 1.0×10 7 N / m, and optionally greater than or equal to 2.0×10 7 N / m, or greater than or equal to 3.0×10 7 N / m.

[0410] In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness of the gearbox support may be greater than or equal to 1.0×10 7 N / m or greater than or equal to 2.7×10 7 In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness of the gearbox support may be greater than or equal to 3.2×10 7 N / m or greater than or equal to 4.0×10 7 N / m.

[0411] In various embodiments, the radial bending stiffness of the gearbox support can be within 1.0×10 7 N / m to 4.0×10 8 N / m, and optionally in the range of 2.0×10 7 N / m to 3×10 8 N / m, or optionally in the range of 3.0×10 7 N / m to 2.0×10 8 In the range of N / m.

[0412] In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness of the gearbox support may be in the range of 1.0×10 7 N / m to 3.1×10 8 N / m, and optionally in the range of 2.7×10 7 N / m to 3.7×10 7 N / m (and optionally may be equal to 3.2×10 7 N / m).

[0413] In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness of the gearbox support may be in the range of 3.2×10 7 N / m to 4.0×10 8N / m, and optionally in the range of 4.0×10 7 N / m to 5.0×10 7 N / m range (and optionally may be equal to 4.5×10 7 N / m).

[0414] Gearbox support anti-tilting stiffness

[0415] refer to Figure 10 and Figure 11 The anti-tilt stiffness of the gearbox support 40 is defined. This anti-tilt stiffness can be considered to represent the resistance of the gearbox support to an applied moment. The anti-tilt stiffness is determined by treating the gearbox support 40 as a free body fixed at its connection point with the fixed support structure 24. To measure the anti-tilt stiffness, a moment M1 is applied at the connection point between the gearbox support 40 and the gearbox 30, as Figure 10 exemplified. exist Figure 11 The deformation of the gearbox support 40 caused by the applied moment M1 is shown in FIG, wherein the shape of the support without the applied moment is shown in dashed lines. The anti-tilting stiffness is defined based on the angular displacement θ1 of the gearbox support 40 at the location where the moment M1 is applied. Therefore, the anti-tilting stiffness is given by M1 / θ1. Even if the gearbox support 40 to which the gearbox 30 is connected is deformed, Figure 11 The gearbox 30 is shown again for reference held in a stationary position.

[0416] In various embodiments, the anti-tilt stiffness of the gearbox support can be greater than or equal to 1.2×10 5 Nm / rad, and optionally greater than or equal to 2.4×10 5 Nm / rad, or greater than or equal to 3.9×10 5 Nm / rad.

[0417] In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the anti-tilt stiffness of the gearbox support may be greater than or equal to 1.2×10 5 Nm / rad or greater than or equal to 4.5×10 5 In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the anti-tilt stiffness of the gearbox support may be greater than or equal to 5.0×10 5 Nm / rad or greater than or equal to 6.0×10 5 Nm / rad.

[0418] In various embodiments, the anti-tilt stiffness of the gearbox support can be 1.2×10 5 Nm / rad to 2.1×107 Nm / rad, and optionally in the range of 2.4×10 5 Nm / rad to 1.6×10 7 Nm / rad, or optionally also in the range of 3.9×10 5 Nm / rad to 9.0×10 6 In the range of Nm / rad.

[0419] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the anti-tilt stiffness of the gearbox support may be in the range of 1.2×10 5 Nm / rad to 7.0×10 6 Nm / rad, and optionally in the range of 4.5×10 5 Nm / rad to 6.5×10 5 Nm / rad (and optionally may be equal to 5.5×10 5 Nm / rad).

[0420] In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the anti-tilt stiffness of the gearbox support may be in the range of 5.0×10 5 Nm / rad to 2.1×10 7 Nm / rad, and optionally in the range of 6.0×10 5 Nm / rad to 2.6×10 6 Nm / rad range (and can be equal to 1.6×10 6 Nm / rad).

[0421] Torsional shear stress and torsional strength of gearbox supports

[0422] As described elsewhere herein, a gearbox support is provided to position the gearbox within the engine. In order to resist relative rotation of the gearbox compared to the stationary structure of the engine, torque is transmitted through the gearbox support. This varies during the operating cycle of the engine, as different levels of torque are transmitted through the gearbox (as defined elsewhere herein) at different stages of the flight cycle of the aircraft to which the engine is mounted.

[0423] The torque transmitted through the gearbox support is defined as the torque at the connection point between the gearbox support 40 and the gearbox 30 .

[0424] The torsional strength of the gearbox support 40 is defined as the level of torque applied at the connection point between the gearbox support 40 and the gearbox 20 that would cause the gearbox support to fail.

[0425] When measuring the torsional strength or torque transmitted through the gearbox support 40, the gearbox support is considered to be a free body fixed at its connection point to the stationary support structure 24 of the engine, where the torque is applied at the connection point to the gearbox (i.e., in a manner similar to the force or moment applied to determine the radial bending stiffness and anti-tilting stiffness of the gearbox support).

[0426] exist Figure 7 In the arrangement shown, the connection point between the gearbox support 40 and the gearbox 30 is the connection point to the gearbox's planet carrier 34. In the case of a planetary gearbox, this connection point would be between the gearbox support 40 and the ring gear 38.

[0427] In various embodiments, the torsional strength of the gearbox support may be greater than or equal to 1.60×10 5 Nm, and optionally greater than or equal to 1.8×10 5 Nm.

[0428] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torsional strength of the gearbox support may be greater than or equal to 1.8×10 5 In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the torsional strength of the gearbox support may be greater than or equal to 4.0×10 5 Nm or greater than or equal to 5.5×10 5 Nm.

[0429] In various embodiments, the torsional strength of the gearbox support can be between 1.60×10 5 Nm to 2.00×10 7 Nm, and optionally in the range of 1.8×10 5 Nm to 1.5×10 6 Nm range.

[0430] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torsional strength of the gearbox support may be in the range of 1.8×10 5 Nm to 7.0×10 5 Nm, and optionally in the range of 1.8×10 5 Nm to 2.6×10 5 Nm range (and can be equal to 2.2×10 5 Nm).

[0431] In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the torsional strength of the gearbox support may be in the range of 4.0×10 5 Nm to 2.0×10 7 Nm, and optionally in the range of 5.5×10 5 Nm to 7.5×10 5 Nm range (and can be equal to 6.5×10 5 Nm).

[0432] Torque transmitted through the gearbox support under MOT conditions:

[0433] In various embodiments, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 6.00×10 4 Nm, and optionally greater than or equal to 7.2×10 4 Nm.

[0434] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 7.0 × 10 4 In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 1.8 × 10 5 Nm.

[0435] In various embodiments, the torque transmitted through the gearbox support at maximum takeoff conditions may be 6.00 × 10 4 Nm to 5.00×10 5 Nm, and optionally in the range of 7.2×10 4 Nm to 4.2×10 5 Nm range.

[0436] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torque transmitted through the gearbox support at maximum takeoff conditions may be in the range of 7.0 × 10 4 Nm to 1.9×10 5 Nm range.

[0437] In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the torque transmitted through the gearbox support at maximum takeoff conditions may be in the range of 1.8 × 10 5 Nm to 4.5×10 5 Nm range.

[0438] The values ​​in the above paragraphs can be applied to any gearbox configuration (ie, sun or planetary or other gearbox arrangements).

[0439] In various embodiments, such as those in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 1.10×10 5 Nm, and optionally greater than or equal to 1.3×10 5 Nm.

[0440] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm and / or in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 1.2×10 5 Nm or greater than or equal to 1.4×10 5 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm and / or in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 3.0×10 5 Nm or greater than or equal to 3.4×10 5 Nm.

[0441] In various embodiments, such as those in which the gearbox is of a star configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be between 1.10×10 5 Nm to 5.00×10 5 Nm, and optionally in the range of 1.3×10 5 Nm to 4.2×10 5 Nm range.

[0442] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm and / or in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be in the range of 1.2×10 5 Nm to 1.9×10 5 Nm, and optionally in the range of 1.4×10 5 Nm to 1.8×10 5 Nm range (and can be equal to 1.6×10 5 Nm).

[0443] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm and / or in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be 3.0×10 5 Nm to 4.5×10 5 Nm, and optionally in the range of 3.4×10 5 Nm to 4.2×10 5 Nm range (and can be equal to 3.8×10 5 Nm).

[0444] In various embodiments, such as those in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 6.00×10 4 Nm, and optionally greater than or equal to 7.2×10 4 Nm.

[0445] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm and / or in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 7.0×10 4 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm and / or in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be greater than or equal to 1.8×10 5 Nm.

[0446] In various embodiments, such as those in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be 6.00×10 4 Nm to 3.00×10 5 Nm, and optionally in the range of 7.2×10 4 Nm to 2.6×10 5 Nm range.

[0447] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm and / or in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be in the range of 7.0×10 4 Nm to 1.1×10 5 Nm range.

[0448] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm and / or in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support at maximum takeoff conditions may be in the range of 1.8×10 5 Nm to 2.5×10 5 Nm range.

[0449] Torque transmitted through the gearbox support under cruising conditions:

[0450] In various embodiments, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 2.00×10 4 Nm, and optionally greater than or equal to 2.2×10 4 Nm.

[0451] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 2.2×10 4 In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 6.8 × 10 4 Nm.

[0452] In various embodiments, the torque transmitted through the gearbox support under cruise conditions may be 2.00×10 4 Nm to 2.00×10 5 Nm, and optionally in the range of 2.2×10 4 Nm to 1.7×10 5 Nm range.

[0453] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torque transmitted through the gearbox support under cruise conditions may be in the range of 2.2×10 4 Nm to 6.6×10 4 Nm range.

[0454] In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the torque transmitted through the gearbox support under cruise conditions may be in the range of 6.8×10 4 Nm to 1.8×10 5 Nm range.

[0455] The values ​​in the above paragraphs can be applied to any gearbox configuration (ie, sun or planetary or other gearbox arrangements).

[0456] In various embodiments, such as those in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 4.00×10 4 Nm, and optionally greater than or equal to 4.8×10 4 Nm.

[0457] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm and / or in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 4.5×10 4 Nm or greater than or equal to 5.0×10 4 In some embodiments, such as in embodiments where the fan diameter is in the range of 330 cm to 380 cm and / or in embodiments where the gearbox is of a star-type configuration, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 1.2×10 5 Nm.

[0458] In various embodiments, such as those in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support under cruise conditions may be 4.00×10 4 Nm to 2.00×10 5 Nm, and optionally in the range of 4.8×10 4 Nm to 1.7×10 5 Nm range.

