Gas turbine engine with reduced vibration

Abstract

The present invention discloses a gas turbine engine 10 for an aircraft, the gas turbine engine comprising: an engine core 11, the engine core comprising a turbine 19, a compressor 14 and a spindle 26 connecting the turbine to the compressor, the compressor being the lowest pressure compressor of the engine, and the turbine being the lowest pressure turbine and having lowest pressure group blades 19c, 19d, each blade of the lowest pressure group blades having a mass m, a radius r at the middle height of the blade and a cruising angular velocity ω; a fan 23, the fan being located upstream of the engine core; and a gearbox 30, the gearbox receiving input from the spindle and outputting a drive to the fan. The engine core 11 also comprises three bearings 26a-c arranged to support the spindle, the three bearings comprising a front bearing 26a and two rear bearings 26b, 26c, wherein the secondary span (S) is defined as the axial distance between the two rear bearings. The ratio of the first blade to the bearing: #imgabs0# may have a value between 2.0x10 ‑6 to 7.5×10 ‑6 kg ‑1 .rad ‑2 .s 2 The second blade to bearing ratio: #imgabs1# may have a value in the range of 0.8 to 6.0 kg ‑1 value in the range of .

Description

Technical Field

[0001] The present disclosure relates to the installation of a spindle within a gas turbine engine for an aircraft, and in particular to bearing positioning and how such a shaft may be arranged and mounted in order to manage vibration and resonance effects. Background Art

[0002] As used herein, a range "from value X to value Y" or "between value X and value Y" or the like means an inclusive range; including the bounding values ​​of X and Y. As used herein, the term "axial plane" means a plane extending along the length of the engine, parallel to and including the axial centerline of the engine, and the term "radial plane" means a plane extending perpendicular to the axial centerline of the engine, thus including all radial lines at the axial position of the radial plane. Axial planes may also be referred to as longitudinal planes because they extend along the length of the engine. Thus, a radial distance or an axial distance is a distance in a radial or axial plane, respectively. Summary of the invention

[0003] According to a first aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, wherein the compressor is the lowest pressure compressor of the engine, and the turbine is the lowest pressure turbine of the engine and has a lowest pressure group of blades, each blade of the lowest pressure group of blades having a mass m, a radius r at the mid-height of the blade, and a cruising angular velocity ω. The engine also comprises a fan, the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox, the gearbox receiving input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0004] The engine core also includes three bearings arranged to support the spindle, the three bearings including a front bearing and two rear bearings, wherein the secondary span (S) is defined as the axial distance between the two rear bearings. The first blade to bearing ratio:

[0005]

[0006] With 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 value in the range of .

[0007] The inventors have recognized that the engine should not scale linearly with increasing force (mrω 2The invention relates to a method for reducing the length of the subspan (S) to reduce the engine length and weight, thereby allowing efficiency gains to be increased by avoiding additional weight, and in order to avoid the formation of undesirable vortex patterns within the subspan. Although conventional wisdom suggests that a larger subspan is required to improve the reaction of the force from the low pressure turbine, the inventors have found that the risk of introducing vortex patterns and the introduction of more length and weight offset the benefits of the force reaction, and therefore the first blade to bearing ratio should be maintained within a specified range.

[0008] The first blade to bearing ratio can be 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 Up to 4.5×10 -6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0009] Ratio of second blade to bearing:

[0010]

[0011] Can have values ​​in the following range: 0.8 to 6.0 kg -1 In the range of 0.8 to 5.0 kg -1 In the range of 0.8 to 4.0 kg -1 In the range of 0.8 to 3.0 kg -1 In the range of 0.8 to 2.0 kg -1 within the range.

[0012] According to a second aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, wherein the turbine is a minimum pressure turbine of the engine and has a minimum pressure set of blades, each blade of the minimum pressure set of blades having a mass m, a radius r at the mid-height of the blade, and the compressor is a minimum pressure compressor of the engine. The engine also comprises a fan, the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox, the gearbox receiving input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0013] The engine core also includes three bearings arranged to support the spindle, the three bearings including a front bearing and two rear bearings, wherein the secondary span (S) is defined as the axial distance between the two rear bearings. Second blade to bearing ratio:

[0014]

[0015] With 0.8 to 6.0kg -1 value in the range of .

[0016] The ratio of the second blade to the bearing can be between 0.8 and 5.0 kg -1 In the range of 0.8 to 4.0 kg -1 In the range of 0.8 to 3.0 kg -1 In the range of 0.8 to 2.0 kg -1 within the range.

[0017] Each blade in the lowest pressure group of blades has a cruising angular velocity ω, and the first blade to bearing ratio is:

[0018]

[0019] can have values ​​in the following range: 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s2 In the range of 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg - 1 .rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0020] In the first aspect or the second aspect, one or more of the following features may be present:

[0021] The secondary span S may be in the range of 250 mm to 350 mm. In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0022] The length (L) of the mandrel may be in the range of 1800 mm to 2900 mm, optionally in the range of 2000 mm to 2900 mm, further optionally in the range of 2300 mm to 2800 mm, and further optionally in the range of 2400 mm to 2750 mm.

[0023] The value of the blade mass m multiplied by the blade radius r at mid-height may be in the range of 180 kg.mm to 280 kg.mm.

[0024] The gearbox may have a gear ratio greater than 3, and optionally in the range of 3.1 to 3.8.

[0025] The spindle may have an operating speed in the range of 1500 rpm to 6200 rpm.

[0026] When cruising, the cruising operating speed of the spindle may be in the range of 5400 rpm to 5700 rpm, and optionally in the range of 5500 rpm to 5600 rpm.

[0027] Under maximum take-off (MTO) conditions, the operating speed of the spindle may be in the range of 5800 rpm to 6200 rpm, and optionally in the range of 5900 rpm to 6100 rpm.

[0028] The length ratio (S / L) of the secondary span between the two rear bearings to the spindle length may be equal to or less than 0.14, or equal to or less than 0.13, or equal to or less than 0.12. The length ratio S / L may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio S / L may be in the range 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.

[0029] The mass m of the blades in the lowest pressure group of blades may be in the range of 0.2 kg to 0.6 kg.

[0030] The radius r of the blades in the lowest pressure group of blades may be in the range of 400 mm to 600 mm.

[0031] Each blade in the lowest pressure group of blades may have a pressure between 560 and 600 rad.s -1 The cruising angular velocity ω within the range.

[0032] The length of the mandrel may be in the range of 1800mm to 2900mm or 2750mm. The secondary span may be in the range of 250mm to 350mm, and optionally in the range of 260mm to 350mm.

[0033] The fan may have a fan diameter in the range of 330 cm to 380 cm.

[0034] The turbine has a turbine length defined between the leading edge of its most upstream blade and the trailing edge of its most downstream blade. The ratio of the subspan to the turbine length is:

[0035]

[0036] The ratio of the subspan to the turbine length may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example, the ratio of the subspan to the turbine length may be in the range of 0.70 to 1.05, optionally in the range of 0.70 to 1.00, optionally in the range of 0.70 to 0.95, optionally in the range of 0.80 to 1.05, optionally in the range of 0.80 to 1.00, optionally in the range of 0.80 to 0.95, optionally in the range of 0.85 to 1.05, optionally in the range of 0.85 to 1.00, and also optionally in the range of 0.85 to 0.95.

[0037] The rear bearing may be located axially flush with or behind the leading edge of the lowest pressure turbine blade of the turbine at the root of the blade.

[0038] The rear bearing may be positioned axially flush with or behind the trailing edge of a turbine blade in the third set of turbine blades from the front of the turbine at the root of the blade. In such embodiments, the turbine may include four sets of turbine blades, and optionally may have a total of four sets of turbine blades.

[0039] A forwardmost bearing of the rear bearings may have a bearing stiffness in the range of 30 kN / mm to 100 kN / mm.

