The present disclosure relates to the mounting of a core shaft within a gas turbine engine for an aircraft, and in particular to bearing positioning and how such a shaft may be arranged and mounted so as to manage vibrational and resonance effects.

As used herein, a range “of value X to value Y”, “from value X to value Y” or “between value X and value Y”, or the likes, denotes an inclusive range; including the bounding values of X and Y. As used herein, the term “axial plane” denotes a plane extending along the length of an engine, parallel to and containing an axial centreline of the engine, and the term “radial plane” denotes a plane extending perpendicular to the axial centreline of the engine, so including all radial lines at the axial position of the radial plane. Axial planes may also be referred to as longitudinal planes, as they extend along the length of the engine. A radial distance or an axial distance is therefore a distance in a radial or axial plane, respectively.

According to a first aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, and wherein the forwardmost rearward bearing has a bearing stiffness in the range of 30 kN/mm to 100 kN/mm, the bearing stiffness being defined by the radial displacement caused by the application of a radial force at the axial centrepoint of the bearing.

The inventor appreciated that controlling the bearing stiffness to lie within the specified range may allow or facilitate management of vibrational modes, so potentially reducing damage to the engine in use caused by whirl mode displacements of the core shaft.

The core shaft has a length (L) between the forward bearing and the rearmost rearward bearing and a minor span (S) between the rearward bearings. The bearings may be arranged such that a length ratio (S/L) of the minor span to the core shaft length is equal to or less than 0.14, or equal to or less than 0.13, or equal to or less than 0.12. The length ratio may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio may be in the range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13, optionally in the range from 0.06 to 0.13, and further optionally in the range from 0.08 to 0.13.

The core shaft may have a running speed range with a lower bound of 1500 rpm and an upper bound of 6200 rpm.

The bearing stiffness of the forwardmost rearward bearing may be equal to or around 50 kN/mm.

The gas turbine engine may further comprise a stationary supporting structure and a first bearing support structure. The forwardmost rearward bearing may be mounted to the stationary supporting structure by the first bearing support structure. The first bearing support structure may be attached to the stationary supporting structure at a first position located axially rearward of the forwardmost rearward bearing.

In such embodiments, the first bearing support structure may comprise a plurality of connecting members, which may be spaced circumferentially around the engine axis, connecting the bearing to the stationary support structure. The first bearing support structure may be described as a spring bar-type support structure.

The first bearing support structure may comprise an outer race of the forwardmost rearward bearing.

The gas turbine engine of embodiments with a first bearing support structure may further comprise a second bearing support structure. The second bearing support structure may be mounted to the stationary supporting structure, optionally at a second position located forward of, and at a larger radial distance from the engine axis than, the first position. The second bearing support structure may be connected to the first bearing support structure by a squeeze film damper in the region of the forwardmost rearward bearing.

A stiffness ratio of the bearing stiffness at the forwardmost rearward bearing to the minor span may be in the range from 0.08 to 0.5 kN/mm2, optionally in the range from 0.08 to 0.40 kN/mm2, optionally in the range from 0.08 to 0.30 kN/mm2, optionally in the range from 0.08 to 0.20 kN/mm2, optionally in the range from 0.09 to 0.40 kN/mm2, optionally in the range from 0.15 to 0.50 kN/mm2, optionally in the range from 0.15 to 0.40 kN/mm2, and further optionally in the range from 0.15 to 0.30 kN/mm2.

The length of the core shaft (L) may be in the range from 1800 mm to 2900 mm, optionally from 2000 mm to 2900 mm, further optionally from 2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The length of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. In such embodiments, the minor span may be in the range from 250 mm to 350 mm, optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The length of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. The minor span may be in the range from 250 mm to 350 mm. The running speed of the core shaft may be in the range from 1500 rpm to 6200 rpm; and/or a diameter of the fan may be in the range from 330 cm to 380 cm and the gear ratio of the gearbox is in the range from 3.1 to 3.8. The length, minor span and/or running speed may be selected such that no primary resonance of the core shaft lies within the running range of the engine.

The rearward bearings may be positioned axially level with or rearward of a leading edge of a lowest pressure turbine blade of the turbine at the root of the blade.

The rearward bearings may be positioned axially level with or rearward of a trailing edge of a turbine blade of a third set of turbine blades from the front of the turbine, at the root of the blade, wherein the turbine comprises four sets of turbine blades.

The lowest pressure turbine of the engine has a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, co.

A first blade to bearing ratio of:

the⁢minor⁢span⁢(S)mr⁢ω2⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set⁢19⁢c,19⁢d)

may have a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

A second blade to bearing ratio of:

the⁢minor⁢span⁢(S)m×r⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set⁢19⁢c,19⁢d)

may have a value in the range from 0.8 to 6.0 kg−1, optionally from 0.8 to 5.0 kg−1, optionally from 0.8 to 4.0 kg−1, optionally from 0.8 to 3.0 kg−1, optionally from 0.8 to 2.0 kg−1.

The turbine has a length between the leading edge of the forwardmost turbine blade of the turbine and a trailing edge of the rearmost turbine blade of the turbine. A minor span to turbine length ratio may be equal to or less than 1.05.