[0459] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm and / or in embodiments in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support under cruise conditions may be in the range of 4.5×10 4 Nm to 6.6×10 4 Nm, and optionally in the range of 5.0×10 4 Nm to 6.0×10 4 Nm range (and can be equal to 5.5z×10 4 Nm).

[0460] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm and / or in embodiments in which the gearbox is of a star-type configuration, the torque transmitted through the gearbox support under cruise conditions may be in the range of 1.2×10 5 Nm to 1.8×10 5 Nm, and optionally in the range of 1.2×10 5 Nm to 1.8×105 Nm range (and can be equal to 1.5×10 5 Nm).

[0461] In various embodiments, such as those in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 2.00×10 4 Nm, and optionally greater than or equal to 2.2×10 4 Nm.

[0462] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm and / or in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 2.2×10 4 In some embodiments, such as in embodiments where the fan diameter is in the range of 330 cm to 380 cm and / or in embodiments where the gearbox is of a planetary configuration, the torque transmitted through the gearbox support under cruise conditions may be greater than or equal to 6.8×10 4 Nm.

[0463] In various embodiments, such as those in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support under cruise conditions may be between 2.00×10 4 Nm to 1.30×10 5 Nm, and optionally in the range of 2.2×10 4 Nm to 1.1×10 5 Nm range.

[0464] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm and / or in embodiments in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support under cruise conditions may be in the range of 2.2×10 4 Nm to 3.3×10 4 Nm range.

[0465] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm and / or in which the gearbox is of a planetary configuration, the torque transmitted through the gearbox support under cruise conditions may be in the range of 6.8×10 4 Nm to 1.1×10 5 Nm range.

[0466] Torsional shear stress in gearbox supports

[0467] The gearbox support 40 has a torsional shear stress which also represents the resistance of the gearbox support 40 to the torque applied by the gearbox 30. The torsional shear stress is as defined elsewhere herein.

[0468] In various embodiments, the torsional shear stress of the gearbox support at maximum takeoff conditions may be less than or equal to 4.90×10 8 N / m 2 , and optionally less than or equal to 3.5×10 8 N / m 2 .

[0469] In various embodiments, the torsional shear stress of the gearbox support at maximum takeoff conditions may be 1.40×10 8 N / m 2 to 4.90×10 8 N / m 2 in the range of, and optionally in the range of 2.0×10 8 N / m 2 to 3.5×10 8 N / m 2 within the range.

[0470] In some embodiments, such as those 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 torsional shear stress of the gearbox support at maximum takeoff conditions may be in the range of 2.3×10 8 N / m 2 to 3.7×10 8 N / m 2 within the range (and can be equal to 2.5×10 8 N / m 2 ).

[0471] In various embodiments, the torsional shear stress of the gearbox support under cruise conditions may be greater than or equal to 7.00×10 7 N / m 2 , and optionally greater than or equal to 8.2×10 7 N / m 2 .

[0472] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the torsional shear stress of the gearbox support under cruise conditions may be greater than or equal to 8.0×10 7 N / m 2In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the torsional shear stress of the gearbox support under cruise conditions may be greater than or equal to 8.5×10 7 N / m 2 or greater than or equal to 9.0×10 7 N / m 2 .

[0473] In various embodiments, the torsional shear stress of the gearbox support under cruise conditions may be 7.00×10 7 N / m 2 to 1.90×10 8 N / m 2 in the range of 8.2×10 7 N / m 2 to 1.5×10 8 N / m 2 within the range.

[0474] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torsional shear stress of the gearbox support under cruise conditions may be in the range of 8.0×10 7 N / m 2 to 1.5×10 8 N / m 2 in the range of 8.0×10 7 N / m 2 to 9.2×10 7 N / m 2 within the range (and can be equal to 8.6×10 7 N / m 2 ).

[0475] In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the torsional shear stress of the gearbox support under cruise conditions may be in the range of 8.5×10 7 N / m 2 to 1.9×10 8 N / m 2 in the range of 9.0×10 7 Nm 2 to 1.2×10 8 N / m 2 within the range (and can be equal to 9.6×10 7 N / m 2 ).

[0476] Maximum takeoff conditions and cruise conditions referred to in this section may be as defined elsewhere herein.

[0477] This shear stress can be measured through the highest loaded flat portion of the component (i.e., the gearbox support). It does not include the effects of stress concentrations at small radii and small holes. A skilled person will understand that for the shear stress of the gearbox support defined herein, since the primary load is torque, the effective radius also affects the shear measurement and area.

[0478] Fan moment of inertia

[0479] 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. the total mass including the plurality of fan blades, the fan hub and any support arms or other connecting rods provided for connecting the fan to the fan shaft. Therefore, 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 relative to the rotation about the main axis of rotation 9 of the engine. The rotation of the fan will cause a gyroscopic effect, which means that the fan shaft will tend to maintain a stable direction of its axis of rotation. However, during maneuvering of an aircraft equipped with a gas turbine engine, the orientation of the axis of rotation 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 produced and therefore have an impact on the design of the fan shaft and the fan shaft mounting structure, as discussed elsewhere herein.

[0480] In various embodiments, the fan's moment of inertia may be greater than or equal to 7.40×10 7 kgm 2 , and optionally greater than or equal to 8.3×10 7 kgm 2 .

[0481] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the fan's moment of inertia may be greater than or equal to 7.4×10 7 kgm 2 or 8.6×10 7 kgm 2 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan's moment of inertia may be greater than or equal to 3.0×10 8 kgm 2 or 4.0×10 8 kgm 2 .

[0482] In various embodiments, the fan's moment of inertia may be 7.40×10 7 kgm2 to 9.00×10 8 kgm 2 in the range of 8.3×10 7 kgm 2 to 6.5×10 8 kgm 2 within the range.

[0483] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the fan's moment of inertia may be 7.4×10 7 kgm 2 to 1.5×10 8 kgm 2 in the range of 8.6×10 7 kgm 2 to 9.6×10 7 kgm 2 within the range (and can be equal to 9.1×10 7 kgm 2 ).

[0484] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan's moment of inertia may be 3.0×10 8 kgm 2 to 9.0×10 8 kgm 2 range, and optionally can be in the range of 4.0×10 8 kgm 2 to 5.0×10 8 kgm 2 within the range (and can be equal to 4.5×10 8 kgm 2 ).

[0485] Relative position of fan and gearbox

[0486] See again Figure 7 , the fan-gearbox axial distance 110 is defined as the gearbox output position (at Figures 15 to 18 The axial distance between an axial position P (also marked as X in FIG) and the fan axial centerline Q is defined as the axial midpoint of the fan blades forming the fan. The gearbox output position is defined as the connection point between the fan shaft 36 and the gearbox. This can be defined differently for different types of gearboxes, as described below.

[0487] In various embodiments, the fan-gearbox axial distance can be greater than or equal to 0.35 m, and optionally greater than or equal to 0.37 m.

[0488] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the fan-gearbox axial distance can be greater than or equal to 0.38 m, or greater than or equal to 0.40 m. In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the fan-gearbox axial distance can be greater than or equal to 0.48 m, or greater than or equal to 0.50 m.

[0489] In various embodiments, the fan-gearbox axial distance may be in the range of 0.35 m to 0.8 m, and optionally in the range of 0.37 m to 0.75 m.

[0490] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the fan-gearbox axial distance may be in the range of 0.38 m to 0.65 m, and optionally may be in the range of 0.40 m to 0.44 m (and may be equal to 0.42 m).

[0491] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan-gearbox axial distance may be in the range of 0.48 m to 0.8 m, and optionally may be in the range of 0.50 m to 0.66 m (and may be equal to 0.58 m).

[0492] The fan shaft 36 is defined as the torque transfer component that extends from the output end of the gearbox to the input end of the fan. Therefore, it includes any gearbox output shaft and part or all of the fan input shaft that may be provided between these points. For the purpose of defining the stiffness of the fan shaft 36, the fan shaft is considered to extend between the fan input location and the gearbox output location and include all torque transfer components between these points. Therefore, it does not include any components of the gearbox that transfer discrete forces other than fan shaft torque (for example, a planet carrier or a connecting plate coupled thereto). Therefore, the gearbox output location can be defined as the connection point between the fan shaft 36 and the gearbox 30. The fan input location can be defined as the connection point between the fan shaft 36 and the fan.

[0493] refer to Figure 12 , where the gearbox is in a sun-type configuration, the gearbox output position is defined as the connection point 702 between the ring gear 38 and the fan shaft 36. More specifically, it is the connection point to the ring of the ring gear 38 (any connection extending from the outer surface of the ring is considered to be part of the ring gear). In the case where the connection point is formed by an interface extending in a direction having an axial component, the connection point is considered to be the axial centerline of the interface, such as Figure 7 shown.

[0494] The fan shaft 36 includes all torque transmitting components up to the connection point 702 with the ring gear 38. Thus, it includes any flexible portions or links 704 of the fan shaft 36 that may be provided, and any connections 706 therebetween (e.g., splined connections).

[0495] In the case where the gearbox 30 is of a planetary configuration, the gearbox output position is also defined as the connection point between the fan shaft 36 and the gearbox 30. Figure 13 An example of such a configuration is shown in FIG, which shows a frame including 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 thereon. The fan shaft 36 is connected to the front plate 34a via a keyed connection 708. In an arrangement such as this, the gearbox output location is considered to be any point on the junction 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.

[0496] Figure 13 Only one example of one type of connection between the fan shaft and the planet carrier 34 is shown. In embodiments having different connection arrangements, the gearbox output location is still considered to be at the junction between a component that transmits torque (i.e., a portion of the fan shaft) and a component that transmits discrete forces (e.g., a portion of the gearbox). The spline connection 708 is only one example of a connection that can be formed between the fan shaft and the gearbox (i.e., in the presently described embodiment, between the fan shaft and the front plate 34b). In other embodiments, the junction forming the gearbox output location can be formed by, for example, a curved connection, a bolted connection, or other toothed or mechanically fixed arrangements.

