CN216381626U

CN216381626U - Gas turbine engine and propeller for aircraft

Info

Publication number: CN216381626U
Application number: CN202022871955.8U
Authority: CN
Inventors: M·斯普鲁斯
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2019-12-05
Filing date: 2020-12-04
Publication date: 2022-04-26
Anticipated expiration: 2030-12-04
Also published as: GB201917783D0; DE102020131802A1; FR3104203A1

Abstract

A gas turbine engine and propeller for an aircraft is provided. The engine includes: 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 an input from a gearbox input shaft (26a) portion of the spindle (26) and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle; and a gearbox support (40) arranged to support the gearbox (30) in a fixed position within the engine (10). The planet carrier (34) and gearbox support (40) of the gearbox each have a torsional stiffness and the ratio of the carrier to gearbox support torsional stiffness is:

greater than or equal to 2.3.

Description

Gas turbine engine and propeller for aircraft

Technical Field

The present disclosure relates to gearboxes for aircraft engines, to aircraft engines comprising such gearboxes, and to methods of operating such aircraft. Such a gearbox may be an epicyclic gearbox comprising a planet carrier with a stiffness that meets specified criteria.

Background

As used herein, a range "value X to value Y" or "between value X and value Y" and the like means an inclusive range; including the boundary values of X and Y. As used herein, the term "axial plane" means a plane extending along the length of the engine, parallel to and encompassing the axial centerline of the engine, and the term "radial plane" means a plane extending perpendicular to the axial centerline of the engine, thus including all radial lines at the axial location of the radial plane. The 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 extending in a radial direction in a radial plane or in an axial direction in an axial plane, respectively.

SUMMERY OF THE UTILITY MODEL

According to a first aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet gear mounted thereonAnd an upper planet carrier. The radial bending stiffness of the planet carrier is equal to or greater than 1.20 x 10⁹N/m。

The radial bending stiffness of the planet carrier may be less than or equal to 1.00 x 10¹²N/m。

The radial bending stiffness of the planet carrier may be equal to or greater than 2.0 x 10⁹N/m, and/or optionally at 1.20X 10⁹N/m to 1.00X 10¹²N/m or 2.0X 10⁹N/m to 1.5X 10¹¹N/m.

The tilt stiffness of the planet carrier may be greater than or equal to 6.00 x 10⁸Nm/radian, and optionally may be in the range of 1.3X 10⁹Nm/radian to 1.2X 10¹¹Nm/radian.

The fan may have a fan diameter in the range of 240cm to 280 cm. In such embodiments, the radial bending stiffness of the planet carrier may be equal to or greater than 1.5 x 10⁹N/m, and optionally less than or equal to 5X 10¹⁰N/m。

Alternatively, the fan may have a fan diameter in the range of 330cm to 380 cm. In such embodiments, the radial bending stiffness of the planet carrier may be equal to or greater than 2.0 x 10⁹N/m, and optionally less than or equal to 1.6X 10¹¹N/m。

According to a second aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The tilt stiffness of the carrier is greater than or equal to 6.0 x 10⁸Nm/rad。

The tilt stiffness of the planet carrier may be less than or equal to 2.8 x 10¹¹Nm/rad。

Optionally, the planet carrier tilt stiffness may be greater than or equal to 1.3 × 10⁹Nm/radian, and optionally at 1.3X 10⁹Nm/radian to 1.2X 10¹¹Nm/radian.

The radial bending stiffness of the planet carrier may be equal to or greater than 1.20 x 10⁹N/m, and optionally at 1.20X 10⁹N/m to 1X 10¹²N/m or 2.0X 10⁹N/m to 1.5X 10¹¹N/m.

The fan may have a fan diameter in the range of 240cm to 280 cm. In such embodiments, the planet carrier's tilt stiffness may be greater than or equal to 2.2 × 10⁹Nm/radian, and optionally less than or equal to 1.4 x 10¹¹Nm/radian.

Alternatively, the fan may have a fan diameter in the range of 330cm to 380 cm. In such embodiments, the planet carrier's tilt stiffness may be greater than or equal to 2.3 × 10⁹Nm/radian, and optionally less than or equal to 2.8 x 10¹¹Nm/radian.

According to a third aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The torsional rigidity of the planet carrier is greater than or equal to 1.60 multiplied by 10⁸Nm/rad, and optionally less than or equal to 1.00X 10¹¹Nm/rad。

The inventors have found that maintaining one or more stiffnesses of the carrier within the ranges claimed herein may allow for compensation of gear misalignment (e.g., due to variations within manufacturing tolerances and/or wear during operation) while avoiding significant deformation of the gearbox. This compensation may result in reduced variability in the load on and/or between gears (e.g., more uniform load sharing). This, in turn, may allow for the reduction of the mass of the gears while still maintaining the life and efficiency required for aircraft applications.

Thus, one or more of the stiffnesses of the carrier may be selected to be relatively high to reduce or avoid unwanted rolling or deformation of the carrier and/or misalignment of gears carried by the carrier. It may be beneficial to keep one or more of the stiffnesses low enough to allow sufficient flexibility to correct any slight gear misalignment due to manufacturing issues. The inventors have found that in some arrangements, maintaining one or more of the stiffnesses within the corresponding specified ranges provides a beneficial combination of these effects.

It has been found that maintaining an even distribution of load between the planet gears is desirable to improve the life and reliability of the gearbox. The inventors have found that maintaining one or more of the carrier stiffnesses within the specified ranges applicable allows the carrier to be sufficiently flexible to facilitate more uniform load sharing (i.e., improved load sharing factor) by allowing the planet gears to move relative to each other and/or relative to the carrier. In some arrangements, if the frame stiffness is too high, the load sharing factor may be reduced by failing to accommodate any pre-existing misalignment or any misalignment that may occur during use.

Accordingly, the frame design of various embodiments having one or more defined stiffnesses may help to obtain and/or maintain proper gear alignment.

One or more of the following features may be applicable to any of the three aspects described above:

the torsional stiffness of the planet carrier may be greater than or equal to 1.60 x 10⁸Nm/rad, and optionally at 1.60X 10⁸Nm/rad to 1.00X 10¹¹Nm/rad, or in the range of 2.7X 10⁸Nm/rad to 1X 10¹⁰Nm/rad.

The pitch circle diameter of the pin on which the planet gear is mounted may be in the range 0.38m to 0.65m, and optionally may be equal to 0.4m or 0.55 m.

According to a fourth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and to output 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted, and wherein the radial bending stiffness of the planet carrier is equal to or greater than 1.20 x 10⁹N/m, and optionally less than or equal to 1.00X 10¹²N/m. The method includes operating a gas turbine engine to provide propulsion at cruise conditions.

According to a fifth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted, and wherein the planet carrier has a pitch stiffness greater than or equal to 6.0 x 10⁸Nm/rad. Additionally or alternatively, in another aspect, the torsional stiffness of the planet carrier may be greater than or equal to 1.60 × 10⁸Nm/rad, and optionally less than or equal to 1.00X 10¹¹Nm/rad. The method includes operating a gas turbine engine to provide propulsion at cruise conditions.

The method of the fourth or fifth aspect may comprise driving the gearbox with the following input torques:

(i) greater than or equal to 10000Nm, and optionally from 10000Nm to 50000Nm, at cruise; and/or

(ii) Greater than or equal to 28000Nm at MTO, and optionally from 28000Nm to 135000 Nm.

According to a sixth aspect, there is provided a propeller for an aircraft, the propeller: comprising a fan comprising a plurality of fan blades; a gear case; and a power unit for driving the fan via the gearbox. The gearbox is an epicyclic gearbox arranged to receive input from a spindle driven by a power unit and to output drive to the fan so as to drive the fan at a lower rotational speed than the spindle, and comprises a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The radial bending stiffness of the planet carrier is equal to or greater than 1.20 x 10⁹N/m. Optionally, the radial bending stiffness of the planet carrier may be less than or equal to 1.00 x 10¹²N/m。

According to a seventh aspect, there is provided a thruster for an aircraft, the thruster comprising: a fan comprising a plurality of fan blades; a gear case; and a power unit for driving the fan via the gearbox. The gearbox is an epicyclic gearbox arranged to receive an input from the spindle and to output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle, and comprises a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The tilt stiffness of the carrier is greater than or equal to 6.00 x 10⁸Nm/rad. The tilt stiffness of the planet carrier may be less than or equal to 2.8 x 10¹¹Nm/rad。

Additionally or alternatively, in another aspect, the torsional stiffness of the planet carrier may be greater than or equal to 1.60 × 10⁸Nm/rad, and optionally less than or equal to 1.00X 10¹¹Nm/rad。

The propeller of the sixth or seventh aspect may have any or all of the features as described for the gas turbine engine of the first, second and/or third aspect.

According to an eighth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The ratio of the radial bending stiffness to the torsional frame stiffness is:

greater than or equal to 0.030.

The ratio of radial bending stiffness to torsional frame stiffness may be less than or equal to 2.0 x 10⁰(i.e., 2.0).

The radial bending stiffness to torsion frame stiffness ratio may be in the range of 0.030 to 2.0, and optionally in the range of 0.060 to 1.0.

The radial bending stiffness of the planet carrier may be equal to or greater than 1.20 x 10⁹N/m, and optionally at 1.20X 10⁹N/m to 1.00X 10¹²N/m or 2.0X 10⁹N/m to 1.5X 10¹¹N/m.

The effective linear torsional stiffness of the planet carrier may be greater than or equal to 7.00 x 10⁹N/m, and optionally at 7.00X 10⁹N/m to 1.20X 10¹¹N/m or 9.1X 10⁹N/m to 8.0X 10¹⁰N/m.

The tilt stiffness of the planet carrier may be greater than or equal to 6.00 x 10⁸Nm/radian, andand optionally at 1.3X 10⁹Nm/radian to 1.2X 10¹¹Nm/radian.

The ratio of radial bending stiffness to torsional frame stiffness may be in the range of 0.060 to 0.30. Alternatively, the radial bending stiffness to torsional frame stiffness ratio may be in the range of 0.30 to 2.0.

The ratio of the tilt stiffness to the torsional frame stiffness is:

may be in the range of 0.7 to 20.

According to a ninth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The ratio of the tilt stiffness to the torsional frame stiffness is:

in the range of 0.7 to 20.

The ratio of the tilt stiffness to the torsional frame stiffness may be in the range of 0.7 to 7.3.

The tilt stiffness of the planet carrier may be greater than or equal to 6.00 x 10⁸Nm/radian, and optionally at 1.3X 10⁹Nm/radian to 1.2X 10¹¹Nm/radian.

The radial bending stiffness of the planet carrier may be equal to or greater than 1.20 x 10⁹N/m, and optionally at 1.20×10⁹N/m to 1X 10¹²N/m or 2.0X 10⁹N/m to 1.5X 10¹¹N/m.

The torsional stiffness of the planet carrier may be greater than or equal to 1.60 x 10⁸Nm/radian, and optionally at 1.60X 10⁸Nm/radian to 1.00X 10¹¹Nm/radian, or 2.7X 10⁸Nm/radian to 1 × 10¹⁰Nm/radian.

The fan may have a fan diameter in the range of 240cm to 280 cm. In such embodiments, the pitch to torsional frame stiffness ratio may be in the range of 2.5 to 8.0.

Alternatively, the fan may have a fan diameter in the range of 330cm to 380 cm. In such embodiments, the pitch to torsional frame stiffness ratio may be in the range of 1.5 to 7.9.

The ratio of the radial bending stiffness to the torsional frame stiffness is:

and may be in the range of 0.030 to 2.0.

The inventors have found that maintaining the ratio of the radial bending or tilting stiffness of the carrier to the torsional stiffness of the carrier within a specified range allows gear tooth damage to be better avoided (due to the relative increase in carrier torsional stiffness). It has been found that any further relative increase in torsional stiffness does not provide further benefit in terms of tooth protection, and instead there may be a risk of reducing overall performance by adding unnecessary size and/or weight to the shelf. Thus, the torsional stiffness of the carrier is arranged to be relatively high to reduce or avoid the risk of deformation of the gear teeth. In particular, the inventors have recognized that the rolling up of the rack when the gear teeth mesh may cause the gear teeth to chip, grind, or deform as the gear teeth are forced against the opposing gear teeth.

According to a tenth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The ratio of the radial bending stiffness to the torsional frame stiffness is:

in the range of 0.030 to 2.0. Additionally or alternatively, the ratio of the tilt stiffness to the torsional frame stiffness is:

in the range of 0.7 to 20.

The method includes operating a gas turbine engine to provide propulsion at cruise conditions.

The method may comprise driving the gearbox at an input torque of greater than or equal to 10000Nm, and optionally from 10000Nm to 50000Nm, whilst cruising.

The method may include driving the gearbox at an MTO with an input torque of greater than or equal to 28000Nm, and optionally 28000Nm to 135000 Nm.

The engine for use in the method of the tenth aspect may have any or all of the features of the engine of the eighth or ninth aspect.

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 gear case; and a power unit for driving the fan via the gearbox. The gearbox is an epicyclic gearbox arranged to receive an input from the spindle and to output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle, and comprises a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The ratio of the radial bending stiffness to the torsional frame stiffness is:

in the range of 0.030 to 2.0.

According to a twelfth aspect, there is provided a thruster for an aircraft, the thruster being: comprising a fan comprising a plurality of fan blades; a gear case; and a power unit for driving the fan via the gearbox. The gearbox is an epicyclic gearbox arranged to receive an input from the spindle and to output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle, and comprises a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The ratio of the tilt stiffness to the torsional frame stiffness is:

in the range of 0.7 to 20.

The propeller of the eleventh or twelfth aspect may have any or all of the features of the engine of the eighth or ninth aspect.

According to a thirteenth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox receiving an input from the gearbox input shaft portion of the spindle and outputting drive to the fan 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The torsional rigidity ratio of the frame to the input shaft of the gear box is as follows:

greater than or equal to 70, and optionally less than or equal to 5.0 x 10³。

The inventors have found that the torsional stiffness of the gearbox system (including in particular the gearbox input shaft and the carrier) should be distributed in the required proportions, since as mentioned above the relatively high torsional stiffness of the carrier has been found to reduce or avoid the risk of gear tooth deformation, but the relatively increased torsional stiffness of the spindle (the input shaft of the gearbox) does not improve this effect, but detrimentally increases the size and/or weight of the shaft without a corresponding benefit.

While torsional compliance of the gearbox input shaft has less of an effect on gearbox performance than carrier compliance, those skilled in the art will appreciate that too low a torsional stiffness of the gearbox input shaft may result in wind-up of the sun gear, which may result in gear misalignment. A relative increase in the stiffness of the gearbox input shaft beyond the required ratio range may not provide further benefit and may instead unnecessarily increase the size and/or weight of the fan shaft.

The carrier to gearbox input shaft torsional stiffness ratio can be equal to or greater than 75, and optionally at 7.5 x 10¹To 3X 10³Within the range of (1).

The torsional stiffness of the gearbox input shaft may be equal to or greater thanAt 1.4X 10⁶Nm/radian, and optionally at 1.4X 10⁶Nm/radian to 2.5X 10⁸Nm/radian.

The fan may have a fan diameter in the range of 240cm to 280 cm. In such embodiments, the carrier to gearbox input shaft torsional stiffness ratio can be greater than or equal to 7.3 x 10¹And optionally less than or equal to 1.0X 10³。

The fan may have a fan diameter in the range of 330cm to 380 cm. In such embodiments, the carrier to gearbox input shaft torsional stiffness ratio can be greater than or equal to 1.0 x 10²And optionally less than or equal to 5.0 x 10³。

The gas turbine engine may also include a gearbox support arranged to support the gearbox in a fixed position within the engine and having torsional stiffness. The frame to gearbox support torsional stiffness ratio is:

can be greater than or equal to 2.3, and optionally greater than or equal to 2.6.

The gas turbine engine may also include a fan shaft connecting the output of the gearbox to the fan. The ratio of the stiffness of the frame to the stiffness of the fan shaft is:

can be greater than or equal to 8, and optionally greater than or equal to 9.

The product of the torsional stiffness of the planet carrier and the torsional stiffness of the input shaft of the gearbox may be greater than or equal to 1.5 x 10¹⁴N²m²rad^-2And optionally greater than or equal to 2.2X 10¹⁴N²m²rad^-2。

The fan may have a fan diameter in the range of 240cm to 280cm, and the torsional stiffness of the planet carrier and the gearboxThe product of the torsional stiffness of the input shaft may be greater than or equal to 1.5 x 10¹⁴N²m²rad^-2。

The fan may have a fan diameter in the range of 330cm to 380cm, and the product of the torsional stiffness of the planet carrier and the torsional stiffness of the gearbox input shaft may be greater than or equal to 3.0 x 10¹⁵N²m²rad^-2。

The gearbox may be a sun gearbox in which the planet carrier does not rotate in use.

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 including 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; and a gearbox arranged to receive an input from the gearbox input shaft portion of the spindle and to output 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted.

The torsional rigidity ratio of the frame to the input shaft of the gear box is as follows:

greater than or equal to 70. The method includes operating a gas turbine engine to provide propulsion at cruise conditions. The carrier to gearbox input shaft torsional stiffness ratio may be less than or equal to 5000.

According to a fifteenth aspect, there is provided a thruster for an aircraft, the thruster comprising: a fan comprising a plurality of fan blades; a gear case; and a power unit for driving the fan via the gearbox. The gearbox is an epicyclic gearbox arranged to receive input from a gearbox input shaft portion of a spindle driven by the power unit and to output drive to the fan so as to drive the fan at a lower rotational speed than the spindle, and comprises a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The torsional rigidity ratio of the frame to the input shaft of the gear box is as follows:

greater than or equal to 70. The carrier to gearbox input shaft torsional stiffness ratio may be less than or equal to 5000.