The minor span to turbine length ratio may be equal to or less than 1.00, optionally equal to or less than 0.95. The minor span to turbine length ratio may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example the minor span to turbine length ratio may be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to 0.95 in the range from 0.85 to 0.95.

According to a second aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The turbine has a turbine length defined as the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the turbine at its trailing edge. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, with a minor span (S) defined as the distance between the two rearward bearings. A minor span to turbine length ratio of:

minor⁢spanturbine⁢length

is equal to or less than 1.05.

The inventor appreciated that keeping the minor span to turbine length ratio within this range, and more generally smaller than that in known aircraft, may help to reduce or avoid deleterious whirl modes in operation.

The minor span to turbine length ratio may be equal to or less than 1.00, optionally equal to or less than 0.95. The minor span to turbine length ratio may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example the minor span to turbine length ratio may be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to 0.95 in the range from 0.85 to 0.95.

The inventor appreciated that the increased length of a core shaft when a gas turbine engine is scaled up may lead to a resonance frequency of the core shaft lying in or near the engine running range. Simply scaling up a known engine may therefore lead to increased risks of resonance-induced damage—the longer core shaft may be problematic. The inventor discovered that arranging the minor span to be approximately equal to, and optionally slightly smaller than, the turbine length (of the lowest pressure turbine, in embodiments with more than one turbine), and more specifically within the claimed range, may facilitate avoidance of deleterious whirl modes, and may reduce damage to the engine in use.

The turbine may comprise a total of four sets of turbine blades. In such embodiments, both of the two rearward bearings may be located downstream of the trailing edge of a turbine blade of the third set of turbine blades from the front of the turbine, at the root of the blade.

The two rearward bearings may be located downstream of the leading edge of the lowest pressure (most downstream) turbine blade of the turbine at the root of the blade.

The length of the core shaft (L) may be in the range from 1800 mm to 2900 mm, optionally from 2000 mm to 2900 mm, further optionally from 2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The core shaft may have a running speed range with a lower bound of 1500 rpm and an upper bound of 6200 rpm. The upper bound on the core shaft running speed range may be the upper bound on maximum take-off (MTO) running speed. The lower bound on the running speed of the core shaft may be the minimum running speed under ground idle conditions.

The core shaft has a length, L, between the forwardmost bearing and the rearmost bearing and a distance (the minor span, S) between the two rearward bearings. The bearings may be arranged such that a length ratio (S/L) of the distance between the two rearward bearings to the core shaft length is equal to or less than 0.14, or equal to or less than 0.13, or equal to or less than 0.12. The length ratio, S/L, of the distance between the two rearward bearings (S) to the core shaft length (L) may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio, S/L, of the distance between the two rearward bearings (S) to the core shaft length (L) may be in the range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13, optionally in the range from 0.06 to 0.13, and optionally in the range from 0.08 to 0.13.

The forwardmost bearing of the rearward bearings may have a bearing stiffness in the range of 30 kN/mm to 100 kN/mm. A stiffness ratio of the bearing stiffness at the forwardmost rearward bearing divided by the distance between the two rearward bearings may be in the range from 0.08 to 0.5 kN/mm2, optionally in the range from 0.08 to 0.40 kN/mm2, optionally in the range from 0.08 to 0.30 kN/mm2, optionally in the range from 0.08 to 0.20 kN/mm2, optionally in the range from 0.09 to 0.40 kN/mm2, optionally in the range from 0.15 to 0.50 kN/mm2, optionally in the range from 0.15 to 0.40 kN/mm2, and further optionally in the range from 0.15 to 0.30 kN/mm2.

The length (L) of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. The distance (S) between the two rearward bearings may be in the range from 250 mm to 350 mm, optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The gas length of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm; and the minor span may be in the range from 250 mm to 350 mm, optionally from 260 mm to 350 mm. The running speed of the core shaft may be in the range from 1500 rpm to 6200 rpm; and/or a diameter of the fan may be in the range from 330 cm to 380 cm and the gear ratio of the gearbox may be in the range from 3.1 to 3.8. The length, minor span and/or running speed may be selected such that no primary resonance of the core shaft lies within the running range of the engine.

The lowest pressure turbine of the engine has a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, co.

A first blade to bearing ratio of:

the⁢minor⁢span⁢(S)mr⁢ω2⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

may have a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

A second blade to bearing ratio of:

the⁢minor⁢span⁢(S)m×r⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

may have a value in the range from 0.8 to 6.0 kg−1, optionally from 0.8 to 5.0 kg−1, optionally from 0.8 to 4.0 kg−1, optionally from 0.8 to 3.0 kg−1, optionally from 0.8 to 2.0 kg−1.

According to a third aspect there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine, and comprises turbine blades, and the compressor is the lowest pressure compressor of the engine. The engine additionally comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising two rearward bearings located downstream of the leading edge of the lowest pressure turbine blades of the turbine at the root of the blades. The two rearward bearings may therefore be described as being located downstream of the leading edge of the last/rearmost turbine blade.

The inventor appreciated that the core shaft generally moves least (in the radial direction), and is most level (parallel to the engine axis) at the axial position of the bearings—whilst whirl modes and other bends or displacements may occur between bearings, the bearings serve to limit radial shaft movement. The inventor appreciated that careful control of shaft length and bearing position may therefore allow whirl modes of the engine to be managed, so reducing the risk of damage to the engine.