[0497] The fan input location is defined as a point on the fan shaft that is 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 (e.g., as in FIG. 1 ) that is arranged to connect the fan 23 to the fan shaft 36. Figure 7 ). The support arm 23a is connected to the fan shaft by a keyed coupling 36c that extends along a portion of the length of the fan shaft 36. The fan input location is defined as the axial midpoint of the keyed coupling. The keyed coupling is only one example of a coupling that can form a joint between the fan and the fan shaft. In other embodiments, for example, a curved connection, a bolted connection, or other toothed or mechanically fixed arrangement can be used. For example, a flange coupling can be provided between the support arm 23a and the fan shaft 36. In such an embodiment, the support arm can be connected at the rear of the fan hub. In this embodiment, the fan input location is the axial midpoint of the flange coupling. Figure 14An arrangement is shown in which an alternative coupling is provided between the fan 23 and the fan shaft 36. Figure 7 , the fan 23 is coupled to the fan shaft 36 via the support arm 23a. However, in this arrangement, a flange connector 36d is provided between the support arm 23a and the fan shaft 36. In this embodiment, the support arm 23a may be connected at the rear of the fan hub. The flange connector 36d may be a curved connector. In other embodiments, other forms of flange connectors may be provided. Figure 14 In the embodiment of FIG. 1 , the fan input location is the axial midpoint of the flange connection (marked Y).

[0498] The fan shaft 36 has a certain degree of flexibility, which is characterized by its radial bending stiffness and its anti-tilting stiffness.

[0499] Fan shaft end stiffness at the gearbox output:

[0500] refer to Figures 15 to 18 Defines the stiffness of the fan shaft, which is coupled to the gearbox 30 .

[0501] By defining the gearbox output position (such as Figure 15 The radial bending stiffness of the fan shaft 36 at the output end of the gearbox 30 is measured by applying a force F2 to the fan shaft at the output end of the gearbox 30. The fan shaft 36 is considered as a free body and is held fixed at all bearing locations where it is supported (i.e., Figure 15 Due to the force F2, the fan shaft 36 is deformed, causing the gearbox output position to shift by a distance δ2 (as shown in FIG. Figure 16 ). The radial bending stiffness of the fan shaft 36 at the output end of the gearbox is then given by F2 / δ2.

[0502] In various embodiments, the radial bending stiffness of the fan shaft at the output end of the gearbox may be greater than or equal to 4.00×10 6 N / m, and optionally greater than or equal to 3.7×10 7 N / m.

[0503] In some embodiments, such as those 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 output end of the gearbox may be greater than or equal to 3.7×10 7 In some embodiments, such as those 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 end of the gearbox may be greater than or equal to 3.9×10 7 N / m or greater than or equal to 5.0×10 7 N / m.

[0504] In various embodiments, the radial bending stiffness of the fan shaft at the output end of the gearbox may be 4.00×10 6 N / m to 1.5×10 9 N / m, and optionally in the range of 3.7×10 7 N / m to 1.0×10 9 In the range of N / m.

[0505] In some embodiments, such as those 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 output end of the gearbox may be in the range of 3.7×10 7 N / m to 5.0×10 8 N / m, and optionally in the range of 3.7×10 7 N / m to 4.3×10 7 N / m range (and can be equal to 4.0×10 7 N / m).

[0506] In some embodiments, such as those 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 output end of the gearbox may be 3.9×10 7 N / m to 1.5×10 9 N / m, and optionally in the range of 5.0×10 7 N / m to 9.0×10 7 N / m range (and can be equal to 7.0×10 7 N / m).

[0507] By defining the gearbox output position (such as Figure 17 The anti-tilt stiffness of the fan shaft 36 at the output of the gearbox 30 is measured by applying a torque M2 to the fan shaft at the output of the gearbox 30. The fan shaft 36 is again considered as a free body and is held fixed at all bearing locations where it is supported (i.e., Figure 15 Due to the moment M2, the fan shaft 36 is deformed, causing the gearbox output position to shift by an angular displacement θ2, as shown in FIG. Figure 17 The anti-tilt stiffness of the fan shaft 36 at the output end of the gearbox is then given by M2 / θ2.

[0508] In various embodiments, the anti-tilt stiffness of the fan shaft at the output end of the gearbox can be greater than or equal to 7.00×10 4 Nm / rad, and optionally greater than or equal to 9.5×10 5 Nm / rad.

[0509] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the anti-tilt stiffness of the fan shaft at the output end of the gearbox may be greater than or equal to 9.5×10 5 Nm / rad or greater than or equal to 9.5×10 5 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the anti-tilt stiffness of the fan shaft at the output end of the gearbox may be greater than or equal to 1.1×10 6 Nm / rad or greater than or equal to 2.0×10 6 Nm / rad.

[0510] In various embodiments, the anti-tilt stiffness of the fan shaft at the output end of the gearbox may be 7.00×10 4 Nm / rad to 7.00×10 7 Nm / rad, and optionally in the range of 9.5×10 5 Nm / rad to 3.5×10 7 In the range of Nm / rad.

[0511] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the anti-tilt stiffness of the fan shaft at the output end of the gearbox may be in the range of 9.5×10 5 Nm / rad to 2.0×10 7 Nm / rad, and optionally in the range of 9.5×10 5 Nm / rad to 2.4×10 6 Nm / rad range (and can be equal to 1.2×10 6 Nm / rad).

[0512] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the anti-tilt stiffness of the fan shaft at the output end of the gearbox may be in the range of 1.1×10 6 Nm / rad to 7.0×10 7 Nm / rad, and optionally in the range of 2.0×10 6 Nm / rad to 5.2×10 6 Nm / rad range (and can be equal to 3.6×10 6 Nm / rad).

[0513] Torque transmission in gearboxes

[0514] The gearbox provides torque conversion between the torque at its input (i.e., the spindle) and its output (i.e., the fan shaft). The torque transmitted by the gearbox is defined as the torque at the output of the gearbox (the output position being defined elsewhere herein). The torque transmitted by the gearbox varies with the operating cycle of the engine.

[0515] In various embodiments, the torque transmitted through the gearbox at maximum takeoff conditions may be greater than or equal to 7.00 × 10 4 Nm, and optionally greater than or equal to 1.0×10 5 Nm.

[0516] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torque transmitted through the gearbox at maximum takeoff conditions may be greater than or equal to 1.1×10 5 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the torque transmitted through the gearbox at maximum takeoff conditions may be greater than or equal to 1.5×10 5 Nm or greater than or equal to 2.0×10 5 Nm.

[0517] In various embodiments, the torque transmitted through the gearbox at maximum takeoff conditions may be 7.00 × 10 4 Nm to 5.00×10 5 Nm, and optionally in the range of 1.0×10 5 Nm to 3.5×10 5 Nm range.

[0518] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the torque transmitted through the gearbox at maximum takeoff conditions may be in the range of 1.1×10 5 Nm to 1.5×10 5 Nm, and optionally in the range of 1.1×10 5 Nm to 1.3×10 5 Nm range (and can be equal to 1.2×10 5 Nm).

[0519] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the torque transmitted through the gearbox at maximum takeoff conditions may be in the range of 1.5×10 5 Nm to 5.0×10 5 Nm, and optionally in the range of 2.0×10 5 Nm to 3.8×105 Nm range (and can be equal to 2.9×10 5 Nm).

[0520] In various embodiments, the torque transmitted through the gearbox under cruise conditions may be greater than or equal to 2.30×10 4 Nm, and optionally greater than or equal to 3.1×10 4 Nm.

[0521] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the torque transmitted through the gearbox under cruise conditions may be greater than or equal to 3.2×10 4 Nm or greater than or equal to 3.8×10 4 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the torque transmitted through the gearbox under cruise conditions may be greater than or equal to 7.3×10 4 Nm or greater than or equal to 9.8×10 4 Nm.

[0522] In various embodiments, the torque transmitted through the gearbox under cruise conditions may be 2.30×10 4 Nm to 1.80×10 5 Nm, and optionally in the range of 3.1×10 4 Nm to 1.5×10 5 Nm range.

[0523] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the torque transmitted through the gearbox under cruise conditions may be in the range of 3.2×10 4 Nm to 7.2×10 4 Nm, and optionally in the range of 3.8×10 4 Nm to 4.6×10 4 Nm range (and can be equal to 4.2×10 4 Nm).

[0524] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the torque transmitted through the gearbox under cruise conditions may be in the range of 7.3×10 4 Nm to 1.8×10 5 Nm, and optionally in the range of 9.8×10 4 Nm to 1.4×10 5 Nm range (and can be equal to 1.1×10 5 Nm).

[0525] Gearbox CSA

[0526] The cross-sectional area (CSA) of a gearbox is defined as the area of ​​the ring gear's pitch circle. A gear's pitch circle is an imaginary circle that rolls with the pitch circles of any other gears meshing with the first gear without slipping. The pitch circle passes through the point where the two meshing gears meet during rotation; it typically passes through the midpoint of the gear teeth's length. The gearbox's CSA can be determined by measuring the gear teeth's pitch circle diameter (PCD). A rough estimate of the PCD can be obtained by taking the average of the diameter between the tips of the gear teeth and the diameter between the bases of the gear teeth.

[0527] In various embodiments, the CSA of the gearbox may be greater than or equal to 2.4×10 -1 m 2 , and optionally greater than or equal to 2.6×10 -1 m 2 .

[0528] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the CSA of the gearbox may be greater than or equal to 2.4×10 -1 m 2 or greater than or equal to 2.5×10 -1 m 2 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the CSA of the gearbox may be greater than or equal to 4.5×10 -1 m 2 or greater than or equal to 5.5×10 -1 m 2 .

[0529] In various embodiments, the CSA of the gearbox may be 2.4×10 -1 m 2 Up to 1.10m 2 in the range of, and optionally in the range of 2.6×10 -1 m 2 to 9.0×10 -1 m 2 within the range.

[0530] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the CSA of the gearbox may be in the range of 2.4×10 -1 m 2 to 5.0×10 -1 m 2 in the range of 2.5×10-1 m 2 to 3.4×10 -1 m 2 within the range (and can be equal to 2.9×10 -1 m 2 ).

[0531] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the CSA of the gearbox may be in the range of 4.5×10 -1 m 2 Up to 1.1m 2 in the range of 5.5×10 -1 m 2 to 6.4×10 -1 m 2 within the range (and can be equal to 5.9×10 -1 m 2 ).

[0532] Angle between adjacent planetary gears

[0533] The planet gear pitch angle (β) in radians is defined as 2π / N, where N is the number of planet gears 32 provided in the gearbox. Figure 3 The planet gear pitch angle is shown in . The planet gearbox pitch angle corresponds to the average angle (in radians) between all adjacent planet gear pairs.

[0534] In various embodiments, the planet gear spacing angle (β) may be greater than or equal to 9.0×10 -1rad , and optionally at 9.0×10 -1rad to 2.1 rad, and optionally also in the range between 1.1 rad and 1.3 rad (and may be equal to 1.26 rad).