The impeller may have some or all of the features described above with respect to the gas turbine engine, and may be a gas turbine engine in some embodiments.

In various other aspects of the invention, the specified boundary on the carrier to gearbox input shaft torsional stiffness ratio may be replaced or otherwise provided by a specified boundary on the product of the components of the carrier to gearbox input shaft torsional stiffness ratio (i.e. the boundary on the torsional stiffness of the planet carrier times the torsional stiffness of the gearbox input shaft). In various aspects, the value of the product may be greater than or equal to 1.5 × 10¹⁴N²m²rad^-2And optionally less than 1.0 x 10¹⁷N²m²rad^-2. Optionally, the value may be greater than or equal to 2.2 × 10¹⁴N²m²rad^-2And optionally less than 5.0 x 10¹⁶N²m²rad^-2。

For example, according to another aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox receiving an input from the gearbox input shaft portion of the spindle and outputting drive to the fan 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The torsional stiffness of the planet carrier multiplied by the torsional stiffness of the input shaft of the gearbox is greater than or equal to 1.5 x 10¹⁴N²m²rad^-2。

Those skilled in the art will appreciate that the method and propeller aspects may be formulated accordingly. Optional features of aspects of corresponding ratios may also be applied to these aspects.

According to a sixteenth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting a drive to the fan 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted; and a gearbox support arranged to support the gearbox in a fixed position within the engine and having torsional stiffness. The frame to gearbox support torsional stiffness ratio is:

greater than or equal to 2.3. The carrier to gearbox support torsional stiffness ratio may be less than or equal to 300.

The inventors have found that the torsional stiffness of the gearbox system (including in particular the gearbox support and the planet carrier) should be distributed in the required proportions to provide an improvement as described above in relation to the thirteenth aspect. In particular, the inventors have found that maintaining this ratio within a specified range allows reducing the risk of gear tooth damage, while still maintaining sufficient stiffness to avoid very large amplitudes of torsional vibration modes of the gearbox support.

The planet carrier to gearbox support torsional stiffness ratio may be greater than or equal to 2.6, and optionally in the range of 2.6 to 50.

The torsional stiffness of the input shaft portion of the gearbox of the spindle may be greater than or equal to 1.4 x 10⁶Nm/radian, and optionally greater than or equal to 1.6 x 10⁶Nm/radian.

The fan may have a fan diameter in the range of 240cm to 280cm, and the planet carrier to gearbox support torsional stiffness ratio may be greater than or equal to 2.3.

The fan may have a fan diameter in the range of 330cm to 380cm, and the planet carrier to gearbox support torsional stiffness ratio may be greater than or equal to 3.5.

The torsional stiffness of the planet carrier multiplied by the torsional stiffness of the gearbox may be greater than or equal to 5 x 10¹⁵N²m²rad^-2And optionally less than 1.0 x 10¹⁹N²m²rad^-2。

The spindle may comprise a gearbox input shaft portion arranged to provide an input to the gearbox. The torsional rigidity ratio of the frame to the input shaft of the gear box is as follows:

and may be greater than or equal to 70.

can be greater than or equal to 8, and optionally greater than or equal to 9, and can be less than or equal to 1100.

According to a seventeenth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox receiving an input from the spindle and outputting a drive to the fan to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted; and a gearbox support arranged to support the gearbox in a fixed position within the engine and having torsional stiffness. The frame to gearbox support torsional stiffness ratio is:

greater than or equal to 2.3. The method includes operating a gas turbine engine to provide propulsion at cruise conditions. The carrier to gearbox support torsional stiffness ratio may be less than or equal to 300.

The method may further comprise driving the gearbox at an input torque greater than or equal to 10000Nm, and optionally from 10000Nm to 50000Nm, while cruising.

The method may further include driving the gearbox at an MTO at an input torque of greater than or equal to 28000Nm, and optionally 28000Nm to 135000 Nm.

According to an eighteenth aspect, there is provided a thruster for an aircraft, the thruster comprising: a fan comprising a plurality of fan blades; a gear case; and a power unit for driving the fan via the gearbox. The gearbox is arranged to receive input from a spindle driven by the power unit and to output 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The thruster further comprises a gearbox support arranged to support the gearbox in a fixed position within the thruster and having a torsional stiffness. The frame to gearbox support torsional stiffness ratio is:

greater than or equal to 2.3, and optionally less than or equal to 300.

In various other aspects of the invention, the specified boundary on the carrier to gearbox support torsional stiffness ratio may be replaced or otherwise provided by a specified boundary on the product of the components of the carrier to gearbox support torsional stiffness ratio (i.e. the boundary on the torsional stiffness of the planet carrier times the torsional stiffness of the gearbox support). In various aspects, the value of the product may be greater than or equal to 5 × 10¹⁵N²m²rad^-2And optionally less than 1.0 x 10¹⁹N²m²rad^-2. Optionally, the value may be greater than or equal to 8.0 × 10¹⁵N²m²rad^-2And optionally less than 2.0 x 10¹⁸N²m²rad^-2。

For example, according to another aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting a drive to the fan 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted; and a gearbox support arranged to support the gearbox in a fixed position within the engine and having torsional stiffness. The torsional stiffness of the planet carrier multiplied by the torsional stiffness of the gearbox support is greater than or equal to 5 x 10¹⁵N²m²rad^-2。

According to a nineteenth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox receiving an input from the spindle and outputting a drive to the fan 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted; and a fan shaft connecting the output of the gearbox to the fan. The ratio of the stiffness of the frame to the stiffness of the fan shaft is:

greater than or equal to 8. The ratio of case to fan shaft stiffness may be less than or equal to 1100.

The inventors have found that the torsional stiffness of the gearbox system, in particular comprising the carrier and the gearbox output shaft (fan shaft), should be distributed in the required proportions to provide the improvements as described above in relation to the thirteenth and sixteenth aspects. In particular, the inventors have found that maintaining this ratio within a specified range allows reducing the risk of gear tooth damage, while still maintaining sufficient stiffness of the fan shaft to avoid very large displacements of the gears within the gearbox.

Those skilled in the art will appreciate that a fan shaft stiffness that is too low can cause wind-up of the carrier (for planetary gearboxes) or ring gear (for sun gearboxes), resulting in gear misalignment. However, a relative increase in fan shaft stiffness beyond the required ratio range may not provide further benefit, and may conversely detrimentally increase the size and/or weight of the fan shaft.

The ratio of shelf to fan shaft torsional stiffness can be greater than or equal to 9, and optionally between 9 and 1.9 x 10²Within the range of (1).

The fan shaft may include: two shaft portions; a gearbox output shaft portion extending from the gearbox and a fan portion extending between the gearbox output shaft portion and the fan.

The torsional stiffness of the planet carrier may be greater than or equal to 1.60 x 10⁸Nm/rad, and optionally at 1.60X 10⁸Nm/rad to 1.00X 10¹¹Nm/rad, or in the range of 2.7X 10⁸Nm/radian to 1 × 10¹⁰Nm/radian.

The fan may have a fan diameter in a range of 240cm to 280cm, and the rack-to-fan shaft torsional stiffness ratio may be greater than or equal to 9.

The fan may have a fan diameter in a range of 330cm to 380cm, and the rack-to-fan shaft torsional stiffness ratio may be greater than or equal to 12.

The carrier to gearbox support torsional stiffness ratio may be in the range of 2.3 to 300, and optionally in the range of 2.6 to 50.

can be greater than or equal to 70, and optionally less than or equal to 5000.

According to a twentieth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades. The engine core further comprising a gearbox that receives an input from the spindle and outputs a drive to the fan 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted; and a fan shaft connecting the output of the gearbox to the fan. The ratio of the stiffness of the frame to the stiffness of the fan shaft is:

greater than or equal to 8, and optionally may be less than or equal to 1100.

The engine may be as described in the nineteenth aspect.

According to a twenty-first aspect, there is provided a thruster for an aircraft, the thruster comprising: a fan comprising a plurality of fan blades; a gear case; and a power unit for driving the fan via the gearbox. The gearbox being arranged to receive input from a spindle driven by the power unit and to output 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted; and a fan shaft connecting the output of the gearbox to the fan. The ratio of the stiffness of the frame to the stiffness of the fan shaft is:

greater than or equal to 8, and optionally may be less than or equal to 1100.

In various other aspects of the invention, the specified boundary on the ratio of the stiffness of the carrier to the fan shaft may be replaced or otherwise provided by a specified boundary on the product of the components of the ratio of the stiffness of the carrier to the fan shaft (i.e., the boundary on the torsional stiffness of the planet carrier multiplied by the torsional stiffness of the fan shaft). In various aspects, the value of the product may be greater than or equal to 1.5 × 10¹⁵N²m²rad^-2And optionally less than 3.0 x 10¹⁸N²m²rad^-2. Optionally, the value may be greater than or equal to 2.0 × 10¹⁵N²m²rad^-2And optionally less than 7.0 x 10¹⁷N²m²rad^-2。

For example, according to another aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor and connecting the turbine to the compressorThe mandrel of (1); 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 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 planet gears, a ring gear and a planet carrier on which the planet gears are mounted; and a fan shaft connecting the output of the gearbox to the fan. The torsional rigidity of the planet carrier multiplied by the torsional rigidity of the fan shaft is greater than or equal to 1.5 multiplied by 10¹⁵N²m²rad^-2。

According to a twenty-second aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier comprising a plurality of pins, each pin being arranged to have a planet gear of the plurality of planet gears mounted thereon. The first frame to pin stiffness ratio is:

greater than or equal to 10. The first shelf to pin stiffness ratio may be less than or equal to 40.

The first carrier to pin stiffness ratio can be greater than or equal to 15, and optionally in the range of 15 to 30.

The radial bending stiffness of each pin may be greater than or equal to 3.00 x 10⁸N/m, and optionally greater than or equal to 6.3X 10⁸N/m。

The second frame to pin stiffness ratio is:

greater than or equal to 24, and optionally greater than or equal to 34, and optionally less than or equal to 180.

The fan may have a fan diameter in the range of 240cm to 280 cm. In such embodiments, the first carrier to pin stiffness ratio can be greater than or equal to 15, and optionally in the range of 15 to 25. Alternatively, the fan may have a fan diameter in the range of 330cm to 380 cm. In such embodiments, the first carrier to pin stiffness ratio can be greater than or equal to 16, and optionally in the range of 16 to 35.

According to a twenty-third aspect, there is provided a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier comprising a plurality of pins, each pin being arranged to have a planet gear of the plurality of planet gears mounted thereon. The second frame to pin stiffness ratio is:

greater than or equal to 24, and optionally may be less than or equal to 180.

The second planet carrier to pin stiffness ratio may be greater than or equal to 34, and optionally in the range of 34 to 140.

The tilt stiffness of each pin can be greater than or equal to 4.0 x 10⁶Nm/rad, and may optionally be greater than or equal to 8.7X 10⁶Nm/rad。

The first frame to pin stiffness ratio is:

may be greater than or equal to 10 or 15.

The first shelf to pin stiffness ratio can be less than or equal to 40, and optionally in the range of 15 to 30.

The fan may have a fan diameter in the range of 240cm to 280 cm. In such embodiments, the second shelf to pin stiffness ratio can be greater than or equal to 34, and optionally in the range of 34 to 120.

Alternatively, the fan may have a fan diameter in the range of 330cm to 380 cm. In such embodiments, the second shelf to pin stiffness ratio can be greater than or equal to 40, and optionally in the range of 40 to 180.

The inventors have found that the torsional stiffness of the carrier and the radial bending stiffness and/or the tilt stiffness of each pin of the carrier should be selected to match the required relationship in order to improve engine life and/or efficiency, for example by protecting gear teeth and/or improving planetary load sharing.

Thus, the torsional stiffness of the carrier is arranged to be relatively high compared to the radial bending stiffness and/or the tilt stiffness of each individual pin to reduce or avoid the risk of gear tooth deformation as described above, while also reducing differential loading/improving load sharing. Furthermore, the inventors have found that pin tilt stiffness has a more pronounced effect than pin radial bending stiffness-for the same amount of deflection, excessive tilt deflection of the pin is more disruptive than radial bending deflection, as tilt deflection can produce two compound effects, firstly, load sharing can worsen, some planet gears bear greater load sharing than others, and secondly, face load distribution shifts. Thus, the greater force on a particular planet gear is concentrated on one side of the gear, rather than being evenly distributed across the entire tooth. The increased load on the gear and the increased concentration of the load may thus damage the gear teeth. Thus, in some embodiments, the holding pin tilt stiffness is higher than 4.0 x 10⁶Nm/rad, and optionally higher than 8.7X 10⁶Nm/rad or 1.4X 10⁷Nm/rad is particularly important.

Relatively increasing the torsional stiffness of the frame beyond a specified relationship may provide a reduced benefit, or indeed negatively impact performance, since the unnecessary size and/or weight of the stiffer frame offsets the performance gain from reduced roll-up.

Each pin may have a flexible connection to the frame. The flexible connection may be provided by one or more of a portion of the pin, a portion of the shelf, and/or a separate component. For the purpose of evaluating pin stiffness, the soft connection may be classified as a part of the pin.

According to a twenty-fourth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the engine comprising: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted.

The first frame to pin stiffness ratio is:

greater than or equal to 10, and optionally greater than or equal to 15 (and/or optionally less than or equal to 40); or

The second frame to pin stiffness ratio is:

greater than or equal to 24 (and optionally less than or equal to 180).

The method includes operating a gas turbine engine to provide propulsion at cruise conditions. The engine may be as described in the twenty-second or twenty-third aspect.

The method may include driving the gearbox with the following input torques:

(ii) Greater than or equal to 28000Nm at maximum takeoff, and optionally from 28000Nm to 135000 Nm.

According to a twenty-fifth aspect, there is provided a thruster for an aircraft, the thruster comprising: a fan comprising a plurality of fan blades; a gear case; and a power unit for driving the fan via the gearbox. The gearbox is an epicyclic gearbox arranged to receive input from a spindle driven by a power unit and to output drive to the fan so as to drive the fan at a lower rotational speed than the spindle, and comprises a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The first frame to pin stiffness ratio is:

greater than or equal to 10, and optionally less than or equal to 40.

According to a twenty-sixth aspect, there is provided a thruster for an aircraft, the thruster comprising: a fan comprising a plurality of fan blades; a gear case; and a power unit for driving the fan via the gearbox. The gearbox is an epicyclic gearbox arranged to receive input from a spindle driven by a power unit and to output drive to the fan so as to drive the fan at a lower rotational speed than the spindle, and comprises a sun gear, a plurality of planet gears, a ring gear and a planet carrier on which the planet gears are mounted. The second frame to pin stiffness ratio is:

greater than or equal to 24, and optionally less than or equal to 180.

In various other aspects of the invention, the specified boundaries on the first and second carrier-to-pin stiffness ratios may be replaced or otherwise provided by specified boundaries on the product of the components of the respective stiffness ratios (i.e., boundaries on the torsional stiffness of the planet carrier times the tilt stiffness of the pins, or boundaries on the effective linear torsional stiffness of the planet carrier times the radial bending stiffness of the pins). In various aspects, the value of the product of the first carrier to pin stiffness ratio (i.e., the effective linear torsional stiffness of the carrier multiplied by the pin radial bending stiffness) may be greater than or equal to 2.1 × 10¹⁸N²m^-2And optionally less than 3.6 x 10²⁰N²m^-2. Optionally, the value may be greater than or equal to 5.8 × 10¹⁸N²m^-2And optionally less than 1.7 x 10²⁰N²m^-2. In various aspects, the value of the product of the second carrier and pin stiffness ratio 9 (i.e., the torsional stiffness of the carrier multiplied by the pin tilt stiffness) may be greater than or equal to 1.0 × 10¹⁵N²m²rad^-2And optionally less than 4.7 x 10¹⁷N²m²rad^-2. Optionally, the value may be greater than or equal to 2.5 × 10¹⁵N²m²rad^-2And optionally less than 2.0 x 10¹⁷N²m²rad^-2。

Those skilled in the art will appreciate that gas turbine engine, method, and propulsor aspects may be formulated accordingly. Optional features of aspects of corresponding ratios may also be applied to these aspects.

In any of the preceding aspects, any one or more of the following may apply:

the turbine may be a first turbine, the compressor is a first compressor, and the spindle is a first spindle. The engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor. The second turbine, the second compressor and the second spindle may be arranged to rotate at a higher rotational speed than the first spindle.

The planet carrier may include a front plate and a rear plate and a pin extending between the front plate and the rear plate. Each pin may be arranged to have a planet gear mounted thereon. The planet carrier may further comprise lugs extending between the front and rear plates, the lugs being arranged to pass between adjacent planet gears.

The gearbox may comprise an odd number of planet gears and optionally may comprise 3, 5 or 7 planet gears.

The fan may have a fan diameter greater than 240cm and less than or equal to 380cm, and optionally greater than 300cm and less than or equal to 380 cm.

The gearbox input shaft may provide a soft mount for the sun gear, facilitating some movement of the sun gear. The spindle may comprise a stiffer part and a less stiff part, the less stiff part providing the gearbox input shaft and being arranged to be located between the stiffer part and the sun gear and being arranged to provide or contribute to a soft mount for the sun gear.

The gear ratio of the gearbox may be within any range disclosed herein, for example, within a range of 3.2 to 4.5, and optionally within a range of 3.3 to 4.0.

The specific thrust of the engine during cruising can be 70NKg^-1s to 90NKg^-1s is in the range of.