The inventor additionally discovered that positioning the bearings nearer the bigger and larger turbine stages, towards the rear of the turbine, provides improved turbine support as shaft movements relative to the turbine position may have more of a deleterious effect on these larger turbine stages.

The gas turbine engine may further comprise a disc arranged to support the lowest pressure turbine blades of the turbine. The two rearward bearings may be located downstream of a centreline of the disc.

The length of the core shaft may be in the range from 1800 to 2900 mm, optionally in the range from 2000 to 2900 mm, optionally in the range from 2300 to 2800 mm, and further optionally in the range from 2400 to 2750 mm.

The core shaft may have a running speed range with a lower bound of 1500 rpm and an upper bound of 6200 rpm.

The upper bound on the core shaft running speed range may be the upper bound on maximum take-off (MTO) running speed. The lower bound on the running speed of the core shaft may be the minimum running speed under ground idle conditions.

The core shaft has a length, L, between the forwardmost bearing and the rearmost bearing, and a distance, S, between the two rearward bearings. The bearings may be arranged such that a length ratio, S/L, of the distance between the two rearward bearings (S) to the core shaft length (L) may be equal to or less than 0.14, or equal to or less than 0.13, or equal to or less than 0.12. The length ratio, S/L, of the distance between the two rearward bearings (S) to the core shaft length (L) may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio, S/L, of the distance between the two rearward bearings (S) to the core shaft length (L) may be in the range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13, optionally in the range from 0.06 to 0.13, and optionally in the range from 0.08 to 0.13.

The length (L) of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. In such embodiments, the distance between the two rearward bearings (S) may be in the range from 250 mm to 350 mm, or optionally from 260 mm and 350 mm.

According to a fourth aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine, and comprises four sets of turbine blades, and the compressor is the lowest pressure compressor of the engine. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising two rearward bearings located downstream of the trailing edge of a turbine blade of the third set of turbine blades from the front of the turbine, at the root of the blade.

One or more of the following features may apply for a gas turbine engine of the third and/or fourth aspect.

The forwardmost bearing of the rearward bearings may have a bearing stiffness in the range of 30 kN/mm to 100 kN/mm.

A stiffness ratio of the stiffness at the forwardmost rearward bearing to the distance between the two rearward bearings (S) may be in the range from 0.08 to 0.5 kN/mm2, optionally in the range from 0.08 to 0.40 kN/mm2, optionally in the range from 0.08 to 0.30 kN/mm2, optionally in the range from 0.08 to 0.20 kN/mm2, optionally in the range from 0.09 to 0.40 kN/mm2, optionally in the range from 0.15 to 0.50 kN/mm2, optionally in the range from 0.15 to 0.40 kN/mm2, and further optionally in the range from 0.15 to 0.30 kN/mm2.

The length (L) of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. In such embodiments, the distance between the two rearward bearings (S) may be in the range from 250 mm to 350 mm, or optionally from 260 mm and 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The length of the core shaft (L) may be in the range from 1800 mm to 2900 mm, optionally from 2000 mm to 2900 mm, further optionally from 2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

The distance between the two rearward bearings (S), which may be referred to as the minor span, may be in the range from 250 mm to 350 mm. In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The running speed of the core shaft may be in the range from 1500 rpm to 6200 rpm. Additionally or alternatively, the diameter of the fan may be in the range from 330 cm to 380 cm. The gear ratio of the gearbox may be in the range from 3.1 to 3.8.

The length, minor span and/or running speed may be selected such that no primary resonance of the core shaft lies within the running range of the engine.

The lowest pressure turbine of the engine has a lowest pressure set of blades. Each blade of the lowest pressure set of blades has a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, co. A minor span (S) is defined, as mentioned above, as the axial distance between the two rearward bearings.

A first blade to bearing ratio of:

the⁢minor⁢span⁢(S)mr⁢ω2⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

may have a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

Additionally or alternatively, a second blade to bearing ratio of:

the⁢minor⁢span⁢(S)m×r⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

may have a value in the range from 0.8 to 6.0 kg−1, optionally from 0.8 to 5.0 kg−1, optionally from 0.8 to 4.0 kg−1, optionally from 0.8 to 3.0 kg−1, optionally from 0.8 to 2.0 kg−1.

The turbine has a length between the leading edge of the forwardmost turbine blade of the turbine and a trailing edge of the rearmost turbine blade of the turbine. A minor span to turbine length ratio (i.e. S divided by the turbine length) may be equal to or less than 1.05, optionally equal to or less than 1.00, optionally equal to or less than 0.95. The minor span to turbine length ratio may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example the minor span to turbine length ratio may be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to 0.95.

The turbine may comprise a total of four sets of turbine blades. In an alternative embodiment, the turbine may comprise a total of three sets of turbine blades.

According to a fifth aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The core shaft has a running speed range with a lower bound of 1500 rpm and an upper bound of 6200 rpm. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings. The core shaft has a length (L) between the forward bearing and the rearmost rearward bearing, and a minor span (S) between the rearward bearings. The bearings are arranged such that a length ratio of the minor span to the core shaft length (S/L) is equal to or less than 0.14.

The length ratio may be equal to or less than 0.13, or equal to or less than 0.12. The length ratio may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio may be in the range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13, optionally in the range from 0.06 to 0.13, and further optionally in the range from 0.08 to 0.13.