[0535] Parameter ratio

[0536] The inventors have discovered that the ratios (and / or products) of certain properties have a considerable impact on the operation of the gearbox and its input / output / support structure. Some or all of the following may apply to any embodiment:

[0537] In various embodiments, the ratio of radial bending stiffness to moment of inertia may be defined as:

[0538]

[0539] In various embodiments, the ratio of the radial bending stiffness to the moment of inertia may be greater than or equal to 2.5×10 -2 Nkg -1 m-3 (i.e., (N / m) / (kg.m 2 )), and optionally greater than or equal to 0.05 Nkg -1 m -3 .

[0540] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of radial bending stiffness to moment of inertia may be greater than or equal to 0.05 Nkg -1 m -3 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of radial bending stiffness to moment of inertia may be greater than or equal to 0.025 Nkg -1 m -3 .

[0541] In various embodiments, the ratio of radial bending stiffness to moment of inertia may be between 2.5x10 -2 Nkg -1 m -3 Up to 6.0Nkg -1 m -3 in the range of, and optionally in the range of 0.05 Nkg -1 m -3 Up to 3.0Nkg -1 m -3 In the range of, and optionally also in the range of 0.05Nkg -1 m -3 Up to 0.6Nkg -1 m -3 within the range.

[0542] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of radial bending stiffness to moment of inertia may be 0.05 Nkg -1 m -3 Up to 4.0Nkg -1 m -3 within the range.

[0543] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of radial bending stiffness to moment of inertia may be 0.025 Nkg -1 m -3 Up to 2.0Nkg -1 m -3 within the range.

[0544] In various embodiments, in addition to or instead of the ratio of the radial bending stiffness to the moment of inertia, a product of the parameters that make up the ratio can be defined. This product (referred to as the radial bending stiffness and moment of inertia product) can be defined as:

[0545] The radial bending stiffness of at least one of the following

[0546] The anti-tilt stiffness of at least one of the fan shaft (36) and the gearbox support (40) at the output end of the gearbox is equal to the moment of inertia of the fan (23).

[0547] In various embodiments, the product of the radial bending stiffness and the moment of inertia may be greater than or equal to 2.0×10 14 Nkgm(ie, (N / m).(kg.m 2 )), and optionally greater than or equal to 4.0×10 14 Nkgm, and optionally greater than or equal to 2.0×10 15 Nkgm.

[0548] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness multiplied by the moment of inertia may be greater than or equal to 1.5×10 15 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness multiplied by the moment of inertia may be greater than or equal to 2.0×10 15 Nkgm.

[0549] In various embodiments, the radial bending stiffness multiplied by the moment of inertia may be 2.0×10 14 Nkgm to 1.4×10 18 Nkgm, and optionally in the range of 4.0×10 14 Nkgm to 7.0×10 17 Nkgm, and optionally also within the range of 2.0×10 15 Nkgm to 7.0×10 17 within the range of Nkgm.

[0550] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the radial bending stiffness multiplied by the moment of inertia may be in the range of 1.5×10 15 Nkgm to 1.3×10 17 within the range of Nkgm.

[0551] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the radial bending stiffness multiplied by the moment of inertia may be in the range of 2.0×1015 Nkgm to 1.4×10 18 within the range of Nkgm.

[0552] In various embodiments, the ratio of the fan shaft radial bending stiffness to the moment of inertia may be defined as:

[0553]

[0554] In various embodiments, the ratio of the fan shaft radial bending stiffness to the moment of inertia may be greater than or equal to 2.5×10 -2 Nkg -1 m -3 (i.e., (N / m) / (kg.m 2 )), and optionally greater than or equal to 0.05 (N / m) / (kg.m 2 ).

[0555] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of the fan shaft radial bending stiffness to the moment of inertia may be greater than or equal to 0.05 Nkg -1 m -3 or greater than or equal to 0.4Nkg -1 m -3 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of the fan shaft radial bending stiffness to the moment of inertia may be greater than or equal to 0.025 Nkg -1 m -3 or greater than or equal to 0.06Nkg -1 m -3 .

[0556] In various embodiments, the ratio of the fan shaft radial bending stiffness to the moment of inertia may be 2.5x10 -2 Nkg -1 m -3 Up to 6.0Nkg -1 m -3 in the range of, and optionally in the range of 0.05 Nkg -1 m -3 Up to 3.0Nkg -1 m -3 In the range of, and optionally also in the range of 0.05Nkg -1 m -3 Up to 0.6Nkg -1 m -3 within the range.

[0557] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of the fan shaft radial bending stiffness to the moment of inertia may be in the range of 0.05 Nkg -1 m -3 Up to 0.6Nkg -1 m -3 within the range of 0.4 Nkg and optionally at 0.4 Nkg -1 m -3 Up to 0.5Nkg -1 m -3 Within the range (and can be equal to 0.44Nkg -1 m -3 ).

[0558] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of the fan shaft radial bending stiffness to the moment of inertia may be 0.025 Nkg -1 m -3 Up to 0.6Nkg -1 m -3 within the range of 0.06 Nkg and optionally at 0.06 Nkg -1 m -3 Up to 0.26Nkg -1 m -3 Within the range (and can be equal to 0.16Nkg -1 m -3 ).

[0559] In various embodiments, in addition to or instead of the ratio of the fan shaft radial bending stiffness to the moment of inertia, a product of the parameters constituting the ratio may be defined. This product (referred to as the fan shaft radial bending stiffness and moment of inertia product) may be defined as:

[0560] The radial bending stiffness of the fan shaft (36) at the output end of the gearbox×the moment of inertia of the fan (23).

[0561] In various embodiments, the fan shaft radial bending stiffness and the moment of inertia product may be greater than or equal to 3.0×10 14 Nkgm(ie, (N / m).(kg.m 2 )), and optionally greater than or equal to 6.0×10 14 Nkgm, and optionally greater than or equal to 2.0×10 15 Nkgm.

[0562] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft radial bending stiffness multiplied by the moment of inertia may be greater than or equal to 2.0×10 15In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft radial bending stiffness multiplied by the moment of inertia may be greater than or equal to 2.3×10 15 Nkgm.

[0563] In various embodiments, the fan shaft radial bending stiffness and the moment of inertia product can be 3.0×10 14 Nkgm to 1.4×10 18 Nkgm, and optionally in the range of 6.0×10 14 Nkgm to 7.0×10 17 Nkgm, and optionally also within the range of 2.0×10 15 Nkgm to 7.0×10 17 within the range of Nkgm.

[0564] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the fan shaft radial bending stiffness multiplied by the moment of inertia may be in the range of 2.0×10 15 Nkgm to 7.5×10 16 within the range of Nkgm.

[0565] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft radial bending stiffness multiplied by the moment of inertia may be 2.3×10 15 Nkgm to 1.4×10 18 within the range of Nkgm.

[0566] In various embodiments, the ratio of the gearbox support radial bending stiffness to the moment of inertia may be defined as:

[0567]

[0568] In various embodiments, the ratio of the gearbox support radial bending stiffness to the moment of inertia may be greater than or equal to 3.0×10 -2 Nkg -1 m -3 (i.e., (N / m) / (kg.m 2 )), and optionally greater than or equal to 0.06 Nkg -1 m -3 .

[0569] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of the gearbox support radial bending stiffness to moment of inertia may be greater than or equal to 0.06 Nkg -1 m -3or greater than or equal to 0.25Nkg -1 m -3 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of the gearbox support radial bending stiffness to moment of inertia may be greater than or equal to 0.03 Nkg -1 m -3 or greater than or equal to 0.06Nkg -1 m -3 .

[0570] In various embodiments, the ratio of the gearbox support radial bending stiffness to the moment of inertia may be between 3.0×10 - 2 Nkg -1 m -3 Up to 4.0Nkg -1 m -3 in the range of, and optionally in the range of 0.06 Nkg -1 m -3 Up to 2.0Nkg -1 m -3 In the range of, and optionally also in the range of 0.06Nkg -1 m -3 Up to 0.48Nkg -1 m -3 within the range.

[0571] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the gearbox support radial bending stiffness to moment of inertia ratio may be 0.06 Nkg -1 m -3 Up to 4.0Nkg -1 m -3 within the range of 0.25 Nkg -1 m -3 Up to 0.45Nkg -1 m -3 Within the range (and can be equal to 0.35Nkg -1 m -3 ).

[0572] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the gearbox support radial bending stiffness to moment of inertia ratio may be 0.03 Nkg -1 m -3 Up to 2.0Nkg -1 m -3 within the range of 0.06 Nkg and optionally at 0.06 Nkg -1 m -3 Up to 0.6Nkg-1 m -3 Within the range (and can be equal to 0.1Nkg -1 m -3 ).

[0573] In various embodiments, in addition to or instead of the ratio of the gearbox support radial bending stiffness to the moment of inertia, a product of the parameters that make up the ratio can be defined. This product (referred to as the gearbox support radial bending stiffness and moment of inertia product) can be defined as:

[0574]

[0575] In various embodiments, the product of the radial bending stiffness and the moment of inertia of the gearbox support can be greater than or equal to 2.0×10 14 Nkgm(ie, (N / m).(kg.m 2 )), and optionally greater than or equal to 4.0×10 14 Nkgm, or optionally greater than or equal to 2.0×10 15 Nkgm.

[0576] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the gearbox support radial bending stiffness multiplied by the moment of inertia may be greater than or equal to 1.5×10 15 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the gearbox support radial bending stiffness multiplied by the moment of inertia may be greater than or equal to 2.0×10 15 Nkgm.

[0577] In various embodiments, the product of the radial bending stiffness and the moment of inertia of the gearbox support can be 2.0×10 14 Nkgm to 3.0×10 17 Nkgm, and optionally in the range of 4.0×10 14 Nkgm to 1.3×10 17 Nkgm, and optionally also within the range of 2.0×10 15 Nkgm to 1.3×10 17 within the range of Nkgm.

[0578] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the gearbox support radial bending stiffness multiplied by the moment of inertia may be in the range of 1.5×10 15 Nkgm to 1.3×10 17 within the range of Nkgm.

[0579] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the gearbox support radial bending stiffness multiplied by the moment of inertia may be in the range of 2.0×10 15 Nkgm to 3.0×10 17 within the range of Nkgm.

[0580] In various embodiments, the ratio of anti-tilt stiffness to moment of inertia may be defined as:

[0581]

[0582] In various embodiments, the ratio of the anti-tilt stiffness to the moment of inertia may be greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 (i.e., (Nm / rad) / (kg.m 2 )), and optionally greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m -1 .