The bypass ratio at cruise may be in the range of 12.5 to 18; and optionally in the range of 13 to 16.

For any parameter or ratio of parameter X claimed or disclosed herein, a limit on the value that X may take to be represented as "X greater than or equal to Y" may alternatively be represented as "1/X less than or equal to 1/Y". Thus, any ratio or parameter defined in the above aspects and statements may be expressed as "1/X less than or equal to 1/Y" instead of "X greater than or equal to Y". Zero can be considered as the lower limit of the value 1/X.

Various parameters of the gearbox and/or more generally the engine may be adjusted to allow the engine to meet the specifications of the various aspects outlined above. Comments on various such parameters are provided below, with further examples of the ways in which these parameters may be adjusted being provided later in the description of the components.

One or more of the factors of gearbox size, gearbox geometry (including whether or not there are lugs in the carrier, and the number, size and/or shape of any lugs present), and material selection may be selected or adjusted to achieve the desired carrier stiffness. The young's modulus of the material from which the shelf is made (typically steel) may for example be in the range 100GPa to 250GPa or 105GPa to 215GPa, and optionally about 210 GPa. Different grades of steel or other types of metal may be selected to achieve different stiffness for the same size and geometry. For example, steel having a Young's modulus in the range of 190GPa to 215GPa, a titanium alloy having a Young's modulus in the range of 105GPa to 120GPa, or a metal (such as titanium) having a Young's modulus of about 110GPa may be used in various embodiments.

The flexibility of the carrier (which is in fact the inverse of the stiffness) allows the alignment of the gears and bearings to be varied. The inventors have recognized that having a certain amount of flexibility in some places may advantageously allow for correction of manufacturing misalignment in use, may tolerate a certain misalignment, but a larger misalignment may have a detrimental effect on the operation of the engine, and have discovered various stiffness relationships to obtain the advantages of an appropriate stiffness range.

One or more of material selection, pin geometry (e.g., diameter), pin mounting design, and internal pin structure (e.g., solid or hollow) may be selected or adjusted to achieve a desired pin stiffness. The pin material is typically steel (typically having a young's modulus of 100GPa to 250GPa, and optionally about 210GPa), and one or more different steel grades may be selected to tune the stiffness.

Some flexibility of the pins may be provided to allow correction of planetary misalignment, but too much flexibility may produce destructive misalignment. Excessive pin stiffness may result in excessive size and/or weight, thereby reducing overall performance.

Turning to the gearbox input shaft, the inventors have found that the torsional stiffness of the gearbox input shaft has an effect on the torsional stiffness of the overall transmission, but has a relatively minimal effect on the gearbox operation, as torsional deflection results only in windup and does not result in gear misalignment. Thus, the torsional stiffness of the gearbox input shaft may be lower than the carrier without detrimental effects.

Similar considerations apply to the fan shaft (gearbox output shaft).

The present inventors have recognized that lowering the torsional stiffness of the shaft below the range defined herein may produce detrimental torsional vibrations at low modal frequencies (those skilled in the art will appreciate that lower modal frequency rotational modes have greater amplitude/deflection than higher modes, and thus avoid the lower modes being more important), while increasing the torsional stiffness above the range defined herein may result in an excessive shaft size and/or weight without a corresponding improvement in performance. One or more of the shaft diameter, material, and wall thickness may be adjusted in order to obtain a desired range of shaft stiffness.

Turning to the gearbox size, particularly the ring gear pitch diameter (PCD), which is a measure of the gearbox size, the inventors have recognized that the optimal PCD may be selected by considering the relationship between improved performance due to the use of improved leverage for larger gearbox sizes and the effect of increasing the drag for larger gearbox sizes (reduced return for improved leverage from larger sizes larger than a certain PCD, and increased size and weight for larger sizes). The ring gear material may be selected to ensure that the maximum expected torque density for the PCD dimensions is well within tolerance limits.

One or more of the gearbox support materials and geometry may be adjusted to achieve a desired torsional stiffness. The inventors have realized that the gearbox support torsional stiffness may be selected to be high enough to suppress the effect of torque ripple, thereby keeping the gearbox movement within an acceptable range, while avoiding adding unnecessary size and/or weight.

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

The gas turbine engine may include a gearbox that receives an input from the spindle and outputs a drive to the fan to drive the fan at a lower rotational speed than the spindle. The input to the gearbox may come directly from the spindle or indirectly from the spindle, for example via spur gear shafts and/or gears. The spindle may rigidly connect the turbine and compressor such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

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 shaft, two shafts, or three shafts. By way of example only, the turbine connected to the spindle may be a first turbine, the compressor connected to the spindle may be a first compressor, and the spindle may be a first spindle. The engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor. The second turbine, the second compressor and the second spindle may be arranged to rotate at a higher rotational speed than the first spindle.

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 (e.g. directly, e.g. via a substantially annular conduit).

The gearbox may be arranged to be driven by a spindle (e.g. the first spindle in the above example) which is configured (e.g. in use) to rotate at the lowest rotational speed. For example, the gearbox may be arranged to be driven only by the spindles (e.g. only the first spindle, not the second spindle in the above example) that are configured to rotate (e.g. in use) at the lowest rotational speed. Alternatively, the gearbox may be arranged to be driven by any one or more shafts, such as the first shaft and/or the second shaft in the above examples.

The gearbox may be a reduction gearbox (since the output to the fan is lower than the rotational rate of the input from the spindle). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "sun" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range 3 to 4.2, or 3.2 to 3.8, for example around 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 may be between any two values in the preceding sentence. By way of example only, the gearbox may be a "sun" gearbox having a gear ratio in the range of 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside of these ranges.

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

The or each compressor (e.g. the first and second compressors as described above) may comprise any number of stages, for example 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 (as 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 each other.

The or each turbine (e.g. the first and second turbines as described above) may comprise any number of stages, for example 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.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas wash 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 about) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26 or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be within a range of inclusion defined by any two values in the preceding sentence (i.e., the values may form an upper or lower limit), for example, within a range of 0.28 to 0.32. These ratios may be generally referred to as hub-to-tip ratios. Both the radius at the hub and the radius at the tip may be measured at the leading (or axially forwardmost) portion of the blade. Of course, the hub-to-tip ratio refers to the gas scrubbing portion of the fan blade, i.e., the portion radially outside of any platform.

The radius of the fan may be measured between the engine centerline and the tip at the leading edge of the fan blade. The fan diameter (which may be only twice the fan radius) may be greater than (or about) any one of: 220cm, 230cm, 240cm, 250cm (about 100 inches), 260cm, 270cm (about 105 inches), 280cm (about 110 inches), 290cm (about 115 inches), 300cm (about 120 inches), 310cm, 320cm (about 125 inches), 330cm (about 130 inches), 340cm (about 135 inches), 350cm, 360cm (about 140 inches), 370cm (about 145 inches), 380cm (about 150 inches), 390cm (about 155 inches), 400cm, 410cm (about 160 inches), or 420cm (about 165 inches). The fan diameter 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), for example, within a range of 240cm to 280cm or 330cm to 380 cm.

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

In use of the gas turbine engine, a fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tips of the fan blades to rotate at a speed U_{Tip end}And (4) moving. The work done by the fan blades 13 on the flow results in an enthalpy rise dH for the flow. Fan tip load may be defined as dH/U_{Tip end} ²Where dH is the enthalpy rise across the fan (e.g., 1-D mean enthalpy rise), and U_{Tip end}Is the (translational) speed of the fan tip, e.g. at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular speed). The fan tip load at cruise conditions may be greater than (or about) any one of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. The fan tip load 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), such as within a range of 0.28 to 0.31 or 0.29 to 0.3.

A gas turbine engine 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 at cruise conditions. In some arrangements, the bypass ratio may be greater than (or about) any one of: 10. 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be within an inclusive range defined by any two values in the preceding sentence (i.e., the values may form an upper or lower limit), for example, within a range of 12 to 16, or a range of 13 to 15, or a range of 13 to 14. The bypass conduit may be substantially annular. The bypass duct may be located radially outward of the core engine. The radially outer surface of the bypass duct may be defined by the nacelle and/or the fan casing.

The overall pressure ratio of the gas turbine engine described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the highest pressure compressor outlet (before entering the combustor). By way of non-limiting example, the overall pressure ratio at cruise of a gas turbine engine as described and/or claimed herein may be greater than (or about) any one of: 35. 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within the inclusive range defined by any two values in the preceding sentence (i.e., these values may form an upper or lower limit), for example, within the range of 50 to 70.

The specific thrust of the engine may be defined as the net thrust of the engine divided by the thrust passing throughTotal mass flow of the engine. At cruise conditions, the specific thrust of the engines described and/or claimed herein may be less than (or about) any of the following: 110Nkg^-1s、105Nkg^-1s、100Nkg^-1s、95Nkg^-1s、 90Nkg^-1s、85Nkg^-1s or 80Nkg^-1And s. The specific 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), for example at 80Nkg^-1s to 100Nkg^-1s, or 85Nkg^-1s to 95Nkg^-1s is in the range of. Such engines may be particularly efficient compared to conventional gas turbine engines.

A 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 about) any of: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN or 550 kN. The maximum thrust may be within the 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 a range of 330kN to 420kN, for example 350kN to 400 kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions, at sea level, plus 15 ℃ (ambient pressure 101.3kPa, temperature 30 ℃), with the engine at rest.

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

As used herein, Maximum Takeoff (MTO) conditions have conventional meaning. Maximum takeoff conditions may be defined as operating the engine at maximum takeoff thrust at the end of the runway under International Standard Atmospheric (ISA) sea level pressure and temperature conditions +15 ℃, which is generally defined as an aircraft speed of about 0.25Mn, or between about 0.24Mn and 0.27 Mn. Thus, the maximum takeoff condition of the engine may be defined as operating the engine at its maximum takeoff thrust (e.g., maximum throttle) at International Standard Atmospheric (ISA) sea level pressure and temperature +15 ℃, with a fan inlet speed of 0.25 Mn.

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

A fan as described and/or claimed herein may include a central portion from which fan blades may extend, for example, in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fastener that may engage a corresponding slot in the hub (or disk). By way of example only, such fasteners may be of dovetail form that may be inserted into and/or engage corresponding slots in the hub/disk to secure the fan blade to the hub/disk. By way of further example, the fan blade may be integrally formed 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 disc 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/disk by welding (such as linear friction welding).

The gas turbine engines described and/or claimed herein may or may not be provided with Variable Area Nozzles (VANs). Such variable area nozzles may allow the outlet area of the bypass duct to vary in use. The general principles of the present disclosure may be applied to engines with or without a VAN.

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.

As used herein, cruise conditions have the conventional meaning and will be readily understood by the skilled artisan. Thus, for a given gas turbine engine of an aircraft, the technician will immediately recognize a cruise condition as meaning the operating point at which that gas turbine engine is designed for intermediate cruising of the engine attached to the aircraft at a given mission (which may be referred to in the industry as an "economic mission"). In this regard, intermediate cruise is a critical point in the aircraft flight cycle at which 50% of the total fuel burned between the peak of ascent and the beginning of descent has been burned (which may be approximated in time and/or distance to the midpoint between the peak of ascent and the beginning of descent). Thus, the cruise conditions define an operating point of the gas turbine engine that, taking into account the number of engines provided to the aircraft, provides a 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 at intermediate cruising. For example, if the engine is designed to be attached to an aircraft having two engines of the same type, the engines provide half of the total thrust required for steady state operation of the aircraft at intermediate cruise under cruise conditions.

In other words, for a given gas turbine engine of an aircraft, the cruise condition is defined as the operating point of the engine that provides a specified thrust at an intermediate cruise atmospheric condition (defined by the international standard atmosphere according to ISO 2533 at an intermediate cruise altitude) (it is necessary to provide, in combination with any other engine on the aircraft, a steady state operation of the aircraft to which the gas turbine engine is designed to be attached, at a given intermediate cruise mach number). For any given gas turbine engine of an aircraft, the intermediate cruise thrust, atmospheric conditions and mach number are known, so at cruise conditions the operating point of the engine is well defined.

By way of example only, the forward speed at cruise conditions may be at any point in the range from 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 about mach 0.8, about mach 0.85, or 0.8 to 0.85. Any single speed within these ranges may be part of the cruise conditions. For some aircraft, cruise conditions may be outside of these ranges, such as below mach 0.7 or above mach 0.9.

By way of example only, the cruise conditions may correspond to standard atmospheric conditions (according to the international standard atmospheric ISA) at an altitude within the following ranges: 10000m to 15000m, for example in the range 10000m to 12000m, for example in the range 10400m to 11600m (about 38000 feet), for example in the range 10500m to 11500m, for example in the range 10600m to 11400m, for example in the range 10700m (about 35000 feet) to 11300m, for example in the range 10800m to 11200m, for example in the range 10900m to 11100m, for example about 11000 m. Cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

By way of example only, the cruise conditions may correspond to an operating point of the engine providing a known desired thrust level (e.g., a value in the range of 30kN to 35 kN) at a forward mach number (Mn) of 0.8 and standard atmospheric conditions (according to international standard atmosphere) at an altitude of 38000ft (11582 m). By way of another example only, the cruise conditions may correspond to an operating point of the engine providing a known desired thrust level at a forward mach number of 0.85 (e.g., a value in the range of 50kN to 65 kN) and standard atmospheric conditions at an altitude of 35000ft (10668m) (according to international standard atmosphere).

In use, the gas turbine engine described and/or claimed herein may be operated at cruise conditions as defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft on which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

According to one aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is an aircraft to which the gas turbine engine has been designed for attachment. Thus, the cruise condition according to this aspect corresponds to an intermediate cruise of the aircraft, as defined elsewhere herein.

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

According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. Operations according to this aspect may include (or may be) operations at intermediate cruising of the aircraft, as defined elsewhere herein.

Although in the arrangement described herein the drive source for the pusher fan is provided by a gas turbine engine, it will be appreciated by those skilled in the art that the gearbox configuration disclosed herein may be applied to other forms of aircraft propellers including alternative drive types. For example, the above described gearbox arrangement may be used in an aircraft propeller comprising a propeller fan driven by an electric motor. In such cases, the electric motor may be configured to operate at a higher rotational speed, and thus may have a smaller rotor diameter, and may be more power intensive. The gearbox configuration of the foregoing aspect 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 propulsive fan via a gearbox, the gearbox and/or an input/output/support thereof being as described and/or claimed herein.

Those skilled in the art will appreciate that features or parameters described in relation to any one of the above aspects are applicable to any other aspect unless mutually exclusive. Furthermore, unless mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

Drawings

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

FIG. 1 is a cross-sectional side view of a gas turbine engine;

FIG. 2 is a close-up cross-sectional side view of an upstream portion of a gas turbine engine;

FIG. 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

FIG. 4 is a schematic diagram illustrating automatic load sharing adjustment within a gearbox;

FIG. 5 is a schematic diagram showing the radial bending stiffness of a cantilever beam;

FIG. 6 is a schematic diagram showing the tilt stiffness of a cantilever beam;

FIG. 7 is a schematic diagram illustrating torsional stiffness of a cantilever beam;

FIG. 8 is a schematic diagram showing the radial bending stiffness of a cantilever beam with its movable end slidably mounted on a flat surface;

FIG. 9 is a schematic diagram illustrating radial bending stiffness of a planet carrier;

FIG. 10 is a schematic diagram illustrating the tilt stiffness of the gantry, and more particularly illustrating the determination of the effective linear tilt stiffness of the gantry;

FIG. 11 is a schematic view showing the tilt stiffness of the frame;

FIG. 12 is a schematic diagram showing torsional stiffness of the frame in side view;

FIG. 13 is a schematic diagram showing the torsional stiffness of an alternative frame in elevation;

FIG. 14 is a schematic diagram illustrating torsional stiffness of the frame of FIG. 13;

FIG. 15 is a schematic diagram showing a front view of a cage including lugs;

FIG. 16 is a schematic view showing the radial bending stiffness of unbonded pins;

FIG. 17 is a schematic view showing the radial bending stiffness of the dowel pin;

FIG. 18 is a schematic diagram showing the tilt stiffness of the pin;

FIG. 19 is a schematic view showing a spindle, in particular a gearbox input shaft;

FIG. 20 is a schematic diagram illustrating torsional stiffness of a gearbox input shaft;

FIG. 21 includes side and radial views of the gearbox support, illustrating the torsional stiffness of the gearbox support;

FIG. 22 is a schematic diagram showing a portion of an engine having a sun gearbox;

FIG. 23 is a schematic diagram showing the connection of the fan shaft to the sun gear box;

FIG. 24 is a schematic view showing the connection of the fan shaft to the planetary gearbox;

FIG. 25 is a schematic diagram illustrating fan shaft torsional stiffness in an engine having a sun gear box;

FIG. 26 is a graph of displacement versus load illustrating a spring region in which component stiffness may be determined; and is

Fig. 27 illustrates a method.

Detailed Description

Fig. 1 shows a gas turbine engine 10 having a main axis of rotation 9. The engine 10 comprises an air intake 12 and a propeller fan 23 which generates two air flows: core stream a and bypass stream B. The gas turbine engine 10 includes a core 11 that receives a core gas flow A. The engine core 11 includes, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, a combustion apparatus 16, a high pressure turbine 17, a low pressure turbine 19, and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is led into a combustion device 16, where the compressed air is mixed with fuel and the mixture is combusted. The resulting hot combustion products are then expanded by the high and

low pressure turbines

17, 19 before being discharged through the nozzle 20, thereby driving the high and

low pressure turbines

17, 19 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by means of a suitable interconnecting shaft 27. The fan 23 typically provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement of the geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see fig. 1) drives a shaft 26, which shaft 26 is coupled to a sun gear or sun gear 28 of an epicyclic gear arrangement 30. Radially outward of and intermeshed with the sun gear 28 is a plurality of planet gears 32 that are coupled together by a carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously about the sun gear 28, while rotating each planet gear 32 about its own axis. The planet carrier 34 is coupled to the fan 23 via a connecting rod 36 for driving the fan in rotation about the engine axis 9. Radially outward of and intermeshes with the planet gears 32 is a ring gear or ring gear 38 that is coupled to the fixed support structure 24 via a linkage 40.

