Operation mode of gas turbine engine

专利摘要:
The invention relates to a gas turbine engine 10 for an aircraft comprising an engine core 11 comprising a turbine 19, a compressor 14, and a core shaft 26 connecting the turbine to the compressor, the turbine 19 being the lowest turbine. engine pressure, the core shaft having a running speed in the range of 1500 rpm to 6200 rpm, and the supercharger being the lowest pressure compressor in the engine; a fan 23 located upstream of the engine core; and a gear 30 which receives input from the core shaft and outputs drive to the fan. The motor core 11 further comprises three bearings 26a-c arranged to support the core shaft 26, the three bearings comprising a front bearing 26a) and two rear bearings 26b, 26c, the core shaft having a length between the front bearing and the rearmost rear bearing 26c in the range of 1800mm to 2900mm, and a minor span between the two rear bearings in the range of 250mm to 350mm, such that it There is no core shaft primary resonance between the front bearing and the most forward rear bearing within the core shaft running speed range. Figure for abstract: Figure 6
公开号:FR3105305A1
申请号:FR2013082
申请日:2020-12-11
公开日:2021-06-25
发明作者:Jillian C Gaskell；Chathura K Kannangara；Punitha Kamesh
申请人:Rolls Royce PLC；
IPC主号:

专利说明:

[0001] The present description relates to the mounting of a heart shaft within a gas turbine engine for an aircraft, and in particular to a bearing positioning and how such a shaft can be arranged and mounted so as to manage effects. vibratory and resonant.
[0002] As used herein, a range of "X-value to Y-value", "X-value to Y-value" or "between X-value and Y-value", or the like, refers to an inclusive range; including the bounding values of X and Y. As used herein, the term "axial plane" means a plane extending the length of an engine, parallel to and containing an axial center line of the engine, and the term "radial plane" denotes a plane extending perpendicular to the axial centerline of the engine, thus including all radial lines at the axial position of the radial plane. Axial planes can also be referred to as longitudinal planes because they extend the length of the engine. A radial distance or an axial distance is therefore a distance in a radial or axial plane, respectively.
[0003] 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 heart shaft connecting the turbine to the compressor, and in which the turbine is the lowest pressure turbine. engine, the heart shaft has a running speed in the range of 1500rpm to 6200rpm, 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 reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft.
[0004] The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, the heart shaft having a length (L) between the front bearing and the rear bearing on the. more rear in the range 1800mm to 2900mm or 2750mm, and a minor reach (S) between the two rear bearings in the range 250mm to 350mm.
[0005] As a result, there may be no primary resonance of the heart shaft between the front bearing and the forward-most rear bearing within the running speed range of the heart shaft.
[0006] The length (L) and minor span (S) can be chosen as appropriate for the desired walking speed range to avoid such primary resonances.
[0007] In various embodiments the length of the core shaft (L) can be in the range of 2000mm to 2900mm, optionally further from 2300mm to 2800mm, and optionally further from 2400mm to 2750mm.
[0008] In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0009] The inventor was aware that the increased length of a heart shaft when a gas turbine engine is scaled up can result in a resonant frequency of the heart shaft being in or near the range of engine operation. Simply upscaling a known motor can therefore lead to increased risks of resonance-induced damage - the longer heart shaft can be problematic. The inventor has found that selecting the length, L, and minor reach, S, as appropriate for a given running speed range can facilitate the avoidance of damaging vortex modes, and can reduce engine damage. in use.
[0010] The lower limit of 1500rpm on the heart shaft running speed may be the minimum running speed under ground idle conditions and / or the upper limit of 6200rpm on the shaft running speed of heart may be the upper limit on the maximum take-off thrust walking speed.
[0011] 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 heart shaft connecting the turbine to the compressor, and in which the turbine is the lower pressure turbine. of the engine, and the compressor is the engine's lowest pressure compressor. The engine further comprises a blower located upstream of the engine core, the blower comprising a plurality of blower blades and having a blower diameter in the range of 330cm to 380cm; and a reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft, the reducer having a gear ratio in the range of 3.1 to 3.8.
[0012] The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, the heart shaft having a length between the front bearing and the rearmost bearing. 'rear in the range of 1800 to 2900mm or 2750mm, and a minor span between the two rear bearings in the range of 250mm to 350mm, so that there is no primary resonance of the heart shaft within a heart tree running speed range.
[0013] Those skilled in the art will understand that the second aspect shares a novel concept with the first aspect, as the combination of fan diameter and gear ratio is related to the heart shaft running speed. For a particular aircraft design, a given combination of fan diameter and gear ratio can be used to derive an expected range of heart shaft running speed.
[0014] In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0015] The heart shaft running speed range can be from 1500rpm to 6200rpm.
[0016] In the first or second aspects, one or more of the following characteristics may be present:
[0017] 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 ratio of minor throw to turbine length of:
[0018]
[0019] may be equal to or less than 1.05. The ratio of the minor seat to the turbine length can be equal to or less than 1.00, optionally equal to or less than 0.95. The ratio of minor seat to turbine length may be 0.70 or greater, or 0.75 or greater, or 0.80 or greater, or 0.85 or greater. For example the ratio of minor span to turbine length may be in the range 0.70 to 1.05, optionally 0.70 to 1.00, optionally 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 optionally further from 0.85 to 0.95 in the range of 0.85 to 0.95.
[0020] The turbine can include four sets of turbine blades. The two rear 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 level of the turbine. dawn. The turbine can include a total of four sets of turbine blades.
[0021] The two rear bearings may be located downstream of the leading edge of the lower pressure turbine blade (the most downstream) of the turbine at the root of the blade.
[0022] The length of the heart shaft (L) can be in the range from 1800mm to 2900mm, optionally from 2000mm to 2900mm, optionally further from 2300mm to 2800mm, and optionally further from 2400mm to 2750mm.
[0023] In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0024] A length ratio (S / L) of the minor bearing between the two rear bearings to the heart shaft length may be equal to or less than 0.14.
[0025] The aspect ratio can be equal to or less than 0.13, or equal to or less than 0.12. The aspect ratio may be equal to or greater than 0.05, or equal or greater than 0.06, or equal or greater than 0.07, or equal or greater than 0.08. For example, the length ratio can be in the range of 0.05 to 0.14, optionally in the range of 0.05 to 0.13, optionally in the range of 0.06 to 0.13, and optionally further in the range of 0.08 to 0.13
[0026] The forward-most bearing of the rear bearings can have a bearing stiffness in the range of 30kN / mm to 100kN / mm.
[0027] A stiffness ratio of the bearing stiffness at the forward-most rear bearing to the distance between the two rear bearings may be in the range of 0.08 to 0.5kN / mm ² , optionally in the range of 0.08 to 0.40kN / mm ² , optionally in the range from 0.08 to 0.30kN / mm ² , optionally in the range from 0.08 to 0.20kN / mm ² , optionally in the range from from 0.09 to 0.40 kN / mm ² , optionally in the range from 0.15 to 0.50 kN / mm ² , optionally in the range from 0.15 to 0.40 kN / mm ² , and optionally in addition in the range from 0.15 to 0.30kN / mm ² .
[0028] The engine's lower pressure turbine has a lower pressure blade set, each blade of the lower pressure blade set having a mass, m , a mid-blade radius, r , and an angular speed in cruising, ω .
[0029] A first report from dawn to the landing of:
[0030]
[0031] can have a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly 3.0 x10 ^-6 to 7.5 x10 ^-6 kg ^{- 1} .rad ^-2 .s ² , possibly from 4.0x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and possibly in addition from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0032] A second report from dawn to the landing of:
[0033]
[0034] can have a value in the range of 0.8 to 6.0kg ^-1 , optionally 0.8 to 5.0kg ^-1 , optionally 0.8 to 4.0kg ^-1 , optionally 0.8 to 3 , 0kg ^-1 , optionally 0.8 to 2.0kg ^-1 .
[0035] In a third aspect, there is provided a method of operating a gas turbine engine for an aircraft. The gas turbine engine comprises an engine core comprising a turbine, a compressor and a heart shaft connecting the turbine to the compressor, and in which the turbine is the lowest pressure turbine of the engine. The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, the heart shaft having a length (L) between the front bearing and the rear bearing on the. more rear in the range 1800mm to 2900mm, and a minor reach (S) between the two rear bearings in the range 250mm to 350mm. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reduction gear arranged to receive an input from the heart shaft and to output a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft
[0036] The method includes operating the motor such that the heart shaft has a running speed in the range of 1500rpm to 6200rpm, and in which there is no primary resonance of the motor shaft. heart within the walking speed range of the heart tree.
[0037] The length of the heart shaft (L) can be in the range of 1800mm to 2750mm, or 2000 to 2750, or 2400mm to 2750mm.
[0038] In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0039] The engine used to implement the method can be as described in the first and / or second aspects.
[0040] In a fourth aspect, there is provided a method of designing a gas turbine engine for an aircraft. The engine comprises an engine heart comprising a turbine, a compressor and a heart shaft connecting the turbine to the compressor, and in which the turbine is the lowest pressure turbine of the engine, the heart shaft has a running speed in the range is from 1500rpm to 6200rpm, 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 reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft. The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, the heart shaft having a length (L) between the front bearing and the rear bearing on the. more rear in the range 1800mm to 2900mm.
[0041] The process includes:
[0042] selection of positions for the front bearing and the foremost bearing of the rear bearings; and
[0043] the elongation of the heart shaft at the rear of the forward-most bearing of the rear bearings such that a minor span (S) defined between the two rear bearings is in the range of 250mm to 350mm, and that there is no primary resonance of the heart shaft between the front bearing and the forward-most rear bearing within the operating speed range of the heart shaft.
[0044] The length of the heart shaft (L) can be in the range of 1800mm to 2750mm, or 2000 to 2750, or 2400mm to 2750mm.
[0045] In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0046] The selection of positions for the front bearing and the foremost bearing of the rear bearings may include positioning the two rear bearings downstream of a leading edge of the lower pressure turbine blade (the most downstream ) of the turbine at the level of a blade root.
[0047] The turbine may include four sets of turbine blades, and the selection of positions for the front bearing and the foremost bearing of the rear bearings may include positioning the two rear bearings downstream of a trailing edge of a. turbine blade of a third set of turbine blades from the front of the turbine at a root of the blade.
[0048] The engine designed by the method of the fourth aspect may be the engine of the first and / or the second aspect.
[0049] 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 heart shaft connecting the turbine to the compressor, and in which the turbine is the lowest pressure turbine of the aircraft. engine, and includes turbine blades, and the compressor is the engine's lowest pressure compressor. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft. The engine heart further comprises three bearings arranged to support the heart shaft, the three bearings comprising two rear bearings located downstream of the leading edge of the lower pressure turbine blades of the turbine at the level of the root of the engine. dawn. The two rear bearings can therefore be described as being located downstream of the leading edge of the last / rearmost turbine blade.
[0050] The inventor was aware that the heart shaft generally moves less (in the radial direction), and is the most level (parallel to the motor axis) at the axial position of the bearings - while modes Vortices and other deflections or displacements can occur between bearings, the bearings serve to limit radial shaft movement. The inventor was aware that careful control of shaft length and bearing position can therefore allow the vortex modes of the motor to be managed, thereby reducing the risk of damage to the motor.
[0051] The inventor has further found that positioning the bearings closer to the larger and taller turbine stages, toward the rear of the turbine, provides improved turbine support, as relative shaft movements. to the turbine position can have more than a detrimental effect on these larger turbine stages.
[0052] The gas turbine engine may further include a disk arranged to support the lower pressure turbine blades of the turbine. The two rear bearings may be located downstream of a center line of the disc.
[0053] The length of the heart shaft can be in the range of 1800 to 2900mm, optionally in the range of 2000 to 2900mm, optionally in the range of 2300 to 2800mm, and optionally further in the range of 2400 to 2750mm.
[0054] The heart shaft can have a running speed range with a lower limit of 1500rpm and an upper limit of 6200rpm.
[0055] The upper limit on the heart shaft running speed range may be the upper limit on the maximum take-off thrust (PMD) walking speed. The lower limit on the heart shaft running speed may be the minimum running speed under ground idle conditions.
[0056] The heart shaft has a length, L, between the front-most bearing and the rear-most bearing, and a distance, S, between the two rear bearings. The bearings may be arranged such that a length ratio, S / L, of the distance between the two rear bearings (S) to the heart 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 rear bearings (S) to the heart 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 rear bearings (S) to the heart shaft length (L) can be in the range of 0.05 to 0.14, possibly in the range of 0.05 to 0.13, optionally in the range of 0.06 to 0.13, and optionally in the range of 0.08 to 0.13.
[0057] The length (L) of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. In such embodiments, the distance between the two rear bearings (S) can be in the range from 250mm to 350mm, or possibly 260mm and 350mm.
[0058] 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 heart shaft connecting the turbine to the compressor, and in which the turbine is the lower pressure turbine. engine, and includes four sets of turbine blades, and the compressor is the engine's lowest pressure compressor. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft. The engine core further comprises three bearings arranged to support the heart shaft, the three bearings comprising two rear bearings located downstream of the trailing edge of a turbine blade of the third set of turbine blades from there. 'front of the turbine, at the level of the root of the blade.
[0059] One or more of the following characteristics may apply for a gas turbine engine of the fifth and / or sixth aspect:
[0060] The forward-most bearing of the rear bearings can have a bearing stiffness in the range of 30kN / mm to 100kN / mm.
[0061] A stiffness ratio of the stiffness at the forward-most rear bearing to the distance between the two rear bearings (S) may be in the range of 0.08 to 0.5kN / mm ² , optionally in the range of from 0.08 to 0.40kN / mm ² , optionally in the range from 0.08 to 0.30kN / mm ² , optionally in the range from 0.08 to 0.20kN / mm ² , optionally in the range ranging from 0.09 to 0.40 kN / mm ² , optionally in the range ranging from 0.15 to 0.50 kN / mm ² , optionally in the range ranging from 0.15 to 0.40 kN / mm ² , and optionally in further in the range from 0.15 to 0.30kN / mm ² .
[0062] The length (L) of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. In such embodiments, the distance between the two rear bearings (S) can be in the range from 250mm to 350mm, or possibly 260mm and 350mm.
[0063] The blower can have a blower diameter in the range of 330cm to 380cm.
[0064] The length of the heart shaft (L) can be in the range from 1800mm to 2900mm, optionally from 2000mm to 2900mm, optionally further from 2300mm to 2800mm, and optionally further from 2400mm to 2750mm.
[0065] The distance between the two rear bearings (S), which can be referred to as minor span, can be in the range from 250mm to 350mm. In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any of 350mm, 345mm, 340mm, or 335mm.
[0066] The running speed of the heart shaft can be in the range of 1500rpm to 6200rpm. Additionally or alternatively, the diameter of the blower can be in the range of 330cm to 380cm. The gear ratio of the reducer can be in the range of 3.1 to 3.8.
[0067] The length, minor reach and / or run speed can be selected such that no primary resonance of the heart shaft is within the running range of the motor.
[0068] The engine's lower pressure turbine has a lower pressure blade set. Each vane in the set of lower pressure vanes has a mass, m, a radius at mid-vane height, r, and a cruising angular speed, ω. A minor bearing (S) is defined, as mentioned previously, as the axial distance between the two rear bearings.
[0069] A first report from dawn to the landing of:
[0070]
[0071] can have a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly 3.0 x10 ^-6 to 7.5 x10 ^-6 kg ^{- 1} .rad ^-2 .s ² , possibly from 4.0x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and possibly in addition from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0072] In addition or as a variant, a second ratio of the vane to the bearing of:
[0073]
[0074] can have a value in the range of 0.8 to 6.0kg ^-1 , optionally 0.8 to 5.0kg ^-1 , optionally 0.8 to 4.0kg ^-1 , optionally 0.8 to 3 , 0kg ^-1 , optionally 0.8 to 2.0kg ^-1 .
[0075] 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 ratio of minor span to turbine length (i.e. S divided by turbine length) can be equal to or less than 1.05, optionally equal to or less than 1.00, optionally equal to or less at 0.95. The ratio of the minor seat to the turbine length can be equal to or greater than 0.70, or equal or greater than 0.75, or equal or greater than 0.80, or equal or greater than 0.85. For example the ratio of minor span to turbine length may be in the range 0.70 to 1.05, optionally 0.70 to 1.00, optionally 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 optionally further from 0.85 to 0.95.
[0076] The turbine can include a total of four sets of turbine blades. In an alternate embodiment, the turbine may include a total of three sets of turbine blades.
[0077] 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. The turbine is the engine's lowest pressure turbine and the compressor is the engine's lowest pressure compressor. The heart shaft has a running speed range with a lower limit of 1500rpm and an upper limit of 6200rpm. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft.
[0078] The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings. The heart shaft has a length (L) between the front bearing and the rearmost rear bearing, and a minor bearing (S) between the rear bearings. The bearings are arranged such that a ratio of the length of the minor bearing to the length of the heart shaft (S / L) is equal to or less than 0.14.
[0079] The aspect ratio can be equal to or less than 0.13, or equal to or less than 0.12. The aspect ratio may be equal to or greater than 0.05, or equal or greater than 0.06, or equal or greater than 0.07, or equal or greater than 0.08. For example, the length ratio can be in the range of 0.05 to 0.14, optionally in the range of 0.05 to 0.13, optionally in the range of 0.06 to 0.13, and optionally further in the range of 0.08 to 0.13.
[0080] The inventor was aware that the increased length of a heart shaft when a gas turbine engine is scaled up can result in a resonant frequency of the heart shaft being in or near the range of. engine operation. Simply upscaling a known motor can therefore lead to increased risks of resonance-induced damage - the longer heart shaft can be problematic.
[0081] However, it was found that making the heart shaft even longer - extending to the rear of a first rear bearing and to a second rear bearing, with the spacing between the first and second rear bearings. being within a defined range of the total heart shaft length - increases the stiffness of the heart shaft and shifts the resonant frequency away from the motor operating range in some embodiments .
[0082] The upper limit on the heart shaft walking speed range may be the upper limit on the maximum take-off thrust walking speed.
[0083] The lower limit on the heart shaft running speed may be the minimum running speed under ground idle conditions.
[0084] The running speed of the heart shaft in cruising conditions can be in the range of 5400 to 5700rpm, and optionally in the range of 5500 to 5600rpm.
[0085] The running speed of the heart shaft under maximum take-off thrust (PMD) conditions may be in the range of 5800 to 6200rpm, and optionally in the range of 5900 to 6100rpm.
[0086] The length of the heart shaft (L) can be in the range from 1800mm to 2900mm, optionally from 2000mm to 2900mm, optionally further from 2300mm to 2800mm, and optionally further from 2400mm to 2750mm.
[0087] In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0088] The forward-most bearing of the rear bearings can have a bearing stiffness in the range of 30kN / mm to 100kN / mm. A stiffness ratio of the stiffness at the forward-most rear bearing to the minor bearing can be in the range of 0.08 to 0.5kN / mm ² , optionally in the range of 0.08 to 0, 40kN / mm ² , optionally in the range from 0.08 to 0.30kN / mm ² , optionally in the range from 0.08 to 0.20kN / mm ² , optionally in the range from 0.09 to 0 , 40kN / mm ² , optionally in the range from 0.15 to 0.50 kN / mm ² , optionally in the range from 0.15 to 0.40 kN / mm ² , and optionally further in the range from 0 , 15 to 0.30kN / mm ² .