[0583] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of the anti-tilt stiffness to the moment of inertia may be greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m -1 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of the anti-tilt stiffness to the moment of inertia may be greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 .

[0584] In various embodiments, the ratio of the anti-tilt stiffness to the moment of inertia may be between 4.0×10 -4 Nrad -1 kg -1 m -1 to 2.7×10 -1 Nrad -1 kg -1 m -1 in the range of, and optionally in the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 0.1Nrad -1 kg -1 m-1 in the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 1.5×10 -2 Nrad -1 kg -1 m -1 within the range.

[0585] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of the anti-tilt stiffness to the moment of inertia may be in the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 1.45×10 -2 Nrad -1 kg -1 m -1 within the range.

[0586] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of the anti-tilt stiffness to the moment of inertia may be in the range of 4.0×10 -4 Nrad -1 kg -1 m -1 to 3.0×10 -2 Nrad -1 kg -1 m -1 within the range.

[0587] In various embodiments, in addition to or instead of the ratio of the anti-tilting stiffness to the moment of inertia, a product of the parameters that make up the ratio can be defined. This product (referred to as the anti-tilting stiffness and moment of inertia product) can be defined as:

[0588] The anti-tilt stiffness of at least one of the fan shaft (36) and the gearbox support (40) at the output end of the gearbox is equal to the moment of inertia of the fan (23).

[0589] In various embodiments, the product of the anti-tilting stiffness and the moment of inertia may be greater than or equal to 3.0×10 12 Nm 3 rad -1 kg(i.e., (Nm / rad).(kg.m 2 )), and optionally greater than or equal to 6.0×10 12 Nm 3 rad -1 kg, and optionally greater than or equal to 2.5×10 13 Nm3 rad -1 kg.

[0590] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the product of the anti-tilt stiffness and the moment of inertia may be greater than or equal to 2.0×10 13 Nm 3 rad -1 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the product of the anti-tilting stiffness and the moment of inertia may be greater than or equal to 2.5×10 13 Nm 3 rad -1 kg.

[0591] In various embodiments, the product of the anti-tilting stiffness and the moment of inertia may be 3.0×10 12 Nm 3 rad -1 kg to 6.0×10 16 Nm 3 rad -1 kg, and optionally in the range of 6.0×10 12 Nm 3 rad -1 kg to 3.0×10 16 Nm 3 rad -1 kg, and optionally also in the range of 2.5×10 13 Nm 3 rad -1 kg to 3.0×10 16 Nm 3 rad -1 kg range.

[0592] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the product of the anti-tilt stiffness and the moment of inertia may be in the range of 2.0×10 13 Nm 3 rad -1 kg to 4.0×10 15 Nm 3 rad -1 kg range.

[0593] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the product of the anti-tilt stiffness and the moment of inertia may be in the range of 2.5×10 13 Nm 3 rad -1 kg to 6.0×1016 Nm 3 rad -1 kg range.

[0594] In various embodiments, the ratio of the fan shaft anti-tilt stiffness to the moment of inertia may be defined as:

[0595]

[0596] In various embodiments, the ratio of the fan shaft anti-tilt stiffness to the moment of inertia may be greater than or equal to 4.0×10 - 4 Nrad -1 kg -1 m -1 (i.e., (Nm / rad) / (kg.m 2 )), and optionally greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m -1 .

[0597] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of the fan shaft anti-tilt stiffness to the moment of inertia may be greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m -1 or greater than or equal to 3.0×10 -3 Nrad -1 kg -1 m -1 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of the fan shaft anti-tilt stiffness to the moment of inertia may be greater than or equal to 4.0×10 -4 Nrad -1 kg -1 m -1 or greater than or equal to 0.7×10 -3 Nrad -1 kg -1 m -1 .

[0598] In various embodiments, the fan shaft anti-tilt stiffness to moment of inertia ratio may be 4.0×10 -4 Nrad -1 kg - 1 m -1 to 0.27Nrad -1 kg -1 m -1in the range of, and optionally in the range of 1.0×10 -3 Nrad- 1 kg -1 m -1 to 0.1Nrad -1 kg -1 m -1 in the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 1.5×10 -2 Nrad -1 kg -1 m -1 within the range.

[0599] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of the fan shaft anti-tilt stiffness to the moment of inertia may be in the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 1.45×10 -2 Nrad -1 kg - 1 m -1 in the range of, and optionally in the range of 3.0×10 -3 Nrad -1 kg -1 m -1 to 2.3×10 -2 Nrad -1 kg -1 m -1 within the range (and can be equal to 1.3×10 -2 Nrad -1 kg -1 m -1 ).

[0600] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of the fan shaft anti-tilt stiffness to moment of inertia may be in the range of 4.0×10 -4 Nrad -1 kg -1 m -1 to 1.4×10 -2 Nrad -1 kg -1 m -1 in the range of, and optionally in the range of 7.0×10 -3 Nrad -1 kg -1 m-1 to 9.0×10 -3 Nrad -1 kg -1 m -1 within the range (and can be equal to 8.0×10 -3 Nrad -1 kg -1 m -1 ).

[0601] In various embodiments, in addition to or instead of the ratio of the fan shaft anti-tilt stiffness to the moment of inertia, a product of the parameters that make up the ratio can be defined. This product (referred to as the fan shaft anti-tilt stiffness and moment of inertia product) can be defined as:

[0602] The anti-tilt stiffness of the fan shaft (36) at the output end of the gearbox×the moment of inertia of the fan (23).

[0603] In various embodiments, the fan shaft anti-tilt stiffness and the moment of inertia product may be greater than or equal to 5.0×10 12 Nm 3 rad -1 kg(i.e., (Nm / rad).(kg.m 2 )), and optionally greater than or equal to 1.0×10 13 Nm 3 rad -1 kg, and optionally greater than or equal to 6.0×10 13 Nm 3 rad -1 kg.

[0604] 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-tilt stiffness multiplied by the moment of inertia may be greater than or equal to 6.0×10 13 Nm 3 rad -1 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft anti-tilt stiffness multiplied by the moment of inertia may be greater than or equal to 1.0×10 14 Nm 3 rad -1 kg.

[0605] In various embodiments, the fan shaft anti-tilt stiffness and the moment of inertia product can be 5.0×10 12 Nm 3 rad -1 kg to 6.0×10 16 Nm 3 rad -1kg, and optionally in the range of 1.0×10 13 Nm 3 rad -1 kg to 3.0×10 16 Nm 3 rad -1 kg, and optionally also in the range of 6.0×10 13 Nm 3 rad -1 kg to 3.0×10 16 Nm 3 rad -1 kg range.

[0606] 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-tilt stiffness multiplied by the moment of inertia may be in the range of 6.0×10 13 Nm 3 rad -1 kg to 3.0×10 15 Nm 3 rad -1 kg range.

[0607] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan shaft anti-tilt stiffness multiplied by the moment of inertia may be in the range of 1.0×10 14 Nm 3 rad -1 kg to 6.0×10 16 Nm 3 rad -1 kg range.

[0608] In various embodiments, the ratio of the gearbox support anti-tilt stiffness to the moment of inertia may be defined as:

[0609]

[0610] In various embodiments, the ratio of the gearbox support anti-tilt stiffness to the moment of inertia may be greater than or equal to 1.0×10 -3 Nrad -1 kg -1 m -1 (i.e., (Nm / rad) / (kg.m 2 )), and optionally greater than or equal to 2.0×10 -3 Nrad -1 kg -1 m -1 .

[0611] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the ratio of the gearbox support anti-tilt stiffness to moment of inertia may be greater than or equal to 2.0×10 -3 Nrad -1 kg -1 m -1 or greater than or equal to 5.0×10 -3 Nrad -1 kg -1 m -1 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the ratio of the gearbox support anti-tilt stiffness to moment of inertia may be greater than or equal to 1.0×10 - 3 Nrad -1 kg -1 m -1 or greater than or equal to 2.6×10 -3 Nrad -1 kg -1 m -1 .

[0612] In various embodiments, the gearbox support anti-tilt stiffness to moment of inertia ratio can be between 1.0×10 - 3 Nrad -1 kg -1 m -1 to 7.0×10 -2 Nrad -1 kg -1 m -1 in the range of, and optionally in the range of 2.0×10 -3 Nrad -1 kg -1 m -1 to 3.0×10 -2 Nrad -1 kg -1 m -1 in the range of 2.0×10 -3 Nrad -1 kg -1 m -1 to 7.0×10 -3 Nra d - 1 kg -1 m -1 within the range.

[0613] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the gearbox support anti-tilt stiffness to moment of inertia ratio may be in the range of 2.0×10 -3 Nrad -1 kg -1 m -1 to 7.2×10 -3 Nrad -1 kg -1 m -1 in the range of, and optionally in the range of 5.0×10 -3 Nrad -1 kg -1 m -1 to 7.0×10 -3 Nrad -1 kg -1 m -1 within the range (and can be equal to 6.0×10 -3 Nrad -1 kg -1 m -1 ).

[0614] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the gearbox support anti-tilt stiffness to moment of inertia ratio may be in the range of 1.0×10 -3 Nrad -1 kg -1 m -1 to 3.0×10 -2 Nrad -1 kg -1 m -1 in the range of, and optionally in the range of 2.6×10 -3 Nrad -1 kg -1 m -1 to 4.6×10 -3 Nrad -1 kg -1 m -1 within the range (and can be equal to 3.6×10 -3 Nrad -1 kg -1 m -1 ).

[0615] In various embodiments, in addition to or instead of the ratio of the gearbox support anti-tilt stiffness to the moment of inertia, a product of the parameters that make up the ratio can be defined. This product (referred to as the gearbox support anti-tilt stiffness and moment of inertia product) can be defined as:

[0616] The anti-tilting stiffness of the gearbox support (40)×the moment of inertia of the fan (23).

[0617] In various embodiments, the product of the anti-tilting stiffness and the moment of inertia of the gearbox support can be greater than or equal to 3.0×10 12 Nm 3 rad -1 kg(i.e., (Nm / rad).(kg.m 2 )), and optionally greater than or equal to 6.0×10 12 Nm 3 rad -1 kg, and optionally greater than or equal to 2.5×10 13 Nm 3 rad -1 kg.

[0618] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the product of the gearbox support anti-tilt stiffness and the moment of inertia may be greater than or equal to 2.0×10 13 Nm 3 rad -1 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the product of the anti-tilting stiffness and the moment of inertia of the gearbox support may be greater than or equal to 2.5×10 13 Nm 3 rad -1 kg.