The connecting rod 36 may be referred to as a fan shaft 36, the fan shaft 36 optionally including 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. In the embodiment shown in fig. 1 and 2, the gearbox 30 is a planetary gearbox and the gearbox output shaft portion 36a is connected to the planet carrier 34, and thus may be referred to as a carrier output shaft 36 a. In the sun gearbox 30, the gearbox output shaft portion 36a may be connected to the ring gear 38, and thus may be referred to as the ring output shaft 36 a. In the embodiment shown in fig. 1 and 2, the fan section 36b of the fan shaft 36 connects the gearbox output shaft section 36a to the fan 23. Accordingly, the output of the gear box 30 is transmitted to the fan 23 via the fan shaft 36 to rotate the fan. In alternative embodiments, the fan shaft 36 may comprise 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 the sun gearbox 30 may be equally applicable to an engine having the planetary gearbox 30, and vice versa.

It is noted that the terms "low pressure turbine" and "low pressure compressor" as used herein may refer to the lowest pressure turbine stage and lowest pressure compressor stage, respectively (i.e., not including the fan 23), and/or the turbine and compressor stages that are connected together by an interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan 23). In some documents, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be referred to as an "intermediate pressure turbine" and an "intermediate pressure compressor". Where such alternative nomenclature is used, the fan 23 may be referred to as the first or lowest pressure compression stage.

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

The epicyclic gearbox 30 shown by way of example in fig. 2 and 3 is planetary, with a planet carrier 34 coupled to the output shaft via a connecting rod 36, with a ring gear 38 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 sun arrangement in which the planet carrier 34 is held stationary, 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.

It should be understood that the arrangements shown in fig. 2 and 3 are exemplary only, and that 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 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of another example, the connections (such as the

links

36, 40 in the example of FIG. 2) between the gearbox 30 and other components of the engine 10 (such as the input shaft 26, the output shaft, and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of another example, any suitable arrangement of bearings between rotating and stationary components of the engine (e.g., between input and output shafts from a gearbox and a stationary structure such as a gearbox housing) may be used, and the present disclosure is not limited to the exemplary arrangement of fig. 2. For example, where the gearbox 30 has a sun arrangement (as described above), the skilled person will readily appreciate that the arrangement of the output and support links and the bearing locations will generally differ from that shown by way of example in figure 2.

Accordingly, the present disclosure extends to gas turbine engines having any of a gearbox type (e.g., sun or planetary gear), support structure, input and output shaft arrangements, and bearing locations.

Optionally, the gearbox may drive additional and/or alternative components (e.g., a medium pressure compressor and/or a booster compressor).

Other gas turbine engines to which the present disclosure is applicable may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in fig. 1 has a

splitter nozzle

18, 20, which means that the flow through the bypass duct 22 has its own nozzle 18, which is separate from and radially outside of the core engine nozzle 20. However, this is not limiting and any aspect of the disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed or combined before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split) may have a fixed or variable area. Although the described examples relate to turbofan engines, the present disclosure is applicable to any type of gas turbine engine, such as an open rotor (where the fan stages are not surrounded by a nacelle) or, for example, a turboprop engine.

The geometry of the gas turbine engine 10 and its components are defined by conventional shafting, including axial (aligned with the axis of rotation 9), radial (in the direction from bottom to top in fig. 1), and circumferential (perpendicular to the page in the view of fig. 1). The axial, radial and circumferential directions are mutually perpendicular. In the arrangement depicted, the shelf 34 includes two

plates

34a, 34 b; specifically a front plate 34a and a rear plate 34b (see, e.g., fig. 9). Each

plate

34a, 34b extends in a radial plane, with the front plate 34a being further forward/closer to the fan 23 in the engine 10 than the rear plate 34 b.

The shelf 34 may take any suitable form. For example, the cage may be symmetrical or asymmetrical about its axial midpoint. By way of example only, in the arrangement depicted, the carrier 34 is asymmetric about its axial midpoint, but the rear plate 34b is stiffer (e.g., 50% to 300% stiffer) than the front plate 34a to compensate for asymmetric torque variations on the gearbox 30. In some embodiments, no front panel 34a may be provided, or only a smaller front panel 34a may be provided. In some embodiments, the

plates

34a, 34b of the carrier 34 may be equally rigid (e.g., in various planetary gearbox arrangements; in some sun gearbox arrangements, a stiffer back plate 34b may be preferred).

A plurality of pins 33 extend across the shelf 34 (in the arrangement depicted, between the front plate 34a and the rear plate b), as shown for example in fig. 9-18. The pin 33 forms part of a shelf 34. Each pin 33 has a planetary gear 32 mounted thereon. As mentioned herein, reference to the planetary gear 32 includes the gear 32 mounted on the pin 33, regardless of whether the gearbox is a so-called "sun" arrangement (such as shown in fig. 22) or a "planetary gearbox" (such as shown in fig. 2).

In the described embodiment, the stiffness of the shelf in the region at each of the front and rear ends of each pin 33 may be arranged to be relatively low to facilitate more even load distribution; i.e. to increase the load sharing factor. This can be described as a soft mount for each pin 33. The

soft mounts

33a, 33b may allow a degree of movement of the pins 33 relative to each other and to the

carrier plates

34a, 34b, allowing for accommodation of differences or other manufacturing imperfections between the planet gears 34 without creating significant load differences between different planet gears 34.

In various embodiments, such

soft mounts

34a, 34b may be provided by a portion of the pin 33, by a separate component, and/or by a portion of the respective

planet carrier plate

34a, 34 b. The

soft mounts

34a, 34b can be designed to accommodate movement to address one or more of the following issues: carrier bearing positioning accuracy and clearance, planet pin run-out from bearing surface to mounting feature, planet gear tooth to bearing run-out, planet gear tooth spacing and thickness variation/manufacturing tolerance, sun gear tooth spacing and thickness variation/manufacturing tolerance, and/or gearbox input shaft main line bearing positioning accuracy and clearance, among others. For example, in various embodiments, the

soft mounts

34a, 34b may be arranged to allow the pins to move about 500 μm.

The size, design, and/or material of the pins may be adjusted to provide the appropriate stiffness for the shelf 34.

In some arrangements such as that shown in fig. 15, a lug 34c may be provided which extends between the

carrier plates

34a, 34b and past the planet gears 32. The presence/absence of lugs 34c, as well as the number, shape, and/or material of the lugs, may vary in various embodiments, and may be adjusted to provide the proper stiffness for the cage 34.

The use of compliance within the gearbox 30 to improve load sharing is schematically illustrated in fig. 4, which shows a planetary gearbox 30 with three

planetary gears

32a, 32b, 32c (misalignment is exaggerated in this figure for clarity of explanation). In this example, the sun gear 28 is slightly off-center relative to the ring gear 38, specifically closer to the two

planet gears

32a and 32b than to the third planet gear 32 c. In the illustrated exemplary embodiment, there is no contact between the third planet gear 32c and the sun gear 28, so that the other two

planet gears

32a, 32b each bear 50% of the load, rather than one third of the load, as would be expected for a uniform load distribution. Such relatively extreme examples are provided merely for ease of reference, and in practice, it would be more likely that contact with one of the planet gears 32c would be reduced but not completely eliminated, for example, so that the percent load sharing is 20:40:40, 26:37:37 or 31:34:34, etc., rather than the ideal uniform load sharing fraction 1/3:1/3:1/3 (i.e., rounded to the nearest integer as the percent load sharing is 33:33: 33).

In the example shown in fig. 4, each of the two

planet gears

32a, 32b in contact with the sun gear 28 exerts a force F on the sun gear 28_a、F_b. Resultant force F on sun gear 28_RPushing the sun gear 28 toward the third planet gears 32c reestablishes contact and makes the load sharing between the planets 32 more uniform. This rebalancing is facilitated by the flexibility in the soft mount/core input shaft 26 of the sun gear 28. Such soft mounts for the sun gear 28 may be designed to accommodate movement to address one or more of the following issues: carrier bearing positioning accuracy and clearance, planet and/or sun gear tooth spacing and thickness variation/manufacturing tolerances, and/or gearbox input shaft main line bearing positioning accuracy and clearance, among others. For example, in eachIn one embodiment, such a soft mount may be arranged to allow the sun gear to move approximately 1000 μm.

Those skilled in the art will appreciate that a similar effect will apply if one of the planet gears 32 is closer to the sun gear 28 than the other planet gears; or if one of the planet gears 32 is larger or smaller than the other planet gears, the associated planet gear 32 is pushed back towards the ring gear 38. The flexibility in the soft mount/bracket 34 of the pin 33 assists in this rebalancing. Thus, minor variations between the planet gears 32 and/or misalignment of the pins 33 or shafts 26 may be accommodated by flexibility within the gearbox 30. Having an odd number of planet gears 32 (e.g., 3, 5, or 7 planet gears) may facilitate such automatic redistribution of load sharing, although an even number of planet gears may be used in some arrangements.

The following general definition of stiffness may be used herein:

radial bending stiffness

Radial bending stiffness is a measure of the deformation caused by a given force applied in any one selected radial direction (i.e., any direction perpendicular to and through the engine axis). The radial bending stiffness is defined in terms of the deformation of the cantilever beam 401 with reference to figure 5. As shown in fig. 5, 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 rotational axis 9 of the engine, and thus relates to the resistance to linear deformation in the radial direction of the engine caused by radial forces. The beam or equivalent cantilever beam member extends along the axis of rotation of the engine, the force F is applied perpendicular to the axis of the engine in any radial direction, 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) 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 an individual component in the cantilever configuration, without the presence of other components that may affect its stiffness. When a force is applied perpendicular to the cantilevered beam and at the free end of the beam, the resultant curvature is not constant, but increases toward the fixed end of the beam.

For some components, the beam may be more appropriately constrained from moving in a particular manner, as described in more detail below for a specific example of pin 33.

Tilt stiffness

The tilt stiffness is defined with reference to figure 6, which shows the resulting deformation of the cantilever beam 401 under a moment M applied at its free end. The tilt stiffness is a measure of the resistance to rotation at the location where the moment is applied to the component. As can be seen in fig. 6, the moment applied at the free end of the cantilever beam produces a constant curvature along the length of the beam between its free and fixed ends. The applied moment M causes a rotation theta at the point where the moment is applied. Thus, the tilt stiffness as defined herein has SI units of Nm/radian.

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

Torsional rigidity

Fig. 7 illustrates the definition of torsional stiffness of a shaft 401 or other body. A torque τ applied to the free end of the beam causes a rotational deformation θ (e.g., torsion) along the length of the beam. Torsional stiffness is the torque applied for a given torsion angle, i.e., τ/θ. The SI unit of torsional stiffness is Nm/rad.

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

For ease of understanding, the following provides a more specific definition of stiffness in relation to the embodiments described herein.

Frame stiffness

The planet carrier 34 holds the planet gears 32 in place. In various arrangements including the described embodiment, the planet carrier 34 includes a front plate 34a and a rear plate 34b, and pins 33 extending between the plates, as shown in fig. 9-17. The pin 33 is arranged parallel to the engine axis 9. In alternative embodiments, only plate 34b may be provided on one side of shelf 34, no plate or only a portion of the plate may be provided on the other side. In the embodiment shown in fig. 15, the shelf 34 also includes a ledge 34c (also referred to as a wedge or web) located between the front plate 34a and the back plate 34 b. The lugs 34c may increase the overall stiffness of the cage 34.

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

Radial flexural rigidity of the frame

In the depicted embodiment, the shelf radial bending stiffness is determined by treating the shelf 34 as a free body fixedly mounted at one plate 34b, and applying a (radial) force F at an axial location of the axial center point of the other plate 34 a. This is illustrated in fig. 9, where the arrow F indicates a (radial) force on the plate 34a, while δ shows a (radial) displacement of the plate 34 a. The force F is shown acting along a line through the engine axis 9. In embodiments having only one plate 34a, the end of the pin 33 remote from the single plate 34a may alternatively be held in place.

In various embodiments, the radial bending stiffness of the carrier 34 may be equal to or greater than 1.20×10⁹N/m, and optionally equal to or greater than 2.0X 10⁹N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the radial bending stiffness of the shelf 34 may be equal to or greater than 1.5 x 10⁹N/m, and optionally may be equal to or greater than 2X 10⁹N/m (and optionally may be equal to 2.30X 10)⁹N/m or 3.85X 10⁹N/m). In some embodiments, such as embodiments with fan diameters in the range of 330cm to 380cm, the radial bending stiffness of the shelf 34 may be equal to or greater than 2.0 x 10⁹N/m, and optionally may be equal to or greater than 3X 10⁹N/m (and optionally may be equal to 3.92 × 10)⁹N/m or 7.70X 10⁹N/m)。

In various embodiments, the radial bending stiffness of the carrier 34 is at 1.20 × 10⁹N/m to 1X 10¹²N/m, and optionally in the range of 2.0X 10⁹N/m to 1.5X 10¹¹N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the radial bending stiffness of the shelf 34 may be in the range of 1.5 x 10⁹N/m to 5X 10¹⁰N/m, and optionally may be in the range of 2X 10⁹N/m to 5X 10⁹N/m or 1.9X 10⁹N/m to 2.7X 10⁹In the range of N/m (and optionally may be equal to 3.85X 10)⁹N/m or equal to 2.30X 10⁹N/m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the radial bending stiffness of the shelf 34 may be 2.0 x 10⁹N/m to 1.6X 10¹¹N/m, and optionally may be in the range of 3.0X 10⁹N/m to 9.0X 10⁹In the range of N/m (and optionally may be equal to 7.70X 10)⁹N/m or 3.92X 10⁹N/m)。

The carrier 34 serves to position the planet gears 32 within the gearbox 30 and reduce or avoid misalignment. Those skilled in the art will appreciate that the desired stiffness of the shelf may be achieved in a variety of different ways, such as by appropriately adjusting one or more of the shelf material and the shelf geometry. For a given material, the stiffness may be a function of, for example, the frame and gearbox dimensions and gearbox configuration. The number of planets and gearbox ratios can also be adjusted to achieve the desired stiffness. For example, the gear ratio of the gearbox 30 may be changed, causing the planetary gear spacing and, in some cases, the number of planetary gears 32 to change. The variation in planet gear spacing may provide more (or less) space for the lugs 34c between the gears 32, and the size and shape of these lugs 34c may be adjusted to achieve the desired carrier stiffness.

Stiffness of rack tilt

The shelf tilt stiffness is a measure of the resistance of the shelf 34 to applied torque M, as shown in fig. 10. The axis of the moment is perpendicular to the engine axis 9. Two points of the planet carrier 34 are selected to measure the tilt stiffness: a forward point at the axial position of the axial center point of the front plate 34a, and a rearward point at the axial position of the axial center point of the rear plate 34 b. As shown by oblique lines in fig. 10, the rear plate 34b is held rigid and is not rotated.

In response to the applied moment M (a counterclockwise moment in the illustrated example, but may be a clockwise moment in other examples), the shelf 34 bends through an angle θ that is not constant at every point along the length of the shelf 34 because the shelf does not have a constant cross-section. θ can thus be measured between a line parallel to the engine axis 9 before deformation and passing through the rear and front points (perpendicular to the front and rear faces of the frame plate) and a line passing through the front point and perpendicular to the front and rear faces of the front frame plate after deformation (no longer parallel to the engine axis). This is shown in fig. 11.

The shelf 34 bends through the total angle a resulting in an arc displacement δ. The angle alpha is measured between a line parallel to the engine axis 9 and passing through the rear and front points before deformation (e.g. theta) and a line passing through the front and rear points after deformation (opposite to theta). Therefore, the values of θ and α may be different.

Thus, as described above, an effective linear tilt stiffness may be defined for the shelf 34. The radius r selected for the definition of effective linear tilt stiffness is the radius of a circle centered at a point on the surface of the plate 34b that is held rigid and passes through the original axial center point of the front plate 34a and the same point after deformation, as shown in the close-up portion of fig. 10. The schematic force diagram shown at the bottom right of fig. 10 marks two equal and opposite forces F, one at the center of the circle of radius r and one at the far end of the radius — the magnitude of F being selected based on the applied moment. The total angle alpha is measured between the first radius before deformation and the second radius after deformation. The axial center point at the radial location of the lower edge of the front plate 34a is selected for ease of reference — any other point along the axial centerline of the front plate 34a (e.g., the point of application of the moment) may be equivalently selected. The arc displacement (arc distance) δ is equal to r α.

In various embodiments, the shelf 34 has a tilt stiffness of greater than or equal to 6.00 x 10⁸Nm/radian, and optionally greater than or equal to 1.3 x 10⁹Nm/radian. In some embodiments, such as embodiments having a fan diameter in the range of 240cm to 280cm, the tilt stiffness of the shelf 34 may be greater than or equal to 2.2 x 10⁹Nm/radian, and optionally may be greater than or equal to 2.4 x 10⁹Nm/radian (and optionally may be equal to 2.71X 10)⁹Nm/radian). In some embodiments, such as embodiments having a fan diameter in the range of 330cm to 380cm, the tilt stiffness of the shelf 34 may be greater than or equal to 2.3 x 10⁹Nm/radian, and optionally may be greater than or equal to 3.0 x 10⁹Nm/radian (and optionally may be equal to 5.70 × 10)⁹Nm/radian).