[0089] The length (L) of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. The minor span (S) can be in the range of 250mm to 350mm, possibly 260mm to 350mm.
[0090] The blower can have a blower diameter in the range of 330cm to 380cm.
[0091] In some embodiments, the length of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. The minor span can be in the range of 250mm to 350mm, possibly 260mm to 350mm. The running speed of the heart shaft can be in the range of 1500rpm to 6200rpm; and / or a diameter of the blower can be in the range of 330cm to 380cm and the gear ratio of the reducer can be in the range of 3.1 to 3.8. In such embodiments, the length, minor range and / or run speed can be selected such that no primary resonance of the heart shaft is (is within the running range of the motor. .
[0092] The rear bearings can be positioned axially flush with or behind:
[0093] (i) a leading edge of a lower pressure turbine blade of the turbine at the root of the blade; and or
[0094] (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, in which the turbine comprises four sets turbine blades.
[0095] The inventor was aware that positioning the rear bearings closer to the larger and larger turbine stages, towards the rear of the turbine, provides improved turbine support, as shaft movements relative to the turbine. The turbine position may have a more detrimental effect on these larger turbine stages.
[0096] The engine's lower pressure turbine has a lower pressure blade set, each blade of the lower pressure blade set having a mass, m , a mid-blade radius, r , and an angular speed in cruising, ω . A minor bearing (S) is defined as the axial distance between the two rear bearings, as mentioned previously.
[0097] A first report from dawn to the landing of:
[0098]
[0099] can have a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly 3.0 x10 ^-6 to 7.5 x10 ^-6 kg ^{- 1} .rad ^-2 .s ² , possibly from 4.0x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and possibly in addition from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0100] In addition or as a variant, a second ratio of the vane to the bearing of:
[0101]
[0102] can have a value in the range of 0.8 to 6.0kg ^-1 , optionally 0.8 to 5.0kg ^-1 , optionally 0.8 to 4.0kg ^-1 , optionally 0.8 to 3 , 0kg ^-1 , optionally 0.8 to 2.0kg ^-1 .
[0103] 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 ratio of minor span to turbine length can be in the range 0.70 to 1.05, optionally 0.70 to 1.00, optionally 0.70 to 0.95, optionally 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 optionally in addition from 0.85 to 0.95.
[0104] According to an eighth aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor and a heart shaft connecting the turbine to the compressor, and in which the turbine is the lower pressure turbine. of the engine, and the compressor is the engine's lowest pressure compressor. 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 the level of its leading edge. of 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 reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft.
[0105] The motor core further includes three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, with a minor bearing (S) defined as the distance between the two rear bearings. A ratio of minor throw to turbine length of:
[0106]
[0107] is equal to or less than 1.05.
[0108] The inventor was aware that keeping the ratio of minor reach to turbine length within this range, and more generally smaller than that in known aircraft, can help reduce or avoid vortex modes. detrimental in operation.
[0109] The ratio of the minor seat to the turbine length can be equal to or less than 1.00, optionally equal to or less than 0.95. The ratio of minor seat to turbine length may be 0.70 or greater, or 0.75 or greater, or 0.80 or greater, or 0.85 or greater. For example the ratio of minor span to turbine length may be in the range 0.70 to 1.05, optionally 0.70 to 1.00, optionally 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 optionally further from 0.85 to 0.95 in the range of 0.85 to 0.95.
[0110] The inventor was aware that the increased length of a heart shaft when a gas turbine engine is scaled up can result in a resonant frequency of the heart shaft being in or near the range of. engine operation. Simply upscaling a known motor can therefore lead to increased risks of resonance-induced damage - the longer heart shaft can be problematic. The inventor has found that an arrangement of the minor span approximately equal to, and possibly slightly smaller than, the turbine length (of the lower pressure turbine, in embodiments with more than one turbine), and more specifically within the claimed range, can facilitate the avoidance of detrimental vortex modes, and can reduce damage to the engine in use.
[0111] The turbine can include a total of four sets of turbine blades. In such embodiments, either of the two rear 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 level of the root of the dawn.
[0112] The two rear bearings may be located downstream of the leading edge of the lower pressure turbine blade (the most downstream) of the turbine at the root of the blade.
[0113] The length of the heart shaft (L) can be in the range from 1800mm to 2900mm, optionally from 2000mm to 2900mm, optionally further from 2300mm to 2800mm, and optionally further from 2400mm to 2750mm.
[0114] In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0115] The heart shaft can have a running speed range with a lower limit of 1500rpm and an upper limit of 6200rpm. The upper limit on the heart shaft running speed range may be the upper limit on the maximum take-off thrust (PMD) walking speed. The lower limit on the heart shaft running speed may be the minimum running speed under ground idle conditions.
[0116] The heart shaft has a length, L, between the forward-most bearing and the rear-most bearing, and a distance (the minor span, S) between the two rear bearings. The bearings can be arranged such that a length ratio (S / L) of the distance between the two rear bearings to the heart 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 rear bearings (S) to the heart 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 rear bearings (S) to the heart shaft length (L) can be in the range of 0.05 to 0.14, possibly in the range of 0.05 to 0.13, optionally in the range of 0.06 to 0.13, and optionally in the range of 0.08 to 0.13.
[0117] The forward-most bearing of the rear bearings can have a bearing stiffness in the range of 30kN / mm to 100kN / mm. A stiffness ratio of the bearing stiffness at the forward-most rear bearing divided by the distance between the two rear bearings may be in the range of 0.08 to 0.5kN / mm ² , optionally in the range of from 0.08 to 0.40kN / mm ² , optionally in the range from 0.08 to 0.30kN / mm ² , optionally in the range from 0.08 to 0.20kN / mm ² , optionally in the range ranging from 0.09 to 0.40 kN / mm ² , optionally in the range ranging from 0.15 to 0.50 kN / mm ² , optionally in the range ranging from 0.15 to 0.40 kN / mm ² , and optionally in further in the range from 0.15 to 0.30kN / mm ² .
[0118] The length (L) of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. The distance (S) between the two rear bearings can be in the range from 250mm to 350mm, possibly from 260mm to 350mm.
[0119] The blower can have a blower diameter in the range of 330cm to 380cm.
[0120] The gas length of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm; and the minor span can be in the range of 250mm to 350mm, possibly 260mm to 350mm. The running speed of the heart shaft can be in the range of 1500rpm to 6200rpm; and / or a diameter of the blower can be in the range of 330cm to 380cm and the gear ratio of the reducer can be in the range of 3.1 to 3.8. The length, minor reach and / or run speed can be selected such that no primary resonance of the heart shaft is within the running range of the motor.
[0121] The engine's lower pressure turbine has a lower pressure blade set, each blade of the lower pressure blade set having a mass, m , a mid-blade radius, r , and an angular speed in cruising, ω .
[0122] A first report from dawn to the landing of:
[0123]
[0124] can have a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly 3.0 x10 ^-6 to 7.5 x10 ^-6 kg ^{- 1} .rad ^-2 .s ² , possibly from 4.0x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and possibly in addition from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0125] A second report from dawn to the landing of:
[0126]
[0127] can have a value in the range of 0.8 to 6.0kg ^-1 , optionally 0.8 to 5.0kg ^-1 , optionally 0.8 to 4.0kg ^-1 , optionally 0.8 to 3 , 0kg ^-1 , possibly from 0.8 to 2.0kg ^-1 ..
[0128] According to a ninth aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor and a heart shaft connecting the turbine to the compressor, and in which the turbine is the lowest pressure turbine. of the engine, and the compressor is the engine's lowest pressure compressor. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft. The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, and wherein the forward-most rear bearing has a bearing stiffness in the range of 30kN. / mm at 100kN / mm, the bearing stiffness being defined by the radial displacement caused by the application of a radial force at the axial center point of the bearing.
[0129] The inventor was aware that controlling the stiffness of the bearing to be within the specified range can allow or facilitate the management of vibration modes, thus potentially reducing damage to the motor in use caused by moving in motion. heart tree swirl mode.
[0130] The heart shaft has a length (L) between the front bearing and the rearmost rear bearing, and a minor bearing (S) between the rear bearings. The bearings may be arranged such that a length ratio (S / L) of the minor seat 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 aspect ratio may be equal to or greater than 0.05, or equal or greater than 0.06, or equal or greater than 0.07, or equal or greater than 0.08. For example, the length ratio can be in the range of 0.05 to 0.14, optionally in the range of 0.05 to 0.13, optionally in the range of 0.06 to 0.13, and optionally further in the range of 0.08 to 0.13.
[0131] The heart shaft can have a running speed range with a lower limit of 1500rpm and an upper limit of 6200rpm.
[0132] The bearing stiffness of the forward-most rear bearing may be equal to or about 50kN / mm.
[0133] The gas turbine engine may further include a stationary support structure and a first bearing support structure. The forward-most rear bearing may be mounted to the stationary support structure by the first bearing support structure. The first bearing support structure is attachable to the stationary support structure at a first position located axially aft of the forward-most rear bearing.
[0134] In such embodiments, the first bearing support structure may include a plurality of connecting members, which may be circumferentially spaced around the motor axis, connecting the bearing to the stationary support structure. The first bearing support structure can be described as a spring bar type support structure.
[0135] The first bearing support structure may include an outer cage of the forward-most rear bearing.
[0136] The gas turbine engine of embodiments with a first bearing support structure may further include a second bearing support structure. The second bearing support structure may be mounted to the stationary support structure, optionally at a second position located forward of, and at a greater radial distance from the motor shaft than, the first position. . The second bearing support structure may be connected to the first bearing support structure by a damping fluid pad in the region of the forward-most rear bearing.
[0137] A stiffness ratio of the bearing stiffness at the forward-most rear bearing to the minor bearing can be in the range of 0.08 to 0.5kN / mm ² , optionally in the range of 0.08 to 0.40kN / mm ² , optionally in the range from 0.08 to 0.30kN / mm ² , optionally in the range from 0.08 to 0.20kN / mm ² , optionally in the range from 0.09 to 0.40kN / mm ² , optionally in the range from 0.15 to 0.50 kN / mm ² , optionally in the range from 0.15 to 0.40 kN / mm ² , and optionally in addition in the range from from 0.15 to 0.30kN / mm ² ...
[0138] The length of the heart shaft (L) can be in the range from 1800mm to 2900mm, optionally from 2000mm to 2900mm, optionally further from 2300mm to 2800mm, and optionally further from 2400mm to 2750mm.
[0139] In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0140] The length of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. In such embodiments, the minor span may be in the range of 250mm to 350mm, optionally 260mm to 350mm.
[0141] The blower can have a blower diameter in the range of 330cm to 380cm.
[0142] The length of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. The minor span can be in the range of 250mm to 350mm. The running speed of the heart shaft can be in the range of 1500rpm to 6200rpm; and / or a diameter of the blower may be in the range of 330cm to 380cm and the gear ratio of the reducer is in the range of 3.1 to 3.8. The length, minor reach and / or run speed can be selected such that no primary resonance of the heart shaft is within the running range of the motor.
[0143] The rear bearings may be positioned axially flush with or behind a leading edge of a lower pressure turbine blade of the turbine at the root of the blade.
[0144] The rear bearings may be positioned axially level with or behind a trailing edge of a turbine blade of a third set of turbine blades from the front of the turbine at the level of blade root, in which the turbine comprises four sets of turbine blades.
[0145] The engine's lower pressure turbine has a lower pressure blade set, each blade of the lower pressure blade set having a mass, m , a mid-blade radius, r , and an angular speed in cruising, ω .
[0146] A first report from dawn to the landing of:
[0147]
[0148] can have a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly 3.0 x10 ^-6 to 7.5 x10 ^-6 kg ^{- 1} .rad ^-2 .s ² , possibly from 4.0x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and possibly in addition from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0149] A second report from dawn to the landing of:
[0150]
[0151] can have a value in the range of 0.8 to 6.0kg ^-1 , optionally 0.8 to 5.0kg ^-1 , optionally 0.8 to 4.0kg ^-1 , optionally 0.8 to 3 , 0kg ^-1 , optionally 0.8 to 2.0kg ^-1 .
[0152] 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 ratio of minor seat to turbine length may be equal to or less than 1.05. The ratio of the minor seat to the turbine length can be equal to or less than 1.00, optionally equal to or less than 0.95. The ratio of minor seat to turbine length may be 0.70 or greater, or 0.75 or greater, or 0.80 or greater, or 0.85 or greater. For example the ratio of minor span to turbine length may be in the range 0.70 to 1.05, optionally 0.70 to 1.00, optionally 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 optionally further from 0.85 to 0.95 in the range of 0.85 to 0.95.
[0153] 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 heart shaft connecting the turbine to the compressor, and in which the compressor is the lower pressure compressor. of the engine and the turbine is the lowest pressure turbine of the engine and has one set of lower pressure vanes each vane of the lower pressure set of vanes having a mass, m , a mid-height radius vane, r , and a cruising angular speed, ω . The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft.
[0154] The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, with a minor bearing (S) being defined as the axial distance between the two rear bearings. A first report from dawn to the landing of:
[0155]
[0156] has a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0157] fournissant une mesure de force), mais plutôt que la longueur de portée mineure(S) doit être augmentée aussi peu que possible de façon à réduire relativement la longueur et le poids de moteur, permettant ainsi aux gains de rendement d'être accrus en évitant le poids supplémentaire, et pour éviter le développement de modes tourbillonnaires indésirables au sein de la portée mineure. Tandis que le sens commun suggère qu'une portée mineure plus grande est souhaitable pour améliorer une réaction de forces à partir de la turbine basse pression, l'inventeur a trouvé que le risque d'introduire des modes tourbillonnaires, et l'introduction de davantage de longueur et de poids, contrebalançaient les bénéfices de réaction de force et que le premier rapport de l'aube au palier doit pour cette raison être maintenu au sein de la plage spécifiée.The inventor was aware that the motor should not be linearly scaled up with an increase in force (providing a measure of force), but rather that the minor span length (S) should be increased as little as possible so as to relatively reduce motor length and weight, thus allowing efficiency gains to be increased by avoiding the additional weight, and to avoid the development of unwanted vortex modes within the minor staff. While common sense suggests that a greater minor range is desirable to improve a reaction of forces from the low pressure turbine, the inventor has found that the risk of introducing vortex modes, and the introduction of more length and weight, outweighed the force feedback benefits and therefore the vane-to-bearing first ratio should be kept within the specified range.
[0158] The first vane-to-bearing ratio can be in the range of 3.0x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , optionally 4.0 x10 ^-6 to 7, 5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0 x10 ^{- 6} to 4.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and possibly additionally from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0159] A second report from dawn to the landing of:
[0160]
[0161] can have a value in the range of 0.8 to 6.0kg ^-1 , optionally 0.8 to 5.0kg ^-1 , optionally 0.8 to 4.0kg ^-1 , optionally 0.8 to 3 , 0kg ^-1 , optionally 0.8 to 2.0kg ^-1 .
[0162] 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 heart shaft connecting the turbine to the compressor, and in which the turbine is the lowest pressure turbine. engine and has a lower pressure vane set, each vane in the lower pressure vane set having a mass, m , a mid-vane radius, r , and the compressor is the lower engine pressure compressor. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft.
[0163] The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, with a minor bearing (S) being defined as the axial distance between the two rear bearings. A second report from dawn to the landing of:
[0164]
[0165] has a value in the range of 0.8 to 6.0kg ^-1 .
[0166] The second vane to bearing ratio can be in the range 0.8 to 5.0kg ^-1 , optionally 0.8 to 4.0kg ^-1 , optionally 0.8 to 3.0kg ^-1 , possibly from 0.8 to 2.0kg ^-1 .
[0167] Each vane of the lowest pressure vane assembly has a cruising angular speed, ω, and a first vane-to-step ratio of:
[0168]
[0169] can have a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly 3.0 x10 ^-6 to 7.5 x10 ^-6 kg ^{- 1} .rad ^-2 .s ² , possibly from 4.0x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and possibly in addition from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0170] In the tenth or eleventh aspects, one or more of the following characteristics may be present:
[0171] The minor span, S, can be in the range of 250mm to 350mm. In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm.
[0172] The length of the heart shaft (L) can be in the range from 1800mm to 2900mm, optionally from 2000mm to 2900mm, optionally further from 2300mm to 2800mm, and optionally further from 2400mm to 2750mm.
[0173] The vane mass value, m, multiplied by the mid-height vane radius, r, can be in the range of 180 to 280kg.mm
[0174] The reducer can have a gear ratio greater than 3, and optionally in the range of 3.1 to 3.8.
[0175] The heart shaft can have a running speed in the range of 1500rpm to 6200rpm.
[0176] The driving speed of the heart shaft in cruising can be in the range of 5400 to 5700rpm, and optionally from 5500 to 5600rpm, in cruising.
[0177] The running speed of the heart shaft under maximum take-off thrust (PMD) conditions can be in the range of 5800 to 6200rpm, and optionally 5900 to 6100rpm.
[0178] A ratio of the length (S / L) of the minor bearing between the two rear bearings to the heart 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 can be equal to or greater than 0.05, or equal or greater than 0.06, or equal or greater than 0.07, or equal or greater than 0.08. For example, the length ratio S / L can be in the range of 0.05 to 0.14, optionally in the range of 0.05 to 0.13, optionally in the range of 0.06 to 0 , 13, and optionally in the range of 0.08 to 0.13.
[0179] The mass, m , of a vane of the lower pressure vane assembly can be in the range of 0.2 to 0.6 kg.
[0180] The radius, r , of a vane of the lower pressure vane assembly can be in the range of 400 to 600mm.
[0181] Each vane of the lower pressure vane assembly can have a cruising angular speed, ω in the range of 560 to 600rad.s ^-1 .
[0182] The length of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. The minor span can be in the range of 250mm to 350mm, and possibly 260mm to 350mm.
[0183] The blower can have a blower diameter in the range of 330cm to 380cm.
[0184] 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 ratio of minor throw to turbine length of:
[0185]
[0186] may be equal to or less than 1.05, optionally equal or less than 1.00, optionally equal or less than 0.95. The ratio of minor seat to turbine length may be 0.70 or greater, or 0.75 or greater, or 0.80 or greater, or 0.85 or greater. For example the ratio of minor span to turbine length may be in the range 0.70 to 1.05, optionally 0.70 to 1.00, optionally 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 optionally further from 0.85 to 0.95.
[0187] The rear bearings may be positioned axially flush with or behind a leading edge of a lower pressure turbine blade of the turbine at the root of the blade.