[0040] The ratio of the bearing stiffness at the foremost rear bearing to the distance between the two rear bearings can be between 0.08 and 0.5 kN / mm. 2 In the range of, optionally in the range of 0.08 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.30 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.20 kN / mm 2 In the range of, optionally in the range of 0.09 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.50 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.40 kN / mm 2 and optionally in the range of 0.15 to 0.30 kN / mm 2 within the range.

[0041] According to a third aspect, there is provided a method of operating a gas turbine engine for an aircraft, the engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, and wherein the compressor is the lowest pressure compressor of the engine, and the turbine is the lowest pressure turbine of the engine and has a lowest pressure group of blades, each blade of the lowest pressure group of blades having a mass m, a radius r at the mid-height of the blade, and a cruising angular velocity ω. The engine core also comprises three bearings arranged to support the spindle, the three bearings comprising a front bearing and two rear bearings, wherein the secondary span (S) is defined as the axial distance between the two rear bearings. The engine also comprises a fan, the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox, the gearbox being arranged to receive input from the spindle and output drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0042] The method includes operating the engine such that a first blade to bearing ratio is:

[0043]

[0044] When cruising, it has a 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 value in the range of .

[0045] In various embodiments, the method may include operating the engine so that the first blade to bearing ratio may be within 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 6.5×10-6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0046] The method may further comprise operating the mandrel at an operating speed in the range of 1500 rpm to 6200 rpm, and optionally wherein the operating speed of the mandrel may be one or more of:

[0047] (i) at cruising, in the range of 5400 rpm to 5700 rpm, and optionally in the range of 5500 rpm to 5600 rpm; and / or

[0048] (ii) at maximum take-off conditions, in the range of 5800 rpm to 6200 rpm, and optionally in the range of 5900 rpm to 6100 rpm.

[0049] The engine for implementing the method of the third aspect may be as described in the first aspect and / or the second aspect.

[0050] According to a fourth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, wherein the turbine is the lowest pressure turbine of the engine, the spindle has an operating speed in the range of 1500 rpm to 6200 rpm, and the compressor is the lowest pressure compressor of the engine. The engine also comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox receiving an input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0051] The engine core also includes three bearings arranged to support a core shaft, the three bearings including a front bearing and two rear bearings, the core shaft having a length (L) in the range of 1800mm to 2900mm or 2750mm between the front bearing and the last rear bearing, and a secondary span (S) in the range of 250mm to 350mm between the two rear bearings.

[0052] Thus, the first resonance of the spindle between the front bearing and the foremost rear bearing may not be within the operating speed range of the spindle.

[0053] The length (L) and subspan (S) may be selected to be appropriate for the expected operating speed range to avoid such a first resonance.

[0054] In various embodiments, the length (L) of the mandrel may be in the range of 2000 mm to 2900 mm, further optionally in the range of 2300 mm to 2800 mm, and further optionally in the range of 2400 mm to 2750 mm.

[0055] In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0056] The inventors have recognized that when a gas turbine engine is scaled up, an increase in the mandrel length may result in a resonant frequency of the mandrel being within or close to the engine operating range. Therefore, simply scaling up known engines may result in an increased risk of damage caused by resonance, and longer mandrels may be problematic. The inventors have found that selecting a length L and a subspan S appropriate for a given operating speed range may be advantageous in avoiding harmful vortex modes and may reduce damage to the engine in use.

[0057] The lower limit of the spindle operating speed of 1500 rpm may be the minimum operating speed under ground idling conditions, and / or the upper limit of the spindle operating speed of 6200 rpm may be the upper limit of the maximum takeoff operating speed.

[0058] According to a fifth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, wherein the turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The engine also comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades and having a fan diameter in the range of 330 cm to 380 cm; and a gearbox receiving an input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle, the gearbox having a gear ratio in the range of 3.1 to 3.8.

[0059] The engine core also includes three bearings arranged to support a core shaft, the three bearings including a front bearing and two rear bearings, the core shaft having a length in the range of 1800mm to 2900mm or 2750mm between the front bearing and the last rear bearing, and a secondary span in the range of 250mm to 350mm between the two rear bearings, so that the first resonance of the core shaft is not within the operating speed range of the core shaft.

[0060] The skilled person will appreciate that the fifth aspect shares the inventive concept with the fourth aspect, because the combination of fan diameter and gear ratio is related to the spindle operating speed. For a particular design of an aircraft, a given combination of fan diameter and gear ratio can be used to infer the expected spindle operating speed range.

[0061] In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0062] The spindle can run at speeds ranging from 1500 rpm to 6200 rpm.

[0063] In the fourth or fifth aspects, one or more of the following features may be present:

[0064] The turbine has a turbine length defined between the leading edge of its most upstream blade and the trailing edge of its most downstream blade. The ratio of the subspan to the turbine length is:

[0065]

[0066] The ratio of the subspan to the turbine length may be equal to or less than 1.05. The ratio of the subspan to the turbine length may be equal to or less than 1.00, optionally equal to or less than 0.95. The ratio of the subspan to the turbine length may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example, the ratio of the subspan to the turbine length may be in the range of 0.70 to 1.05, optionally in the range of 0.70 to 1.00, optionally in the range of 0.70 to 0.95, optionally in the range of 0.80 to 1.05, optionally in the range of 0.80 to 1.00, optionally in the range of 0.80 to 0.95, optionally in the range of 0.85 to 1.05, optionally in the range of 0.85 to 1.00, and also optionally in the range of 0.85 to 0.95.

[0067] The turbine may comprise four sets of turbine blades. Both rear bearings may be located downstream of the trailing edge of a turbine blade in the third set of turbine blades from the front of the turbine (at the root of the blade). The turbine may comprise a total of four sets of turbine blades.

[0068] The two rear bearings may be located downstream of the leading edge of the lowest pressure (most downstream) turbine blade of the turbine (at the root of the blade).

[0069] The length (L) of the mandrel may be in the range of 1800 mm to 2900 mm, optionally in the range of 2000 mm to 2900 mm, further optionally in the range of 2300 mm to 2800 mm, and further optionally in the range of 2400 mm to 2750 mm.

[0070] In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0071] The length ratio (S / L) of the secondary span to the spindle length between the two rear bearings may be equal to or less than 0.14.

[0072] The length ratio may be equal to or less than 0.13, or equal to or less than 0.12. The length ratio may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio may be in the range 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 further optionally in the range of 0.08 to 0.13.

[0073] A forwardmost bearing of the rear bearings may have a bearing stiffness in the range of 30 kN / mm to 100 kN / mm.

[0074] The ratio of the bearing stiffness at the foremost rear bearing to the distance between the two rear bearings can be between 0.08 and 0.5 kN / mm. 2 In the range of, optionally in the range of 0.08 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.30 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.20 kN / mm 2 In the range of, optionally in the range of 0.09 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.50 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.40 kN / mm 2 and optionally in the range of 0.15 to 0.30 kN / mm 2 within the range.

[0075] The lowest pressure turbine of the engine has a lowest pressure set of blades, each blade of which has a mass m, a radius r at mid-height of the blade, and a cruising angular velocity ω.

[0076] Ratio of first blade to bearing:

[0077]

[0078] can have values ​​in the following range: 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg - 1 .rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0079] Ratio of second blade to bearing:

[0080]

[0081] Can have values ​​in the following range: 0.8 to 6.0 kg -1 In the range of 0.8 to 5.0 kg -1 In the range of 0.8 to 4.0 kg -1 In the range of 0.8 to 3.0 kg -1 In the range of 0.8 to 2.0 kg-1 within the range.

[0082] According to a sixth aspect, a method for operating a gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine. The engine core also comprises three bearings arranged to support a spindle, the three bearings comprising a front bearing and two rear bearings, the spindle having a length (L) in the range of 1800mm to 2900mm between the front bearing and the last rear bearing, and having a secondary span (S) in the range of 250mm to 350mm between the two rear bearings. The engine also comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the spindle and output a drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0083] The method includes operating the engine such that the spindle has an operating speed within a range of 1500 rpm to 6200 rpm, and wherein a first resonance of the spindle is not within the operating speed range of the spindle.