In various embodiments, the tilt stiffness of the shelf 34 is at 6.00 × 10⁸Nm/radian to 2.80X 10¹¹Nm/radian, and optionally 1.3X 10⁹Nm/radian to 1.2X 10¹¹Nm/radian. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the tilt stiffness of the shelf 34 may be in the range of 2.2 x 10⁹Nm/radian to 1.4X 10¹¹Nm/radian, and optionally may be in the range of 2.4X 10⁹To 5.0X 10⁹In the range of Nm/radian (and optionally may be equal to 2.71X 10)⁹Nm/radian). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the tilt stiffness of the shelf 34 may be in the range of 2.3 x 10⁹Nm/Radian to 2.8X 10¹¹Nm/radian, and optionally may be in the range of 3.0 × 10⁹Nm/radian to 9.0 × 10⁹Nm/radian (and optionally may be equal to 5.70 × 10)⁹Nm/radian).

In various embodiments, the effective linear tilt stiffness of the shelf 34 is greater than or equal to 3.40 x 10⁹N/m, and optionally greater than or equal to 8.0X 10⁹N/m. In some embodiments, such as embodiments having a fan diameter in the range of 240cm to 280cm, the effective linear tilt stiffness of the shelf 34 may be greater than or equal to 1.4 x 10¹⁰N/m, and optionally may be greater than or equal to 1.42X 10¹⁰N/m (and optionally may be equal to 1.68X 10)¹⁰N/m). In some embodiments, such as embodiments having a fan diameter in the range of 330cm to 380cm, the effective linear tilt stiffness of the shelf 34 may be greater than or equal to 1.5 x 10¹⁰N/m, and optionally may be greater than or equal to 3.0X 10¹⁰N/m, and optionally greater than or equal to 7.0X 10¹⁰N/m (and optionally may be equal to 8.36X 10)¹⁰N/m)。

In various embodiments, the effective linear tilt stiffness of the shelf 34 is at 3.40 × 10⁹N/m to 4.20X 10¹²N/m, and optionally 8.0X 10⁹N/m to 1.7X 10¹²N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the effective linear tilt stiffness of the shelf 34 may be at 1.4 x 10¹⁰N/m to 8.4X 10¹¹N/m, and optionally may be in the range of 1.42X 10¹⁰N/m to 2.72X 10¹⁰In the range of N/m (and optionally may be equal to 1.68X 10)¹⁰N/m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the effective linear tilt stiffness of the shelf 34 may be 1.5 x 10¹⁰N/m to 4.2X 10¹²N/m, and optionally may be in the range of 3.0X 10¹⁰N/m to 1.0X 10¹¹N/m, and optionally in the range of 7.0X 10¹⁰N/m to 1.0X 10¹¹In the range of N/m (and optionally may be equal to 8.36X 10)¹⁰N/m)。

As mentioned above, for deflections of the same magnitude, excessive tilt deflection of the carrier 34 is more disruptive than radial buckling or torsional deflection, as tilt deflection can produce two compound effects, firstly, load sharing can deteriorate, with some planet gears 32 taking on greater load sharing than others, and secondly facing load distribution shifting. Thus, the greater force on a particular planet gear 32 is concentrated on one side of the gear, rather than being evenly distributed across the entire tooth. The increased load on the gear 32 and the increased concentration of the load may thus damage the gear teeth.

Torsional stiffness of the frame

The carrier torsional stiffness is a measure of the resistance of the carrier 34 to an applied torque τ, as shown in fig. 12 (axial cross-section) and fig. 13-15 (radial cross-section). The axis of torque is parallel to the engine axis 9. The diagonal line of the plate 34b at the rear end of the shelf 30 indicates that the plate 34b is considered rigid and non-rotating (as with a cantilever beam mount). In embodiments having only one plate 34a, the end of the pin 33 (and the lug 34c (if present)) remote from the single plate 34a may alternatively be held in place.

A torque τ is applied to the carrier 34 (at the axial midpoint of the front plate 34a) and induces a rotational deformation θ (e.g., torsion) along the length of the carrier 34. This twisting causes the shelf 34 to "wind up" when the ends of the pin 33 (and the lugs 34c if present) are held at a fixed radius on the

shelf plates

34a, 34 b.

The angle through which a point on an imaginary circle 902 on the front plate 34a passing through the longitudinal axis of each pin 33 moves is θ, where θ is the angle measured in radians. The imaginary circle 902 may be referred to as a pin pitch circle diameter (pin PCD). The pin PCD may be in the range of 0.38m to 0.65m, for example equal to 0.4m or 0.55 m. Thus, as described above, the radius r of the imaginary circle 902 (e.g., as shown in FIG. 13) may be used to define an effective linear torsional stiffness for the frame 34.

In various embodiments, the torsional stiffness of the shelf 34 is greater than or equal to 1.60 x 10⁸Nm/rad, and optionally greater than or equal to 2.7X 10⁸Nm/rad. In some embodiments, for example, in the range of 240cm to 280cm in fan diameterIn other embodiments, the torsional stiffness of the carrier 34 may be greater than or equal to 1.8 x 10⁸Nm/rad, and optionally may be greater than or equal to 2.5X 10⁸Nm/rad (and optionally may be equal to 4.83X 10)⁸Nm/rad). In some embodiments, such as embodiments having a fan diameter in the range of 330cm to 380cm, the torsional stiffness of the shelf 34 may be greater than or equal to 6.0 x 10⁸Nm/rad, and optionally may be greater than or equal to 1.1X 10⁹Nm/rad (and optionally may be equal to 2.17X 10⁹Nm/rad)。

In various embodiments, the torsional stiffness of the frame 34 is 1.60 × 10⁸To 1.00X 10¹¹Nm/rad, and optionally in the range of 2.7X 10⁸To 1X 10¹⁰Nm/rad. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the torsional stiffness of the shelf 34 may be in the range of 1.8 x 10⁸Nm/rad to 4.8X 10⁹Nm/rad, and optionally may be in the range of 2.5X 10⁸Nm/rad to 6.5X 10⁸Nm/rad (and optionally may be equal to 4.83X 10)⁸Nm/rad). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the torsional stiffness of the shelf 34 may be 6.0 x 10⁸Nm/rad to 2.2X 10¹⁰Nm/rad, and optionally in the range of 1.1X 10⁹Nm/rad to 3.0X 10⁹Nm/rad (and optionally may be equal to 2.17 × 10)⁹Nm/rad)。

In various embodiments, the effective linear torsional stiffness of the carrier 34 can be greater than or equal to 7.00 x 10⁹N/m, and optionally greater than or equal to 9.1X 10⁹N/m. In some embodiments, such as embodiments having a fan diameter in the range of 240cm to 280cm, the effective linear torsional stiffness of the shelf 34 may be greater than or equal to 7.70 x 10⁹N/m. In other such embodiments, the effective linear torsional stiffness of the carrier 34 may be greater than or equal to 9.1 x 10⁹N/m, optionally greater than or equal to 1.1X 10¹⁰N/m (and optionally may be equal to 1.26X 10)¹⁰N/m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cmThe effective linear torsional stiffness of the carrier 34 may be greater than or equal to 1.2 x 10¹⁰N/m, and optionally may be greater than or equal to 2.1X 10¹⁰N/m (and optionally may be equal to 2.88X 10)¹⁰N/m)。

In various embodiments, the effective linear torsional stiffness of the carrier 34 may be in the range of 7.00 x 10⁹N/m to 1.20X 10¹¹N/m, and optionally in the range of 9.1X 10⁹N/m to 8.0X 10¹⁰N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the effective linear torsional stiffness of the shelf 34 may be 9.1 x 10⁹N/m to 6.0X 10¹⁰N/m, and optionally may be in the range of 7X 10⁹N/m to 2X 10¹⁰N/m or 8.5X 10⁹N/m to 2X 10¹⁰In the range of N/m (and optionally may be equal to 1.26X 10)¹⁰N/m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the effective linear torsional stiffness of the shelf 34 may be 1.2 x 10¹⁰N/m to 1.2X 10¹¹N/m, and optionally may be in the range of 1.0X 10¹⁰N/m to 5.0X 10¹⁰In the range of N/m (and optionally may be equal to 2.88X 10)¹⁰N/m)。

The torsional stiffness of the shelf 34 may be controlled to be within a desired range by adjusting one or more parameters, including the shelf material, the shelf geometry, and the presence or absence of lugs.

Stiffness of planet pin

The planet carrier 34 holds the planet gears 32 in place. The pins 33 serve to locate each planet gear 32 on the carrier 34 and reduce or avoid misalignment of the planet gears 32. In the embodiment shown in fig. 16, five planet gears 32 are provided in the gear box 30, one pin 33 for each planet gear 32. In alternative embodiments, a different number of planet gears 32 may be provided, for example 3, 4, 5, 6, 7, 8 or 9 planet gears 32. The skilled person will appreciate that an odd number (e.g. 3, 5, 7, 9) of planet gears may improve load and/or stress sharing within the gearbox 30. In particular, those skilled in the art will appreciate that the use of an odd number of planet gears may improve dynamic loading in a gearbox having a sun gear tooth count that is not divisible by the number of planet gears (e.g., 41 sun gear teeth, 5 planet gears), as in the described embodiment. In alternative embodiments where the number of sun gear teeth may be fully divided by the number of planet gears (e.g., sun gear 40 teeth, 5 planet gears), there may be little or no dynamic benefit.

Each planet gear 32 is mounted to the remainder of the planet carrier 34 by pins 33 (also referred to as planet pins). The planet pins 33 are mounted on the carrier 34 such that, in the planetary arrangement, the planet pins move with the carrier 34 as the carrier rotates; or in a star arrangement, the planet pins and carrier 34 which rotates are held in place. The pin 33 may be referred to as an axle/support for the planet gear 32. The lower stiffness of the planet pins 33 relative to the higher stiffness of the planet pins 33 may reduce differential loads and increase the load sharing factor.

The pins 33 may be connected to the

shelves

34a, 34b in any desired manner. For example, in the described embodiment, the pin 33 is provided with a

soft connection

31a, 31b to each

shelf

34a, 34 b. The connections are described as soft because they are arranged to facilitate some movement of the pin 33, which may help improve load sharing. Such a soft connection may be formed by the

plates

34a, 34b themselves (e.g., with a cut-out of material to provide some movement of the plates), or by a portion of the pin or by a separate component. The

flexible connections

31a, 31b are classified as part of the pin 33 for evaluation of the stiffness described herein. The

flexible connections

33a, 33b are shown in fig. 16 and 17 only for the pins of interest, but those skilled in the art will appreciate that each pin 33 will have equivalent connections.

Radial bending stiffness of planet pin

The radial bending stiffness of the pin 33 may be measured in different ways, e.g. depending on the pin design. The

flexible connections

31a, 31b (if present) of the pin 33 are classified as part of the pin 33. For unengaged pins 33 as shown in fig. 16, each pin 33 may be considered to be rigidly mounted on one shelf 34b (shown by diagonal lines on shelf 34 b). The other end of each pin may be considered to be slidably mounted on the other shelf 34a so that the end of pin 33 can slide along plate 34a but cannot move away from plate 34 a. Then, a radial force F is applied on the pin at a position corresponding to the sliding plane, and the resultant radial displacement δ is measured at this axial position.

For a pin 33 having two

separate shaft portions

33a, 33b and a joint 33c (e.g. a ball joint) between the two shaft portions as shown in fig. 17, the pin 33 may alternatively be considered to be rigidly mounted on both

shelves

34a, 34b (shown by diagonal lines on

shelves

34a, 34 b). Then, a radial force F is applied on the pin 33 at a position corresponding to the axial center point of the pin, and the resultant radial displacement δ at that axial position is measured. In embodiments having a non-central joint 33c, the axial center point of the joint 33c may be selected in place of the axial center point of the pin 33.

In various embodiments, the pin 33 has a radial bending stiffness greater than or equal to 3.00 x 10⁸N/m, and optionally greater than or equal to 6.3X 10⁸N/m. In some embodiments, such as embodiments in which the fan diameter is in the range of 240cm to 280cm, the radial bending stiffness of the pin 33 axis may be greater than or equal to 6.3 x 10⁸N/m, and optionally may be greater than or equal to 6.7X 10⁸N/m (and optionally may be equal to 7.70X 10)⁸N/m). In some embodiments, such as embodiments in which the fan diameter is in the range of 330cm to 380cm, the radial bending stiffness of the pins 33 may be greater than or equal to 9.0 x 10⁸N/m, and optionally may be greater than or equal to 1.0X 10⁹N/m (and optionally may be equal to 1.54X 10)⁹N/m)。

In various embodiments, the pin 33 has a radial bending stiffness of 3.00 x 10⁸N/m to 3.00X 10⁹N/m, and optionally 6.3X 10⁸N/m to 2.5X 10⁹N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the radial bending stiffness of the pin 33 axis may be 6.3 x 10⁸N/m to 1.5X 10⁹In the range of N/m, the ratio of the carbon atoms,and optionally may be at 6.7X 10⁸N/m to 8.7X 10⁸In the range of N/m (and optionally may be equal to 7.70X 10)⁸N/m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the radial bending stiffness of the pins 33 may be 9.0 x 10⁸N/m to 3.0X 10⁹N/m, and optionally may be in the range of 1.0X 10⁹N/m to 2.0X 10⁹In the range of N/m (and optionally may be equal to 1.54 x 10)⁹N/m)。

Those skilled in the art will appreciate that the desired pin stiffness may be achieved in a variety of different ways, such as by appropriately adjusting one or more of the pin material (typically steel) and the gearbox and/or pin geometry. For a given material, the stiffness may be a function of the pin diameter, and for example whether the pin is solid or hollow. By way of further example, the gear ratio of the gearbox 30 may be adjusted and the planet gear size may be changed accordingly, for example allowing a larger diameter pin 33 for a larger planet gear 32, thereby achieving higher stiffness.

Planet pin tilt stiffness

Each pin 33 (including its

soft connection

31a, 31b, if present) is modeled as a free body rigidly mounted on one shelf 34b (as indicated by the diagonal lines on shelf 34b shown in fig. 18) to calculate the tilt stiffness. A moment is then applied at the axial centre point of the pin 33. The axis of the moment is perpendicular to the engine axis 9. Two points of the pin 33 were selected to measure the tilt stiffness: a center point at the axial position of the axial center point of the pin 33, and a rear point at the axial position of the rigid connection to the rear plate 34 b. As shown by oblique lines in fig. 17 and 18, the rear plate 34b is held rigid and is not rotated.

In response to an applied moment M (a counterclockwise moment in the illustrated example, but may be a clockwise moment in other examples), the pin 33 bends through an angle θ, resulting in an arc displacement δ of the center point (the point at which the moment is applied).

If the pin 33 is asymmetric, a second measurement can be made, modeling the pin 33 as a free body rigidly mounted on the other shelf 34 a. Two points of the pin 33 are chosen to make this measurement: a centre point at the axial position of the axial centre point of the pin 33, and a front point at the axial position of the rigid connection to the front plate 34 a. The front plate 34a is held rigid and non-rotating. The two tilt stiffness values may then be averaged.

Thus, as described above, an effective linear tilt stiffness may be defined for the pin 33. The radius r chosen for the definition of the effective linear tilt stiffness is the radius of a circle centred on a point on the surface of the plate 34b which is held rigid and passes through the original axial centre point of the pin 33 (the point where the moment is applied) and the same point after deformation. The arc displacement δ is equal to r θ.

The same method can be used for any design of the pin 33.

In various embodiments, the pin 33 has a tilt stiffness greater than or equal to 4.00 x 10⁶Nm/rad, and optionally greater than or equal to 8.7X 10⁶Nm/rad. In some embodiments, such as embodiments in which the fan diameter is in the range of 240cm to 280cm, the tilt stiffness of the pin 33 may be greater than or equal to 8.7 x 10⁶Nm/rad, and optionally may be greater than or equal to 9.8X 10⁶Nm/rad (and optionally may be equal to 1.02X 10)⁷Nm/rad). In some embodiments, such as embodiments having a fan diameter in the range of 330cm to 380cm, the pin 33 may have a tilt stiffness greater than or equal to 1.4 x 10⁷Nm/rad, and optionally may be greater than or equal to 2.5X 10⁷Nm/rad (and optionally may be equal to 3.14X 10)⁷Nm/rad)。

In various embodiments, the pin 33 has a tilt stiffness of 4.00 × 10⁶Nm/rad to 6.30X 10⁷Nm/rad, and optionally in the range of 8.7X 10⁶Nm/rad to 4.5X 10⁷Nm/rad. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the pin 33 may have a tilt stiffness of 8.7 x 10⁶Nm/rad to 2.1X 10⁷Nm/rad, and optionally may be in the range of 9.8X 10⁶Nm/rad to 1.9X 10⁷Nm/rad (and optionally may be equal to 1.02 × 10)⁷Nm/rad). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the pin 33 may have a tilt stiffness of 1.4 x 10⁷Nm/rad to 6.3X 10⁷Nm/rad, and optionally in the range of 2.5X 10⁷Nm/rad to 4.2X 10⁷Nm/rad (and optionally may be equal to 3.14 × 10)⁷Nm/rad)。

To obtain the desired pin tilt stiffness, a similar option to the planet pin radial bending stiffness described above may be applied.