[0619] In various embodiments, the product of the anti-tilting stiffness and the moment of inertia of the gearbox support can be 3.0×10 12 Nm 3 rad -1 kg to 9.0×10 15 Nm 3 rad -1 kg, and optionally in the range of 6.0×10 12 Nm 3 rad -1 kg to 4.0×10 15 Nm 3 rad -1 kg, and optionally also in the range of 2.5×10 13 Nm 3 rad -1 kg to 4.0×10 15 Nm 3 rad -1 kg range.

[0620] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the gearbox support anti-tilt stiffness and moment of inertia product may be in the range of 2.0×10 13 Nm 3 rad -1 kg to 4.0×10 15 Nm 3 rad - 1 kg range.

[0621] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the gearbox support anti-tilt stiffness and moment of inertia product may be in the range of 2.5×10 13 Nm 3 rad -1 kg to 9.0×10 15 Nm 3 rad - 1 kg range.

[0622] In various embodiments, the product of the fan-gearbox axial distance multiplied by the fan's moment of inertia may be defined.

[0623] In various embodiments, the product of the fan-gearbox axial distance and the fan moment of inertia may be greater than or equal to 1.9×10 7 kgm 3 , and optionally greater than or equal to 2.9×10 7 kgm 3 .

[0624] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the product of the fan-gearbox axial distance and the fan moment of inertia may be greater than or equal to 2.0×10 7 kgm 3 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the product of the fan-gearbox axial distance and the fan moment of inertia may be greater than or equal to 1.2×10 8 kgm 3 .

[0625] In various embodiments, the product of the fan-gearbox axial distance and the fan moment of inertia may be 1.9×10 7 kgm 3 to 6.2×10 8 kgm 3 in the range of, and optionally in the range of 2.9×10 7 kgm 3to 3.9×10 8 kgm 3 within the range.

[0626] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the product of the fan-gearbox axial distance and the fan moment of inertia may be 2.0×10 7 kgm 3 to 8.0×10 7 kgm 3 within the range.

[0627] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the product of the fan-gearbox axial distance and the fan moment of inertia may be 1.2×10 8 kgm 3 to 6.2×10 8 kgm 3 within the range.

[0628] In various embodiments, a ratio given by the fan-gearbox axial distance divided by the fan's moment of inertia may also be defined.

[0629] In various embodiments, the fan-gearbox axial distance divided by the fan moment of inertia may be less than or equal to 8.8×10 -9 m / kgm 2 (i.e., kg -1 m -1 ), and optionally less than or equal to 6.2×10 -9 m / kgm 2 .

[0630] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the fan-gearbox axial distance divided by the fan moment of inertia may be less than or equal to 6.5×10 -9 m / kgm 2 or less than or equal to 5.2×10 -9 m / kgm 2 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan-gearbox axial distance divided by the fan moment of inertia may be less than or equal to 2.8×10 -9 m / kgm 2 or less than or equal to 1.8×10 -9 m / kgm 2 .

[0631] In various embodiments, the fan-gearbox axial distance divided by the fan moment of inertia may be 5.3×10-10 m / kgm 2 to 8.8×10 -9 m / kgm 2 in the range of 8.8×10 -10 m / kgm 2 to 6.2×10 -9 m / kgm 2 within the range.

[0632] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the fan-gearbox axial distance divided by the fan moment of inertia may be in the range of 2.1×10 -9 m / kgm 2 to 6.5×10 -9 m / kgm 2 range, and optionally can be in the range of 4.0×10 -9 m / kgm 2 to 5.2×10 -9 m / kgm 2 within the range (and can be equal to 4.6×10 - 9 m / kgm 2 ).

[0633] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the fan-gearbox axial distance divided by the fan moment of inertia may be within the range of 5.3×10 -10 m / kgm 2 to 2.8×10 -9 m / kgm 2 in the range of 8.0×10 -10 m / kgm 2 to 1.8×10 -9 m / kgm 2 within the range (and can be equal to 1.3×10 - 9 m / kgm 2 ).

[0634] In various embodiments, the first gearbox support strength ratio may be defined as:

[0635]

[0636] In various embodiments, the first gearbox support member strength ratio may be greater than or equal to 7.0×10 -3 , and optionally greater than or equal to 1.0×10 -2 , or greater than or equal to 2.0×10 -2.

[0637] In various embodiments, the first gearbox support strength ratio may be 7.0×10 -3 to 2.5×10 -1 in the range of, and optionally in the range of 1.0×10 -2 to 1.0×10 -1 In the range of 7.0×10 -3 to 2.0×10 -2 In the range of 2.0×10 -2 to 2.5×10 -1 within the range.

[0638] In some embodiments, such as in embodiments where the fan diameter is in the range of 240 cm to 280 cm or in the range of 330 cm to 380 cm, the first gearbox support strength ratio may be in the range of 1.9×10 -2 to 2.9×10 -2 within the range (and can be equal to 2.4×10 -2 ).

[0639] In various embodiments, the second gearbox support strength ratio may be defined as:

[0640]

[0641] In various embodiments, the second gearbox support strength ratio may be greater than or equal to 1.0×10 -1 , and optionally greater than or equal to 1.5×10 -1 , greater than or equal to 1.0×10 -1 or greater than or equal to 2.5×10 -1 .

[0642] In various embodiments, the second gearbox support strength ratio may be 1.0×10 -1 to 3.5, and optionally in the range of 1.5×10 -1 to 1.7, and within 1.0×10 -1 to 2.5×10 -1 in the range of 2.5×10 -1 to 3.5.

[0643] In some embodiments, such as in embodiments where the fan diameter is in the range of 240 cm to 280 cm or in the range of 330 cm to 380 cm, the second gearbox support strength ratio may be 2.8×10 -1 to 3.8×10 -1 within the range (and can be equal to 3.3×10-1 ).

[0644] In various embodiments, the first gearbox support shear stress ratio may be defined as:

[0645]

[0646] In various embodiments, the first gearbox support shear stress ratio may be less than or equal to 4.9×10 1 m -1 , and optionally less than or equal to 20m -1 .

[0647] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the first gearbox support shear stress ratio may be less than or equal to 35 m -1 or less than or equal to 10.0m -1 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the first gearbox support shear stress ratio may be less than or equal to 12 m -1 or less than or equal to 6.0m -1 .

[0648] In various embodiments, the first gearbox support shear stress ratio may be 0.35 m -1 to 4.9×10 1 m -1 within the range of, and optionally within 0.70 m -1 Up to 20m -1 within the range.

[0649] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the first gearbox support shear stress ratio may be 0.70 m -1 Up to 35m -1 within the range, and optionally within 6.0 m -1 Up to 10.0m -1 Within the range (and can be equal to 7.8m -1 ).

[0650] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the first gearbox support shear stress ratio may be 0.50 cm to 1.50 cm. -1 Up to 12m -1 within the range, and optionally within 3.0m -1 Up to 6.0m -1 within the range (and can be equal to 5.4m -1).

[0651] In various embodiments, the second gearbox support shear stress ratio may be defined as:

[0652]

[0653] In various embodiments, the second gearbox support shear stress ratio may be less than or equal to 4.1×10 3 rad / m 3 , and optionally less than or equal to 1.4×10 3 rad / m 3 .

[0654] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the second gearbox support shear stress ratio may be less than or equal to 2.9×10 3 rad / m 3 or less than or equal to 6.5×10 2 rad / m 3 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the second gearbox support shear stress ratio may be less than or equal to 7.0×10 2 rad / m 3 or less than or equal to 2.5×10 2 rad / m 3 .

[0655] In various embodiments, the second gearbox support shear stress ratio may be 6.6 rad / m 3 to 4.1×10 3 rad / m 3 in the range of, and optionally in the range of 1.25×10 1 rad / m 3 to 1.4×10 3 rad / m 3 within the range.

[0656] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the second gearbox support shear stress ratio may be 2.9×10 1 rad / m 3 to 2.9×10 3 rad / m 3 in the range of 2.5×10 2 rad / m 3 to 6.5×10 2 rad / m 3within the range (and can be equal to 4.5×10 2 rad / m 3 ).

[0657] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the second gearbox support shear stress ratio may be in the range of 1.0×10 1 rad / m 3 to 7.0×10 2 rad / m 3 in the range of 50 rad / m 3 to 2.5×10 2 rad / m 3 within the range (and can be equal to 1.5×10 2 rad / m 3 ).

[0658] In various embodiments, the flight cycle ratio may be defined as:

[0659]

[0660] In various embodiments, the flight cycle ratio may be less than or equal to 3.20, and optionally less than or equal to 2.95, optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85, or optionally less than or equal to 2.75.

[0661] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the flight cycle ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the flight cycle ratio may be less than or equal to 2.8 or less than or equal to 2.7.

[0662] In various embodiments, the flight cycle ratio may be in the range of 2.10 to 3.20, and optionally in the range of 2.3 to 2.9.

[0663] In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the flight cycle ratio may be in the range of 2.3 to 3.2, and optionally in the range of 2.8 to 3.0 (and may be equal to 2.9).

[0664] In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the flight cycle ratio may be in the range of 2.1 to 2.8, and optionally in the range of 2.3 to 2.7 (and may be equal to 2.6).

[0665] In various embodiments, in addition to or instead of the flight cycle ratio, a product of the parameters that make up the ratio may be defined. This product (referred to as the flight cycle product) may be defined as:

[0666] The torsional shear stress of the gearbox support (40) under maximum takeoff conditions×the torsional shear stress of the gearbox support (40) under cruise conditions.

[0667] In various embodiments, the flight cycle product may be greater than or equal to 1.00×10 16 (N / m 2 ) 2 , and optionally greater than or equal to 2.05×10 16 (N / m 2 ) 2 .

[0668] In some embodiments, such as those where the fan diameter is in the range of 240 cm to 280 cm, the flight cycle product may be greater than or equal to 2.0×10 16 (N / m 2 ) 2 In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the flight cycle product may be greater than or equal to 2.2×10 16 (N / m 2 ) 2 .

[0669] In various embodiments, the flight cycle product may be 1.00×10 16 (N / m 2 ) 2 to 7.50×10 16 (N / m 2 ) 2 in the range of, and optionally in the range of 2.05×10 16 (N / m 2 ) 2 to 4.9×10 16 (N / m 2 ) 2 within the range.

[0670] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the flight cycle product may be 2.0×10 16 (N / m 2 ) 2 to 4.9×10 16 (N / m 2 ) 2 within the range.