The inventor appreciated that the increased length of a core shaft when a gas turbine engine is scaled up may lead to a resonance frequency of the core shaft lying in or near the engine running range. Simply scaling up a known engine may therefore lead to increased risks of resonance-induced damage—the longer core shaft may be problematic.

However, making the core shaft still longer—extending rearward of a first rearward bearing and to a second rearward bearing, with the spacing between the first and second rearward bearings being within a set range with respect to the total core shaft length—was found to increase stiffness of the core shaft and to move the resonant frequency away from the engine running range in some embodiments.

The upper bound on the core shaft running speed range may be the upper bound on maximum take-off running speed.

The lower bound on the core shaft running speed may be the minimum running speed under ground idle conditions.

The running speed of the core shaft under cruise conditions may be in the range from 5400 to 5700 rpm, and optionally in the range from 5500 to 5600 rpm.

The running speed of the core shaft under maximum take-off (MTO) conditions may be in the range from 5800 to 6200 rpm, and optionally in the range from 5900 to 6100 rpm.

The length of the core shaft (L) may be in the range from 1800 mm to 2900 mm, optionally from 2000 mm to 2900 mm, further optionally from 2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The forwardmost bearing of the rearward bearings may have a bearing stiffness in the range of 30 kN/mm to 100 kN/mm. A stiffness ratio of the stiffness at the forwardmost rearward bearing to the minor span may be in the range from 0.08 to 0.5 kN/mm2, optionally in the range from 0.08 to 0.40 kN/mm2, optionally in the range from 0.08 to 0.30 kN/mm2, optionally in the range from 0.08 to 0.20 kN/mm2, optionally in the range from 0.09 to 0.40 kN/mm2, optionally in the range from 0.15 to 0.50 kN/mm2, optionally in the range from 0.15 to 0.40 kN/mm2, and further optionally in the range from 0.15 to 0.30 kN/mm2.

The length (L) of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. The minor span (S) may be in the range from 250 mm to 350 mm, optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

In some embodiments, the length of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. The minor span may be in the range from 250 mm to 350 mm, optionally from 260 mm to 350 mm. The running speed of the core shaft may be in the range from 1500 rpm to 6200 rpm; and/or a diameter of the fan may be in the range from 330 cm to 380 cm and the gear ratio of the gearbox may be in the range from 3.1 to 3.8. In such embodiments, the length, minor span and/or running speed may be selected such that no primary resonance of the core shaft (lies within the running range of the engine.

The rearward bearings may be positioned axially level with or rearward of:(i) a leading edge of a lowest pressure turbine blade of the turbine at the root of the blade; and/or(ii) a trailing edge of a turbine blade of a third set of turbine blades from the front of the turbine, at the root of the blade, wherein the turbine comprises four sets of turbine blades.

The inventor appreciated that positioning the rearward bearings nearer the bigger and larger turbine stages, towards the rear of the turbine, provides improved turbine support as shaft movements relative to the turbine position may have more of a deleterious effect on these larger turbine stages.

The lowest pressure turbine of the engine has a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, co. A minor span (S) is defined as the axial distance between the two rearward bearings, as stated above.

A first blade to bearing ratio of:

the⁢minor⁢span⁢(S)mr⁢ω2⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

may have a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

Additionally or alternatively, a second blade to bearing ratio of:

the⁢minor⁢span⁢(S)m×r⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

may have a value in the range from 0.8 to 6.0 kg−1, optionally from 0.8 to 5.0 kg−1, optionally from 0.8 to 4.0 kg−1, optionally from 0.8 to 3.0 kg−1, optionally from 0.8 to 2.0 kg−1.

The turbine has a length between the leading edge of the forwardmost turbine blade of the turbine and a trailing edge of the rearmost turbine blade of the turbine. A minor span to turbine length ratio may be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to 0.95.

According to a sixth aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine, the core shaft has a running speed in the range from 1500 rpm to 6200 rpm, and the compressor is the lowest pressure compressor of the engine. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, the core shaft having a length (L) between the forward bearing and the rearmost rearward bearing in the range from 1800 mm to 2900 mm or 2750 mm, and a minor span (S) between the two rearward bearings in the range from 250 mm to 350 mm.

As a result, there may be no primary resonance of the core shaft between the forward bearing and the forwardmost rearward bearings within the running speed range of the core shaft.

The length (L) and minor span (S) may be selected as appropriate for the desired running speed range to avoid such primary resonances.

In various embodiments the length of the core shaft (L) may be in the range from 2000 mm to 2900 mm, further optionally from 2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The inventor appreciated that the increased length of a core shaft when a gas turbine engine is scaled up may lead to a resonance frequency of the core shaft lying in or near the engine running range. Simply scaling up a known engine may therefore lead to increased risks of resonance-induced damage—the longer core shaft may be problematic. The inventor discovered that selecting the length, L, and minor span, S, as appropriate for a given running speed range may facilitate avoidance of deleterious whirl modes, and may reduce damage to the engine in use.

The lower bound of 1500 rpm on the core shaft running speed may be the minimum running speed under ground idle conditions and/or the upper bound of 6200 rpm on the core shaft running speed may be the upper bound on maximum take-off running speed.