Stiffness of shaft and support

Torsional stiffness of input shaft of gearbox

In the arrangement depicted, the gearbox input shaft 26a drives the sun gear 28. Thus, the gearbox input shaft 26a may be referred to as the sun input shaft 26 a. The gearbox input shaft 26a may be the sun input shaft 26a in a sun arrangement (as well as a planetary arrangement). Gearbox input shaft 26a may also be referred to as part of spindle 26, with forward portion 26a of spindle 26 providing input to gearbox 30.

Thus, the spindle 26 includes a gearbox input shaft 26a that rotates with the remainder of the spindle 26, but may have a different stiffness than the remainder of the spindle. In the arrangement described with respect to fig. 1 and 2, the spindle extends between the turbine 19 and the gearbox 30, thereby connecting the turbine 19 to the compressor 14, and connecting the turbine and compressor to the gearbox 30. The rear portion 26b of the spindle 26 extends between the turbine 19 and the compressor 14, connecting the turbine to the compressor. The forward portion 26a extends between the compressor 14 and the gearbox, connecting the turbine and compressor to the gearbox 30. This forward section is referred to as the gearbox input shaft because it provides torque to the gearbox 30. In the arrangement shown, there is a bearing 26c on the spindle 26 at or near the axial location where the rear portion 26b meets the gearbox input shaft 26 a.

In some gearboxes 30, the planet carrier 34 may be driven by the spindle 26, and more specifically, for example, the gearbox input shaft 26 a. In such embodiments, the gearbox input shaft 26a may not be the sun input shaft 26. However, this may make installation of the sun gear 28 more difficult.

In the arrangement depicted, as shown in fig. 19, the mandrel 26 is divided into two parts: a first portion 26a (gearbox input shaft) extending from the gearbox 30 and connected to the sun gear 28, and a second portion 26b extending rearwardly from the first portion and connected to the turbine 19. In the arrangement depicted, the first portion 26a is designed to have a lower stiffness than the second portion 26b, and thus, the gearbox input shaft 26a can provide a soft mount for the sun gear 28 while maintaining stiffness elsewhere in the engine 10. In the arrangement described, the second portion 26b is designed to be effectively rigid (compared to the rigidity of the first portion 26a) -thus the torsional rigidity of the spindle 26 can be effectively equal to that of its gearbox input shaft portion. The second portion 26b connecting the turbine and the compressor may be referred to as a turbine shaft 26 b. Turbine shaft 26b is arranged to transmit torsional loads to drive the compressor and gearbox 30, as well as to transmit compressor and turbine axial loads.

In alternative embodiments, the mandrel 26 may not be divided into sections of differing stiffness, but may have a constant stiffness. In alternative or additional embodiments, the mandrel 26 may be divided into more sections.

The spindle 26 is mounted using a bearing 26c, the bearing 26c being the first bearing on the spindle 26 axially downstream of the gear box 30. In the arrangement depicted, the bearing 26c is on the second portion 26b of the shaft 26, which in other embodiments may be on a different shaft portion, or on a single shaft portion. The stiffness of the gearbox input shaft 26a is measured in a state where the bearing 26c is kept rigid and the connection of the bearing 26c with the rest of the spindle 26b is considered rigid, so that only the stiffness of the first part 26a is taken into account. To determine the torsional stiffness, the gearbox input shaft 26a is considered free at the end where the torque τ is applied.

Gearbox input shaft torsional stiffness is a measure of the resistance of shaft 26a to applied torque τ, as shown in FIG. 20. Which can be described as the resistance of the shaft 26a to twisting or winding. The axis of the moment is parallel to the engine axis 9. A slash box 402 is shown at the location of the bearing 26c of the shaft 26a to indicate that the connection to the shaft 26 at the location of the bearing 26 c/bearing is considered rigid and non-rotating (as with a cantilever beam mount). Otherwise, the shaft 26a is considered to be a free body (not including sun-planet meshing stiffness).

A torque τ is applied to the shaft 26a (at a forward position, i.e., an axial midpoint of the sun gear 28) and causes a rotational deformation θ (e.g., torsion) along the length of the shaft 26 a. Theta is measured at the location where the torque is applied. As described above, the spindle 26 is held against rotation at the location of the bearing 26c such that the twist value increases from zero to θ along the length of the first shaft portion 26 a. The point on the shaft circumference at the forward position is moved by an angle theta, where theta is the angle measured in radians. r is the radius of the shaft 26 a. In embodiments where the radius of the gearbox input shaft 26 varies, the radius of the shaft 26a at the interface with the sun gear 28 may be used as the radius r to calculate the effective linear torsional stiffness (i.e., the radius at the forward end of the shaft for which torque is applied in the illustrated embodiment). Thus, as described above, an effective linear torsional stiffness may be defined for the gearbox input shaft 26 a.

In the illustrated embodiment, the axial midpoint location of the sun gear 28 is also at or near the forward end of the shaft 26. In an alternative embodiment, the shaft 26 may extend further forward of the sun gear 28; in such embodiments, the forward position for applying torque, force or moment is still considered to be the position of the axial midpoint of the sun gear 28.

In various embodiments, the torsional stiffness of the gearbox input shaft 26a is greater than or equal to 1.4 x 10⁶Nm/radian, and optionally greater than or equal to 1.6 x 10⁶Nm/radian. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the torsional stiffness of the gearbox input shaft may be greater than or equal to 1.4 x 10⁶Nm/radian or 1.8X 10⁶Nm/radian. In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the torsional stiffness of the gearbox input shaft may be greater than or equal to 3 x 10⁶Nm/radian or 5X 10⁶Nm/radian.

In various embodiments, the torsional stiffness of the gearbox input shaft 26a is at 1.4 × 10⁶Nm/radian to 2.5X 10⁸Nm/radian, and optionally 1.6X 10⁶Nm/radian to 2.5X 10⁷Nm/radian. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the torsional stiffness of the gearbox input shaft may be 1.4 x 10⁶Nm/radian to 2.0X 10⁷Nm/radian, and optionally may be in the range of 1.8 × 10⁶Nm/radian to 3X 10⁶Nm/radian (and optionally may be equal to 2.0 × 10)⁶Nm/radian). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the torsional stiffness of the gearbox input shaft may be 3 x 10⁶Nm/radian to 1 × 10⁸Nm/radian, and optionally may be in the range of 5 × 10⁶Nm/radian to 6X 10⁶Nm/radian (and optionally may be equal to 5.7 × 10)⁶Nm/radian).

In various embodiments, the effective linear torsional stiffness of the gearbox input shaft 26a is greater than or equal to 4.0 x 10⁸N/m, and optionally greater than or equal to 4.3X 10⁸N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the effective linear torsional stiffness of the gearbox input shaft may be greater than or equal to 4.0 x 10⁸N/m or 4.4X 10⁸N/m. In some embodiments, such as embodiments with fan diameters in the range of 330cm to 380cm, the effective linear torsional stiffness of the gearbox input shaft may be greater than or equal to 4.3 x 10⁸N/m or 6.8X 10⁸N/m。

In various embodiments, the effective linear torsional stiffness of the gearbox input shaft is 4.0 x 10⁸N/m to 3.0X 10¹⁰N/m, and optionally in the range of 4.3X 10⁸N/m to 9.0X 10⁹N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the effective linear torsional stiffness of the gearbox input shaft may be 4.0 x 10⁸N/m to 1.5X 10¹⁰N/m, and optionally 4.4X 10⁸N/m to 5.4X 10⁹In the range of N/m (and optionally may be equal to 4.9X 10)⁸N/m, and optionally equal to 4.92X 10⁸N/m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the effective linear torsional stiffness of the gearbox input shaft may be 4.3 x 10⁸N/m to 3.0X 10¹⁰N/m, and optionally may be in the range of 5.0X 10⁸N/m to 8.0X 10¹⁰In the range of N/m (and optionally may be equal to 6.8X 10)⁸N/m, and optionally equal to 6.84X 10⁸N/m)。

One or more of the gearbox input shaft 26a material, diameter, and structure (e.g., hollow or solid wall thickness) may be adjusted to achieve a desired range of stiffness.

Torsional stiffness of fan shaft

The fan shaft 36 is defined as the torque transfer component that extends from the output of the gearbox 30 to the fan input. It therefore includes some or all of any gearbox output shaft and fan input shaft that may be provided between these points. The fan shaft 36 is considered to extend between a fan input position and a gearbox output position for the purpose of defining the stiffness of the fan shaft, and includes all torque transfer components between these points. Thus, the fan shaft does not include any components of the gearbox 30 (e.g., the planet carrier 34 or any connecting plates coupled thereto) that transmit discrete forces, but not fan shaft torque. Thus, the gearbox output position may be defined as the connection point between the fan shaft 36 and the gearbox 30. The fan input location Y may be defined as the connection point between the fan axis 36 and the fan. The torsional stiffness of the fan shaft 36 is measured between the forward and aft ends of the fan shaft; the front end is the interface (Y) with the fan 23 and the rear end is the interface (X) with the gearbox 30.

Fan shaft torsional stiffness is a measure of the resistance of shaft 36 to applied torque τ, as shown in fig. 25. Which may be described as the resistance to twisting or winding of the shaft 36 pair. The axis of the moment is parallel to the engine axis 9.

Referring to fig. 23 and 25, where the gearbox 30 is a sun gearbox, the gearbox output position (X) 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 member 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 regarded as an axial center line (X) of the interface, as shown in fig. 25. The fan shaft 36 includes all of the torque transmitting components up to the connection point 702 with the ring gear 38. It therefore includes any flexible portion or link 704 of the fan shaft 36 that may be provided, as well as any connection 706 (e.g., a splined connection) therebetween.

In the case of a planetary configuration of the gearbox 30, the gearbox output position is also defined as the connection point between the fan shaft 36 and the gearbox 30. An example of such a configuration is shown in fig. 24, which shows a carrier comprising a front plate 34a and a rear plate 34b, with a plurality of pins 33 extending between the two plates, and on which the planet gears are mounted. The fan shaft 36 is connected to the front plate 34a via a splined connection 708. In such embodiments, the gearbox output position X is considered to be any point on the interface between the fan shaft 36 and the front plate 34 a. The front plate 34a is considered to transmit discrete forces rather than a single torque and is therefore considered to be part of the gear box 30 rather than the fan shaft. Fig. 24 shows only one example of one type of connection between the fan shaft and the planet carrier 34. In embodiments with different connection arrangements, the gearbox output position is still considered to be at the interface between the component that transmits torque (i.e., a portion of the fan shaft) and the component that transmits the discrete force (e.g., a portion of the gearbox). The splined connection 708 is just one example of a connection that may be formed between the fan shaft 36 and the gear box 30 (i.e., between the fan shaft and the front plate 34b in the presently described embodiment). In other embodiments, the interface forming the output position of the gearbox may be formed by, for example, a curvilinear connection, a bolted joint, or other toothed or mechanically fixed arrangement.

The fan input position Y is defined as a point on the fan axis 36 that is at the axial midpoint of the interface between the fan 23 and the fan axis 36. In the presently described embodiment, the fan 23 includes a support arm 23a arranged to connect the fan 23 to the fan shaft 36. The support arm 23a is connected to the fan shaft by a splined coupling 36a that extends along the length of a portion of the fan shaft 36. The fan input position is defined as the axial midpoint of the spline coupler, as shown by the Y-axis in fig. 25. The splined coupling shown in fig. 25 is only one example of a coupling that can interface between a fan and a fan shaft. In other embodiments, curvilinear connections, bolted joints, or other toothed or mechanically fixed arrangements, for example, may be used. The fan input location Y may be independent of the gearbox type.

The fan shaft 36 has a degree of flexibility that is characterized in part by its torsional stiffness. For the purpose of determining the torsional stiffness, the fan shaft 36 is considered free at the end impulse applied by the torque τ.

For the purpose of evaluating the torsional stiffness of the fan shaft 36, the slashed ring gear 38 in fig. 25 indicates that the ring gear 38 is considered to be rigid and non-rotating. The torque τ is applied to the shaft 36 at the fan input location Y and causes a rotational deformation θ (e.g., a twist) along the length of the shaft 36. The angle through which a point on the shaft circumference at the fan input location moves is θ, where θ is the angle measured in radians. The radius r of the fan shaft 36 can be used to determine the effective linear torsional stiffness. In embodiments where the radius of the fan shaft 36 varies (such as the described embodiments), the radius of the shaft 36 at the fan input location may be used as the radius r (i.e., the radius at the forward end of the shaft for the illustrated embodiment). Thus, as described above, an effective linear torsional stiffness may be defined for the fan shaft 36.

In various embodiments, the torsional stiffness of the fan shaft 36 is equal to or greater than 1.3 x 10⁷Nm/rad, and optionally equal to or greater than 1.4X 10⁷Nm/rad. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the torsional stiffness of the fan shaft 36 may be equal to or greater than 1.3 x 10⁷Nm/radian, and optionally may be equal to or greater than 1.4 x 10⁷Nm/radian (and optionally may be equal to 1.8 × 10)⁷Nm/radian). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the torsional stiffness of the fan shaft 36 may be equal to or greater than 2.5 x 10⁷Nm/radian, and optionally may be equal to or greater than 3.5 x 10⁷Nm/radian (and optionally may be equal to 5.2 × 10)⁷Nm/radian).

In various embodiments, the torsional stiffness of the fan shaft 36 is at 1.3 × 10⁷Nm/rad to 2.5X 10⁹Nm/rad, and optionally in the range of 1.4X 10⁷Nm/rad to 3.0X 10⁸Nm/rad. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the torsional stiffness of the fan shaft 36 may be in the range of 1.3 x 10⁷Nm/radian to 2.0X 10⁸Nm/radian, and optionally may be in the range of 1.3X 10⁷Nm/radian or 1.4X 10⁷Nm/radian to 2.3X 10⁷Nm/radian (and optionally may be equal to 1.8 × 10)⁷Nm/radian). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the torsional stiffness of the fan shaft 36 may be 2.5 x 10⁷Nm/radian to 2.5X 10⁹Nm/radian, and optionally may be in the range of 3.5 × 10⁷Nm/radian to 7.5 × 10⁷Nm/radian (and optionally may be equal to 5.2 × 10)⁷Nm/radian).

In various embodiments, the effective linear torsional stiffness of the fan shaft 36 may be greater than 1.2 x 10⁹N/m, and optionally greater than 1.35X 10⁹N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the effective linear torsional stiffness of the fan shaft 36 may be greater than 1.2 x 10⁹N/m, and optionally may be greater than 1.3X 10⁹Nm/radian (and optionally may be equal to 1.5 × 10)⁹N/m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the effective linear torsional stiffness of the fan shaft 36 may be greater than 1.5 x 10⁹N/m, and optionally may be greater than 1.8X 10⁹Nm/radian (and optionally may be equal to 2.1×10⁹N/m)。

In various embodiments, the effective linear torsional stiffness of the fan shaft 36 is 1.2 × 10⁹N/m to 2.0X 10¹⁰N/m, and optionally in the range of 1.35X 10⁹N/m to 1.0X 10¹⁰N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the effective linear torsional stiffness of the fan shaft 36 may be in the range of 1.2 x 10⁹N/m to 1.5X 10¹⁰N/m, and optionally may be in the range of 1.3X 10⁹N/m to 2.3X 10⁹Nm/radian (and optionally may be equal to 1.5 × 10)⁹N/m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the effective linear torsional stiffness of the fan shaft 36 may be 1.5 x 10⁹N/m to 2.0X 10¹⁰N/m, and optionally may be in the range of 1.8X 10⁹N/m to 3.5X 10⁹Nm/radian (and optionally may be equal to 2.1 × 10)⁹N/m)。

One or more of the fan shaft 36 material, diameter, and structure (e.g., hollow or solid wall thickness) may be adjusted to achieve a desired range of stiffness.

In various embodiments, including the arrangements described, for example with respect to fig. 21 and 22, the mounting structure of the fan shaft 36 includes a fan shaft support 504. The fan shaft support structure includes two bearings, a first bearing 506a and a second bearing 506b, via which the fan shaft support structure is coupled to the fan shaft 36. The

bearings

506a, 506b are spaced along the axial length of the fan shaft 36. In the depicted embodiment, both

bearings

506a, 506b are disposed at a location forward of the gearbox 30. In other embodiments, one of the two

bearings

506a, 506b for supporting the fan shaft 36 may be located at a position rearward of the gear box 30. In other embodiments, more than two bearings may be provided as part of the fan shaft support structure.

Gearbox support torsional stiffness

An exemplary embodiment of a gas turbine engine is shown in FIG. 22, which provides an enlarged view of the area of the engine core 11 around the gearbox 30. The same reference numerals are used for parts corresponding to those shown in fig. 1 to 3. In the arrangement shown in fig. 22, the gearbox 30 has a sun arrangement in which the ring gear 38 is coupled to the fan shaft 36 and the carrier 34 is held in a fixed position relative to the static structure 24 of the engine core (also referred to as a fixed support structure). As noted elsewhere herein, all of the features and characteristics described herein are applicable to sun and planetary gearboxes, unless explicitly stated otherwise.

The engine core 11 comprises a gearbox support 40 (corresponding to the connecting rod described with reference to fig. 2) arranged to at least partially support the gearbox 30 in a fixed position within the engine 10. The gearbox support is coupled at a first end to a fixed support structure 24 that extends across the core duct 20 carrying the core airflow a, as shown in fig. 22. In the presently described arrangement, the stationary support structure 24 is or includes an engine-part stator (ESS) that serves both as a structural component to provide a stationary mount for a core component, such as the gearbox support 40, and as guide vanes provided to direct the airflow from the fan 23. In other embodiments, the fixed support structure 24 may include struts that extend across the core airflow path and individual stator vanes provided to direct the airflow. In the present embodiment, 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 24 of the engine core. 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 24 of the engine core.