[0084] The length (L) of the mandrel may be in the range of 1800 mm to 2750 mm, or in the range of 2000 mm to 2750 mm, or in the range of 2400 mm to 2750 mm.

[0085] In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0086] The engine used to implement the method may be as described in the fourth aspect and / or the fifth aspect.

[0087] According to a seventh aspect, a method for designing a gas turbine engine for an aircraft is provided. The engine comprises an engine core, the engine core comprising a turbine, a compressor, and a mandrel connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine, the mandrel having an operating speed in the range of 1500rpm to 6200rpm, and the compressor is the lowest pressure compressor of the engine. The engine also comprises a fan, the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox, the gearbox receiving an input from the mandrel and outputting a drive to the fan so as to drive the fan at a rotational speed lower than that of the mandrel. The engine core also comprises three bearings arranged to support a mandrel, the three bearings comprising a front bearing and two rear bearings, the mandrel having a length (L) in the range of 1800mm to 2900mm between the front bearing and the last rear bearing.

[0088] The method includes:

[0089] selecting a position for a forward-most bearing of the front bearing and the rear bearing; and

[0090] The spindle is elongated behind the foremost of the rear bearings so that the secondary span defined between the two rear bearings is in the range of 250 mm to 350 mm and the first resonance of the spindle between the front bearings and the foremost rear bearing is not within the operating speed range of the spindle.

[0091] The length (L) of the mandrel may be in the range of 1800 mm to 2750 mm, or in the range of 2000 mm to 2750 mm, or in the range of 2400 mm to 2750 mm.

[0092] In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0093] The engine designed by the method of the seventh aspect may be the engine of the fourth aspect and / or the fifth aspect.

[0094] According to an eighth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine, and the compressor is the lowest pressure compressor of the engine. The engine also comprises a fan, the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox, the gearbox receiving input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The engine core also comprises three bearings arranged to support the spindle, the three bearings comprising a front bearing and two rear bearings, and wherein the frontmost rear bearing has a bearing stiffness in the range of 30 kN / mm to 100 kN / mm, the bearing stiffness being defined by a radial displacement caused by applying a radial force at an axial center point of the bearing.

[0095] The present inventors have recognised that controlling bearing stiffness within a specified range may allow or facilitate management of vibration modes, thereby potentially reducing damage to the engine caused by vortex mode displacement of the spindle in use.

[0096] The spindle has a length (L) between the front bearing and the last rear bearing and a minor span (S) between the rear bearings. The bearings may be arranged so that the length ratio of the minor span to the spindle length (S / L) is equal to or less than 0.14, or equal to or less than 0.13, or equal to or less than 0.12. The length ratio may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio may be in the range 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 also optionally in the range of 0.08 to 0.13.

[0097] The spindle may have an operating speed range with a lower limit of 1500 rpm and an upper limit of 6200 rpm.

[0098] The bearing stiffness of the foremost rear bearing may be equal to or approximately 50 kN / mm.

[0099] The gas turbine engine may further include a fixed support structure and a first bearing support structure. The forward rear bearing may be mounted to the fixed support structure via the first bearing support structure. The first bearing support structure may be attached to the fixed support structure at a first position axially rearward of the forward rear bearing.

[0100] In such embodiments, the first bearing support structure may include a plurality of connection members that may be circumferentially spaced about the engine axis to connect the bearing to the fixed support structure.The first bearing support structure may be described as a spring bar type support structure.

[0101] The first bearing support structure may comprise an outer race of the foremost aft bearing.

[0102] The gas turbine engine with the first bearing support structure according to the embodiment may further include a second bearing support structure. The second bearing support structure may be mounted to the fixed support structure, optionally at a second position that is forward of the first position and at a greater radial distance from the engine axis than the first position. The second bearing support structure may be connected to the first bearing support structure via a squeeze film damper in the forwardmost aft bearing region.

[0103] The ratio of the bearing stiffness at the foremost rear bearing to the stiffness of the secondary span can be between 0.08 and 0.5 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.30 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.20 kN / mm 2 In the range of, optionally in the range of 0.09 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.50 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.40 kN / mm 2 and optionally in the range of 0.15 to 0.30 kN / mm 2 within the range.

[0104] The length (L) of the mandrel may be in the range of 1800 mm to 2900 mm, optionally in the range of 2000 mm to 2900 mm, further optionally in the range of 2300 mm to 2800 mm, and further optionally in the range of 2400 mm to 2750 mm.

[0105] In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0106] The length of the mandrel may be in the range of 1800mm to 2900mm or 2750mm. In such embodiments, the secondary span may be in the range of 250mm to 350mm, optionally in the range of 260mm to 350mm.

[0107] The fan may have a fan diameter in the range of 330 cm to 380 cm.

[0108] The length of the spindle may be in the range of 1800 mm to 2900 mm or 2750 mm. The subspan may be in the range of 250 mm to 350 mm. The operating speed of the spindle may be in the range of 1500 rpm to 6200 rpm; and / or the diameter of the fan may be in the range of 330 cm to 380 cm, and the gear ratio of the gearbox in the range of 3.1 to 3.8. The length, subspan and / or operating speed may be selected so that the first resonance of the spindle is not within the operating range of the engine.

[0109] The rear bearing may be located axially flush with or behind the leading edge of the lowest pressure turbine blade of the turbine at the root of the blade.

[0110] The rear bearing 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, wherein the turbine comprises four sets of turbine blades, at the root of the blade.

[0111] The lowest pressure turbine of the engine has a lowest pressure set of blades, each blade of which has a mass m, a radius r at mid-height of the blade, and a cruising angular velocity ω.

[0112] Ratio of first blade to bearing:

[0113]

[0114] can have values ​​in the following range: 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg - 1 .rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s2 In the range of 3.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0115] Ratio of second blade to bearing:

[0116]

[0117] Can have values ​​in the following range: 0.8 to 6.0 kg -1 In the range of 0.8 to 5.0 kg -1 In the range of 0.8 to 4.0 kg -1 In the range of 0.8 to 3.0 kg -1 In the range of 0.8 to 2.0 kg -1 within the range.

[0118] The turbine has a length between the leading edge of the foremost turbine blade of the turbine and the trailing edge of the rearmost turbine blade of the turbine. The ratio of the subspan to the turbine length may be equal to or less than 1.05. The ratio of the subspan to the turbine length may be equal to or less than 1.00, optionally equal to or less than 0.95. The ratio of the subspan to the turbine length may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example, the ratio of the secondary span to the turbine length may be in the range of 0.70 to 1.05, optionally in the range of 0.70 to 1.00, optionally in the range of 0.70 to 0.95, optionally in the range of 0.80 to 1.05, optionally in the range of 0.80 to 1.00, optionally in the range of 0.80 to 0.95, optionally in the range of 0.85 to 1.05, optionally in the range of 0.85 to 1.00, and also optionally in the range of 0.85 to 0.95.

[0119] According to a ninth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The turbine has a turbine length, the turbine length being defined as the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the turbine at its trailing edge. The engine also comprises a fan, the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox, the gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0120] The engine core also includes three bearings arranged to support the spindle, the three bearings including a front bearing and two rear bearings, wherein the subspan (S) is defined as the distance between the two rear bearings. The ratio of subspan to turbine length:

[0121]

[0122] Equal to or less than 1.05.

[0123] The present inventors have recognized that maintaining the subspan to turbine length ratio within this range, and more generally less than that in known aircraft, may help reduce or avoid detrimental vortex modes in operation.

[0124] The ratio of the subspan to the turbine length may be equal to or less than 1.00, optionally equal to or less than 0.95. The ratio of the subspan to the turbine length may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example, the ratio of the subspan to the turbine length may be in the range of 0.70 to 1.05, optionally in the range of 0.70 to 1.00, optionally in the range of 0.70 to 0.95, optionally in the range of 0.80 to 1.05, optionally in the range of 0.80 to 1.00, optionally in the range of 0.80 to 0.95, optionally in the range of 0.85 to 1.05, optionally in the range of 0.85 to 1.00, and also optionally in the range of 0.85 to 0.95.