[0671] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the flight cycle product may be 2.2×10 16 (N / m 2 ) 2 to 6.2×10 16 (N / m 2 ) 2 within the range.

[0672] In various implementations, the first torque transfer ratio may be defined as:

[0673]

[0674] In various implementations, the first torque transfer ratio may be less than or equal to 3.2, and optionally less than or equal to 2.95, optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85, or optionally less than or equal to 2.75.

[0675] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the first torque transmission ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the first torque transmission ratio may be less than or equal to 2.8 or less than or equal to 2.7.

[0676] In various implementations, the first torque transfer ratio may be in the range of 2.1 to 3.2, and optionally in the range of 2.3 to 2.9.

[0677] In some embodiments, such as embodiments where the fan diameter is in the range of 240 cm to 280 cm, the first torque transfer ratio may be in the range of 2.3 to 3.2, and optionally in the range of 2.8 to 3.0 (and may be equal to 2.9).

[0678] In some embodiments, such as embodiments where the fan diameter is in the range of 330 cm to 380 cm, the first torque transfer ratio may be in the range of 2.1 to 2.8, and optionally in the range of 2.5 to 2.7 (and may be equal to 2.6).

[0679] In various embodiments, in addition to or in lieu of the first torque transfer ratio, a product of the parameters constituting the ratio may be defined. This product (referred to as the first torque transfer product) may be defined as:

[0680] The torque transmitted through the gearbox (30) at maximum takeoff conditions is multiplied by the torque transmitted through the gearbox (30) at cruise conditions.

[0681] In various implementations, the first torque transfer product may be greater than or equal to 2.1×10 9 (Nm) 2 , and optionally greater than or equal to 3.5×10 9 (Nm) 2 .

[0682] In some embodiments, such as embodiments where the fan diameter is in the range of 240 cm to 280 cm, the first torque transfer product may be greater than or equal to 4.0×10 9 (Nm) 2 In some embodiments, such as those where the fan diameter is in the range of 330 cm to 380 cm, the first torque transfer product may be greater than or equal to 9.0×10 9 (Nm) 2 .

[0683] In various embodiments, the first torque transfer product may be 2.1×10 9 (Nm) 2 to 9.0×10 10 (Nm) 2 in the range of, and optionally in the range of 3.5×10 9 (Nm) 2 to 5.2×10 10 (Nm) 2 within the range.

[0684] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the first torque transfer product may be in the range of 4.0×10 9 (Nm) 2 to 9.0×10 9 (Nm) 2 within the range.

[0685] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the first torque transfer product may be between 9.0×10 9 (Nm) 2 to 9.0×10 10 (Nm) 2 within the range.

[0686] In various implementations, the second torque transfer ratio may be defined as:

[0687]

[0688] In various implementations, the second torque transfer ratio may be less than or equal to 3.2, and optionally less than or equal to 2.95, optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85, or optionally less than or equal to 2.75.

[0689] In some embodiments, such as embodiments in which the fan diameter is in the range of 240 cm to 280 cm, the second torque transmission ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, such as embodiments in which the fan diameter is in the range of 330 cm to 380 cm, the second torque transmission ratio may be greater than or equal to 2.8 or less than or equal to 2.7.

[0690] In various implementations, the second torque transfer ratio may be in the range of 2.1 to 3.2, and optionally in the range of 2.3 to 2.9.

[0691] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the second torque transfer ratio may be in the range of 2.3 to 3.2, and optionally in the range of 2.8 to 3.0 (and may be equal to 2.9).

[0692] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the second torque transfer ratio may be in the range of 2.1 to 2.8, and optionally in the range of 2.3 to 2.7 (and may be equal to 2.6).

[0693] In various embodiments, in addition to or in lieu of the second torque transfer ratio, a product of the parameters constituting the ratio may be defined. This product (referred to as the second torque transfer product) may be defined as:

[0694] The torque transmitted through the gearbox support (40) at maximum takeoff conditions × the torque transmitted through the gearbox support (40) at cruise conditions.

[0695] In various implementations, the second torque transfer product may be greater than or equal to 4.1×10 9 (Nm) 2 , and optionally greater than or equal to 6.1×10 9 (Nm) 2 .

[0696] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the second torque transfer product may be greater than or equal to 4.1×10 9 (Nm)2 In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the second torque transfer product may be greater than or equal to 1.1×10 10 (Nm) 2 .

[0697] In various embodiments, the second torque transfer product may be 4.1×10 9 (Nm) 2 to 9.0×10 10 (Nm) 2 in the range of, and optionally in the range of 6.1×10 9 (Nm) 2 to 8.0×10 10 (Nm) 2 within the range.

[0698] In some embodiments, such as those in which the fan diameter is in the range of 240 cm to 280 cm, the second torque transfer product may be in the range of 4.1×10 9 (Nm) 2 to 3.9×10 10 (Nm) 2 within the range.

[0699] In some embodiments, such as those in which the fan diameter is in the range of 330 cm to 380 cm, the second torque transfer product may be in the range of 1.1×10 10 (Nm) 2 to 9.0×10 10 (Nm) 2 within the range.

[0700] In the ratios defined above, maximum takeoff conditions and cruise conditions are as defined anywhere herein.

[0701] Figure 19 An exemplary aircraft 1000 is shown having a gas turbine engine 10 attached to each of its wings 1002a, 1002b. Each gas turbine engine 10 is attached via a respective pylon 1004a, 1004b. The gas turbine 10 may be a gas turbine of any embodiment described herein. Figure 19 The aircraft shown should be understood as the aircraft to which the gas turbine engine 10 of any embodiment or aspect disclosed herein has been designed to be attached. The aircraft 1000 has a cruise condition corresponding to the cruise condition defined elsewhere herein and an MTO condition corresponding to the maximum takeoff condition defined elsewhere herein.

[0702] The present disclosure also relates to the use of a vehicle (e.g., Figure 19Method 2000 for operating a gas turbine engine on an aircraft). Figure 20 2000 is shown in FIG. Method 2000 includes operating 2010 a gas turbine engine 10 as described elsewhere herein to provide propulsion to 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 parameters or ratios defined herein are within specified ranges. Cruise conditions and maximum takeoff conditions are as defined elsewhere herein.

[0703] The torque on the spindle 26 may be referred to as input torque because it is the torque input to the gearbox 30. The torque provided to the spindle by the turbine 19 under cruise conditions (i.e., the torque on the spindle) may 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 with a fan diameter in the range of 240 cm to 280 cm, the torque on the spindle 26 under cruise conditions may be greater than or equal to 10,000 Nm or 11,000 Nm (and optionally may be equal to 12,760 Nm). In some embodiments, such as embodiments with a fan diameter in the range of 330 cm to 380 cm, the torque on the spindle 26 under cruise conditions may be greater than or equal to 25,000 Nm, and optionally greater than or equal to 30,000 Nm (and optionally may be equal to 34,000 Nm).

[0704] The torque on the spindle 26 under cruise conditions may 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 with a fan diameter in the range of 240 cm to 280 cm, the torque on the spindle 26 under cruise conditions may 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 may be equal to 12,760 Nm). In some embodiments, such as embodiments with a fan diameter in the range of 330 cm to 380 cm, the torque on the spindle 26 under cruise conditions may 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 may be equal to 34,000 Nm).

[0705] At maximum takeoff (MTO) conditions, the torque on the spindle 26 may be greater than or equal to 28,000 Nm, and optionally greater than or equal to 30,000 Nm. In some embodiments, such as those with fan diameters in the range of 240 cm to 280 cm, the torque on the spindle 26 at MTO conditions may be greater than or equal to 28,000 Nm, and optionally greater than or equal to 35,000 Nm (and optionally may be equal to 36,300 Nm). In some embodiments, such as those with fan diameters in the range of 330 cm to 380 cm, the torque on the spindle 26 at MTO conditions may 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 may be equal to 87,000 Nm).

[0706] At maximum takeoff (MTO) conditions, the torque on the spindle 26 may 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, such as embodiments with fan diameters in the range of 240 cm to 280 cm, the torque on the spindle 26 at MTO conditions may 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 may be equal to 36,300 Nm). In some embodiments, such as in embodiments where the fan diameter is in the range of 330 cm to 380 cm, the torque on the spindle 26 under MTO conditions may 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 may be equal to 87,000 Nm).

[0707] Torque has units of [force] x [distance], may be expressed in Newton-meters (Nm), and is defined in a conventional manner as understood by those skilled in the art.

[0708] Figure 21 It is shown how stiffness as defined herein can be measured. Figure 21 A graph showing the displacement δ due to an applied load L (e.g., force, moment, or torque) applied to a component whose stiffness is being measured. At load levels from zero to LR, there is a nonlinear region where the displacement is due to the motion of the component when loaded (or relative motion of individual parts of the component) rather than deformation of the component (e.g., moving within a gap between parts). Above L SAt a load level of 1000, the elastic limit of the component has been exceeded and the applied load no longer causes elastic deformation, but plastic deformation or component failure may occur. Between points R and S, the applied load and the resulting displacement have a linear relationship. The stiffness defined in this article can be determined by measuring the gradient of the linear region between points R and S (where the stiffness is the inverse of the gradient). The gradient of the largest possible area of ​​the linear region can be found by providing a larger displacement for measurement to increase the accuracy of the measurement. For example, by applying a force equal to or slightly greater than L R and equal to or slightly less than L S Although the displacement is referred to as δ in this specification, those skilled in the art will understand that the equivalent principle will apply to linear or angular displacements.

[0709] Unless otherwise stated, the stiffness defined herein refers to the corresponding component with the engine turned off (i.e., at zero speed / on the bench). This stiffness generally does not vary within the operating range of the engine; therefore, the stiffness at cruise conditions of an aircraft using the engine (these cruise conditions are defined elsewhere herein) may be the same as the stiffness when the engine is not in use. However, in cases where the stiffness varies within the operating range of the engine, the stiffness defined herein should be understood as the value when the engine is at room temperature and not in motion. Unless otherwise stated, the values ​​of component strength given herein (e.g., the torsional strength of a gearbox support) are also given at room temperature.

[0710] 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 applies equally to a sun gearbox.