The gearbox support 40 is defined between its point of connection with the gearbox (e.g., planet carrier in the present embodiment with sun gearbox 30, or ring gear 38 in the planetary embodiment) and its point of connection with the fixed support structure 24. The gearbox support may be formed from any number of separate components providing a coupling between these two points. The gearbox support 40 is connected to the gearbox 30 and more particularly to a stationary gear or gear set, i.e. to the ring gear 38 of the planetary gearbox or to the planet carrier 34/planetary gear set 32 of the sun gearbox. The gearbox support 40 has a degree of flexibility. Gearbox support torsional stiffness is a measure of the resistance of the support 40 to an applied torque τ, as shown in FIG. 21. Which may be described as the resistance to twisting or winding of the support member 40. The axis of the moment is parallel to the engine axis 9. The cross-hatching of the fixed support structure 24 provides for an indication of the connection to the support 40 which is considered to be rigid and non-rotating.

For the sun gearbox 30, the torsional stiffness of the gearbox support 40 is defined between a circle 902 passing through the center of each planet gear 32 of the planetary gear set (i.e., passing through the longitudinal axis of each pin 33) and the interface to the fixed support structure 24 (which is considered fixed). Torsional loads are applied at the carrier 34 and are resisted at the fixed support structure 24.

For the planetary gearbox 30, the torsional stiffness of the gearbox support 40 is defined between the Pitch Circle Diameter (PCD) of the ring gear 38 and the interface to the fixed support structure 24 (which is considered fixed). Torsional loads are applied at the ring gear 38 and are resisted at the fixed support structure 24. A torque τ is applied to the teeth of the ring gear 38 and causes rotational deformation θ (e.g., twisting) of the support 40. The point on the PCD moves by an angle theta, where theta is the angle measured in radians. The radius r may be defined as the radius of the ring gear 38 (i.e., half of the PCD of the ring gear). Thus, as described above, the radius r ═ PCD/2 may be used to define an effective linear torsional stiffness for the gearbox support 40 of the planetary gearbox 30.

The pitch circle of the gear is an imaginary circle that rolls without slipping with the pitch circle of any other gear that meshes with the first gear. The pitch circle passes through the point of two gears meeting as the meshed gears rotate, the pitch circle of a gear typically passing through the mid-point of the tooth length of the gear. The PCD may be roughly estimated by taking the average of the diameter between the tips of the gear teeth and the diameter between the bases of the gear teeth. In various embodiments, the PCD of the ring gear 38 (which may also be considered the diameter of the gearbox 30) may be greater than or equal to 0.55m, and optionally greater than or equal to 0.57 m. In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the gearbox diameter may be greater than or equal to 0.75 m.

In various embodiments, the diameter of the gearbox 30 may be in the range of 0.55m to 1.2m, and optionally in the range of 0.57m to 1.0 m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the gearbox diameter may be in the range of 0.55m to 0.70m, and optionally may be in the range of 0.58m to 0.65m (and optionally may be equal to 0.61 m). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the gearbox diameter may be in the range of 0.75m to 1.0m, and optionally in the range of 0.8m to 0.9m (and optionally may be equal to 0.87 m).

Correspondingly, therefore, as described above, the radius r of the circle 902 passing through the longitudinal axis of each pin 33 on the carrier 34 may be used to define an effective linear torsional stiffness for the gearbox support 40 of the sun gearbox 30. The diameter of this circle 902 may be described as the PCD of a planetary gearset or pin PCD, providing r ═ PCD/2 as in the planetary gearbox example. The pin PCD may be in the range of 0.38m to 0.65m, for example equal to 0.4m or 0.55 m.

In various embodiments, the torsional stiffness of the gearbox support 40 is greater than or equal to 4.2 x 10⁷Nm/rad, and optionally greater than or equal to 4.8X 10⁷Nm/rad. In some embodiments, such as embodiments with fan diameters in the range of 240cm to 280cm, the torsional stiffness of the gearbox support 40 may be greater than or equal to 4.2 x 10⁷Nm/rad, and optionally may be greater than or equal to 5X 10⁷Nm/rad (and optionally may be equal to 6.1X 10⁷Nm/rad). In some embodiments, such as embodiments with fan diameters in the range of 330cm to 380cm, the torsional stiffness of the gearbox support 40 may be greater than or equal to 7.0 x 10⁷Nm/rad, and optionally may be greater than or equal to 9X 10⁷Nm/rad (and optionally may be equal to 1.8X 10⁸Nm/rad)。

In various embodiments, the torsional stiffness of the gearbox support 40 is 4.2 x 10⁷Nm/rad to 1.0X 10¹⁰Nm/rad, and optionally in the range of 4.8X 10⁷Nm/rad to 1.0X 10⁹Nm/rad. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the torsional stiffness of the gearbox support 40 may be 4.2 x 10⁷Nm/rad to 6.0X 10⁸Nm/rad, and optionally may be in the range of 5X 10⁷Nm/rad to 7X 10⁷Nm/rad (and optionally may be equal to 6.1X 10)⁷Nm/rad). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the torsional stiffness of the gearbox support 40 may be 7.0 x 10⁷Nm/rad to 1.0X 10¹⁰Nm/rad, and optionally may be in the range of 9X 10⁷Nm/rad to 4X 10⁸Nm/rad (and optionally may be equal to 1.8 × 10)⁸Nm/rad)。

In various embodiments, the effective linear torsional stiffness of the gearbox support 40 is greater than 7.1 x 10⁸N/m, and optionally greater than 8.4X 10⁸N/m. In some embodiments, such as embodiments with fan diameters in the range of 240cm to 280cm, the effective linear torsional stiffness of the gearbox support 40 may be greater than 7.1 x 10⁸N/m, and optionally may be greater than 8X 10⁸N/m (and optionally may be equal to 9.2X 10)⁸N/m). In some embodiments, such as embodiments with fan diameters in the range of 330cm to 380cm, the effective linear torsional stiffness of the gearbox support 40 may be greater than 9.0 x 10⁸N/m, and optionally may be greater than 9.6X 10⁸N/m (and optionally may be equal to 1.2X 10)⁹N/m)。

In various embodiments, the effective linear torsional stiffness of the gearbox support 40 is 7.1 × 10⁸N/m to 6.0X 10¹⁰N/m, and optionally 8.4X 10⁸N/m to 3.0X 10¹⁰N/m. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cmThe effective linear torsional stiffness of the gearbox support 40 may be in the range of 7.1 x 10⁸N/m to 5.0X 10¹⁰N/m, and optionally 8X 10⁸N/m to 1X 10⁹In the range of N/m (and optionally may be equal to 9.2X 10)⁸N/m). In some embodiments, such as embodiments with fan diameters in the range of 330cm to 380cm, the effective linear torsional stiffness of the gearbox support 40 may be 9.0 x 10⁸N/m to 6.0X 10¹⁰N/m, and optionally may be in the range of 9.0X 10⁸N/m to 2.0X 10⁹In the range of N/m (and optionally may be equal to 1.2X 10)⁹N/m)。

Those skilled in the art will appreciate that the stiffness of the gearbox support 40 may be defined in the same manner for embodiments having different epicyclic gearboxes, such as planetary gearboxes.

One or more of the gearbox support 40 geometry, material, and type of connection for connecting to the fixed support structure 24 may be appropriately selected or adjusted to achieve the desired stiffness.

The inventors have found that the specific ratio of the above defined parameters has a significant impact on gearbox performance. In particular, one, some or all of the following may be applicable to any implementation:

in various embodiments, the radial to torsional frame stiffness ratio is:

greater than or equal to 0.030, and optionally in the range of 0.030 to 2.0.

In various embodiments, the radial to torsional frame stiffness ratio is 3.0 x 10^-2To 2.0X 10⁰In the range of (i.e., 0.030 to 2.0), and optionally 6.0 x 10^-2To 1.0. In some embodiments, the radial to torsional frame stiffness ratio may be 6.0 x 10^-2To 3.0X 10^-1And optionally may be in the range of 0.18 to 0.19 (and optionally may be equal to 0.18). In thatIn some embodiments, the radial to torsional frame stiffness ratio may be in the range of 0.30 to 2.0. In alternative such embodiments, the radial to torsional frame stiffness ratio may be in the range of 0.14 to 0.8, and optionally may be in the range of 0.14 to 0.19 (and optionally may be equal to 0.14).

In various embodiments, the product of the components of the radial to torsional frame stiffness ratio, i.e., the radial bending stiffness of the planet carrier 34 multiplied by the effective linear torsional stiffness of the planet carrier 34, may be calculated. In various embodiments, the product value may be greater than or equal to 5.0 × 10¹⁸N²m^-2And optionally less than 1.3 × 10²⁴N²m^-2And optionally may be greater than or equal to 1.6 x 10¹⁹N²m^-2And optionally less than 1.3 × 10²²N²m^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 240cm to 280cm, the product value may be greater than or equal to 1.6 x 10¹⁹N²m^-2And optionally less than 1.3 × 10²²N²m^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 330cm to 380cm, the product value may be greater than or equal to 3.0 x 10¹⁹N²m^-2And optionally less than 1.3 × 10²³N²m^-2。

In various embodiments, the pitch to torsion frame stiffness ratio is:

less than or equal to 20, and optionally less than or equal to 7.3.

In various embodiments, the pitch to torsion frame stiffness ratio is 7.00 x 10^-1To 2.0X 10¹Is in the range of (i.e., 0.7 to 20), and optionally is in the range of 0.7 to 7.3. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the pitch to torsion frame stiffness ratio may be less than or equal to 8.0, whicheverOptionally in the range of 2.5 to 8.0, and further optionally may be in the range of 4 to 7 (and optionally may be equal to 5.60). In some embodiments, for example in embodiments where the fan diameter is in the range of 330cm to 380cm, the pitch to torsion frame stiffness ratio may be less than or equal to 7.9, optionally in the range of 1.5 to 7.9, and further optionally may be in the range of 1.8 to 5.2 (and optionally may be equal to 2.63).

In various embodiments, the product of the tilt and the component of the torsional frame stiffness ratio, i.e., the tilt stiffness of the planet carrier (34) multiplied by the torsional stiffness of the planet carrier (34), may be calculated. In various embodiments, the product value may be greater than or equal to 1.0 × 10¹⁷N²m²rad^-2And optionally less than 2.8 x 10²²N²m²rad^-2And optionally may be greater than or equal to 5.1 × 10¹⁷N²m²rad^-2And optionally less than 3.0 x 10²¹N²m²rad^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 240cm to 280cm, the product value may be greater than or equal to 5.1 x 10¹⁷N²m²rad^-2And optionally less than 3.0 x 10²⁰N²m²rad^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 330cm to 380cm, the product value may be greater than or equal to 1.0 x 10¹⁸N²m²rad^-2And optionally less than 3.1 × 10²¹N²m²rad^-2。

In various embodiments, the carrier to gearbox input shaft torsional stiffness ratio is:

greater than or equal to 7.0 x 10¹。

In various embodiments, the carrier to gearbox input shaft torsional stiffness ratio can be greater than or equal to 7.5 x 10¹. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the carrier to gearbox input shaft torsional stiffness ratio may be greater than or equal to 7.3 x 10¹And optionally may be greater than or equal to 9.5 x 10¹Or 14.0X 10¹(and optionally may be equal to 152). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the carrier to gearbox input shaft torsional stiffness ratio may be greater than or equal to 1.0 x 10²And optionally may be greater than or equal to 1.5 x 10²(and optionally may be equal to 2.0X 10²). In various embodiments, the carrier to gearbox input shaft torsional stiffness ratio can be greater than or equal to 1.4 x 10²And optionally at 1.4X 10²To 5.4X 10²Within the range of (1).

In various embodiments, the carrier to gearbox input shaft torsional stiffness ratio may be in the range of 7 x 10¹To 5X 10³(i.e., 70 to 5000), and optionally 7.5X 10¹To 3.0X 10³Within the range of (1). In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the carrier to gearbox input shaft torsional stiffness ratio may be 7.3 x 10¹To 1.0X 10³And optionally may be in the range of 9.5 x 10¹Or 14.0X 10¹To 3.0X 10²And optionally may be equal to 152). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the carrier to gearbox input shaft torsional stiffness ratio may be in the range of 1.0 x 10²To 5.0X 10³And optionally may be in the range of 1.5X 10²To 2.7X 10²And optionally may be equal to 2.0 x 10²)。

In various embodiments, the product of the components of the carrier to gearbox input shaft torsional stiffness ratio, i.e. the torsional stiffness of the carrier (34) times the torsional stiffness of the gearbox input shaft (26a), may be calculated. In various embodiments, the product value may be greater than or equal to 1.5 × 10¹⁴N²m²rad^-2And optionally less than 1.0 x 10¹⁷N²m²rad^-2And optionally may be greater than or equal to 2.2 x 10¹⁴N²m²rad^-2And optionally less than 5.0 x 10¹⁶N²m²rad^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 240cm to 280cm, the product value may be greater than or equal to 1.5 x 10¹⁴N²m²rad^-2And optionally less than 1.0 x 10¹⁶N²m²rad^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 330cm to 380cm, the product value may be greater than or equal to 3.0 x 10¹⁵N²m²rad^-2And optionally less than 1.0 x 10¹⁷N²m²rad^-2。

In various embodiments, the carrier to gearbox support torsional stiffness ratio is:

greater than or equal to 2.3.

In various embodiments, the carrier to gearbox support torsional stiffness ratio can be greater than or equal to 2.3, and optionally greater than or equal to 2.6. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the rack-to-gearbox support torsional stiffness ratio may be greater than or equal to 2.3, and optionally may be greater than or equal to 2.5 (and optionally may be equal to 4.8). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the rack-to-gearbox support torsional stiffness ratio may be greater than or equal to 3.5, and optionally may be greater than or equal to 4 (and optionally may be equal to 6.5). In various embodiments, the carrier to gearbox support torsional stiffness ratio may be greater than or equal to 4.4, and optionally in the range of 4.4 or 4.5 to 15.5.

In various embodiments, the carrier to gearbox support torsional stiffness ratio may be in the range of 2.3 x 10⁰To 3.0X 10²(i.e., 2.3 to 300), and optionally in the range of 2.6 to 50. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the rack-to-gearbox support torsional stiffness ratio may be in the range of 2.3 to 30, and optionally may be in the range of 2.5 to 5.5 or 4.3 to 5.5 (and optionally may be equal to 4.8). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the rack-to-gearbox support torsional stiffness ratio may be in the range of 3.5 to 300, and optionally may be in the range of 4 to 15 (and optionally may be equal to 6.5).

In various embodiments, the product of the components of the carrier to gearbox support torsional stiffness ratio, i.e., the torsional stiffness of the carrier 34 multiplied by the torsional stiffness of the gearbox support 40, may be calculated. In various embodiments, the product value may be greater than or equal to 5.0 × 10¹⁵N²m²rad^-2And optionally less than 1.0 x 10¹⁹N²m²rad^-2And optionally may be greater than or equal to 8.0 x 10¹⁵N²m²rad^-2And optionally less than 2.0 x 10¹⁸N²m²rad^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 240cm to 280cm, the product value may be greater than or equal to 5.0 x 10¹⁵N²m²rad^-2And optionally less than 1.2 x 10¹⁷N²m²rad^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 330cm to 380cm, the product value may be greater than or equal to 1.0 x 10¹⁷N²m²rad^-2And optionally less than 1.0 x 10¹⁹N²m²rad^-2。

In various embodiments, the ratio of rack to fan shaft stiffness is:

greater than or equal to 8.

In various embodiments, the ratio of shelf to fan shaft stiffness may be greater than or equal to 8.0 x 10⁰(i.e., 8.0), and optionally greater than or equal to 9. In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the ratio of shelf to fan shaft stiffness may be greater than or equal to 8, and optionally may be greater than or equal to 9 or 15.1 (and optionally may be equal to 16.6). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the shelf to fan shaft stiffness ratio can be greater than or equal to 12, and optionally can be greater than or equal to 15 or 18 (and optionally can be equal to 22.2). In various embodiments, the ratio of shelf to fan shaft stiffness may be greater than or equal to 1.50 x 10¹And optionally greater than or equal to 1.6 x 10¹(ii) a In such embodiments, the ratio of shelf to fan shaft stiffness may be less than 8.4 x 10¹。

In various embodiments, the ratio of shelf to fan shaft stiffness may be 8.0 x 10⁰To 1.1X 10³(i.e., 8.0 to 1100), and optionally 9 to 1.9 x 10²Within the range of (1). In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the ratio of shelf to fan shaft stiffness may be in the range of 8 to 5.0 x 10²And optionally may be in the range of 9 to 40 or 15 or 16 to 40 (and optionally may be equal to 17). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the ratio of shelf to fan shaft stiffness may be in the range of 12 to 1.1 x 10³And optionally may be in the range of 15 or 18 to 55 (and optionally may be equal to 22).