[0125] The inventors have recognized that when a gas turbine engine is scaled up, the increased mandrel length may result in a resonant frequency of the mandrel being within or close to the engine operating range. Therefore, simply scaling up known engines may result in an increased risk of damage caused by resonance, and longer mandrels may be problematic. The inventors have found that arranging the subspan to be approximately equal to and optionally slightly less than the turbine length (of the lowest pressure turbine in embodiments with more than one turbine), and more specifically within the stated range, may be advantageous in avoiding harmful vortex modes and may reduce damage to the engine in use.

[0126] The turbine may comprise a total of four sets of turbine blades.In such embodiments, both rear bearings may be located downstream of the trailing edge of a turbine blade in the third set of turbine blades from the front of the turbine (at the root of the blade).

[0127] The two rear bearings may be located downstream of the leading edge of the lowest pressure (most downstream) turbine blade of the turbine (at the root of the blade).

[0128] The length (L) of the mandrel may be in the range of 1800 mm to 2900 mm, optionally in the range of 2000 mm to 2900 mm, further optionally in the range of 2300 mm to 2800 mm, and further optionally in the range of 2400 mm to 2750 mm.

[0129] In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0130] The spindle may have an operating speed range with a lower limit of 1500 rpm and an upper limit of 6200 rpm. The upper limit of the spindle operating speed range may be the upper limit of the maximum take-off (MTO) operating speed. The lower limit of the spindle operating speed may be the minimum operating speed under ground idling conditions.

[0131] The spindle has a length L between the frontmost bearing and the rearmost bearing and a distance (secondary span S) between the two rear bearings. The bearings may be arranged so that the length ratio (S / L) of the distance between the two rear bearings to the spindle 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 (S) between the two rear bearings to the spindle 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 (S) between the two rear bearings to the spindle length (L) may 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.

[0132] The foremost of the rear bearings may have a bearing stiffness in the range of 30 kN / mm to 100 kN / mm. The stiffness ratio of the bearing stiffness at the foremost rear bearing divided by the distance between the two rear bearings may be 0.08 to 0.5 kN / mm. 2 In the range of, optionally in the range of 0.08 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.30 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.20 kN / mm 2 In the range of, optionally in the range of 0.09 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.50 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.40 kN / mm 2 and optionally in the range of 0.15 to 0.30 kN / mm 2 within the range.

[0133] The length (L) of the spindle may be in the range of 1800 mm to 2900 mm or 2750 mm. The distance (S) between the two rear bearings may be in the range of 250 mm to 350 mm, optionally in the range of 260 mm to 350 mm.

[0134] The fan may have a fan diameter in the range of 330 cm to 380 cm.

[0135] The length of the spindle may be in the range of 1800 mm to 2900 mm or 2750 mm; and the subspan may be in the range of 250 mm to 350 mm, optionally in the range of 260 mm to 350 mm. The operating speed of the spindle may be in the range of 1500 rpm to 6200 rpm; and / or the diameter of the fan may be in the range of 330 cm to 380 cm, and the gear ratio of the gearbox may be in the range of 3.1 to 3.8. The length, subspan and / or operating speed may be selected so that the first resonance of the spindle is not within the operating range of the engine.

[0136] The lowest pressure turbine of the engine has a lowest pressure set of blades, each blade of which has a mass m, a radius r at mid-height of the blade, and a cruising angular velocity ω.

[0137] Ratio of first blade to bearing:

[0138]

[0139] can have values ​​in the following range: 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg - 1 .rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0140] Ratio of second blade to bearing:

[0141]

[0142] Can have values ​​in the following range: 0.8 to 6.0 kg -1 In the range of 0.8 to 5.0 kg -1 In the range of 0.8 to 4.0 kg -1 In the range of 0.8 to 3.0 kg -1 In the range of 0.8 to 2.0 kg -1 within the range.

[0143] According to the tenth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine and comprises turbine blades, and the compressor is the lowest pressure compressor of the engine. The engine also comprises a fan, the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox, the gearbox receiving input from the spindle and outputting a drive to the fan so as to drive the fan at a rotational speed lower than that of the spindle. The engine core also comprises three bearings arranged to support the spindle, the three bearings comprising two rear bearings, the two rear bearings being located downstream of the leading edge of the lowest pressure turbine blade of the turbine (at the root of the blade). Therefore, the two rear bearings can be described as being located downstream of the leading edge of the last turbine blade.

[0144] The inventors have recognized that the mandrel generally moves the least (in the radial direction) and is flattest (parallel to the engine axis) at the axial location of the bearings, and while vortex patterns and other bending or displacement may occur between the bearings, the bearings serve to limit radial shaft movement. The inventors have recognized that careful control of shaft length and bearing location may therefore allow management of the engine's vortex patterns, thereby reducing the risk of engine damage.

[0145] The inventors have also discovered that positioning the bearings closer to the larger turbine stages and towards the rear of the turbine provides improved turbine support, since movement of the shaft relative to the turbine position may have a greater detrimental effect on these larger turbine stages.

[0146] The gas turbine engine may further comprise a disk arranged to support the lowest pressure turbine blades of the turbine.The two aft bearings may be located downstream of the centre line of the disk.

[0147] The length of the mandrel may be in the range of 1800 mm to 2900 mm, optionally in the range of 2000 mm to 2900 mm, optionally in the range of 2300 mm to 2800 mm, and further optionally in the range of 2400 mm to 2750 mm.

[0148] The spindle may have an operating speed range with a lower limit of 1500 rpm and an upper limit of 6200 rpm.

[0149] The upper limit of the core shaft operating speed range may be the upper limit of the maximum take-off (MTO) operating speed. The lower limit of the core shaft operating speed may be the minimum operating speed under ground idling conditions.

[0150] The spindle has a length L between the frontmost bearing and the rearmost bearing and a distance S between the two rear bearings. The bearings may be arranged so that the length ratio S / L of the distance (S) between the two rear bearings to the spindle length (L) 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 (S) between the two rear bearings to the spindle 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 (S) between the two rear bearings to the spindle length (L) may 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.

[0151] The length (L) of the spindle may be in the range of 1800 mm to 2900 mm or 2750 mm. In such embodiments, the distance between the two rearward bearings (S) may be in the range of 250 mm to 350 mm, or optionally in the range of 260 mm to 350 mm.

[0152] According to an eleventh aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, and wherein the turbine is the lowest pressure turbine of the engine and comprises four sets of turbine blades, and the compressor is the lowest pressure compressor of the engine. The engine also comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox receiving input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The engine core also comprises three bearings arranged to support the spindle, the three bearings comprising two rear bearings located downstream of the trailing edge of a turbine blade in a third set of turbine blades from the front of the turbine (at the root of the blade).

[0153] One or more of the following features may be applied to the gas turbine engine of the tenth aspect and / or the eleventh aspect:

[0154] A forwardmost bearing of the rear bearings may have a bearing stiffness in the range of 30 kN / mm to 100 kN / mm.

[0155] The stiffness ratio of the frontmost rear bearing to the distance (S) between the two rear bearings can be between 0.08 and 0.5 kN / mm. 2 In the range of, optionally in the range of 0.08 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.30 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.20 kN / mm 2 In the range of, optionally in the range of 0.09 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.50 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.40 kN / mm 2 and optionally in the range of 0.15 to 0.30 kN / mm 2 within the range.

[0156] The length (L) of the spindle may be in the range of 1800 mm to 2900 mm or 2750 mm. In such embodiments, the distance between the two rearward bearings (S) may be in the range of 250 mm to 350 mm, or optionally in the range of 260 mm to 350 mm.

[0157] The fan may have a fan diameter in the range of 330 cm to 380 cm.