[0188] The rear bearings may be positioned axially flush with or behind a trailing edge of a turbine blade of a third set of turbine blades from the front of the turbine at the level of the root of dawn. In such embodiments, the turbine may include four sets of turbine blades, and optionally may have a total of four sets of turbine blades.
[0189] The forward-most bearing of the rear bearings can have a bearing stiffness in the range of 30kN / mm to 100kN / mm.
[0190] A stiffness ratio of the bearing stiffness at the forward-most rear bearing to the distance between the two rear bearings may be in the range of 0.08 to 0.5kN / mm ² , optionally in the range of 0.08 to 0.40kN / mm ² , optionally in the range from 0.08 to 0.30kN / mm ² , optionally in the range from 0.08 to 0.20kN / mm ² , optionally in the range from from 0.09 to 0.40 kN / mm ² , optionally in the range from 0.15 to 0.50 kN / mm ² , optionally in the range from 0.15 to 0.40 kN / mm ² , and optionally in addition in the range from 0.15 to 0.30kN / mm ² .
[0191] According to a twelfth aspect, there is provided a method of operating a gas turbine engine for an aircraft, the engine comprising an engine core comprising a turbine, a compressor and a heart shaft connecting the turbine to the compressor, and in which the compressor is the lowest pressure compressor of the engine and the turbine is the lowest pressure turbine of the engine and has a lower pressure vane set, each vane of the lower pressure vane set having a mass, m , a radius at mid-blade height, r , and a cruising angular speed, ω. The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, with a minor bearing (S) being defined as the axial distance between the two rear bearings. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reduction gear arranged to receive an input from the heart shaft and to output a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft.
[0192] The method includes operating the engine such that a first vane-to-bearing ratio of:
[0193]
[0194] has a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² while cruising.
[0195] In various embodiments the method can include operating the motor such that the first vane to bearing ratio can be in the range of 3.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 . rad ^-2 .s ² , possibly from 4.0 x10 ^-6 to 7.5x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0 x10 ^-6 to 6.5x10 ^-6 kg ^{- 1} .rad ^-2 .s ² , and possibly in addition from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0196] The engine used to implement the method of the twelfth aspect may be as described in the tenth and / or eleventh aspect.
[0197] 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 heart shaft connecting the turbine to the compressor, and in which the turbine is the lowest pressure turbine. of the engine, and the compressor is the engine's lowest pressure compressor. The engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a reducer which receives input from the heart shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the heart shaft.
[0198] The motor core further comprises three bearings arranged to support the heart shaft, the three bearings comprising a front bearing and two rear bearings, the distance between the two rear bearings being defined as the minor bearing, S. The rear bearing the forward most has a bearing stiffness defined by the radial displacement caused by the application of a radial force at the axial center point of the bearing. A stiffness ratio of the bearing stiffness at the forward-most rear bearing to the minor bearing is in the range of 0.08 to 0.5kN / mm ² .
[0199] The bearing stiffness ratio can be in the range of 0.09 to 0.40kN / mm ² , optionally in the range of 0.08 to 0.30kN / mm ² , optionally in the range of 0.08 to 0.20kN / mm ² , optionally in the range from 0.09 to 0.40 kN / mm ² , optionally in the range from 0.15 to 0.50 kN / mm ² , optionally in the range from 0, 15 to 0.40 kN / mm ² , and optionally further in the range from 0.15 to 0.30 kN / mm ² .
[0200] The inventor was aware that controlling the bearing stiffness and the minor bearing so that the ratio of the two is within the specified range can allow or facilitate the management of vibration modes, thereby potentially reducing damage to the shaft. motor in use caused by swirling motions of the heart shaft.
[0201] The heart shaft has a length, L, between the front bearing and the rearmost rear bearing. The bearings may be arranged such that a length ratio (S / L) of the minor seat 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 can be equal to or greater than 0.05, or equal or greater than 0.06, or equal or greater than 0.07, or equal or greater than 0.08. For example, the length ratio S / L can be in the range of 0.05 to 0.14, optionally in the range of 0.05 to 0.13, optionally in the range of 0.06 to 0 , 13, and optionally in the range of 0.08 to 0.13.
[0202] The heart shaft can have a running speed range with a lower limit of 1500rpm and an upper limit of 6200rpm.
[0203] The bearing stiffness of the forward-most rear bearing can be in the range of 30kN / mm to 100kN / mm. Optionally, the bearing stiffness of the forward-most rear bearing may be at least substantially equal to 50kN / mm.
[0204] The gas turbine engine may further include a stationary support structure and a first bearing support structure. The forward-most rear bearing may be mounted to the stationary support structure by the first bearing support structure. The first bearing support structure is attachable to the stationary support structure at a first position located axially aft of the forward-most rear bearing.
[0205] In such embodiments, the first bearing support structure may include a plurality of connecting elements, which may be spaced circumferentially around the motor axis. The connecting elements can connect the forward-most rear bearing to the stationary support structure.
[0206] Additionally or alternatively, in such embodiments, the first bearing support structure may include an outer race of the forward-most rear bearing.
[0207] In embodiments with a first bearing support structure, the motor may further include a second bearing support structure. The second bearing support structure may be mounted to the stationary support structure, optionally at a second position located forward of, and at a greater radial distance from the motor shaft than, the first position. . The second bearing support structure may be connected to the first bearing support structure by a damping fluid pad in the region of the forward-most rear bearing.
[0208] The length, L, of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm. The minor span can be in the range of 250mm to 350mm, possibly 260mm to 350mm.
[0209] The blower can have a blower diameter in the range of 330cm to 380cm.
[0210] The length, L, of the heart shaft can be in the range of 1800mm to 2900mm or 2750mm and the minor reach in the range of 250mm to 350mm, possibly 260mm to 350mm. The running speed of the heart shaft can be in the range of 1500rpm to 6200rpm; and / or a diameter of the blower may be in the range of 330cm to 380cm and the gear ratio of the reducer in the range of 3.1 to 3.8. The length, minor reach and / or run speed can be selected such that no primary resonance of the heart shaft is within the running range of the motor.
[0211] The rear bearings may be positioned axially flush with or behind a leading edge of a lower pressure turbine blade of the turbine at the root of the blade.
[0212] The rear bearings may be positioned axially flush with or behind a trailing edge of a turbine blade of a third set of turbine blades from the front of the turbine at the level of the root of dawn. In such embodiments, the turbine can include four sets of turbine blades.
[0213] The lower pressure turbine of the engine may have a lower pressure blade set, each blade of the lower pressure blade set having a mass, m , a mid-blade radius, r , and an angular speed in cruising, ω . The minor seat (S) is defined as the axial distance between the two rear bearings, as described elsewhere.
[0214] A first report from dawn to the landing of:
[0215]
[0216] can have a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly 3.0 x10 ^-6 to 7.5 x10 ^-6 kg ^{- 1} .rad ^-2 .s ² , possibly from 4.0x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 5.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 2.0 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , possibly from 3.0x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and possibly in addition from 4.5 x10 ^-6 to 6.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0217] A second report from dawn to the landing of:
[0218]
[0219] can have a value in the range of 0.8 to 6.0kg ^-1 , optionally 0.8 to 5.0kg ^-1 , optionally 0.8 to 4.0kg ^-1 , optionally 0.8 to 3 , 0kg ^-1 , optionally 0.8 to 2.0kg ^-1 .
[0220] 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 ratio of minor span to turbine length (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 ratio of minor seat to turbine length may be 0.70 or greater, or 0.75 or greater, or 0.80 or greater, or 0.85 or greater. For example the ratio of minor span to turbine length may be in the range 0.70 to 1.05, optionally 0.70 to 1.00, optionally 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 optionally further from 0.85 to 0.95.
[0221] In any of the aspects described above, one or more of the following characteristics may be present:
[0222] The turbine can be a first turbine, the compressor can be a first compressor, and the heart shaft can be a first heart shaft. The engine core may further include a second turbine, a second compressor, and an interconnection shaft connecting the second turbine to the second compressor. The second turbine, the second compressor and the second core shaft can be arranged to rotate at a higher rotational speed than the first core shaft.
[0223] The motor may further include a tail bearing housing located at the rear of the turbine. The tail bearing housing may include two bearing discs; each bearing disc can be arranged to support one of the two rear 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 rear bearings (eg, the rearmost bearing).
[0224] In embodiments with one or more bearing disks, one or more of the bearing disks may be oriented at least substantially perpendicular to the motor axis (i.e. at least substantially in a radial plane across. engine)
[0225] In the various aspects and embodiments described herein, the engine operating range can be defined as the range of rotational speeds of the heart shaft during standard engine operation (e.g., during ground idle, takeoff, etc. climb and cruise), and can be measured in revolutions per minute (rpm). In this context, "standard engine operation" can exclude transient periods during starting and stopping, for example when the core shaft rotational speed increases from zero to idle rotational speed. on the ground. The engine operating range includes ground idle speed, cruising speed, and maximum take-off thrust speed (PMD). The core shaft rotational speed may be greater than or equal to the ground idle rotational speed over all standard engine operation. The heart shaft rotational speed during standard motor operation may also be referred to as the heart shaft running speed.
[0226] As noted elsewhere in this document, the present description may relate to a gas turbine engine. Such a gas turbine engine can include an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having fan blades) located upstream of the engine core.
[0227] The arrangements of the present disclosure may be particularly, although not exclusively, beneficial for blowers driven through a reducer. Accordingly, the gas turbine engine may include a reduction gear which receives input from the heart shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the heart shaft. . The input to the reducer can be directly from the heart shaft, or indirectly from the heart shaft, for example via a spur shaft and / or gear. The heart shaft can join the turbine and the compressor together, so that the turbine and the compressor rotate at the same speed (with the blower rotating at a lower speed).
[0228] The gas turbine engine as described and / or claimed herein can have any suitable general architecture. For example, the gas turbine engine can have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. By way of example only, the turbine connected to the heart shaft may be a first turbine, the compressor connected to the heart shaft may be a first compressor, and the heart shaft may be a first heart shaft. . The engine core may further include a second turbine, a second compressor and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor and the second core shaft can be arranged to rotate at a higher rotational speed than the first core shaft.
[0229] In such an arrangement, the second compressor can 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) a flow from the first compressor.
[0230] The reducer may be arranged to be driven by the heart shaft which is configured to rotate (eg in use) at the lowest rotational speed (eg the first heart shaft in the example below). above). For example, the reducer may be arranged to be driven only by the heart shaft which is configured to rotate (eg in use) at the lowest rotational speed (eg to be only the first heart shaft. , not the second heart tree, in the example above). Alternatively, the reducer can be arranged to be driven by any or any shafts, for example the first and / or second shafts in the example above.
[0231] The reducer can be a reduction box (in that the output to the fan has a lower rotational speed than the input from the heart shaft). Any type of reducer can be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere in this document. The reducer can have any desired reduction ratio (defined as the input shaft rotational speed divided by the output shaft rotational speed), for example greater than 2.5, for example in the range of 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 can be, for example, between any two of the values in the previous sentence. Strictly by way of example, the reducer may be a "star" reducer having a 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.
[0232] In any gas turbine engine as described and / or claimed herein, a combustion chamber may be provided axially downstream of the fan and the compressor (s). For example, the combustion chamber can be directly downstream of (for example at the outlet of the) second compressor, when a second compressor is supplied. As a further example, the flow at the outlet to the combustion chamber can be supplied to the inlet of the second turbine, when a second turbine is supplied. The combustion chamber can be supplied upstream of the turbine (s).
[0233] The or each compressor (eg the first compressor and the second compressor as described above) may include any number of stages, eg multiple stages. Each stage can include a row of rotor blades and a row of stator vanes, which can be variable stator vanes (in that their angle of incidence can be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.
[0234] The or each turbine (eg the first turbine and the second turbine as described above) may include any number of stages, eg multiple stages. Each stage can 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.
[0235] Each fan blade can be defined as having a radial seat extending from a root (or hub) to a radially internal gas washed location, or 0% reach position, up to one end at a 100% reach position. The ratio of the fan blade radius at the hub to the fan blade radius 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 fan blade radius at the hub to the fan blade radius at the tip may be included within an inclusive range bounded by any two of the values in the preceding sentence (i.e. i.e. the values may form upper or lower limits), for example in the range 0.28 to 0.32. These ratios can be commonly referred to as the hub-to-end ratio. Both the hub radius and the tip radius can be measured at the leading edge (or axially most forward) portion of the vane. The hub to end ratio refers, of course, to the gas washed portion of the fan blade, i.e. the portion radially outward from any platform.
[0236] The fan radius can be measured from the centerline of the engine to the end of a fan blade at its leading edge. The blower diameter (which can be just twice the blower radius) can be larger than (or on the order of) any of: 220cm, 230cm, 240cm, 250cm (about 100inches), 260cm, 270cm (approx 105inch), 280cm (approx 110inch), 290cm (approx 115inch), 300cm (approx 120inch), 310cm, 320cm (approx 125inch), 330cm (approx 130inch), 340cm (approx 135inch), 350cm, 360cm (approx 140inch) , 370cm (around 145inches), 380cm (around 150inches), 390cm (around 155inches), 400cm, 410cm (around 160inches) or 420cm (around 165inches). The diameter of the blower may be within an inclusive range delimited by any two of the values in the previous sentence (i.e. the values may form upper or lower limits), for example in the range from 240cm to 280cm or from 330cm to 380cm.
[0237] The speed of the blower may vary during use. Generally, the rotational speed is lower for blowers with a larger diameter. Strictly by way of nonlimiting example, the speed of rotation of the fan at cruising conditions may be less than 2500 rpm, for example less than 2300 rpm. Strictly as a further non-limiting example, the fan rotational speed at cruising conditions for an engine having a fan diameter in the range of 220cm to 300cm (e.g. 240cm to 280cm or 250cm to 270cm) may be in the range of 1700rpm to 2500rpm, for example in the range of 1800rpm to 2300rpm, for example in the range of 1900rpm to 2100rpm. By way of non-limiting example only, the rotational speed of the blower under cruising conditions for an engine having a blower diameter in the range of 330cm to 380cm may be in the range of 1200rpm to 2000rpm. , for example in the range of 1300rpm to 1800rpm, for example in the range of 1400rpm to 1800rpm.
[0238] In use of the gas turbine engine, the blower (with associated blower vanes) rotates around an axis of rotation. This rotation results in a displacement of the end of the fan blade with a speed U _tip . The work performed by the fan blades13 on the flow leads to an elevation of the enthalpiedH of the flow. A blower end load can be defined by dH / U _tip ² , where dH is the enthalpy increase (e.g. 1-D mean enthalpy increase) across the blower and U _tip is the speed (transition) of the blower end, for example at the leading edge of the end (which can be defined as the blower end radius at the leading edge multiplied by the angular velocity ). Fan end load 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 blower end load can be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values can form upper or lower limits), for example in the range of 0 , 28 to 0.31 or 0.29 to 0.3.
[0239] Gas turbine engines according to the present description can have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through. the heart at cruising 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 report can be in an inclusive range delimited by Any two of the values from the previous sentence (i.e. the values may form upper or lower limits), for example in the range 12 to 16, 13 to 15, or 13 to 14. The bypass duct can be substantially annular. The bypass duct can be radially outside the core motor. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.
[0240] The overall pressure ratio of a gas turbine engine as described and / or claimed herein can be defined as the ratio of the shutdown pressure upstream of the blower to the shutdown pressure at the outlet of the blower. higher pressure compressor (before entering the combustion chamber). By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein in 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 can be in an inclusive range bounded by any two of the values in the preceding sentence (that is, the values may form upper or lower limits), for example in the range 50 to 70.
[0241] The specific thrust of an engine can be defined as the net engine thrust divided by the total mass flow rate through the engine. Under cruising 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: 110Nkg ^-1 s, 105Nkg ^-1 s, 100Nkg ^-1 s, 95Nkg ^-1 s, 90Nkg ^-1 s, 85Nkg ^-1 s or 80Nkg ^-1 s. The specific thrust can be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values can form upper or lower limits), for example in the range of 80Nkg ^-1 s to 100Nkg ^-1 s, or from 85Nkg ^-1 s to 95Nkg ^-1 s. Such engines can be particularly efficient compared to conventional gas turbine engines.
[0242] A gas turbine engine as described and / or claimed herein can have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and / or claimed here may be capable of producing a maximum thrust of at least (or of the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN or 550kN. The maximum thrust can be an inclusive range bounded by any two of the values in the previous sentence (i.e. the values can form upper or lower limits). By way of example only, a gas turbine as described and / or claimed herein may be capable of producing maximum thrust in the range 330kN to 420kN, for example 350kN to 400kN. The above mentioned thrust can be the maximum net thrust under standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.
[0243] In use, the temperature of the flow at the inlet of the high pressure turbine can be particularly high. This temperature, called TET, can be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine blade, which itself can be called the nozzle guide blade. Under cruising conditions, the TET can be at least (or of the order of) any of the following values: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET under cruising conditions may be in an inclusive range delimited by any two of the values in the previous sentence (i.e. the values may form upper or lower limits). The maximum TET in use of the engine can be, for example, at least (or of the order of) any of the following values: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET can be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values can form upper or lower limits), for example in the range from 1800K to 1950K. Maximum TET may occur, for example, in a high thrust condition, for example in a maximum take-off thrust (PMD) condition.