According to a seventh aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine, and the compressor is the lowest pressure compressor of the engine. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades and having a fan diameter in the range from 330 cm to 380 cm; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio in the range from 3.1 to 3.8.

The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, the core shaft having a length between the forward bearing and the rearmost rearward bearing in the range from 1800 to 2900 mm or 2750 mm, and a minor span between the two rearward bearings in the range from 250 mm to 350 mm, such that there is no primary resonance of the core shaft within a running speed range of the core shaft.

The skilled person would appreciate that the seventh aspect shares an inventive concept with the sixth aspect, as the combination of fan diameter and gear ratio is linked to core shaft running speed. For a particular design of aircraft, a given combination of fan diameter and gear ratio may be used to infer an intended core shaft running speed range.

In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The core shaft running speed range may be from 1500 rpm to 6200 rpm.

In the sixth or seventh aspects, one or more of the following features may be present:

The turbine has a turbine length defined between the leading edge of its most upstream blades and a trailing edge of its most downstream blades. A minor span to turbine length ratio of:

minor⁢spanturbine⁢length

may be equal to or less than 1.05. The minor span to turbine length ratio may be equal to or less than 1.00, optionally equal to or less than 0.95. The minor span to turbine length ratio may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example the minor span to turbine length ratio may be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to 0.95 in the range from 0.85 to 0.95.

The turbine may comprise four sets of turbine blades. The two rearward bearings may both be located downstream of the trailing edge of a turbine blade of the third set of turbine blades from the front of the turbine, at the root of the blade. The turbine may comprise a total of four sets of turbine blades.

The two rearward bearings may be located downstream of the leading edge of the lowest pressure (most downstream) turbine blade of the turbine at the root of the blade.

The length of the core shaft (L) may be in the range from 1800 mm to 2900 mm, optionally from 2000 mm to 2900 mm, further optionally from 2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

A length ratio (S/L) of the minor span between the two rearward bearings to the core shaft length may be equal to or less than 0.14.

The length ratio may be equal to or less than 0.13, or equal to or less than 0.12. The length ratio may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio may be in the range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13, optionally in the range from 0.06 to 0.13, and further optionally in the range from 0.08 to 0.13

The forwardmost bearing of the rearward bearings may have a bearing stiffness in the range of 30 kN/mm to 100 kN/mm.

A stiffness ratio of the bearing stiffness at the forwardmost rearward bearing to the distance between the two rearward bearings may be in the range from 0.08 to 0.5 kN/mm2, optionally in the range from 0.08 to 0.40 kN/mm2, optionally in the range from 0.08 to 0.30 kN/mm2, optionally in the range from 0.08 to 0.20 kN/mm2, optionally in the range from 0.09 to 0.40 kN/mm2, optionally in the range from 0.15 to 0.50 kN/mm2, optionally in the range from 0.15 to 0.40 kN/mm2, and further optionally in the range from 0.15 to 0.30 kN/mm2.

The lowest pressure turbine of the engine has a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, co.

A first blade to bearing ratio of:

the⁢minor⁢span⁢(S)mr⁢ω2⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

may have a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

A second blade to bearing ratio of:

the⁢minor⁢span⁢(S)m×r⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

may have a value in the range from 0.8 to 6.0 kg−1, optionally from 0.8 to 5.0 kg−1, optionally from 0.8 to 4.0 kg−1, optionally from 0.8 to 3.0 kg−1, optionally from 0.8 to 2.0 kg−1.

According to an eighth aspect, there is provided a method of operation of a gas turbine engine for an aircraft. The gas turbine engine comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine. The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, the core shaft having a length (L) between the forward bearing and the rearmost rearward bearing in the range from 1800 mm to 2900 mm, and a minor span (S) between the two rearward bearings in the range from 250 mm to 350 mm. The engine further comprises 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 core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft

The method comprises operating the engine such that the core shaft has a running speed in the range from 1500 rpm to 6200 rpm, and wherein there is no primary resonance of the core shaft within the running speed range of the core shaft.

The length of the core shaft (L) may be in the range from 1800 mm to 2750 mm, or from 2000 to 2750, or from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The engine used to implement the method may be as described in the sixth and/or seventh aspects.

According to a ninth aspect, there is provided a method of designing a gas turbine engine for an aircraft. The engine comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine, the core shaft has a running speed in the range from 1500 rpm to 6200 rpm, and the compressor is the lowest pressure compressor of the engine. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, the core shaft having a length (L) between the forward bearing and the rearmost rearward bearing in the range from 1800 mm to 2900 mm.

The method comprises:selecting positions for the forward bearing and the forwardmost bearing of the rearward bearings; andlengthening the core shaft rearward of the forwardmost bearing of the rearward bearings such that a minor span (S) defined between the two rearward bearings is in the range from 250 mm to 350 mm, and there is no primary resonance of the core shaft between the forward bearing and the forwardmost rearward bearing within the running speed range of the core shaft.

The length of the core shaft (L) may be in the range from 1800 mm to 2750 mm, or from 2000 to 2750, or from 2400 mm to 2750 mm.

In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The engine designed by the method of the ninth aspect may be the engine of the sixth and/or seventh aspect.