[0158] The length (L) of the mandrel may be in the range of 1800 mm to 2900 mm, optionally in the range of 2000 mm to 2900 mm, further optionally in the range of 2300 mm to 2800 mm, and further optionally in the range of 2400 mm to 2750 mm.

[0159] The distance (S) between the two rear bearings, which may be referred to as the minor span, may be in the range of 250 mm to 350 mm. In various embodiments, the minor span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the minor span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0160] The operating speed of the spindle may be in the range of 1500 rpm to 6200 rpm. Additionally or alternatively, the diameter of the fan may be in the range of 330 cm to 380 cm. The gear ratio of the gearbox may be in the range of 3.1 to 3.8.

[0161] The length, subspan and / or operating speed may be selected so that the first resonance of the spindle is not within the operating range of the engine.

[0162] The lowest pressure turbine of the engine has the lowest pressure set of blades. Each blade in the lowest pressure set of blades has a mass m, a radius r at the mid-height of the blade, and a cruising angular velocity ω. As mentioned above, the secondary span is defined as the axial distance between the two rear bearings.

[0163] Ratio of first blade to bearing:

[0164]

[0165] can have values ​​in the following range: 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg - 1 .rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg-1 .rad -2 .s 2 In the range of 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0166] Additionally or alternatively, a ratio of the second blade to the bearing:

[0167]

[0168] Can have values ​​in the following range: 0.8 to 6.0 kg -1 In the range of 0.8 to 5.0 kg -1 In the range of 0.8 to 4.0 kg -1 In the range of 0.8 to 3.0 kg -1 In the range of 0.8 to 2.0 kg -1 within the range.

[0169] The turbine has a length between the leading edge of the foremost turbine blade of the turbine and the trailing edge of the rearmost turbine blade of the turbine. The ratio of the subspan to the turbine length (i.e., S divided by the turbine length) may be equal to or less than 1.05, optionally equal to or less than 1.00, optionally equal to or less than 0.95. The ratio of the subspan to the turbine length may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example, the ratio of the secondary span to the turbine length may be in the range of 0.70 to 1.05, optionally in the range of 0.70 to 1.00, optionally in the range of 0.70 to 0.95, optionally in the range of 0.80 to 1.05, optionally in the range of 0.80 to 1.00, optionally in the range of 0.80 to 0.95, optionally in the range of 0.85 to 1.05, optionally in the range of 0.85 to 1.00, and also optionally in the range of 0.85 to 0.95.

[0170] The turbine may comprise a total of four sets of turbine blades. In an alternative embodiment, the turbine may comprise a total of three sets of turbine blades.

[0171] According to a twelfth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a mandrel connecting the turbine to the compressor. The turbine is the lowest pressure turbine of the engine, and the compressor is the lowest pressure compressor of the engine. The mandrel has an operating speed range with a lower limit of 1500rpm and an upper limit of 6200rpm. The engine also includes a fan, which is located upstream of the engine core, and the fan includes a plurality of fan blades; and a gearbox, which receives input from the mandrel and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the mandrel.

[0172] The engine core also includes three bearings arranged to support the spindle, the three bearings including a front bearing and two rear bearings. The spindle has a length (L) between the front bearing and the last rear bearing and a subspan (S) between the rear bearings. The bearings are arranged so that the length ratio of the subspan to the spindle length (S / L) is equal to or less than 0.14.

[0173] The length ratio may be equal to or less than 0.13, or equal to or less than 0.12. The length ratio may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio may be in the range 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 further optionally in the range of 0.08 to 0.13.

[0174] The inventors have recognized that when a gas turbine engine is scaled up, an increase in the mandrel length may result in a resonant frequency of the mandrel being within or close to the engine operating range. Therefore, simply scaling up known engines may result in an increased risk of damage caused by resonance, and a longer mandrel may be problematic.

[0175] However, in some embodiments, it has been found that making the mandrel still longer (the mandrel extending rearward of the first rear bearing and to the second rear bearing, wherein the spacing between the first rear bearing and the second rear bearing is within a set range relative to the total mandrel length) increases the stiffness of the mandrel and moves the resonant frequency away from the engine operating range.

[0176] The upper limit of the spindle operating speed range may be the upper limit of the maximum takeoff operating speed.

[0177] The lower limit of the spindle running speed may be the minimum running speed under ground idling conditions.

[0178] Under cruising conditions, the operating speed of the spindle may be in the range of 5400 rpm to 5700 rpm, and optionally in the range of 5500 rpm to 5600 rpm.

[0179] The operating speed of the spindle may be in the range of 5800 rpm to 6200 rpm, and optionally in the range of 5900 rpm to 6100 rpm under maximum take-off (MTO) conditions.

[0180] The length (L) of the mandrel may be in the range of 1800 mm to 2900 mm, optionally in the range of 2000 mm to 2900 mm, further optionally in the range of 2300 mm to 2800 mm, and further optionally in the range of 2400 mm to 2750 mm.

[0181] In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm.

[0182] The forward most of the rear bearings may have a bearing stiffness in the range of 30 kN / mm to 100 kN / mm. The ratio of the stiffness at the forward most rear bearing to the stiffness of the secondary span may be 0.08 to 0.5 kN / mm. 2 In the range of, optionally in the range of 0.08 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.30 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.20 kN / mm 2 In the range of, optionally in the range of 0.09 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.50 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.40 kN / mm 2 and optionally in the range of 0.15 to 0.30 kN / mm 2 within the range.

[0183] The length (L) of the mandrel may be in the range of 1800mm to 2900mm or 2750mm. The secondary span (S) may be in the range of 250mm to 350mm, optionally in the range of 260mm to 350mm.

[0184] The fan may have a fan diameter in the range of 330 cm to 380 cm.

[0185] In some embodiments, the length of the mandrel may be in the range of 1800 mm to 2900 mm or 2750 mm. The subspan may be in the range of 250 mm to 350 mm, optionally in the range of 260 mm to 350 mm. The operating speed of the mandrel may be in the range of 1500 rpm to 6200 rpm; and / or the diameter of the fan may be in the range of 330 cm to 380 cm, and the gear ratio of the gearbox may be in the range of 3.1 to 3.8. In such embodiments, the length, subspan and / or operating speed may be selected so that the first resonance of the mandrel is not within the operating range of the engine.

[0186] The rear bearing may be positioned axially flush with or behind:

[0187] (i) the leading edge of the lowest pressure turbine blade of the turbine at the root of that blade; and / or

[0188] (ii) A turbine blade in a third set of turbine blades from the front of the turbine, wherein the turbine comprises four sets of turbine blades, has its trailing edge at the root of the blade axially flush or behind the trailing edge.

[0189] The present inventors have recognised that positioning the rear bearing closer to the larger turbine stages and towards the rear of the turbine provides improved turbine support, as movement of the shaft relative to the turbine position may have a greater detrimental effect on these larger turbine stages.

[0190] The lowest pressure turbine of the engine has a lowest pressure set of blades, each of which has a mass m, a radius r at mid-height of the blade, and a cruising angular velocity ω. The secondary span (S) is defined as the axial distance between the two rear bearings, as described above.

[0191] Ratio of first blade to bearing:

[0192]

[0193] can have values ​​in the following range: 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg - 1.rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0194] Additionally or alternatively, a ratio of the second blade to the bearing:

[0195]

[0196] Can have values ​​in the following range: 0.8 to 6.0 kg -1 In the range of 0.8 to 5.0 kg -1 In the range of 0.8 to 4.0 kg -1 In the range of 0.8 to 3.0 kg -1 In the range of 0.8 to 2.0 kg -1 within the range.

[0197] The turbine has a length between a leading edge of a forward turbine blade of the turbine and a trailing edge of a rearward turbine blade of the turbine. The ratio of the subspan to the turbine length may be in the range of 0.70 to 1.05, optionally in the range of 0.70 to 1.00, optionally in the range of 0.70 to 0.95, optionally in the range of 0.80 to 1.05, optionally in the range of 0.80 to 1.00, optionally in the range of 0.80 to 0.95, optionally in the range of 0.85 to 1.05, optionally in the range of 0.85 to 1.00, and further optionally in the range of 0.85 to 0.95.