[0244] As used herein, a maximum take-off thrust (PMD) condition has the classic meaning. Maximum take-off thrust conditions can be defined as engine operation under conditions of pressure and temperature at sea level of the International Standard Atmosphere (ISA) + 15 ° C at maximum take-off thrust at the end of runway, which is typically defined at an aircraft speed of about 0.25Mn (i.e. a Mach number of 0.25), or between about 0.24 and 0.27Mn. The maximum take-off thrust conditions for the engine can therefore be defined as running the engine at maximum take-off thrust for the engine at the pressure and temperature ISA at sea level + 15 ° C at a speed of d. aircraft of 0.25Mn.
[0245] A fan blade and / or airfoil portion of a fan blade described and / or claimed herein can be made from any suitable material or combination of materials. For example at least part of the fan blade and / or of the airfoil can be made at least in part from a composite, for example a metal matrix composite and / or an organic matrix composite, such as 'a carbon fiber. As a further example at least part of the fan blade and / or the airfoil can be fabricated at least in part from a metal, such as a titanium-based metal or a titanium-based material. aluminum (such as an aluminum-lithium alloy) or a steel-based material. The fan blade can include at least two regions made from different materials. For example, the blower blade can have a protective leading edge, which can be made using a material that is more able to withstand impact (e.g. from birds, ice, or other material) than the rest of dawn. Such a leading edge can, for example, be made using titanium or a titanium-based alloy. Thus, strictly by way of example, the fan blade can have a carbon fiber or aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.
[0246] A fan as described and / or claimed herein may include a central portion, from which the fan vanes may extend, for example in a radial direction. The fan blades can be connected to the central part in any desired way. For example, each fan blade can include a fastener which can engage with a corresponding notch in the hub (or disc). Strictly by way of example, such a fastening element may be in the form of a dovetail which can notch into and / or engage with a corresponding notch in the hub / disc in order to secure the hub / disc. hub / disc blower blade. As a further example, the fan blades may be integrally formed with a central portion. Such an arrangement may be referred to as a bladed disc or bladed crown. Any suitable method can be used to manufacture such a bladed disc or such a bladed crown. For example, at least part of the fan blades can be machined from a block and / or at least part of the fan blades can be joined to the hub / disc by welding, such as linear friction welding.
[0247] The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable-section nozzle can make it possible to vary the outlet area of the bypass duct during use. The general principles of this description can be applied to engines with or without VAN.
[0248] The fan of a gas turbine as described and / or claimed herein can have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.
[0249] As used herein, cruising conditions have the conventional meaning and would be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, one skilled in the art would immediately recognize that cruising conditions signify the engine operating point at mid-cruising of a given mission (which may be referred to in the industry as an "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this sense, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between the end of climb and the start of descent has been burned (which can be approximated by the midpoint - in terms of time and / or distance - between the end of climb and the start of descent. Cruising conditions thus define an operating point of the gas turbine engine that provides thrust that would ensure steady-state operation (i.e. the maintenance of a constant altitude and a constant Mach, Mn, number) at mid-cruise of an aircraft to which it is designed to be attached, in taking into account the number of engines supplied on that aircraft.For example, when an engine is designed to be attached to an aircraft which has two engines of the same type, at cruising conditions the engine provides half of the total thrust that would be required for steady-state operation of this aircraft at mid-cruise.
[0250] In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the engine operating point which provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady-state operation of the aircraft at which it is designed to be set at a given mid-cruise Mach number) at mid-cruise atmospheric conditions (defined by the international standard atmosphere according to ISO2533 at mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and therefore the engine operating point at cruise conditions is clearly defined.
[0251] Strictly as an example, the forward speed at cruising condition can be any point in the range of Mach0.7 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 of the order of Mach 0.8, of the order of Mach 0.85 or in the range of 0.8 to 0.85. Any single speed within these ranges can be part of the cruise condition. For some aircraft, cruising conditions may be outside of these ranges, for example below Mach0.7 or above Mach0.9.
[0252] Strictly by way of example, the cruising conditions may correspond to typical atmospheric conditions (according to the international standard atmosphere, ISA) at an altitude which is in the range of 10000m to 15000m, for example in the range of 10000m at 12000m, for example in the range from 10400m to 11600m (roughly 38000ft), for example in the range from 10500m to 11500m, for example in the range from 10600m to 11400m, for example in the range from 10700m (approximately 35000ft) to 11300m, for example in the range from 10800m to 11200m, for example in the range from 10900m to 11100m, for example of the order of 11000m. Cruising conditions can be typical atmospheric conditions at any given altitude within these ranges.
[0253] Strictly as an example, cruising conditions may correspond to an engine operating point that provides a known required level of thrust (e.g. a value in the range 30kN to 35kN) at a forward Mach number of 0 , 8 and typical atmospheric conditions (according to the international standard atmosphere) at an altitude of 38000ft (11582m). Strictly as a further example, cruising conditions may correspond to an engine operating point which provides a known required level of thrust (e.g. a value in the range 50kN to 65kN) at a Mach number before 0.85 and typical atmospheric conditions (according to the international standard atmosphere) at an altitude of 35,000 feet (10668m).
[0254] In use, a gas turbine engine described and / or claimed herein can operate at the cruising conditions defined elsewhere in this document. Such cruise conditions may be determined by the cruise conditions (e.g. mid-cruise conditions) of an aircraft to which at least one (e.g. 2 or 4) gas turbine engine can be mounted in order to provide propulsion thrust.
[0255] In 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 the aircraft for which the gas turbine engine has been designed to be attached. Thus, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere in this document.
[0256] In one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. Operation may be at cruise conditions as defined elsewhere in this document (eg in terms of thrust, atmospheric conditions and Mach number).
[0257] In one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. Operation according to this aspect may include (or may be) mid-cruise operation of the aircraft, as defined elsewhere in this document.
[0258] Those skilled in the art would understand that, except mutually exclusive, a characteristic or a parameter described in relation to any of the above aspects can be applied to any other aspect. Moreover, except mutual exclusivity, any characteristic or any parameter described here can be applied to any aspect and / or associated with any other characteristic or any other parameter described here.
[0259] Embodiments will now be described by way of example only, with reference to the Figures, in which:
[0260] is a sectional side view of a gas turbine engine;
[0261] is a close-up sectional side view of an upstream portion of a gas turbine engine;
[0262] is a partially cut-away view of a reducer for a gas turbine engine;
[0263] is a sectional side view of a gas turbine engine showing three bearings on the heart shaft;
[0264] is a schematic side view of a gas turbine engine showing the major and minor spans between the bearings;
[0265] illustrates core shaft bending modes for two different bearing configurations;
[0266] is a schematic side view of a gas turbine engine illustrating an alternative bearing arrangement to that shown in Figure 5;
[0267] is a schematic view of the mounting of two bearings on a tail bearing housing;
[0268] is a schematic view of the mounting of two bearings on a tail bearing housing of a different arrangement from that shown in Figure 8;
[0269] illustrates a method of operating a gas turbine engine;
[0270] illustrates the first four resonant frequencies of a beam;
[0271] is a close-up view of the bearing housing shown in Figure 5;
[0272] is a perspective view of a first bearing support structure;
[0273] is a cross-sectional view of the first bearing support structure shown in Figure 13 in position within a bearing housing;
[0274] and illustrate a determination of bearing stiffness, in particular showing the application of a radial force and the resulting displacement;
[0275] illustrates a method of obtaining a gas turbine engine as described here;
[0276] is a graph of displacement as a function of load, illustrating an elastic region within which stiffnesses of components can be determined;
[0277] illustrates a method of designing a gas turbine engine as described herein;
[0278] is a schematic view of the mounting of two bearings shown in Figure 9, with bearing disc angles marked; and
[0279] is a schematic view of an arrangement of two different bearings than that shown in Figure 9, with corresponding bearing disc angles marked.
[0280] Figure 1 illustrates a gas turbine engine 10 having a primary axis of rotation 9. The engine10 includes an air intake12 and a propulsion blower23 which generates two air flows: a core air flow A and a bypass air flow B. The gas turbine engine10 includes a core11 which receives the airflow from coreA. The engine core11 comprises, in axial flow series, a low pressure compressor14, a high pressure compressor15, combustion equipment16, a high pressure turbine17, a low pressure turbine19 and a core exhaust nozzle20. A nacelle21 surrounds the gas turbine engine10 and defines a bypass duct22 and a bypass exhaust nozzle18. The bypass airflow B flows through the bypass duct 22. The blower23 is attached to and driven by the low pressure turbine19 via a heart shaft26 and an epicyclic reduction gear30.
[0281] In use, the core air flow A is accelerated and compressed by the low pressure compressor14 and directed into the high pressure compressor15 where further compression takes place. The compressed air discharged from the high pressure compressor15 is directed into the combustion equipment16 where it is mixed with fuel and the mixture is burned. The resulting hot combustion products then expand, and thereby drive the high pressure and low pressure turbines17,19 before being vented through the nozzle 20 to provide some propellant thrust. The high pressure turbine17 drives the high pressure compressor15 through a suitable interconnection shaft27. The blower23 typically provides the majority of the propulsive thrust. The epicyclic reduction gear 30 is a reduction box.
[0282] An exemplary arrangement for a gear blower gas turbine engine is shown in Figure 2. The low pressure turbine19 (see Figure 1) drives the heart shaft 26, which is coupled to a planetary gear, or planetary gear, 28 of the epicyclic gear arrangement30. Radially outwardly of and meshing with planetary gear28, there are a plurality of planet gears32 which are coupled together by a planet carrier34. The planet carrier34 forces the planet gears32 to change orientation around the planetary gear28 in synchronism while allowing each planet gear32 to rotate around its own axis. The planet carrier 34 is coupled via links 36 to the fan 23 in order to cause its rotation around the motor axis9. Radially outward from the planetary gears32 and meshing with them, there is a ring or ring gear38 which is coupled, via links40, to a fixed structure24.
[0283] It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein can be taken to indicate lower pressure turbine stages and lower pressure compressor stages (ie. i.e. not including the blower23) respectively and / or the turbine and compressor stages which are connected together by the interconnecting shaft with the lowest rotational speed in the engine (i.e. i.e. not including the gearbox output shaft which drives the blower23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as "intermediate pressure turbine" and "intermediate pressure compressor". When such an alternative nomenclature is used, the blower 23 may be referred to as the first compression stage or the lower pressure compression stage.
[0284] The epicyclic reduction gear 30 is exemplified in more detail in Figure 3. Each of the planetary gear28, planetary gears32, and ring gear38 includes teeth around its periphery to mesh with the other gears. However, for clarity, only exemplary portions of the teeth are shown in Figure 3. There are four planet gears32 illustrated, it will, however, be apparent to the skilled reader that more or less planet gears32 can be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic reduction gear30 typically include at least three planetary gears32.
[0285] The epicyclic reduction gear 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via links 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic reducer can be used. As a further example, the epicyclic reduction gear 30 can be a star arrangement, in which the planet carrier 34 is held stationary, with the ring gear (or ring) tooth 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, reducer 30 may be a differential reducer in which ring gear38 and planet carrier34 are both allowed to rotate.
[0286] It will be understood that the arrangement shown in Figures 2 and 3 is by way of example only, and that various alternatives are within the scope of the present description. Strictly by way of example, any suitable arrangement can be used to position reducer30 in motor10 and / or to connect reducer30 to motor10. As an additional example, the connections (such as links36, 40 in the example of Figure2) between the reducer30 and other parts of the motor10 (such as the input shaft (heart shaft26), the (output shaft and fixed structure24) can have any desired degree of stiffness or flexibility. As a further example, any suitable arrangement of the bearings between rotating and stationary parts of the motor (for example between the input and output shafts from the reducer and fixed structures, such as the reducer housing) can be used, and the description is not limited to the exemplary arrangement of Figure 2. For example, when the reducer 30 has a star arrangement (described above), one skilled in the art would readily understand that the arrangement of the output and support links and the bearing locations would typically be different from that shown for reference. example in Figure 2.
[0287] Thus, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (eg, star or planetary), support structures, input shaft arrangement, and gear arrangement. exit, and landing locations.
[0288] Optionally, the reducer can drive additional and / or alternative components (for example the intermediate pressure compressor and / or a booster).
[0289] Other gas turbine engines to which the present description can be applied may have alternative configurations. For example, such engines can have another number of compressors and / or turbines and / or another number of interconnection shafts. As a further example, the gas turbine engine shown in Figure 1 has a split-flow nozzle18, 20 which means that the flow through the bypass duct22 has its own nozzle18 which is independent of, and radially to the outside, the core engine nozzle20. However, this is not limiting, and any aspect of the present description can also be applied to engines in which the flow through the bypass 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 and / or the other of the nozzles (whether they are with mixed or divided flow) can have a fixed or variable area. While the example described relates to a turbofan engine, the description can apply, for example, to any type of gas turbine engine, such as an open rotor (in which the stage fan is not surrounded by a nacelle) or a turboprop, for example. In some arrangements, the gas turbine engine10 may not include a reduction gear30.
[0290] The geometry of the gas turbine engine10, and its components, is defined by a conventional axis system, comprising an axial direction (which is aligned with the axis of rotation9), a radial direction (in the direction from bottom to top in Figure1) and a circumferential direction (perpendicular to the page in the view of Figure1). The axial, radial and circumferential directions are mutually perpendicular.
[0291] In the embodiments which are described, the motor 10 is a geared gas turbine engine, having a reduction gear 30.
[0292] The blower23 is attached to and driven by the low pressure turbine19 via the heart shaft26 and an epicyclic reduction gear30. The heart shaft26 is supported by three bearings26a, 26b, 26c, the three bearings comprising a front bearing26a and two rear bearings26b, 26c. In alternative embodiments, more bearings 26a-c can be provided. The front bearing 26a may be designated as the first bearing, the forward-most 26b of the two rear bearings 26b, 26c as the second bearing, and the rearmost 26c of the two rear bearings 26b, 26c as the third bearing.
[0293] The front bearing26a is the positioning bearing26a for the heart shaft26; that is, it is the bearing on the heart shaft arranged to limit axial movement of the heart shaft 26 as well as radial movement, thereby positioning the heart shaft axially. The front bearing 26a is mounted on the fixed structure so as to axially position the heart shaft 26. Those skilled in the art will understand that having multiple locating bearings on a single shaft can cause the shaft to detrimentally flex, or otherwise deform, upon expansion in use. and that the use of a single positioning bearing per shaft is for this reason generally favored.
[0294] In the embodiment which is described, the front bearing 26a is the heart shaft bearing 26a closest to the front of the motor 10. In alternative embodiments, there may be one or more other bearings, for example roller bearings, on the heart shaft26 in front of the front bearing26a, however such possible bearings are not positioning bearings - that is, although they may help to radially position the heart shaft26, the heart shaft can move axially with respect to these bearings.
[0295] In the embodiments which are described, the front bearing 26a is the closest core shaft bearing 26a, and to the rear, of the reducer 30. The front bearing 26a is axially flush with, or near, the outlet of the compressor 14 in the embodiment which is described. In alternative embodiments, the front bearing 26a may be located in front of the reducer 30, however the proximity of the front bearing 26a to a roller bearing of the fan 23 may increase the complexity.
[0296] The front bearing 26a is mounted on the fixed structure.
[0297] The rear bearings 26b, 26c are the next two heart shaft bearings, behind the front bearing 26a. In the embodiment which is described, the rear bearings 26b, 26c are the heart shaft bearings closest to the rear of the motor 10. In alternative embodiments, an additional bearing may be located at the rear of these bearings.
[0298] In the embodiment shown in Fig. 4, both of the rear bearings 26b, 26c are mounted on the tail bearing housing 29. The tail bearing housing 29 is a structure arranged to be non-rotatable with respect to the fixed structure, and to support the bearings 26b, 26c of the heart shaft 26. The tail bearing housing 29 includes two bearing discs 29a, 29b. Each disc 29a, 29b is arranged to support one of the two rear bearings 26b, 26c.
[0299] In the embodiment shown in Figure 4, the rearmost bearing 26c of the two rear bearings 26b, 26c is located axially level with, or near, the outlet of the low pressure turbine 19. More specifically, the rearmost bearing 26c is at least substantially axially level with a rearmost / lower pressure rotor of the low pressure turbine 19 in the embodiment which is described.
[0300] In various alternative embodiments, such as that shown in Figures 5-7, one and / or the other of the two rear bearings 26b, 26c may be provided axially downstream of the leading edge of the blade assembly. lowest pressure19d of the low pressure turbine19, at the level of the root of the blade.
[0301] In embodiments such as that shown in Figures 1 and 4, the low pressure turbine 19 may have three stages; i.e. three sets of rotor blades 19a, 19b, 19c. Each set of rotor blades has a corresponding axial position along the motor axis9, and is offset from the axial position of the other sets. The most upstream set of blades, or the most axially forward, is the highest pressure set of blades of the low pressure turbine19, and may be referred to as the first set of blades. The most downstream, or most axially aft, blade set is the lowest pressure blade set of the low pressure turbine19, and may be referred to as the latter, or in these embodiments. the third, set of blades. The middle blade set 19b of the three sets may be referred to as the second set of blades.