According to a tenth aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the compressor is the lowest pressure compressor of the engine and the turbine is the lowest pressure turbine of the engine and has a lowest pressure set of blades each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, co. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, with a minor span (S) being defined as the axial distance between the two rearward bearings. A first blade to bearing ratio of:

the⁢minor⁢span⁢(S)mr⁢ω2⁢(for⁢a⁢blade⁢of⁢the⁢lowest⁢pressure⁢set)

has a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2.

The inventor appreciated that the engine should not be linearly scaled up with a force increase (mrω2providing a measure of force), but rather that the minor span length (S) should be increased as little as possible so as to relatively reduce engine length and weight, so allowing the efficiency gains to be increased by avoiding the additional weight, and to avoid the development of unwanted whirl modes within the minor span. Whilst conventional wisdom suggests that a larger minor span is desirable to improve reaction of forces from the low pressure turbine, the inventor found that the risk of introducing whirl modes, and the introduction of more length and weight, counterbalanced the force reaction benefits and that the first blade to bearing ratio should therefore be maintained within the specified range.

The first blade to bearing ratio may be in the range from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 4.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

A second blade to bearing ratio of:

the⁢⁢minor⁢⁢span⁢⁢(S)m×r⁢⁢(for⁢⁢a⁢⁢blade⁢⁢of⁢⁢the⁢⁢lowest⁢⁢pressure⁢⁢set)

may have a value in the range from 0.8 to 6.0 kg−1, optionally from 0.8 to 5.0 kg−1, optionally from 0.8 to 4.0 kg−1, optionally from 0.8 to 3.0 kg−1, optionally from 0.8 to 2.0 kg−1.

According to an eleventh aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine and has a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and the compressor is the lowest pressure compressor of the engine. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, with a minor span (S) being defined as the axial distance between the two rearward bearings. A second blade to bearing ratio of:

the⁢⁢minor⁢⁢span⁢⁢(S)m×r⁢⁢(for⁢⁢a⁢⁢blade⁢⁢of⁢⁢the⁢⁢lowest⁢⁢pressure⁢⁢set)

has a value in the range from 0.8 to 6.0 kg−1.

The second blade to bearing ratio may be in the range from 0.8 to 5.0 kg−1, optionally from 0.8 to 4.0 kg−1, optionally from 0.8 to 3.0 kg−1, optionally from 0.8 to 2.0 kg−1.

Each blade of the lowest pressure set of blades has an angular velocity at cruise, co, and a first blade to bearing ratio of:

the⁢⁢minor⁢⁢span⁢⁢(S)mr⁢⁢ω2⁢⁢(for⁢⁢a⁢⁢blade⁢⁢of⁢⁢the⁢⁢lowest⁢⁢pressure⁢⁢set⁢⁢19⁢c,19⁢d)

may have a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

In the tenth or eleventh aspects, one or more of the following features may be present:

The minor span, S, may be in the range from 250 mm to 350 mm. In various embodiments, the minor span may be greater than or equal to any of 250 mm, 255 mm, 260 mm and 265 mm. In various embodiments, the minor span may be smaller than or equal to any of 350 mm, 345 mm, 340 mm, or 335 mm.

The length of the core shaft (L) may be in the range from 1800 mm to 2900 mm, optionally from 2000 mm to 2900 mm, further optionally from 2300 mm to 2800 mm, and further optionally from 2400 mm to 2750 mm.

The value of blade mass, m, multiplied by blade radius at mid-height, r, may be in the range from 180 to 280 kg·mm

The gearbox may have a gear ratio greater than 3, and optionally in the range from 3.1 to 3.8.

The core shaft may have a running speed in the range from 1500 rpm to 6200 rpm.

The running speed of the core shaft at cruise may be in the range of 5400-5700 rpm, and optionally of 5500-5600 rpm, at cruise.

The running speed of the core shaft under maximum take-off (MTO) conditions may be in the range of 5800-6200 rpm, and optionally of 5900-6100 rpm.

A length ratio (S/L) of the minor span between the two rearward bearings to the core shaft length may be equal to or less than 0.14, or equal to or less than 0.13, or equal to or less than 0.12. The length ratio S/L may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio S/L may be in the range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13, optionally in the range from 0.06 to 0.13, and optionally in the range from 0.08 to 0.13.

The mass, m, of a blade of the lowest pressure set of blades may be in the range from 0.2 to 0.6 kg.

The radius, r, of a blade of the lowest pressure set of blades may be in the range from 400 to 600 mm.

Each blade of the lowest pressure set of blades may have an angular velocity at cruise, co in the range from 560 to 600 rad·s−1.

The length of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. The minor span may be in the range from 250 mm to 350 mm, and optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The turbine has a turbine length defined between the leading edge of its most upstream blades and a trailing edge of its most downstream blades. A minor span to turbine length ratio of:

minor⁢⁢spanturbine⁢⁢length

may be equal to or less than 1.05, optionally equal to or less than 1.00, optionally equal to or less than 0.95. The minor span to turbine length ratio may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example the minor span to turbine length ratio may be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to 0.95.

The rearward bearings may be positioned axially level with or rearward of a leading edge of a lowest pressure turbine blade of the turbine at the root of the blade.