[0711] It should be understood that the present invention is not limited to the above-described embodiments and that various modifications and improvements may be made without departing from the concepts described herein. Unless mutually exclusive, any feature may be used alone or in combination with any other feature, and the present disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

1. A gas turbine engine (10) for an aircraft, comprising: an engine core (11), the engine core comprising a turbine (19), a compressor (14), and a spindle (26) connecting the turbine to the compressor; a fan (23), the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox (30) receiving input from the spindle (26) and outputting drive to the fan so as to drive the fan at a slower rotational speed than the spindle, the gearbox (30) being an epicyclic gearbox (30) comprising a sun gear (28), a plurality of planet gears (32), a ring gear (38), and a planet carrier (34) arranged for mounting the planet gears (32) thereon; and A gearbox support (40) arranged to at least partially support the gearbox within the engine, and wherein: Maximum takeoff conditions are defined as those when operating the engine at maximum takeoff thrust for said engine at ISA sea level pressure and temperature +15°C with a fan inlet speed between Mach 0.25 and 0.27; and Cruise conditions are: (i) conditions experienced by an aircraft mid-cruise to which the engine (10) is attached, or (ii) conditions experienced by the aircraft and the engine (10) at a point midway between the top of a climb and the start of a descent; Flight cycle ratio: Less than or equal to 3.

20.

2. The gas turbine engine (10) of claim 1, wherein the flight cycle ratio is less than or equal to 2.95, less than or equal to 2.9, less than or equal to 2.90, less than or equal to 2.85, less than or equal to 2.75, in a range of 2.10 to 3.20, or in a range of 2.3 to 2.

9.

3. The gas turbine engine (10) according to claim 1 or claim 2, wherein: Under the maximum takeoff condition, the torsional shear stress of the gearbox support (40) is less than or equal to 4.90×10 8 N / m 2 , less than or equal to 2.0×10 8 N / m 2 , at 1.40×10 8 N / m 2 to 4.9×10 8 N / m 2 In the range of 2.0×10 8 N / m 2 to 3.5×10 8 N / m 2 within the scope of; and / or Under the most cruising condition, the torsional shear stress of the gearbox support (40) is greater than or equal to 7.00×10 7 N / m 2 , greater than or equal to 8.2×10 7 N / m 2 , at 7.00×10 7 N / m 2 to 1.90×10 8 N / m 2 In the range of 8.2×10 7 N / m 2 to 1.5×10 8 N / m 2 within the range.

4. The gas turbine engine (10) according to claim 1 or claim 2, wherein: a) the fan diameter is in the range of 240 cm to 280 cm, and the flight cycle ratio is less than or equal to 3.2 or in the range of 2.3 to 3.2; or b) The fan diameter is in the range of 330 cm to 380 cm, and the flight cycle ratio is less than or equal to 2.8 or in the range of 2.1 to 2.

8.

5. The gas turbine engine (10) of claim 1 or claim 2, wherein the product: Torsional shear stress of the gearbox support (40) under maximum takeoff conditions × torsional shear stress of the gearbox support (40) under cruise conditions Greater than or equal to 1.00×10 16 (N / m 2 ) 2 , greater than or equal to 2.05×10 16 (N / m 2 ) 2 , at 1.00×10 16 (N / m 2 ) 2 to 7.50×10 16 (N / m 2 ) 2 In the range of 2.05×10 16 (N / m 2 ) 2 to 4.9×10 16 (N / m 2 ) 2 within the range.

6. The gas turbine engine (10) of claim 5, wherein the gearbox is in a star configuration.

7. The gas turbine engine (10) of claim 1 or claim 2, wherein the first torque transfer ratio: Less than or equal to 3.2, less than or equal to 2.95, less than or equal to 2.9, less than or equal to 2.90, less than or equal to 2.85, less than or equal to 2.75, within a range of 2.1 to 3.2, or within a range of 2.3 to 2.

9.

8. The gas turbine engine (10) of claim 1 or claim 2, wherein the second torque transfer ratio: Less than or equal to 3.2, less than or equal to 2.95, less than or equal to 2.9, less than or equal to 2.90, less than or equal to 2.85, less than or equal to 2.75, within a range of 2.1 to 3.2, or within a range of 2.3 to 2.

9.

9. The gas turbine engine (10) according to claim 1 or claim 2, wherein: a) The torsional strength of the gearbox support (40) is greater than or equal to 1.60×10 5 Nm, greater than or equal to 1.8×10 5 Nm, at 1.60×10 5 Nm to 2.00×10 7 Nm, or within the range of 1.8×10 5 Nm to 1.5×10 6 Nm; and / or b) The cross-sectional area of ​​the gearbox is greater than or equal to 2.4×10 -1 m 2 , greater than or equal to 2.6×10 -1 m 2 , at 2.4×10 -1 m 2 Up to 1.10m 2 In the range of 2.6×10 -1 m 2 to 9.0×10 -1 m 2 within the scope of; and / or c) The planet gear pitch angle (β) in radians is defined as 2π / N, where N is the number of planet gears (32), wherein the planet gear pitch angle (β) is greater than or equal to 9.0×10 -1 rad, or at 9.0×10 -1 rad to 2.1rad; and / or d) The radial bending stiffness of the gearbox support (40) is greater than or equal to 1.0×10 7 N / m, greater than or equal to 2.0×10 7 N / m, greater than or equal to 3.0×10 7 N / m, at 1.0×10 7 N / m to 4.0×10 8 N / m range, within 2.0×10 7 N / m to 3×10 8 N / m, or within the range of 3.0×10 7 N / m to 2.0×10 8 in the range of N / m; and / or e) The anti-tilting stiffness of the gearbox support (40) is greater than or equal to 1.2×10 5 Nm / rad, greater than or equal to 2.4×10 5 Nm / rad, greater than or equal to 3.9×10 5 Nm / rad, at 1.2×10 5 Nm / rad to 2.1×10 7 Nm / rad range, within 2.4×10 5 Nm / rad to 1.6×10 7 Nm / rad, or in the range of 3.9×10 5 Nm / rad to 9.0×10 6 In the range of Nm / rad.

10. The gas turbine engine (10) according to claim 1 or claim 2, wherein: The gearbox is in a star configuration; and / or The gear ratio of the wheel box is 3.2 to 4.5, or 3.2 to 4.0; and / or The gas turbine engine has a 70NKg -1 Up to 90NKg -1 specific thrust; and / or The gas turbine engine has a bypass ratio of 12.5 to 18, or 13 to 16, under cruise conditions; and / or The fan (23) has a fan diameter greater than 240 cm and less than or equal to 380 cm or a fan diameter greater than 300 cm and less than or equal to 380 cm.

11. The gas turbine engine (10) according to claim 1 or claim 2, wherein: The turbine is a first turbine (19), the compressor is a first compressor (14), and the spindle is a first spindle (26); The engine core (11) further includes a second turbine (17), a second compressor (15), and a second spindle (27) connecting the second turbine to the second compressor; and The second turbine, the second compressor and the second spindle are arranged to rotate at a higher rotational speed than the first spindle.

12. The gas turbine engine (10) according to claim 1 or claim 2, wherein: The maximum takeoff condition is defined as the condition when operating the engine at its maximum takeoff thrust at ISA sea level pressure and temperature +15°C with a fan inlet speed of Mach 0.

25.

13. The gas turbine engine (10) of claim 1 or claim 2, wherein any one or more of the following are true: a) the forward speed of the gas turbine engine in the cruise condition is in the range of Mach 0.75 to Mach 0.85, or the forward speed of the gas turbine engine in the cruise condition is Mach 0.8; b) the cruise conditions correspond to the atmospheric conditions defined by the International Standard Atmosphere at an altitude of 11,582 m and a forward speed of Mach 0.8; c) the cruising conditions correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 10,668 m and a forward speed of Mach 0.85; and / or d) The cruising conditions correspond to atmospheric conditions at an altitude in the range of 10,500 m to 11,600 m or atmospheric conditions at an altitude of 11,000 m.

14. An aircraft comprising a gas turbine engine according to claim 1 or claim 2 mounted thereon, wherein the aircraft has a maximum takeoff operating condition and a cruise condition.

15. A propeller for an aircraft, comprising: a fan, the fan comprising a plurality of fan blades; Gearbox; a power unit for driving the fan (23) via the gearbox (30), wherein the gearbox (30) is arranged to receive input from the power unit via a spindle (26) and output drive to a fan shaft (36) so as to drive the fan at a lower rotational speed than the spindle (26), the gearbox (30) being an epicyclic gearbox (30) comprising a sun gear (28), a plurality of planetary gears (32), a ring gear (38) and a planet carrier (34) arranged for mounting the plurality of planetary gears (32) thereon; and a gearbox support (40) arranged to at least partially support the gearbox (30) within the propeller, wherein: Maximum takeoff conditions are defined as those when operating the engine at maximum takeoff thrust for said engine at ISA sea level pressure and temperature +15°C with a fan inlet speed between Mach 0.25 and 0.27; and Cruise conditions are: (i) conditions experienced by an aircraft mid-cruise to which the engine (10) is attached, or (ii) conditions experienced by the aircraft and the engine (10) at a point midway between the top of a climb and the start of a descent; Flight cycle ratio: Less than or equal to 3.

20.

16. A method (2000) of operating a gas turbine engine, the gas turbine engine comprising: an engine core (11), the engine core comprising a turbine (19), a compressor (14), and a spindle (26) connecting the turbine to the compressor; a fan (23), the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox (30) receiving input from the spindle (26) and outputting drive to the fan so as to drive the fan at a slower rotational speed than the spindle, the gearbox (30) being an epicyclic gearbox (30) comprising a sun gear (28), a plurality of planet gears (32), a ring gear (38), and a planet carrier (34) arranged for mounting the planet gears (32) thereon; and a gearbox support (40) arranged to at least partially support a gearbox within the engine, and wherein Maximum takeoff conditions are defined as those when operating the engine at maximum takeoff thrust for said engine at ISA sea level pressure and temperature +15°C with a fan inlet speed between Mach 0.25 and 0.27; and Cruise conditions are: (i) conditions experienced by an aircraft mid-cruise to which the engine (10) is attached, or (ii) conditions experienced by the aircraft and the engine (10) at a point midway between the top of a climb and the start of a descent; The method comprises: Operate (2010, 2020) the gas turbine engine so that the flight cycle ratio: Less than or equal to 3.

20.

17. The method (2000) of claim 16, wherein: The method further includes driving the gearbox with the following input torque: i) greater than or equal to 10,000 Nm under cruising conditions; and / or ii) greater than or equal to 28 000 Nm under maximum take-off conditions.

18. The method (2000) of claim 16, wherein: The method further includes driving the gearbox with the following input torque: i) 10,000 Nm to 50,000 Nm under cruising conditions; and / or ii) 28,000 Nm to 135,000 Nm at maximum take-off conditions.

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