[0302] In embodiments such as that shown in Figures 5 and 6, the low pressure turbine 19 may have four stages; i.e. four sets of rotor blades 19a, 19b, 19c, 19d. The most upstream, or most axially forward, set of vanes is the highest pressure set of vanes of the low pressure turbine19, and may be referred to as the first set of vanes19a. The most downstream, or most axially rearward, blade set19d is the lowest pressure blade set of the low pressure turbine1919, and may be referred to as the latter, or in these embodiments. the fourth, set of blades. The middle vane sets 19b, 19c of the four sets may be designated second and third vane sets, respectively.
[0303] In alternative embodiments, the turbine 19 may have less than three or more than four sets of blades, for example having two sets or five sets.
[0304] A length of the low pressure turbine19 can be defined as the distance between a leading edge of a blade19a of the first set of blades of the low pressure turbine and a trailing edge of a blade19c / 19d of the last set of blades. low pressure turbine blades. A low pressure turbine housing19 may extend beyond the span between the first (higher pressure) and last (lower pressure) vanes.
[0305] In various embodiments with the two rear bearings26b, 26c provided axially downstream of the leading edge of the lower pressure vane19c, 19d of the low pressure turbine19 at the root of the vane, the length ratio (as described below) and / or heart shaft running speed may or may not be within the ranges detailed elsewhere in this document. In such embodiments, the two rear bearings 26b, 26c may be provided axially downstream of the centerline of a disc supporting the lower pressure turbine blade19d of the low pressure turbine19, as illustrated in Figure 12. In such embodiments, the length, L, of the core shaft may be in the range of 1800-2900mm, optionally 2300-2800mm, and optionally further 2400-2750mm.
[0306] In various embodiments, the low pressure turbine19 may be a four-stage turbine19, having four sets19a-d of rotor blades, for example as previously described with respect to Figures 5 and 6. The set of rotor blades 19 may be. further forward19a can be defined as the first set and the furthest one19d as the fourth set. The two rear bearings 26b, 26c can be provided axially downstream of the trailing edge of a low pressure turbine blade of the third set 19c at its root. Either of the two rear bearings26b, 26c may be provided axially upstream of the leading edge of the lower pressure turbine blade19d (a blade of the fourth set) of the low pressure turbine19 at the level of the root of the dawn in such embodiments. The pressure decreases across the low pressure turbine19 - the first set of blades 19a can therefore be described as the higher pressure set of blades of the low pressure turbine19, and the fourth set of blades as lower pressure vanes.
[0307] In various such embodiments, the heart shaft length ratio and / or speed may or may not be within the ranges detailed elsewhere in this document. In such embodiments, the length, L, of the core shaft may be in the range of 1800-2900mm, optionally 2300-2800mm, and optionally further 2400-2750mm.
[0308] In the low pressure turbine19, the turbine blades19a-d within each assembly are designed to be identical (within manufacturing tolerances). The vanes may differ, for example in size and / or shape, between different sets. Each vane within a set19d has a mass, m , and a mid-height turbine vane radius, r . Generally, the heavier and larger blades are in the assemblies towards the rear of the turbine. The turbine blades each have a half-height position, at which the half-height radius is measured. The mid-height position is midway between a radially innermost point and a radially outermost point on the blade leading edge. The mid-height turbine blade radius is measured in a radial direction between an axial centerline9 of the engine10 and the mid-height position.
[0309] Each vane 19a-d also has a maximum nominal angular speed, ω, which may also be referred to as a maximum take-off thrust speed (PMD). SpeedPMD can be the maximum angular speed at which the shaft is designed to rotate. ω can be between 5000 and 9000rpm, optionally in the range of 5000-8000rpm or 5000-7000rpm, and possibly around 5500, 6000 (i.e. around 630radians per second) , or 6500rpm.
[0310] A value, Y , can be defined as follows:
[0311]
[0312] Y can have units of kg.m.rad ² .s ^-2 . In some embodiments, the value of Y for the lowest pressure blade assembly19c (for a three-stage turbine), 19d (for a four-stage turbine) of the low pressure turbine19 is in the range of 45000 to 100000kg.m.rad ² .s ^-2 , optionally in the range from 50,000 to 100,000kg.m.rad ² .s ^-2 , optionally in the range from 55000 to 100,000kg.m.rad ² .s ^-2 , and optionally further in the range from 60,000 to 100,000kg.m.rad ² .s ^-2 . In such embodiments, the mass, m , of each vane of the last set may be in the range of 0.2kg to 0.6kg, and optionally may be about 0.4kg. The radius, r , of each vane of the last set at half height can be in the range of 400mm to 600mm, and optionally can be around 500mm (0.5m).
[0313] The value of Y can be represented as providing a measure of the order of magnitude of the centripetal force ( F _c ) acting on the vane as it rotates at the speed PMD:
[0314]
[0315] where a _c is the centripetal acceleration and v is the linear speed, which is equal to ωr .
[0316] Those skilled in the art will understand that a higher centripetal force ( F _c ) acting on the vane can increase the risk of a vane separation event for that vane.
[0317] The minor span, S, is the distance between the two rearmost bearings, and can be in the range of 250mm to 350mm, possibly in the range of 275mm to 325mm, and possibly additionally can be around 300mm.
[0318] A first ratio of the vane to the landing can be defined as:
[0319]
[0320] The first vane-to-bearing ratio can be in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² , and optionally in the range of 3.0 x10 ^{- 6} to 4.5x10 ^-6 kg ^-1 .rad ^-2 .s ² . The value of this ratio may be lower than that for conventional engines, because the minor range, S, is smaller compared to the value of Y. Those skilled in the art will understand that increasing the angular speed in flight can improve motor efficiency, and that speedPMD provides a measure of maximum rotational speed, and therefore maximum force (indicated by Y), available by a motor10. However, the inventor was aware that the motor10 should not be linearly scaled up with an increase in force (Y providing a measure of force), but rather that the minor span length should be increased by as little as possible. so as to relatively reduce the length and weight of the motor, thereby allowing the efficiency gains to be increased by avoiding the extra weight, and to avoid the development of undesirable vortex modes within the minor span. While common sense suggests that a greater minor range is desirable to improve a reaction of forces from the low pressure turbine, 19 the inventor has found that the risk of introducing vortex modes, and the introduction of more length and weight, outweighed the benefits of force reaction.
[0321] In alternative or additional embodiments, a second vane to bearing ratio can be defined as:
[0322]
[0323] The value of the vane mass, m , multiplied by the vane radius, r , can be in the range of 180 to 280 kg.mm. The minor span, S, i.e. the distance between the two rearmost bearings, can be in the range of 250mm to 350mm. In various embodiments, the minor span can be greater than or equal to any of 250mm, 255mm, 260mm, and 265mm. In various embodiments, the minor span can be less than or equal to any one of 350mm, 345mm, 340mm, or 335mm. The second vane-to-bearing ratio may be in the range of 0.8 to 6.0kg ^-1 , and optionally in the range of 0.9 to 3.9kg ^-1 , and optionally further in the range of from 1.2 to 2.6 kg ^-1 . In such embodiments, the gear ratio of the reducer may be greater than 3, and optionally in the range of 3.1 to 3.8.
[0324] In such embodiments, the running speed of the motor10 / core shaft26 may be in the range of 5400 to 5700rpm (i.e. about 565 to 597radians per second), and optionally from 5500 to 5600rpm, cruising. Additionally or alternatively, the running speed of the heart shaft26 may be in the range of 5800 to 6200rpm, and optionally 5900 to 6100rpm, at PMD.
[0325] In such embodiments, the length, L, of the core shaft 26 may be in the range of 1800 to 2900mm, optionally 2000 to 2900mm, optionally further 2300 to 2800mm, and optionally additionally 2400 to 2750mm.
[0326] In particular, in the embodiment shown in Figures 5 and 6, both of the rear bearings 26b, 26c are mounted on the tail bearing housing 29. The tail bearing housing29 is located axially to the rear of the low pressure turbine19. Two bearing discs 29a, 29b extend from the tail bearing housing 29. While the foremost bearing disc 29a of the bearing housing 29 of this embodiment is angled forward of the tail bearing housing 29, extending towards and, in this case, within the housing of, the low pressure turbine19, the forwardmost bearing26b of the rear bearings, which is mounted on the forwardmost bearing disc29a, is nevertheless located axially aft of the leading edge of the lower turbine blade pressure of the low-pressure turbine19, at the level of the root of the blade. The rearmost bearing26c of the rear bearings, which is mounted on the rearmost bearing disc29b, is located axially aft of the low pressure turbine19.
[0327] In the embodiment shown in Fig. 7, the tail bearing housing 29 has just one bearing disc 29c extending therefrom. The single bearing disc 29c is located in a similar position and has a shape similar to the foremost bearing disc 29a of the embodiment shown in Figures 5 and 6, but this time supports the rearmost bearing 26c of the bearings. rear, maintaining this bearing26c axially aft of the leading edge of the lower pressure turbine blade of the low pressure turbine19. The forward-most bearing26b of the rear bearings is mounted on a front bearing housing31 rather than on the tail bearing housing29. The front bearing housing 31 has a single bearing disc 31a arranged to support the forward most one 26b of the rear bearings. The forwardmost one26b of the rear bearings is located forward of, and adjacent to, the low pressure turbine19 in this embodiment. The forwardmost one26b of the rear bearings is maintained axially level with part of the low pressure turbine19, and more specifically axially level with the leading edge of the higher pressure turbine blade of the low pressure turbine19, in this embodiment. The core shaft26 may be shorter overall in such embodiments with a single rear bearing26c behind the leading edge of the lower pressure turbine blade of the low pressure turbine19, because the shaft core26 cannot extend as far aft of the low pressure turbine19 given that the minor seat, S, is located at least partially axially level with the low pressure turbine19, rather than strictly at the rear of the low pressure turbine19 (more specifically, from the leading edge of the lower pressure turbine blades of the low pressure turbine19).
[0328] A length, L, of the heart shaft26 can be defined between the front bearing26a and the rearmost rear bearing26c, as marked in Figures 5 and 6 (i.e. between the first and third bearings ). This length, L, can be referred to as the major span of the heart tree26, or the length between bearings of the heart tree26. In the embodiment which is described, the length, L, of the heart shaft26 is in the range of 1800 to 2900mm, more particularly in the range of 2300 to 2800mm, and even more particularly in the range of 2400 to 2750mm. .
[0329] The minor span, S, of the heart shaft26 can be defined as the distance between the rear bearings26b, 26c (i.e. between the second and third bearings on the heart shaft26).
[0330] The minor bearing, S, is equal to the length, L, of the heart shaft26 minus the distance, D, between the front bearing26a and the first rear bearing26b (i.e. minus the distance between the first and second bearings on the heart shaft26).
[0331] A heart tree length ratio can be defined as:
[0332]
[0333] In various embodiments, the bearings 26a-26c are arranged such that the ratio of the length of the minor seat to the core shaft length is in the range of 0.08 to 0.14, and optionally in the range of. range from 0.08 to 0.13. In various embodiments, the heart tree 26 can have any suitable length ratio, for example in the range of 0.09 to 0.13, or 0.10 to 0.12, or for example being of the order of or at least 0.08, 0.09, 0.10, 0.11, 0.12, 0.13 or 0.14. The heart tree length ratio can be, for example, between any two of the values in the previous sentence.
[0334] The inventor was aware that the ratio of the distance between the two rear bearings26b, 26c to the shaft length, L, can be a significant parameter for controlling the rotodynamic behavior of the heart shaft26.
[0335] The value of this aspect ratio for motors of various embodiments can be relatively low compared to previous motors.
[0336] As the motor10 gets bigger, the heart shaft26 gets longer - however, an increase in the distance, S, between the two rear bearings26b-c (the minor throw) linearly with the increase in shaft length , L, may cause detrimental effects. Once the minor span, S, reaches a certain length, there may be no benefit to extending this distance in terms of effects on the part of the heartshaft26 located between the first26a and second26b bearings and / or in terms of reaction to the moments of the bearing discs 29a-b, 31a. Indeed, if this distance, S, is linearly increased with the length of the heart shaft, L, the minor span can become long enough to have vortex modes (as described below with respect to Figure 11) between them. second and third stages26b, within the engine operating range, thereby potentially worsening unwanted movement.
[0337] Additionally or alternatively, a relatively large tail bearing housing (TBH) 29 can be used to accommodate the larger minor bearing, S, between the bearings 26b, 26c, as shown in Figure 8. A larger TBH29 may add unwanted weight and / or bulk to the motor10. Alternatively or additionally, the bearing discs 29a, 29b may be bent to a greater extent to accommodate the larger minor span between the bearings 26b, 26c, as shown in Figure 9. The bearing discs 29a, 29b being bent further out of a direction perpendicular to the motor axis 9 / parallel to a spoke may be a structurally suboptimal solution - e.g. the timing of the bearings. Bearing discs may not react as well as in the embodiment shown in Figure 8. Alternatively or additionally, one of the bearings 29a, 29b may be located on a different bearing structure 31, instead of the TBH29.
[0338] In various embodiments, the bearing disks29a, b, 31a, whether on the TBH29 or on an independent bearing structure31, may be oriented at least substantially perpendicular to the motor axis 9, for example having an angle (Θ) with respect to a radius of the motor10 between 0 ° and 20 °, and possibly between 0 ° and 15 ° (i.e. an angle with respect to the axis9 of the motor between 90 ° and 70 °, and possibly between 90 ° and 75 °).
[0339] In embodiments such as that shown in Figures 9 and 19, the bearing discs 29a, 29b each extend inwardly toward the motor axis 9 from the rest of the TBH29 at an at least substantially constant angle. (Bearing discs 29a, 29b may, for example, comprise a solid disc or a series of cross members spaced circumferentially around the axis extending between inner and outer rings). An angle, Θ ₁ , Θ ₂ , can therefore be defined between the disc 29a, 29b and a radial direction. In other embodiments, such as those shown in Fig. 20, a bearing disc 29a, 29b may include two or more parts extending at different angles (as shown by way of example only with three parts for the disc. front-most bearing29a and two parts of the rear bearing disc29b in Figure20). In such cases, the angle selected is that for the part with the longest radial extent, as marked for the two examples shown in Figure 20. In embodiments in which there is no single part with a greater radial extent than the other part (s), an average angle may instead be taken across the two or more parts with the. greater radial extent.
[0340] In the embodiments which are described, the heart shaft26 has an operating speed range with a lower limit of 1500rpm (e.g. at ground idle) and an upper limit of 6200rpm (e.g. at ground idle). maximum take-off thrust - PMD). In particular, in various embodiments the cruising heart shaft running speed is in the range of 5400 to 5700rpm, and optionally in the range of 5500 to 5600rpm. At PMD, the heart shaft running speed may be in the range of 5800 to 6200rpm, and possibly in the range of 5900 to 6100rpm. The cruising speed range for a particular aircraft is generally well below a maximum rated rotational speed for the heart shaft26 of that aircraft (the PMD speed). The motor10 can operate within the PMD run speed range for relatively short periods of time - for example five or ten minutes - during normal operation.
[0341] Figure 6 illustrates the stiffening effect on the heart shaft 26 obtained by the arrangement of bearings 26a-26c as described here, and in particular the use of two rear bearings 26b, 26c. Positioning of the bearings is selected to control, or help control, the vortex modes of the motor 10.
[0342] In particular, the boundary condition on the heart shaft26 at the location of the second bearing26b (the foremost rear bearing) can be changed by extending the heart shaft26 beyond the second bearing26b to third bearing26c (the rearmost rear bearing). This can change the boundary condition at the second soar26b from a simply supported / immobilized boundary condition to a cantilever boundary condition, affecting the shape of the heart tree26 as it flexes. This effect can be seen most clearly immediately to the left of the second bearing26b as shown in Figure 6 - in the upper two-bearing configuration (immobilized-immobilized limit conditions), the angle of the heart shaft at the second bearing26b is steeper than in the lower three-bearing configuration (immobilized-cantilever boundary condition), in which the shaft26 is closer to horizontal when entering the second bearing26b.
[0343] In rotodynamics, the critical speed is the theoretical angular speed which excites the natural frequency of a rotating object, such as a shaft (like the heart shaft26). Higher frequency vortex modes can also be induced (eg twice the natural / resonant frequency). Motors10 of the embodiments which are described may be susceptible to vortex modes (resonant frequencies) within or near the motor operating range, which can cause unwanted and potentially damaging movement of the heart shaft26 , as shown in Figure 6. This coincidence of vortex modes with the engine operating range may be a result of the longer heart shaft26 of the larger engine10.
[0344] While the diameter and / or thickness of the heart shaft26 could be changed to increase stiffness and therefore push the vortex modes outside the engine operating range, the resulting increase in size and / or weight , and / or impact on other components, may eliminate or reduce the feasibility of this option. For example, the heart shaft diameter may be constrained by other engine components located radially outward from the heart shaft26 (for example, the heart shaft26 is radially inward of the interconnection shaft 27 connecting the second turbine (high pressure) 17 to the second compressor (high pressure) 15 in the embodiments which are described).
[0345] In the embodiments which are described, the low pressure turbine 19 drives a low pressure compressor 14 directly, and drives a fan 23 indirectly via a reduction box 30. A higher pressure system15, 17, comprising a high pressure compressor15 and a high pressure turbine17, as well as an interconnection shaft27 between them, is located between the low pressure compressor14 and the low pressure turbine19. The heart shaft26 is therefore longer than the interconnect shaft27, as it extends the full length of the higher pressure system15, 17 and the additional length of the low pressure system14, 19.
[0346] The relatively large length, L, of the heart shaft26, compared to previous motor architectures, reduces the natural frequency of the heart shaft, bringing the natural frequency within the operating speed of the motor10, because the frequency natural is inversely proportional to the major span (length of heart shaft L).
[0347] Natural frequencies exist in various modes (which can be referred to as vortex modes), as shown in Figure 11. A primary resonance (mode1) has a total of two nodes (non-moving points), one at each constrained end (the bearings26a, 26b on the shaft26). A secondary resonance (mode2) has an additional node in the center (a total of three nodes). A third resonance mode (mode3) has four nodes, and a fourth resonance mode (mode4) has five nodes, etc. The nodes are evenly spaced along the span of the shaft26 between the constrained ends26a, 26b. The amplitude decreases with increasing mode number - the maximum amplitude of the mode1 resonance is greater than that of the mode2 resonance, etc. Primary resonance can therefore be the most damaging, because it causes the maximum radial displacement of the shaft26. Avoiding operating speeds that could trigger primary resonance - by moving the primary resonance out of the operating speed range - can therefore be of particular interest.
[0348] The relation of the natural frequency of a beam with a simple support (immobilization) to the length of the beam is given by:
[0349]
[0350] or:
[0351] n is the mode number (with mode1 having nodes at each end immobilized only, as shown in Figure6, mode2 having an additional node at the midpoint between the ends, etc.);
[0352] f _n is the frequency of the n ^th mode (the resonant frequency), measured in Hz;
[0353] K _n is a dimensionless factor dependent on the order of the mode, n, and also on the applicable boundary conditions, and can be derived for various boundary conditions as detailed, for example, in "Roark's Formulas for Stress and Strain ", Warren C. Young and Richard G. Budynas, Seventh Edition.