In various embodiments, the product of the components of the ratio of the carrier to the fan shaft stiffness, i.e., the torsional stiffness of the carrier 34 multiplied by the torsional stiffness of the fan shaft 36, may be calculated. In various embodiments, the product value may be greater than or equal to 1.5 × 10¹⁵N²m²rad^-2And optionally less than 3.0 x 10¹⁸N²m²rad^-2And optionally may be greater than or equal to 2.0 x 10¹⁵N²m²rad^-2And optionally less than 7.0 x 10¹⁷N²m²rad^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 240cm to 280cm, the product value may be greater than or equal to 1.5 x 10¹⁵N²m²rad^-2And optionally less than 1.5 x 10¹⁷N²m²rad^-2. In some embodiments, such as embodiments in which the fan diameter is in the range of 330cm to 380cm, the product value may be greater than or equal to 9.0 x 10¹⁵N²m²rad^-2And optionally less than 3.0 x 10¹⁸N²m²rad^-2。

In various embodiments, the first shelf to pin stiffness ratio is:

greater than or equal to 10, and optionally greater than or equal to 15. In some embodiments, such as embodiments having a fan diameter in the range of 240cm to 280cm, the first rack to pin stiffness ratio may be greater than or equal to 1.5 x 10¹And optionally may be equal to 16.3. In some embodiments, such as embodiments having a fan diameter in the range of 330cm to 380cm, the first rack to pin stiffness ratio may be greater than or equal to 1.6 x 10¹And optionally may be greater than or equal to 16.5 (and optionally may be equal to 18.7).

In various embodiments, the first shelf to pin stiffness ratio is at 1.0 x 10¹To 4.0X 10¹In the range of (i.e., 10 to 40), and optionally in the range of 1.5 x 10¹To 3.0X 10¹Within the range of (1). In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the first rack to pin stiffness ratio may be in the range of 1.5 x 10¹To 2.5X 10¹And optionally may be in the range of 15 to 19 (and optionally may be equal to 16.3). In some embodiments, for example, the fan diameter is from 330cm to 380cmIn embodiments within the scope, the first frame to pin stiffness ratio may be 1.6 x 10¹To 3.5X 10¹And optionally may be in the range of 16 or 16.5 to 20 (and optionally may be equal to 18.7).

The product of the effective linear torsional stiffness of the planet carrier 34 and the radial bending stiffness of each pin 33 may be greater than or equal to 2.1 x 10¹⁸N²m^-2And optionally greater than or equal to 5.8 x 10¹⁸N²m^-2. In some embodiments, such as embodiments having a fan diameter in the range of 240cm to 280cm, the product of the effective linear torsional stiffness of the planet carrier 34 and the radial bending stiffness of each pin 33 may be greater than or equal to 5.3 x 10¹⁸N²m^-2. In some embodiments, such as embodiments having a fan diameter in the range of 330cm to 380cm, the product of the effective linear torsional stiffness of the planet carrier 34 and the radial bending stiffness of each pin 33 may be greater than or equal to 1.2 x 10¹⁹N²m^-2。

The product of the effective linear torsional stiffness of the planet carrier 34 and the radial bending stiffness of each pin 33 may be in the range of 2.1 x 10¹⁸N²m^-2To 3.6X 10²⁰N²m^-2And optionally in the range of 5.8 × 10¹⁸N²m^-2To 1.7X 10²⁰N²m^-2Within the range of (1). In some embodiments, such as embodiments with fan diameters in the range of 240cm to 280cm, the product of the effective linear torsional stiffness of the planet carrier 34 and the radial bending stiffness of each pin 33 may be at 5.3 x 10¹⁸N²m^-2To 4.0X 10¹⁹N²m^-2Within the range of (1). In some embodiments, such as embodiments with fan diameters in the range of 330cm to 380cm, the product of the effective linear torsional stiffness of the planet carrier 34 and the radial bending stiffness of each pin 33 may be 1.2 x 10¹⁹N²m^-2To 1.7X 10²⁰N²m^-2Within the range of (1).

In various embodiments, the second shelf to pin stiffness ratio is:

greater than or equal to 2.4 x 10¹And optionally greater than or equal to 3.4 x 10¹. In some embodiments, such as embodiments having a fan diameter in the range of 240cm to 280cm, the second rack to pin stiffness ratio may be greater than or equal to 3.4 x 10¹Optionally greater than or equal to 36, and optionally may be equal to 47.5. In some embodiments, such as embodiments having a fan diameter in the range of 330cm to 380cm, the second rack to pin stiffness ratio may be greater than or equal to 4.0 x 10¹And optionally may be greater than or equal to 45 (and optionally may be equal to 69.1).

In various embodiments, the second shelf to pin stiffness ratio is 2.4 x 10¹To 1.8X 10²In the range of (i.e., 24 to 180), and optionally in the range of 3.4 x 10¹To 1.4X 10²Within the range of (1). In some embodiments, such as embodiments where the fan diameter is in the range of 240cm to 280cm, the second rack to pin stiffness ratio may be in the range of 3.4 x 10¹To 1.2X 10²And optionally may be in the range of 36 to 58 (and optionally may be equal to 47.5). In some embodiments, such as embodiments where the fan diameter is in the range of 330cm to 380cm, the second rack to pin stiffness ratio may be 4.0 x 10¹To 1.8X 10²And optionally may be in the range of 45 to 95 (and optionally may be equal to 69.1).

The product of the torsional rigidity of the carrier 34 and the tilting rigidity of each pin 33 may be greater than or equal to 1.0 × 10¹⁵N²m²rad^-2And optionally greater than or equal to 2.5 x 10¹⁵N²m²rad^-2. In some embodiments, such as embodiments having a fan diameter in the range of 240cm to 280cm, the product of the torsional stiffness of the planet carrier 34 and the tilt stiffness of each pin 33 may be greater than or equal to 2.5 x 10¹⁵N²m²rad^-2. In some implementationsIn an aspect, such as an embodiment having a fan diameter in the range of 330cm to 380cm, the product of the torsional stiffness of the planet carrier 34 and the tilt stiffness of each pin 33 may be greater than or equal to 1.4 x 10¹⁶N²m²rad^-2。

The product of the torsional stiffness of the planet carrier 34 and the tilting stiffness of each pin 33 may be 1.0 x 10¹⁵N²m²rad^-2To 4.7X 10¹⁷N²m²rad^-2And optionally in the range of 2.5X 10¹⁵N²m²rad^-2To 2.0X 10¹⁷N²m²rad^-2Within the range of (1). In some embodiments, such as embodiments with fan diameters in the range of 240cm to 280cm, the product of the torsional stiffness of the planet carrier 34 and the tilt stiffness of each pin 33 may be 2.5 x 10¹⁵N²m²rad^-2To 3.0X 10¹⁶N²m²rad^-2Within the range of (1). In some embodiments, such as embodiments with fan diameters in the range of 330cm to 380cm, the product of the torsional stiffness of the planet carrier 34 and the tilt stiffness of each pin 33 may be 1.4 x 10¹⁶N²m²rad^-2To 4.7X 10¹⁷N²m²rad^-2Within the range of (1).

Fig. 26 shows how the stiffness as defined herein can be measured. Fig. 26 shows a graph of the displacement δ resulting from the application of a load L (e.g., force, moment, or torque) applied to the component whose stiffness is being measured. From zero to L_PThere is a non-linear region where displacement is caused by movement of the component when loaded (or relative movement of individual parts of the component) rather than deformation of the component (e.g., movement within gaps between parts). Above L_QThe elastic limit of the component has been exceeded and the applied load no longer causes elastic deformation but may plastically deform or the component fails. Between points P and Q, the applied load and the resulting displacement have a linear relationship. Can be measured by measuring the linear region between points P and QTo determine the stiffness as defined herein (wherein stiffness is the inverse of the gradient). The gradient of as large a region as possible of the linear region can be found by making a measurement by providing a larger displacement to increase the accuracy of the measurement. For example, L may be increased by applying a quantity equal to or slightly greater than L_PAnd is equal to or slightly less than L_QTo find the gradient. L may be estimated prior to testing based on material properties_PAnd L_QIn order to apply the appropriate load. Although the displacement is referred to as δ in this specification, those skilled in the art will understand that equivalent principles will apply to either linear or angular displacement.

Unless otherwise noted, stiffness as defined herein is for the corresponding component when the engine is off (i.e., at zero speed/on the table). Stiffness typically does not vary significantly over the operating range of the engine; thus, the stiffness of an aircraft using an engine at cruise conditions (which are as defined elsewhere herein) may be the same as the stiffness without the engine. However, where stiffness varies over the operating range of the engine, stiffness as defined herein is to be understood as the value at which the engine is at room temperature and not in motion.

The present disclosure also relates to a method 1300 of operating the gas turbine engine 10 on board an aircraft. A method 1300 is shown in fig. 27. The method 1300 includes starting and operating 1302 the engine 10 (e.g., taxiing, taking off, and climbing on a runway, if appropriate) to achieve cruise conditions.

Once the cruise condition is reached, the method 1300 then includes operating 1304 the gas turbine engine 10 described in one or more embodiments elsewhere herein to provide propulsion at the cruise condition.

The gas turbine engine 10 is such and/or is operated such that any or all of the parameters or ratios defined herein are within specified ranges.

The torque on the spindle 26 may be referred to as the input torque, as this is the torque input to the gearbox 30. The unit of torque is force x distance, which can be expressed in newton meters (n.m), and is defined in a general manner as understood by those skilled in the art.

The torque provided by the turbine 19 to the spindle (i.e. the torque on the spindle) at cruise conditions may be greater than or equal to 10,000Nm, and optionally greater than or equal to 11,000 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range 240cm to 280cm, the torque on the spindle 26 at cruise conditions may be greater than or equal to 10000Nm (and optionally may be equal to 12760 Nm). In some embodiments, for example in embodiments in which the fan diameter is in the range of 330cm to 380cm, the torque on the spindle 26 at cruise conditions may be greater than or equal to 25000Nm, and optionally greater than or equal to 30000Nm (and optionally may be equal to 33970Nm, or 34000 Nm).

The torque on the spindle at cruise conditions may be in the range 10,000Nm to 50,000Nm, and optionally in the range 11,000Nm to 45,000 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range 240cm to 280cm, the torque on the spindle 26 at cruise conditions may be in the range 10,000Nm to 15,000Nm, and optionally in the range 11,000Nm to 14,000Nm (and optionally may be equal to 12,760 Nm). In some embodiments, for example in embodiments in which the fan diameter is in the range of 330cm to 380cm, the torque on the spindle 26 at cruise conditions may be in the range 25000Nm to 50000Nm, and optionally in the range 30000Nm to 40000Nm (and optionally may be equal to 33970Nm, or 34000 Nm).

The torque on the spindle 26 may be greater than or equal to 28,000Nm, and optionally greater than or equal to 30,000Nm, at Maximum Takeoff (MTO) conditions. In some embodiments, such as embodiments in which the fan diameter is in the range of 240cm to 280cm, the torque on the mandrel 26 under MTO conditions may be greater than or equal to 28,000Nm, and optionally greater than or equal to 35,000Nm (and optionally may be equal to 36,300 Nm). In some embodiments, for example embodiments in which the fan diameter is in the range of 330cm to 380cm, the torque on the mandrel 26 under MTO conditions may be greater than or equal to 70000Nm, and optionally greater than or equal to 80000Nm or 82000Nm (and optionally may be equal to 87000Nm or 87100 Nm).

The torque on the spindle 26 may be in the range of 28,000Nm to 135,000Nm, and optionally in the range of 30,000Nm to 110,000Nm, at Maximum Takeoff (MTO) conditions. In some embodiments, for example in embodiments where the fan diameter is in the range of 240cm to 280cm, the torque on the mandrel 26 under MTO conditions may be in the range of 28000Nm to 50000Nm, and optionally in the range of 35000Nm to 38000Nm (and optionally may be equal to 36000Nm or 36300 Nm). In some embodiments, for example in embodiments where the fan diameter is in the range of 330cm to 380cm, the torque on the mandrel 26 under MTO conditions may be in the range of 70000Nm to 135000Nm, and optionally in the range of 80000Nm to 90000Nm or 82000Nm to 92000Nm (and optionally may be equal to 87000Nm or 87100 Nm).

It is to be understood that the present invention is not limited to the above-described embodiments, and various modifications and improvements may be made without departing from the concept described herein. Any feature may be used alone or in combination with any other feature or features unless mutually exclusive, and the 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) 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 an input from a gearbox input shaft portion of 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 comprising a sun gear (28), a plurality of planet gears (32), a ring gear (38) and a planet carrier (34) on which the planet gears (32) are mounted; and

a gearbox support (40) arranged to support the gearbox (30) in a fixed position within the gas turbine engine (10) and having a torsional stiffness, and characterized by a rack-to-gearbox support torsional stiffness ratio of:

greater than or equal to 2.3.

2. The gas turbine engine (10) of claim 1 wherein the carrier to gearbox support torsional stiffness ratio is greater than or equal to 2.6.

3. The gas turbine engine (10) of claim 2 wherein the carrier to gearbox support torsional stiffness ratio is in the range of 2.6 to 50.

4. The gas turbine engine (10) of claim 1 or claim 2, wherein the torsional stiffness of the planet carrier (34) is greater than or equal to 1.60 x 10⁸Nm/rad。

5. The gas turbine engine (10) of claim 4 wherein the torsional stiffness of the planet carrier (34) is at 1.60 x 10⁸Nm/rad to 1.00X 10¹¹Nm/rad.

6. The gas turbine engine (10) of claim 4 wherein the torsional stiffness of the planet carrier (34) is at 2.7 x 10⁸Nm/rad to 1X 10¹⁰Nm/rad.

7. The gas turbine engine (10) of claim 1 or claim 2, wherein:

(i) the torsional stiffness of the gearbox input shaft is greater than or equal to 1.4 x 10⁶Nm/radian; and/or

(ii) The torsional stiffness of the gearbox support (40) is greater than or equal to 4.2 x 10⁷Nm/radian.

8. The gas turbine engine (10) of claim 7, wherein:

(i) the torsional stiffness of the gearbox input shaft is greater than or equal to 1.6 x 10⁶Nm/radian; and/or

(ii) The torsional stiffness of the gearbox support (40) is greater than or equal to 4.8 x 10⁷Nm/radian.

9. The gas turbine engine (10) of claim 1 or claim 2, wherein the fan (23) has a fan diameter in the range of 240cm to 280 cm.

10. The gas turbine engine (10) of claim 1 or claim 2, wherein the fan (23) has a fan diameter in the range of 330cm to 380cm and the carrier to gearbox support torsional stiffness ratio is greater than or equal to 3.5.

11. The gas turbine engine (10) of claim 1 or claim 2, wherein the torsional stiffness of the planet carrier (34) multiplied by the torsional stiffness of the gearbox support (40) is greater than or equal to 5.0 x 10¹⁵N²m²rad^-2。

12. The gas turbine engine (10) of claim 1 or claim 2, wherein the torsional stiffness of the planet carrier (34) multiplied by the torsional stiffness of the gearbox support (40) is less than 1.0 x 10¹⁹N²m²rad^-2。

13. The gas turbine engine (10) of 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 further comprising a second turbine (17), a second compressor (15) and a second spindle (27) connecting the second turbine to the second compressor; and is

The second turbine, the second compressor and the second spindle are arranged to rotate at a higher rotational speed than the first spindle.

14. The gas turbine engine (10) of claim 1 or claim 2, wherein the planet carrier (34) comprises a front plate (34a) and a rear plate (34b) and pins (33) extending between the front plate (34a) and the rear plate (34b), each pin (33) being arranged to have a planet gear (32) mounted thereon.

15. The gas turbine engine (10) of claim 14, wherein the planet carrier (34) further comprises a lug (34c) extending between the front plate (34a) and the back plate (34b), the lug (34c) being arranged to pass between adjacent planet gears (32).

16. The gas turbine engine (10) of claim 1 or claim 2, wherein:

the gearbox (30) comprising an odd number of planet gears (32); and/or

The fan (23) has a fan diameter greater than 240cm and less than or equal to 380 cm; and/or

The gearbox (30) is a sun gearbox, wherein the planet carrier (34) does not rotate in use; and/or

The pitch circle diameter of a pin (33) on which the planetary gear (32) is mounted is in the range of 0.38m to 0.65 m.

17. The gas turbine engine (10) of claim 16 wherein:

the gearbox (30) comprises 3, 5 or 7 planet gears (32); and/or

The diameter of the fan is more than 300cm and less than or equal to 380 cm; and/or

The pitch circle diameter of the pin (33) on which the planetary gear (32) is mounted is equal to 0.4m or 0.55 m.

18. The gas turbine engine (10) according to claim 1 or claim 2, wherein the spindle (26) provides a soft mount for the sun gear (28) such that some movement of the sun gear (28) is facilitated, and wherein the spindle (26) comprises a stiffer portion (26b) and a less stiff portion (26a), the less stiff portion (26a) being arranged to be located between the stiffer portion (26b) and the sun gear (28) and being arranged to provide or facilitate the soft mount of the sun gear (28).

19. The gas turbine engine (10) of claim 1 or claim 2, wherein:

(i) the gear ratio of the gearbox (30) is in the range of 3.2 to 4.5; and/or

(ii) The gas turbine engine (10) has a specific thrust at cruise of 70NKg^-1s to 90NKg^-1s is in the range of; and/or

(iii) The bypass ratio at cruise is in the range of 12.5 to 18.

20. The gas turbine engine (10) of claim 19 wherein:

(i) the gear ratio of the gearbox (30) is in the range of 3.3 to 4.0; and/or

(ii) The bypass ratio at cruising is in the range of 13 to 16.

21. A propeller for an aircraft, comprising:

a fan (23) comprising a plurality of fan blades;

a gear case (30); and

a power unit for driving the fan (23) via the gearbox (30);

wherein the gearbox (30) is arranged to receive input from a gearbox input shaft portion of a spindle (26) driven by the power unit and to output drive to the fan (23) 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 planet gears (32), a ring gear (38) and a planet carrier (34) on which the planet gears (32) are mounted; and

a gearbox support (40) arranged to support the gearbox (30) in a fixed position within the thruster and having a torsional stiffness, and characterized by a frame to gearbox support torsional stiffness ratio of:

greater than or equal to 2.3.