The rearward bearings may be positioned axially level with or rearward of a trailing edge of a turbine blade of a third set of turbine blades from the front of the turbine, at the root of the blade. In such embodiments, the turbine may comprise four sets of turbine blades, and optionally may have a total of four sets of turbine blades.

The forwardmost bearing of the rearward bearings may have a bearing stiffness in the range of 30 kN/mm to 100 kN/mm.

A stiffness ratio of the bearing stiffness at the forwardmost rearward bearing to the distance between the two rearward bearings may be in the range from 0.08 to 0.5 kN/mm2, optionally in the range from 0.08 to 0.40 kN/mm2, optionally in the range from 0.08 to 0.30 kN/mm2, optionally in the range from 0.08 to 0.20 kN/mm2, optionally in the range from 0.09 to 0.40 kN/mm2, optionally in the range from 0.15 to 0.50 kN/mm2, optionally in the range from 0.15 to 0.40 kN/mm2, and further optionally in the range from 0.15 to 0.30 kN/mm2.

According to a twelfth aspect, there is provided a method of operation of a gas turbine engine for an aircraft, the engine comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the compressor is the lowest pressure compressor of the engine and the turbine is the lowest pressure turbine of the engine and has a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, co. The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, with a minor span (S) being defined as the axial distance between the two rearward bearings. The engine further comprises 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 core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The method comprises operating the engine such that a first blade to bearing ratio of:

the⁢⁢minor⁢⁢span⁢⁢(S)mr⁢⁢ω2⁢⁢(for⁢⁢a⁢⁢blade⁢⁢of⁢⁢the⁢⁢lowest⁢⁢pressure⁢⁢set)

has a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2at cruise.

In various embodiments the method may comprise operating the engine such that the first blade to bearing ratio may be in the range from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

The engine used to implement the method of the twelfth aspect may be as described in the tenth and/or eleventh aspect.

According to a thirteenth aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

The engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, the distance between the two rearward bearings being defined as the minor span, S. The forwardmost rearward bearing has a bearing stiffness defined by the radial displacement caused by the application of a radial force at the axial centrepoint of the bearing. A stiffness ratio of the bearing stiffness at the forwardmost rearward bearing to the minor span is in the range from 0.08 to 0.5 kN/mm2.

The bearing stiffness ratio may be in the range from 0.09 to 0.40 kN/mm2, optionally in the range from 0.08 to 0.30 kN/mm2, optionally in the range from 0.08 to 0.20 kN/mm2, optionally in the range from 0.09 to 0.40 kN/mm2, optionally in the range from 0.15 to 0.50 kN/mm2, optionally in the range from 0.15 to 0.40 kN/mm2, and further optionally in the range from 0.15 to 0.30 kN/mm2.

The inventor appreciated that controlling the bearing stiffness and minor span such that the ratio of the two lies within the specified range may allow or facilitate management of vibrational modes, so potentially reducing damage to the engine in use caused by whirl mode displacements of the core shaft.

The core shaft has a length, L, between the forward bearing and the rearmost rearward bearing. The bearings may be arranged such that a length ratio (S/L) of the minor span to the core shaft length is equal to or less than 0.14, or equal to or less than 0.13, or equal to or less than 0.12. The length ratio S/L may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio S/L may be in the range from 0.05 to 0.14, optionally in the range from 0.05 to 0.13, optionally in the range from 0.06 to 0.13, and optionally in the range from 0.08 to 0.13.

The core shaft may have a running speed range with a lower bound of 1500 rpm and an upper bound of 6200 rpm.

The bearing stiffness of the forwardmost rearward bearing may be in the range of 30 kN/mm to 100 kN/mm. Optionally, bearing stiffness of the forwardmost rearward bearing may be at least substantially equal to 50 kN/mm.

The gas turbine engine may further comprise a stationary supporting structure and a first bearing support structure. The forwardmost rearward bearing may be mounted to the stationary supporting structure by the first bearing support structure. The first bearing support structure may be attached to the stationary supporting structure at a first position located axially rearward of the forwardmost rearward bearing.

In such embodiments, the first bearing support structure may comprise a plurality of connecting members, which may be spaced circumferentially around the engine axis. The connecting members may connect the forwardmost rearward bearing to the stationary support structure.

Additionally or alternatively, in such embodiments, the first bearing support structure may comprise an outer race of the forwardmost rearward bearing.

In embodiments with a first bearing support structure, the engine may further comprise a second bearing support structure. The second bearing support structure may be mounted to the stationary supporting structure, optionally at a second position located forward of, and at a larger radial distance from the engine axis than, the first position. The second bearing support structure may be connected to the first bearing support structure by a squeeze film damper in the region of the forwardmost rearward bearing.

The length, L, of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm. The minor span may be in the range from 250 mm to 350 mm, optionally from 260 mm to 350 mm.

The fan may have a fan diameter in the range from 330 cm to 380 cm.

The length, L, of the core shaft may be in the range from 1800 mm to 2900 mm or 2750 mm and the minor span in the range from 250 mm to 350 mm, optionally from 260 mm to 350 mm. The running speed of the core shaft may be in the range from 1500 rpm to 6200 rpm; and/or a diameter of the fan may be in the range from 330 cm to 380 cm and the gear ratio of the gearbox in the range from 3.1 to 3.8. The length, minor span and/or running speed may be selected such that no primary resonance of the core shaft lies within the running range of the engine.