[0354] E is the modulus of elasticity of the beam (N / m ² );
[0355] I is the surface moment of inertia of the beam (m ⁴ );
[0356] g is the acceleration due to gravity (m / s ² );
[0357] l is the length of the beam between the immobilized ends (m); and
[0358] w is the load per unit length on the core shaft (N / m).
[0359] . Par exemple, pour le deuxième mode,n=2etK _n =(2π)²=39,5 (à trois chiffres significatifs).The nodal positions (depending on the length of the beam, l ) and the values of K _n for the first five modes can be as presented below in Table 1. The data in Table1 are for a beam with both ends immobilized (the boundary conditions), and K _n is therefore equal to. For example, for the second mode, n = 2 and K _n = (2π) ² = 39.5 (three significant digits).
[0360] [Table 1]
[0361] Table1: Fashion K _nNodal position / l 1 9.87 0.0 1.00 2 39.5 0.0 0.50 1.00 3 88.8 0.0 0.33 0.67 1.004 158 0.0 0.25 0.50 0.75 1.00 5 247 0.0 0.20 0.40 0.60 0.80 1.00
[0362] The bending stiffness of the heart shaft26 is a function of both the material property and the surface moment of inertia, I (geometric). The surface moment of inertia for a tubular structure, such as the heart shaft26, is:
[0363]
[0364] where r ₁ and r ₂ are the inner and outer radii of the tube, respectively.
[0365] In order to increase the frequency, the options available include using a more rigid material for the shaft26, decreasing the length, L , of the shaft26, or increasing the shaft diameter. Shaft diameter and length are related to frequency by a quadratic function and therefore have the most influence for a change of a given order of magnitude. However, the diameter is limited by a lack of available space within the engine10.
[0366] The provision of a second rear bearing26c changes the boundary conditions to simulate an overhang, as mentioned earlier - this change in boundary condition has a stiffening effect on the shaft26, increasing its natural frequency. The change in boundary condition is reflected in a change in K _n , as shown in Table 2. In particular, the beam now has just one end immobilized (simply supported), and the other fixed, so
[0367]
[0368] For the example of n = 2, K _n is therefore (2.25) ² π ² = 50.0 (three significant digits).
[0369] [Table 2]
[0370] Table2 Fashion K _nNodal position / l 1 15.4 0.0 1,000 2 50.0 0.0 0.557 1,000 3 104 0.0 0.386 0.692 1,0004 178 0.0 0.295 0.529 0.765 1,000 5 272 0.0 0.239 0.428 0.619 0.810 1,000
[0371] Assuming that all parameters except K _n are constants, it can be seen that a simple change to the boundary condition could increase the first natural frequency by more than 50%.
[0372] The relative positioning of the heart shaft bearings 26a-c can therefore be used to adjust the stiffening effect to control the resonant frequencies.
[0373] A ratio of minor throw to turbine length can be defined as:
[0374]
[0375] In various embodiments, the ratio of minor seat to turbine length can be equal to or less than 1.05, optionally in the range of 0.85 to 1.05, and optionally further from 0.85 to 0 , 95.
[0376] In such embodiments, the heart shaft length ratio and / or speed may or may not be within the ranges detailed elsewhere in this document. Similarly, the two rear bearings 26b, 26c may or may not be positioned aft of various turbine blades 19 as detailed elsewhere in this document. Those skilled in the art will appreciate that a variety of different methods for controlling or modifying vibrational modes are described herein, and that these methods can be used individually or in any suitable combination.
[0377] In various additional or alternative embodiments, the motor10 may have a core shaft26 with a length (L) between the forward-most and rear-most bearings26a, 26c in the range 1800 to 2900mm or 2750mm, and a separation axial (minor throw, S) between the two rear bearings26b, 26c in the range 250mm to 350mm, such that there is no natural heart shaft frequency26 with only two nodes (one at each end of the span between the first and second bearings 26a, 26b - a mode1 frequency or primary resonance as shown in Tables 1 and 2 above) within the operating range of the motor 10.
[0378] In such embodiments:
[0379] • the heart shaft26 can have a running speed range with a lower limit of 1500rpm and an upper limit of 6200rpm; and or
[0380] • The blower23 can have a blower diameter in the range of 330cm to 380cm (130 to 150 ’’) and the reducer30 can have a gear ratio between 3.1 and 3.8.
[0381] The heart shaft length, L, and the distance, S, between the two rear bearings26b, 26c, and the running speed can therefore be selected so that the heart shaft26 has no resonance mode. primary between the first26a and second26b bearings within the engine operating range. A fan diameter of the blower23 of the motor10 can be selected to be appropriate for a desired run speed - the blower diameter can therefore be selected such that an appropriate run speed for the blower does not cause a running mode. primary resonance of the heart shaft26 between the first26a and second26b bearings within the engine operating range.
[0382] Those skilled in the art will understand that a mode1 resonance (a primary resonance) has only two nodes, while a mode2 frequency has an additional central node and a smaller maximum amplitude, A.
[0383] When an engine10 is designed3000, prior to manufacture, the designer has more leeway to adjust engine parameters. A design method 3000 may include selecting 3002 positions for the front bearing26a and the foremost bearing26b of the rear bearings26b, 26c - e.g. selecting the location of each based on the spacing between them (e.g. 1450 to 2500mm), and locations of other engine components. The method 3000 may further include increasing (or decreasing) 3004 the length of the heart shaft26 aft of the forward-most bearing26b of the rear bearings26b, 26c so as to provide an appropriate shaft length for the minor range (i.e. the distance between the two rear bearings26b, 26c) to be within a suitable range (e.g. 250mm to 350mm) such that there is no primary resonance of the heart tree26 within the walking speed range of heart tree26. The rearmost bearing26c of the two rear bearings (the third bearing, in the embodiment which is described) may be located at the rear end of the heart shaft26, or adjacent to the end of the heart shaft. heart tree 26.
[0384] In various such embodiments, the heart shaft26 / motor10 may have a cruising speed range of 5400 to 5700rpm, and optionally 5500 to 5600rpm.
[0385] In various such embodiments, the heart shaft26 / motor10 may have a PMD speed range of 5800 to 6200rpm, and optionally 5900 to 6100rpm.
[0386] In various such embodiments, the core shaft 26 may have a length, L, from 1800 to 2900mm, optionally 2300 to 2800mm, and optionally further 2400 to 2750mm.
[0387] In various embodiments, the bearing stiffness at the forward most 26b of the two rear bearings 26b, 26c is controlled. Rigidity control can allow or facilitate the management of vibration modes.
[0388] The one in front of the two rear bearings26b (which may also be referred to as the second bearing) has a bearing stiffness in the range of 30kN / mm to 100kN / mm, and possibly around 50kN / mm, through a spring bar ( as described below).
[0389] The stiffness of the second bearing 26b is determined, in part, by the engine condition. The lower the excitation at this bearing26b, the lower the stiffness offered, but the higher the excitation, the higher the rigidity and also the higher the damping. Bearing stiffness is a variable parameter which depends on the engine condition, and may, for example, vary within the range listed above during normal aircraft operation.
[0390] The forward-most rear bearing26b includes an outer ring51 surrounding the heart shaft26; the ring 51 may be referred to as a cage, and may contain ball bearings 52 or the like, as well as oil arranged to lubricate the cage during use. An inner ring53 (or cage), radially towards the inside of the outer ring51, can serve to retain the ball bearings52 within the channel formed between the cages51, 53.
[0391] The outer race 51 of this rear bearing 26b is mounted on the stationary support structure 24 of the engine by means of one or more bearing support structures 50, 55. In the embodiment which is described, two bearing support structures 50, 55 are present. The bearing support structures 50, 55 may together form a bearing disc 29a, as shown schematically by comparison of Figures 12 and 14. The bearing support structures 50, 55 of the described embodiment are each connected to the same bearing housing. tail29. In the embodiment which is described, the tail bearing housing 29 forms part of the stationary support structure of the engine 10. The disc 29a (comprising the component support structures 50 and 55) extends from the tail bearing housing 29 to the bearing 26b. In the embodiment which is described, the first bearing support structure 50 is mounted on the stationary support structure 24 of the motor 10 at a first position 58a. The first position 24a is axially behind the bearing 26b in the embodiment which is described. The first bearing support structure 50 provides the outer race 51 in the embodiment which is described; in particular, the outer cage 51 is integrally formed with the first bearing support structure 50 as shown in Figure 13. In alternative embodiments, the outer cage 51 may be formed separately from, and mounted on, the first bearing support structure 50.
[0392] In the embodiment which is described, the first bearing support structure50 comprises a plurality of connection elements57, which are circumferentially spaced around the motor axis9, connecting the outer race51 to a stationary support structure58. The connection elements 57 extend axially along part of the motor 10.
[0393] For example, there may be twenty or forty evenly spaced connecting elements in some embodiments - the numbers and / or spacing may vary in other embodiments. The connection element (s) may extend between the outer race of the bearing 26b, 26c and a bearing housing 29, 31. The bearing housing 29, 31 may be secured to the stationary support structure 58, thus effectively becoming part of the structure. support58.
[0394] The or each connection element 57 may comprise a metal bar or cross member, and may be referred to as a spring bar. The or each connection member 57 may be arranged to provide some flexibility to the bearing 26b, 26c, thereby allowing some radial and / or axial movement (eg due to expansion) to be accepted. The flexibility of the first bearing support structure 50 may therefore be referred to as the spring bar stiffness.
[0395] In the embodiment which is described, the second bearing support structure 55 is mounted on the stationary support structure 58 of the motor 10 at a second position 58b. In the embodiment which is described, the first58a and second58b positions on the stationary support structure58 are each located on a portion of the tail bearing housing29 of the fixed structure58.
[0396] The second position 58b is axially to the rear of the bearing (rearmost forward) 26b, but axially to the front of the first position 58a on the stationary support structure 58, in the embodiment which is described, and is radially to the exterior of both the bearing26b and the first position58a. In this embodiment, the second bearing support structure 55 is connected between a radially outer surface of the first bearing support structure 50 and the stationary support structure 58, as shown in Figure 14 - in other embodiments, the connection may be different. In the embodiment which is described, the second bearing support structure 55 has relatively high rigidity (compared to the first bearing support structure 50), and therefore has a negligible contribution to the flexibility of the bearing 26b. It can be considered as effectively rigid.
[0397] A damping fluid cushion 56 is provided between the first bearing support structure 50 and the second bearing support structure 55, in the region of the outer race 51. In the embodiment which is described, a channel 56a bordered by raised lips is provided around the outer circumference of the first bearing support structure 50 to accommodate the damping fluid pad 56 and to position o-rings (not shown; axially spaced for have one at each end of the channel 56a). In alternative embodiments, no such channel or lips may be provided, and / or a channel and lips may be provided on the second bearing support structure 55 instead or as well, and the damping fluid pad 56 may be provided. (in some cases, entirely) contained by O-rings or the like. This cushion is in addition to the hydrodynamic oil layer between the bearing52 and the heart shaft26 (i.e. ball bearings52, or the like, are often lubricated by a layer of oil - the fluid cushion. damper56 is a separate layer from ball bearings52).
[0398] The damping fluid pad 56 includes a layer of film, usually a layer of oil or other lubricant, between the bearing 26b and the housing 30. The damping fluid pad 56 is arranged to soften the bearing support to increase the damping efficiency. The rigidity of the damping fluid cushion 56 generally depends on the temperature and the speed of rotation of the shaft. Those skilled in the art will understand that any eccentricity in the rotation of the bearing 26b can be damped by the fluid cushion 56. The damping fluid pad 56 can provide some structural isolation of the support structure 58 from the heart shaft 26, can reduce the magnitudes of rotor responses to imbalance, and can help to suppress rotodynamic instability. This damping can be particularly useful if a blade assembly has an irregular mass distribution (for example due to damage in use, or a manufacturing defect), or if an adverse event during manufacture, maintenance or operation places a blade assembly out of radial alignment, such that it is slightly tilted with respect to the motor axis9.
[0399] The inventor was aware that the overall stiffness of the bearing can be represented as having three parts, which can be considered as springs, for example arranged in parallel or in a combination in series and in parallel (depending on the arrangements of the components), to know:
[0400] • the rigidity of an oil layer in the bearing (the stiffness of the cushion fluid damping fluid);
[0401] • the rigidity of the bearing bracket50 (generally influenced mainly by the rigidity of the connection element57 - this can be referred to as the spring bar rigidity); and
[0402] • the rigidity of the stationary support structure29, 58 on which the bearing is mounted.
[0403] The stationary support structure58 has a much higher stiffness than the other two contributions to bearing stiffness, and is therefore treated as effectively stiff. Any flexibility in the stationary support structure 58 can only become apparent under extreme conditions such as vane separation events. The stiffness of the oil layer varies significantly with shaft rotational speed and temperature, but at cruising speeds and above, the oil stiffness is much higher than the spring bar stiffness. of the landing. The stiffness of each bearing is therefore considered in terms of its spring bar stiffness.
[0404] The spring bar stiffness of a bearing26b is defined as radial stiffness - i.e. linear deflection, δ, along a radius of the motor10 is measured, the deflection being caused by a force, F . This is illustrated in Figures 15A and 15B. The oblique hatching illustrates that the fixed structure58 is meant to be rigid / immobile.
[0405] Figure 15A shows the bearing support structure50, 56, 55 (with the heart shaft26, inner race53 and ball bearings52 excluded for clarity), with part of a spoke, r , of the motor10 marked with a dotted line. The marked radial line is located at the axial center point of the bearing cage. A force, F, is shown applied along the radius, r , in a radially outward direction (i.e. away from the axial center line). Figure 15B shows the initial position of the first bearing support structure50 (before application of force, F) in a broken line, and a final position of the first bearing support structure50 (during application of force, F) in a solid line. Those skilled in the art will understand that the deflection shown is much greater than one would expect in normal operation, and is provided for ease of understanding only. Further, the first bearing support structure 50 must return to its original position after removing the force, F, during normal operation. A displacement, or deflection, δ, is then measured along the radius, r , at the axial center point of the bearing cage51. Two black dots illustrate the position of the inner surface of the first bearing support structure before and during application of force, F. Displacement, δ, is the distance between these points. The inner surface of the first bearing support structure 50 is chosen for ease of demonstration only - those skilled in the art will appreciate that another point - such as the radially outer surface, or a radial center point of the first support structure. bearing 50, or the like - could be chosen instead. The displacement reflects the combination of compression of the damping fluid cushion56, bending of the first bearing support structure50 (and in particular of the spring bars57), and possible bending of the second bearing support structure55. Bearing stiffness is therefore a measure of the radial displacement caused by the application of a radial force at the axial center point of the bearing26b.
[0406] In the embodiment which is described, a force, F , of 50kN causes a deflection, δ, of 1mm when it is applied to the most forward bearing26b of the two rear bearings26b, 26c, therefore the rigidity of this bearing26b, such than defined here, is 50kN / mm. The bearing stiffness can vary between the two rear bearings 26b, 26c, and between embodiments.
[0407] The forward-most bearing of the rear bearings 26b therefore has a bearing stiffness in the range of 30kN / mm to 100kN / mm in the embodiment which is described.
[0408] A stiffness ratio at the forward-most rear bearing 26b (i.e. the second bearing along the heart shaft, in the embodiment which is described) can be defined as:
[0409]
[0410] In various embodiments, this stiffness ratio can be in the range from 0.08 to 0.5kN / mm ² , and optionally in the range from 0.09 to 0.40kN / mm ² . Optionally, the stiffness ratio can be at least substantially equal to 0.25, 0.30 or 0.35 kN / mm ² . Figure 10 illustrates a method 1000 that may be implemented, method 1000 comprising starting 1002 an engine 10 of an aircraft and obtaining cruise conditions, and operating 1004 of the aircraft in cruise conditions.
[0411] The engine10 can have a cruising speed range of 5400 to 5700rpm, and possibly 5500 to 5600rpm.
[0412] The motor 10 can be operated such that there is no primary resonance mode between the first 26a and second 26b bearings under cruising conditions.
[0413] The engine 10 can be operated such that there is no primary resonance mode between the first 26a and second 26b bearings anywhere within the engine operating range (including both PMD and cruise).
[0414] The lengths defined here, unless otherwise specified, are for the corresponding component (s) when the engine is stopped (i.e. at zero speed / on the bench, at room temperature). These values generally do not vary significantly over the operating range of the engine (eg having only a few mm shaft length expansion at operating temperature, or less); the value at cruising conditions of the aircraft to which the engine is attached (these cruising conditions being as defined elsewhere in this document) may therefore be the same as for the moment when the engine is not running. 'use. However, when the length varies over the operating range of the engine, the values set here should be understood as lengths for when the engine is at room temperature and stationary.
[0415] On the other hand, since the oil layer stiffness is speed dependent, and contributes to the bearing stiffness, the bearing stiffness is set at cruising conditions / with the shaft rotating at a speed suitable for cruising.
[0416] Figure 17 illustrates how the bearing stiffness defined here can be measured. Figure 17 shows a plot of the displacement δ resulting from the application of a load L (eg a force, a moment or a torque) to an arbitrary component for which the stiffness is measured. At load levels from zero to L _P there is a nonlinear region in which displacement is caused by movement of the component (or relative movement of independent parts of the component) as it is loaded, rather than by a deformation of the component; for example moving within the clearance between the parts. For a bearing 26b, the amount of possible displacement in this non-linear region is likely to be very small. At load levels above L _Q the elastic limit of the component has been exceeded and the applied load no longer causes elastic deformation - plastic deformation or component failure may occur instead. Between points P and Q the applied load and the resulting displacement have a linear relationship. The stiffness defined here can be determined by measuring the gradient of the linear region between points P and Q (with the stiffness being the inverse of this gradient). The gradient can be found for a region of the linear region as large as possible to increase the accuracy of the measurement by providing a larger displacement to be measured. For example, the gradient can be found by applying a load equal to or just greater than L _P and equal or just less than L _Q. Values for L _P and L _Q can be estimated prior to testing based on material characteristics in order to apply appropriate loads.
[0417] It will be understood that the invention is not limited to the embodiments described above and that various modifications and improvements can be made without departing from the concepts described herein. Unless mutually exclusive, any feature may be used separately or in combination with other features and the description extends to and includes all combinations and sub-combinations of one or more features described herein.