[0198] According to a thirteenth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising an engine core, the engine core comprising a turbine, a compressor and a spindle connecting the turbine to the compressor, wherein the turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The engine also comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox receiving input from the spindle and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0199] The engine core also includes three bearings arranged to support the spindle, the three bearings including a front bearing and two rear bearings, the distance between the two rear bearings being defined as a subspan S. The front rear bearing has a bearing stiffness defined by a radial displacement caused by applying a radial force at the axial center point of the bearing. The ratio of the bearing stiffness at the front rear bearing to the stiffness of the subspan is between 0.08 and 0.5 kN / mm 2 within the range.

[0200] The bearing stiffness ratio can be between 0.09 and 0.40 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.30 kN / mm 2 In the range of, optionally in the range of 0.08 to 0.20 kN / mm 2 In the range of, optionally in the range of 0.09 to 0.40 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.50 kN / mm 2 In the range of, optionally in the range of 0.15 to 0.40 kN / mm 2 and optionally in the range of 0.15 to 0.30 kN / mm 2 within the range.

[0201] The present inventors have recognised that controlling bearing stiffness and subspan so that the ratio of the two is within a specified range may allow or facilitate management of vibration modes, thereby potentially reducing damage to the engine caused by vortex mode displacement of the spindle in use.

[0202] The spindle has a length L between the front bearing and the last rear bearing. The bearings may be arranged so that the length ratio of the secondary span to the spindle length (S / L) is equal to or less than 0.14, or equal to or less than 0.13, or equal to or less than 0.12. The length ratio S / L may be equal to or greater than 0.05, or equal to or greater than 0.06, or equal to or greater than 0.07, or equal to or greater than 0.08. For example, the length ratio S / L may be in the range 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.

[0203] The spindle may have an operating speed range with a lower limit of 1500 rpm and an upper limit of 6200 rpm.

[0204] The bearing stiffness of the foremost rear bearing may be in the range of 30 kN / mm to 100 kN / mm. Optionally, the bearing stiffness of the foremost rear bearing may be at least substantially equal to 50 kN / mm.

[0205] The gas turbine engine may further include a fixed support structure and a first bearing support structure. The forward rear bearing may be mounted to the fixed support structure via the first bearing support structure. The first bearing support structure may be attached to the fixed support structure at a first position axially rearward of the forward rear bearing.

[0206] In such embodiments, the first bearing support structure may include a plurality of connection members that may be circumferentially spaced about the engine axis.The connection members may connect the forward-most rear bearing to the fixed support structure.

[0207] Additionally or alternatively, in such embodiments, the first bearing support structure may include an outer race of the forward-most aft bearing.

[0208] In embodiments having a first bearing support structure, the engine may further include a second bearing support structure. The second bearing support structure may be mounted to the fixed support structure, optionally at a second location that is forward of the first location and at a greater radial distance from the engine axis than the first location. The second bearing support structure may be connected to the first bearing support structure via a squeeze film damper in the forward-most rear bearing region.

[0209] The length L of the mandrel may be in the range of 1800 mm to 2900 mm or 2750 mm. The secondary span may be in the range of 250 mm to 350 mm, optionally in the range of 260 mm to 350 mm.

[0210] The fan may have a fan diameter in the range of 330 cm to 380 cm.

[0211] The length L of the spindle may be in the range of 1800 mm to 2900 mm or 2750 mm, and the subspan in the range of 250 mm to 350 mm, optionally in the range of 260 mm to 350 mm. The operating speed of the spindle may be in the range of 1500 rpm to 6200 rpm; and / or the diameter of the fan may be in the range of 330 cm to 380 cm, and the gear ratio of the gearbox in the range of 3.1 to 3.8. The length, subspan and / or operating speed may be selected so that the first resonance of the spindle is not within the operating range of the engine.

[0212] The rear bearing may be located axially flush with or behind the leading edge of the lowest pressure turbine blade of the turbine at the root of the blade.

[0213] The rear bearing may be positioned axially flush with or behind the trailing edge of a turbine blade in the third set of turbine blades from the front of the turbine at the root of the blade. In such embodiments, the turbine may comprise four sets of turbine blades.

[0214] The lowest pressure turbine of the engine may have a lowest pressure set of blades, each blade of which has a mass m, a radius r at mid-height of the blade, and a cruising angular velocity ω. The secondary span (S) is defined as the axial distance between the two rear bearings, as described elsewhere.

[0215] Ratio of first blade to bearing:

[0216]

[0217] can have values ​​in the following range: 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 4.0×10 -6 to 7.5×10 -6 kg - 1 .rad -2 .s 2 In the range of 5.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 In the range of 2.0×10-6 to 6.5×10 -6 kg -1 .rad -2 .s 2 In the range of 3.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 in the range of 4.5×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

[0218] Ratio of second blade to bearing:

[0219]

[0220] Can have values ​​in the following range: 0.8 to 6.0 kg -1 In the range of 0.8 to 5.0 kg -1 In the range of 0.8 to 4.0 kg -1 In the range of 0.8 to 3.0 kg -1 In the range of 0.8 to 2.0 kg -1 within the range.

[0221] The turbine has a length between the leading edge of the foremost turbine blade of the turbine and the trailing edge of the rearmost turbine blade of the turbine. The ratio of the subspan to the turbine length (subspan divided by the turbine length) may be equal to or less than 1.05, optionally equal to or less than 1.00, optionally equal to or less than 0.95. The ratio of the subspan to the turbine length may be equal to or greater than 0.70, or equal to or greater than 0.75, or equal to or greater than 0.80, or equal to or greater than 0.85. For example, the ratio of the secondary span to the turbine length may be in the range of 0.70 to 1.05, optionally in the range of 0.70 to 1.00, optionally in the range of 0.70 to 0.95, optionally in the range of 0.80 to 1.05, optionally in the range of 0.80 to 1.00, optionally in the range of 0.80 to 0.95, optionally in the range of 0.85 to 1.05, optionally in the range of 0.85 to 1.00, and also optionally in the range of 0.85 to 0.95.

[0222] In any of the above aspects, one or more of the following features may be present:

[0223] The turbine may be a first turbine, the compressor may be a first compressor, and the mandrel may be a first mandrel. The engine core may also include a second turbine, a second compressor, and a second mandrel connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second mandrel may be arranged to rotate at a higher rotational speed than the first mandrel.

[0224] The engine may also include an aft bearing housing located behind the turbine. The aft bearing housing may include two bearing discs; each bearing disc may be arranged to support one of the two rearward bearings. In an alternative embodiment, the aft bearing housing may include a single bearing disc arranged to support one of the two rearward bearings (e.g., the rearmost bearing).

[0225] In embodiments having one or more bearing discs, one or more of the bearing discs may be oriented at least substantially perpendicular to the engine axis (ie, at least substantially in a radial plane through the engine).

[0226] In various aspects and embodiments described herein, the engine operating range can be defined as the range of the rotation rate of the spindle during standard operation of the engine (e.g., during ground idling, takeoff, ascent, and cruising), and can be measured in revolutions per minute (rpm). In this context, "standard operation of the engine" can exclude transient periods of startup and shutdown, such as the period during which the spindle rotation rate increases from zero to the ground idling rotation rate. The engine operating range includes ground idling speed, cruising speed, and maximum take-off (MTO) speed. Throughout the standard operation of the engine, the spindle rotation rate may be greater than or equal to the ground idle rotation rate. During standard operation of the engine, the spindle rotation rate may also be referred to as the spindle operating speed.

[0227] As described elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may include an engine core including a turbine, a combustor, a compressor, and a spindle connecting the turbine to the compressor. Such a gas turbine engine may include a fan (with fan blades) located upstream of the engine core.