The rearward bearings may be positioned axially level with or rearward of a leading edge of a lowest pressure turbine blade of the turbine at the root of the blade.

The rearward bearings may be positioned axially level with or rearward of a trailing edge of a turbine blade of a third set of turbine blades from the front of the turbine, at the root of the blade. In such embodiments, the turbine may comprise four sets of turbine blades.

The lowest pressure turbine of the engine may have a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, co. The minor span (S) is defined as the axial distance between the two rearward bearings, as described elsewhere.

A first blade to bearing ratio of:

the⁢⁢minor⁢⁢span⁢⁢(S)mr⁢⁢ω2⁢⁢(for⁢⁢a⁢⁢blade⁢⁢of⁢⁢the⁢⁢lowest⁢⁢pressure⁢⁢set)

may have a value in the range from 2.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 4.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 5.0×10−6to 7.5×10−6kg−1·rad−2·s2, optionally from 2.0×10−6to 6.5×10−6kg−1·rad−2·s2, optionally from 3.0×10−6to 6.5×10−6kg−1·rad−2·s2, and further optionally from 4.5×10−6to 6.5×10−6kg−1·rad−2·s2.

A second blade to bearing ratio of:

the⁢⁢minor⁢⁢span⁢⁢(S)m×r⁢⁢(for⁢⁢a⁢⁢blade⁢⁢of⁢⁢the⁢⁢lowest⁢⁢pressure⁢⁢set⁢⁢19⁢c,19⁢d)

may have a value in the range from 0.8 to 6.0 kg−1, optionally from 0.8 to 5.0 kg−1, optionally from 0.8 to 4.0 kg−1, optionally from 0.8 to 3.0 kg−1, optionally from 0.8 to 2.0 kg−1.

The turbine has a length between the leading edge of the forwardmost turbine blade of the turbine and a trailing edge of the rearmost turbine blade of the turbine. A minor span to turbine length ratio (minor span divided by turbine length) may be equal to or less than 1.05, optionally equal to or less than 1.00, optionally equal to or less than 0.95. The minor span to turbine length ratio may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example the minor span to turbine length ratio may be in the range from 0.70 to 1.05, optionally from 0.70 to 1.00, optionally from 0.70 to 0.95, optionally from 0.80 to 1.05, optionally from 0.80 to 1.00, optionally from 0.80 to 0.95, optionally from 0.85 to 1.05, optionally from 0.85 to 1.00, and further optionally from 0.85 to 0.95.

In any of the aspects described above, one or more of the following features may be present:

The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and an interconnecting shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

The engine may further comprise a tail bearing housing located rearward of the turbine. The tail bearing housing may comprise two bearing discs; each bearing disc may be arranged to support one of the two rearward bearings. In alternative embodiments, the tail bearing housing may comprise a single bearing disc, the bearing disc being arranged to support one of the two rearward bearings (for example the rearmost bearing).

In embodiments with one or more bearing discs, one or more of the bearing discs may be oriented at least substantially perpendicular to the engine axis (i.e. at least substantially in a radial plane through the engine)

In the various aspects and embodiments described herein, the engine running range may be defined as the range of rotation rates of the core shaft during standard operation of the engine (e.g. during ground idle, take-off, climb and cruise), and may be measured in rotations per minute (rpm). In this context “standard operation of the engine” may exclude transient periods on start-up and shut down, e.g. as core shaft rotation rate increases from zero to the ground idle rotation rate. The engine running range includes the ground idle speed, cruise speed, and maximum take-off (MTO) speed. The core shaft rotation rate may be greater than or equal to the ground idle rotation rate throughout standard operation of the engine. The core shaft rotation rate during standard operation of the engine may also be referred to as the core shaft running speed.

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the 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 that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

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

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a “planetary” or “star” 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 of from 3 to 4.2, or 3.2 to 3.8, for example on the order of 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. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a “star” gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence 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 (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise 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-washed location, 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 on the order of) 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 in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of:

220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades13on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utipis the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as 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 exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg−1s, 105 Nkg−1s, 100 Nkg−1s, 95 Nkg−1s, 90 Nkg−1s, 85 Nkg−1s or 80 Nkg−1s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg−1s to 100 Nkg−1s, or 85 Nkg−1s to 95 Nkg−1s. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330 kN to 420 kN, for example 350 kN to 400 kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K. The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

As used herein, a maximum take-off (MTO) condition has the conventional meaning. Maximum take-off conditions may be defined as operating the engine at International Standard Atmosphere (ISA) sea level pressure and temperature conditions +15° C. at maximum take-off thrust at end of runway, which is typically defined at an aircraft speed of around 0.25Mn (i.e. a Mach number of 0.25), or between around 0.24 and 0.27 Mn. Maximum take-off conditions for the engine may therefore be defined as operating the engine at a maximum take-off thrust for the engine at ISA sea level pressure and temperature +15° C. at an aircraft speed of 0.25Mn.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the 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 comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply 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, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the “economic mission”) of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint—in terms of time and/or distance—between top of climb and start of descent. Cruise conditions thus define an operating point of the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number, Mn) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide—in combination with any other engines on the aircraft—steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000 ft (11582 m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50 kN to 65 kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

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

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where 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.