权利要求:
Claims (15)
[0001]
Gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a compressor (14) and a heart shaft (26) connecting the turbine (19) to the compressor, and in which the turbine (19) is the lowest turbine engine pressure (10), the heart shaft (26) has an operating speed in the range of 1500rpm to 6200rpm, and the compressor (14) is the lowest pressure compressor of the engine (10 ); a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades; and a reducer (30) which receives an input from the heart shaft (26) and outputs a drive to the fan (23) so as to drive the fan at a lower rotational speed than the heart shaft ( 26), and wherein the motor core (11) further comprises three bearings (26a-c) arranged to support the heart shaft (26), the three bearings comprising a front bearing (26a) and two rear bearings (26b, 26c) ), the heart shaft (26) having a length between the front bearing (26a) and the rearmost rear bearing (26c) in the range 1800mm to 2900mm, and a minor bearing distance between the two bearings rear (26b, 26c) in the range 250mm to 350mm, so that there is no primary resonance of the heart shaft (26) between the front bearing (26a) and the rear most forward (26b) within the operating speed range of the heart shaft (26).
[0002]
A gas turbine engine (10) according to claim 1 wherein the lower limit of 1500rpm on the heart shaft running speed is the minimum running speed under ground idle conditions and the upper limit of 6200rpm. min on the heart shaft run speed is the upper limit on the maximum take-off thrust run speed.
[0003]
A gas turbine engine (10) according to any preceding claim, wherein the fan (23) has a fan diameter in the range of 330cm to 380cm; and the reducer (30) has a gear ratio in the range of 3.1 to 3.8.
[0004]
est égal ou inférieur à 1,05, et éventuellement dans la plage allant de 0,85 à 0,95.A gas turbine engine (10) according to any preceding claim, wherein the turbine (19) has a turbine length defined between the leading edge of its most upstream vanes (19a) and a trailing edge of its. most downstream vanes (19c, 19d), and in which a ratio of the minor throw to the turbine length of: [Math 32] is equal to or less than 1.05, and optionally in the range of 0.85 to 0.95.
[0005]
A gas turbine engine (10) according to any preceding claim, wherein the rear bearings (26b, 26c) are positioned axially level with or aft of: (i) a leading edge of a lower pressure (most downstream) turbine blade (19c, 19d) of the turbine (19) at the root of the blade; and or (ii) a trailing edge of a turbine blade of a third set (19c) of turbine blades from the front of the turbine (19), at the root of the blade, in wherein the turbine (19) comprises four sets of turbine blades (19a-19d), and wherein optionally the turbine (19) comprises a total of four sets of turbine blades (19a-d).
[0006]
A gas turbine engine (10) according to any preceding claim, wherein the length of the heart shaft (26) is in the range of 2000 to 2900mm or 2750mm, optionally in the range of 2300 to 2750mm, and possibly additionally in the range 2400-2750mm.
[0007]
A gas turbine engine (10) according to any preceding claim, wherein a ratio of the length of the minor seat between the two rear bearings (26b, 26c) to the heart shaft length is equal to or less than 0.14 , and optionally in the range of 0.08 to 0.13.
[0008]
A gas turbine engine (10) according to any preceding claim wherein the forward-most bearing of the rear bearings (26b) has a bearing stiffness in the range of 30kN / mm to 100kN / mm; and optionally wherein a stiffness ratio of the bearing stiffness at the forward-most rear bearing (26b) to the distance between the two rear bearings (26b, 26c) is in the range of 0.08 to 0, 5kN / mm ² .
[0009]
a une valeur dans la plage allant de 2,0 x10^-6à 7,5 x10^-6kg^-1.rad^-2.s².A gas turbine engine (10) according to any preceding claim, wherein the lower pressure turbine of the engine (10) has a set of lower pressure vanes (19c, 19d), each vane of the set d. 'lower pressure vanes (19c, 19d) having a mass, m , a radius at mid-height of the vane, r , and a cruising angular speed, ω ; a minor span, S, is defined as the axial distance between the two rear bearings (26b, 26c), and in which a first vane-to-bearing ratio of: has a value in the range of 2.0 x10 ^-6 to 7.5 x10 ^-6 kg ^-1 .rad ^-2 .s ² .
[0010]
a une valeur dans la plage allant de 0,8 à 6,0kg^-1.0 A gas turbine engine (10) according to any preceding claim, wherein the lower pressure turbine of the engine (10) has a set of lower pressure vanes (19c, 19d), each vane of the set. lower pressure vanes (19c, 19d) having a mass, m , a radius at mid-height of the vane, r ; a minor span (S) is defined as the axial distance between the two rear bearings (26b, 26c), and in which a second vane-to-bearing ratio of: has a value in the range of 0.8 to 6.0kg ^-1 .
[0011]
1 gas turbine engine (10) according to any preceding claim further comprising a tail bearing housing (29) located at the rear of the turbine (19) and comprising two bearing discs (29a, 29b), each bearing disc arranged to support one of the two rear bearings (26b, 26c); and optionally in which the bearing discs are oriented at least substantially perpendicular to the motor axis (9).
[0012]
2 gas turbine engine (10) according to any preceding claim, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the heart shaft is a first heart shaft (26); the engine core further comprises a second turbine (17), a second compressor (15), and an interconnection shaft (27) connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.
[0013]
3 Method (1000) of operating a gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a compressor (14) and a heart shaft (26) connecting the turbine to the compressor, and in which the turbine (19) is the lowest pressure turbine of the engine (10), and wherein the motor core (11) further comprises three bearings (26a-c) arranged to support the heart shaft (26), the three bearings comprising a front bearing (26a) and two rear bearings (26b, 26c), the heart shaft (26) having a length between the front bearing (26a) and the rear most rear bearing (26c) in the range 1800mm to 2900mm, and a minor reach between the two rear bearings (26b, 26c) in the range from 250mm to 350mm; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; and a reducer (30) arranged to receive an input from the heart shaft (26) and to output a drive to the fan (23) so as to drive the fan at a lower rotational speed than the drive shaft heart, the process comprising: operating (1004) of the motor (10) such that the heart shaft (26) has a running speed in the range of 1500rpm to 6200rpm, and in which there is no resonance primary of the heart shaft (26) within the operating speed range of the heart shaft (26).
[0014]
4 A method (3000) of designing a gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a compressor (14) and a heart shaft (26) connecting the turbine (19) to the compressor, and in which the turbine (19) is the lowest pressure turbine of the engine (10), the heart shaft (26) has an operating speed in the range of 1500tr / min at 6200rpm, and the compressor (14) is the lowest pressure compressor of the engine (10); a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades; and a reducer (30) which receives an input from the heart shaft (26) and outputs a drive to the fan (23) so as to drive the fan at a lower rotational speed than the heart shaft ( 26), and wherein the motor core (11) further comprises three bearings (26a-c) arranged to support the heart shaft (26), the three bearings comprising a front bearing (26a) and two rear bearings (26b, 26c) ), the heart shaft (26) having a length between the front bearing (26a) and the rear most rear bearing (26c) in the range of 1800mm to 2900mm, and the method (3000) comprising: selecting (3002) positions for the front bearing (26a) and the foremost bearing (26b) of the rear bearings (26b, 26c); the elongation (3004) of the heart shaft (26) aft of the forward-most bearing (26b) of the rear bearings (26b, 26c) such that a minor contact area defined between the two rear bearings (26b, 26c) is in the range of 250mm to 350mm, and that there is no primary resonance of the heart shaft (26) between the front bearing (26a) and the foremost rear bearing (26b) within the operating speed range of the heart shaft (26).
[0015]
The method of claim 14, wherein the selection (3002) of positions for the front bearing (26a) and the foremost bearing (26b) of the rear bearings (26b, 26c) comprises: - the positioning of the two rear bearings (26b, 26c) downstream of a leading edge of the lower pressure turbine blade (the most downstream) (19c, 19d) of the turbine (19) at the level from a dawn root; and or - the positioning of the two rear bearings (26b, 26c) downstream of a trailing edge of a turbine blade of a third set (19c) of turbine blades from the front of the turbine (19 ), at the level of a dawn root

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CN112923021A|2021-06-08|Aircraft engine

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CN112922730A|2021-06-08|Gas turbine engine arrangement

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CN112923024A|2021-06-08|Reliable gearbox for gas turbine engine

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CN112922741A|2021-06-08|Aircraft engine

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CN112923023A|2021-06-08|High power epicyclic gearbox and operation thereof

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FR3095678A1|2020-11-06|GAS TURBINE ENGINE WITH A DOUBLE WALL CORE CASE

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FR3093350A1|2020-09-04|Geared gas turbine engine

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FR3093755A1|2020-09-18|Efficient gas turbine engine installation and operation

同族专利:

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公开号 | 公开日

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CN113006945A|2021-06-22|

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GB201918782D0|2020-02-05|

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US20210189962A1|2021-06-24|

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US11268441B2|2022-03-08|

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DE102020131807A1|2021-06-24|

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法律状态:
2021-12-27| PLFP| Fee payment|Year of fee payment: 2 |

优先权:

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申请号 | 申请日 | 专利标题

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GBGB1918782.2A|GB201918782D0|2019-12-19|2019-12-19|Shaft bearing arrangement|

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GB1918782.2|2019-12-19|

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