[0228] The arrangement of the present disclosure may be particularly, but not exclusively, beneficial to fans driven via a gearbox. Thus, the gas turbine engine may include a gearbox that receives input from a mandrel and outputs a drive to the fan so as to drive the fan at a rotational speed lower than the mandrel. The input to the gearbox may come directly from the mandrel or indirectly from the mandrel, for example via a spur gear shaft and / or gears. The mandrel may rigidly connect the turbine and the compressor so that the turbine and the compressor rotate at the same speed (wherein the fan rotates at a lower speed).

[0229] As described and / or claimed herein, a gas turbine engine may have any suitable general architecture. For example, a gas turbine engine may have any desired number of shafts, such as one, two, or three shafts, connecting a turbine and a compressor. By way of example only, the turbine connected to the mandrel may be a first turbine, the compressor connected to the mandrel may be a first compressor, and the mandrel may be a first mandrel. The engine core may also include a second turbine, a second compressor, and a second mandrel connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second mandrel may be arranged to rotate at a higher rotational speed than the first mandrel.

[0230] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor.The second compressor may be arranged to receive flow (eg directly, such as via a generally annular conduit) from the first compressor.

[0231] The gearbox may be arranged to be driven by the spindle configured to rotate (e.g. in use) at the lowest rotational speed (e.g. the first spindle in the above example). For example, the gearbox may be arranged to be driven only by the spindle configured to rotate (e.g. in use) at the lowest rotational speed (e.g. in the above example, only the first spindle, not the second spindle). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, such as the first and / or second shafts in the above example.

[0232] The gearbox can be a reduction gearbox (because the output to the fan is at a lower rotational rate than the input from the mandrel). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "stellar" gearbox, as described in more detail elsewhere herein. The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or 3.2 to 3.8, for example, about or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. For example, the gear ratio can be between any two values ​​in the previous sentence. By way of example only, the gearbox can be a "stellar" gearbox having a gear ratio in the range of 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio can be outside these ranges.

[0233] In any gas turbine engine as described and / or claimed herein, the burner may be disposed axially downstream of the fan and one or more compressors. For example, where a second compressor is provided, the burner may be located directly downstream of the second compressor (e.g., at its outlet). By way of another example, where a second turbine is provided, the flow at the burner outlet may be provided to the inlet of the second turbine. The burner may be disposed upstream of one or more turbines.

[0234] The compressor or each compressor (e.g., the first and second compressors described above) may include any number of stages, such as a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes, which may be variable stator vanes (because the angle of incidence of the row of stator vanes may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

[0235] The turbine or each turbine (e.g., the first turbine and the second turbine as described above) may include any number of stages, such as a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

[0236] Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas wash position or 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or approximately) any of the following: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 0.28 to 0.32. These ratios may generally be referred to as hub-tip ratios. Both the radius at the hub and the radius at the tip may be measured at the leading edge (or axially forward) portion of the blade. Of course, the hub-to-tip ratio refers to the gas-washed portion of the fan blade, ie, the portion radially outside of any platform.

[0237] The radius of the fan may be measured between the engine centerline and the tips at the leading edges of the fan blades. The fan diameter (which may be just twice the fan radius) may be greater than (or approximately) any of the following: 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches), or 420 cm (about 165 inches). The fan diameter may be within an inclusive range defined by any two values ​​in the previous sentence (ie, these values ​​may form an upper or lower limit), for example, within the range of 240 cm to 280 cm or 330 cm to 380 cm.

[0238] The rotational speed of the fan can vary in use. Generally speaking, for a fan with a larger diameter, the rotational speed is lower. By way of non-limiting example only, the rotational speed of the fan under cruising conditions may be less than 2500rpm, for example, less than 2300rpm. By way of another non-limiting example only, for an engine with a fan diameter in the range of 220cm to 300cm (e.g., 240cm to 280cm or 250cm to 270cm), the rotational speed of the fan under cruising conditions 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 another non-limiting example only, for an engine with a fan diameter in the range of 330cm to 380cm, the rotational speed of the fan under cruising conditions 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.

[0239] When a gas turbine engine is in use, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tips of the fan blades to move at a speed U 尖端 The work done by the fan blades 13 on the flow results in a rise in the enthalpy of the flow, dH. The fan tip load can be defined as dH / U 尖端 2 , where dH is the enthalpy rise across the fan (e.g., 1-D average enthalpy rise), and U 尖端is the (translational) speed of the fan tip, for example at the leading edge of the tip (may be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip load at cruise conditions may be greater than (or approximately) any of the following: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4. The fan tip load may be within an inclusive range bounded by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example in the range of 0.28 to 0.31 or 0.29 to 0.3.

[0240] A gas turbine engine according to the present disclosure may have any desired bypass ratio, wherein the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core under cruise conditions. In some arrangements, the bypass ratio may be greater than (or approximately) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, within a range of 12 to 16, or a range of 13 to 15, or a range of 13 to 14. The bypass duct may be substantially annular. The bypass duct may be located radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan case.

[0241] The total pressure ratio of the gas turbine engine described and / or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the outlet of the highest pressure compressor (before entering the combustor). By way of non-limiting example, the total pressure ratio of the gas turbine engine as described and / or claimed herein at cruise may be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), such as within a range of 50 to 70.

[0242] The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. Under cruise conditions, the specific thrust of the engine described and / or claimed herein may be less than (or approximately) any of the following: 110 Nkg -1 s, 105Nkg -1 s、100Nkg -1 s, 95Nkg -1 s、90Nkg -1 s, 85Nkg-1 s or 80Nkg -1 s. The specific thrust may be within an inclusive range defined by any two values ​​in the previous sentence (ie, these values ​​may form an upper or lower limit), for example, within 80 Nkg -1 s to 100Nkg -1 s, or 85Nkg -1 s to 95Nkg -1 Such engines are likely to be particularly efficient compared to conventional gas turbine engines.

[0243] The gas turbine engine as described and / or claimed herein may have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described herein may produce a maximum thrust of at least (or approximately) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN or 550 kN. The maximum thrust may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit). By way of example only, a gas turbine as described herein may be capable of producing a maximum thrust in the range of 330 kN to 420 kN, such as 350 kN to 400 kN. The thrust mentioned above may be the maximum net thrust at standard atmospheric conditions, at sea level, plus 15°C (ambient pressure 101.3 kPa, temperature 30°C), when the engine is stationary.

[0244] In use, the temperature of the flow at the inlet of the high-pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the outlet of the combustor, for example just upstream of the first turbine vane, which itself may be referred to as the nozzle guide vane. At cruise, the TET may be at least (or approximately) any one of the following: 1400K, 1450K, 1500K, 1550K, 1600K, or 1650K. The TET at cruise may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit). The maximum TET of the engine when in use may be, for example, at least (or approximately) any one of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, or 2000K. The maximum TET may be within an inclusive range defined by any two values ​​in the previous sentence (i.e., these values ​​may form an upper or lower limit), for example, within a range of 1800K to 1950K. Maximum TET may occur, for example, at high thrust conditions, such as at maximum takeoff (MTO) conditions.

[0245] As used herein, maximum takeoff (MTO) conditions have a conventional meaning. Maximum takeoff conditions may be defined as operating the engine at maximum takeoff thrust at the end of a runway at International Standard Atmosphere (ISA) sea level pressure and temperature conditions of +15°C, which is typically defined as an aircraft speed of about 0.25 Mn (i.e., a Mach number of 0.25), or between about 0.24 Mn and 0.27 Mn. Thus, the maximum takeoff condition for an engine may be defined as operating the engine at its maximum takeoff thrust at ISA sea level pressure and temperature of +15°C and an aircraft speed of 0.25 Mn.

[0246] The fan blades and / or airfoil portions of the fan blades described and / or claimed herein may be manufactured from any suitable material or combination of materials. For example, at least a portion of the fan blades and / or airfoils may be manufactured at least in part from a composite material, such as a metal matrix composite material and / or an organic matrix composite material, such as carbon fiber. In another exemplary manner, at least a portion of the fan blades and / or airfoils may be manufactured at least in part from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions manufactured from different materials. For example, the fan blade may have a protective leading edge, which may be manufactured from a material that resists impact (e.g., from birds, ice, or other materials) better than the rest of the blade. Such a leading edge may be manufactured, for example, using titanium or a titanium-based alloy. Therefore, by way of example only, the fan blade may have carbon fiber or an aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0247] A fan as described and / or claimed herein may include a central portion from which fan blades may extend, for example, radially. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fixing that may engage with a corresponding slot in a hub (or disc). By way of example only, such fixings may be in the form of a dovetail that may be inserted into and / or engage with a corresponding slot in the hub / disc so as to fix the fan blade to the hub / disc. In another example, the fan blade may be formed integrally with the central portion. Such an arrangement may be referred to as a blade disk or blade ring. Any suitable method may be used to manufacture such a blade disk or blade ring. For example, at least a portion of the fan blade may be machined from a block, and / or at least a portion of the fan blade may be attached to the hub / disc by welding (such as linear friction welding).

[0248] The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the outlet area of ​​the bypass duct to vary during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0249] A fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24, or 26 fan blades.

[0250] As used herein, cruise condition has a conventional meaning and will be readily understood by the skilled person. Thus, for a given gas turbine engine of an aircraft, the skilled person will immediately recognize that cruise condition refers to the operating point of the engine of the aircraft to which the gas turbine engine is designed to be attached for a given mission (which may be referred to in the industry as an "economic mission") at mid-cruise. In this regard, mid-cruise is a critical point in the flight cycle of an aircraft at which 50% of the total fuel burned between the peak of ascent and the beginning of descent has been burned (which may be approximated in terms of time and / or distance to the midpoint between the peak of ascent and the beginning of descent). Thus, cruise condition defines the operating point of the gas turbine engine that provides a thrust that will ensure steady-state operation of the aircraft at mid-cruise (i.e., maintaining a constant altitude and a constant Mach number Mn), taking into account the number of engines provided to the aircraft to which the gas turbine engine is designed to be attached. For example, if the engine is designed to be attached to an aircraft having two engines of the same type, then under cruise condition, the engine provides half of the total thrust required for steady-state operation of the aircraft at mid-cruise.

[0251] In other words, for a given gas turbine engine of an aircraft, the cruise condition is defined as the operating point of the engine that provides a specified thrust at mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at mid-cruise altitude) (necessary to provide steady-state operation of the aircraft to which the gas turbine engine is designed to be attached, in combination with any other engines on the aircraft, at a given mid-cruise Mach number). For any given gas turbine engine of an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, so the operating point of the engine at cruise conditions is well defined.

[0252] By way of example only, the forward speed at cruise conditions may be anywhere within the range of from Mach 0.7 to Mach 0.9, such as 0.75 to 0.85, such as 0.76 to 0.84, such as 0.77 to 0.83, such as 0.78 to 0.82, such as 0.79 to 0.81, such as about Mach 0.8, about Mach 0.85, or 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise condition may be outside these ranges, such as below Mach 0.7 or above Mach 0.9.

[0253] By way of example only, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere ISA) at an altitude within the following range: 10000 m to 15000 m, for example within the range of 10000 m to 12000 m, for example within the range of 10400 m to 11600 m (about 38000 feet), for example within the range of 10500 m to 11500 m, for example within the range of 10600 m to 11400 m, for example within the range of 10700 m (about 35000 feet) to 11300 m, for example within the range of 10800 m to 11200 m, for example within the range of 10900 m to 11100 m, for example approximately 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

[0254] By way of example only, the cruise condition may correspond to an operating point of the engine that provides a known desired thrust level (e.g., a value in the range of 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 ft (11,582 m). By way of another example only, the cruise condition may correspond to an operating point of the engine that provides a known desired thrust level (e.g., a value in the range of 50 kN to 65 kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 ft (10,668 m).

[0255] In use, the gas turbine engines described and / or claimed herein may be operated at cruise conditions defined elsewhere herein. Such cruise conditions may be determined by a cruise condition (e.g., an intermediate cruise condition) of an aircraft on which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

[0256] According to one aspect, an aircraft is provided, the aircraft comprising a gas turbine engine as described herein. The aircraft according to this aspect is an aircraft to which the gas turbine engine has been designed for attachment. Thus, the cruise condition according to this aspect corresponds to an intermediate cruise of the aircraft, as defined elsewhere herein.

[0257] According to one aspect, there is provided a method of operating a gas turbine engine as described herein. The operation may be carried out at cruise conditions (eg, in terms of thrust, atmospheric conditions and Mach number) as defined elsewhere herein.

[0258] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described herein. Operation according to this aspect may comprise (or may be) operation at an intermediate cruise of the aircraft, as defined elsewhere herein.

[0259] The skilled person will appreciate that, unless mutually exclusive, features or parameters described in any one of the above aspects may be applied to any other aspect. In addition, unless mutually exclusive, any features or parameters described herein may be applied to any aspect and / or combined with any other features or parameters described herein. BRIEF DESCRIPTION OF THE DRAWINGS

[0260] Embodiments will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0261] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0262] Figure 2 is a close-up cross-sectional side view of an upstream portion of a gas turbine engine;

[0263] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

[0264] Figure 4 is a cross-sectional side view of a gas turbine engine showing three bearings on the mandrel;

[0265] Figure 5 is a schematic side elevation view of a gas turbine engine showing the primary span and secondary span between bearings;

[0266] Figure 6 The bending patterns of the mandrel for two different bearing configurations are shown;

[0267] Figure 7 is a schematic side view of a gas turbine engine showing Figure 5 Alternative bearing arrangement shown;

[0268] Figure 8 It is a schematic diagram of mounting two bearings on the tail bearing housing;

[0269] Fig. 9 So with Figure 8 Schematic diagram of different arrangements of mounting two bearings on the aft bearing housing;

[0270] Fig.10 A method of operating a gas turbine engine is shown;

[0271] Fig.11 The first four resonant frequencies of the beam are shown;

[0272] Fig.12 yes Figure 5 A close up view of the bearing housing is shown;

[0273] Fig.13 is a perspective view of a first bearing support structure;

[0274] Fig.14 is in the proper position in the bearing housing Fig.13 a cross-sectional view of a first bearing support structure shown;

[0275] Fig.15A and Fig. 15B shows the bearing stiffness determination, specifically showing the application of radial force and the resulting displacement;

[0276] Fig.16 A method of obtaining a gas turbine engine as described herein is shown;

[0277] Fig.17 is a graph of displacement versus load showing the elastic region within which the stiffness of the components can be determined;

[0278] Fig.18 A method of designing a gas turbine engine as described herein is shown;

[0279] Fig.19 Yes Install Fig. 9 a schematic diagram of the two bearings shown with the bearing disc angles marked; and

[0280] Fig. 20 So with Fig. 9 Schematic diagram of two bearings mounted in different ways, with the corresponding bearing disc angles marked. DETAILED DESCRIPTION

[0281] Figure 1A gas turbine engine 10 having a main axis of rotation 9 is shown. The engine 10 includes an air inlet 12 and a pusher fan 23, which produces two air flows: a core air flow A and a bypass air flow B. The gas turbine engine 10 includes a core 11 that receives the core air flow A. The engine core 11 includes a low-pressure compressor 14, a high-pressure compressor 15, a combustion device 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core exhaust nozzle 20 in axial series. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass air flow B flows through the bypass duct 22. The fan 23 is attached to the low-pressure turbine 19 via a core shaft or a first core shaft 26 and an epicyclic gearbox 30 and is driven by the low-pressure turbine.

[0282] In use, the core air flow A is accelerated and compressed by the low-pressure compressor 14 and directed to the high-pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is directed to the combustion device 16, in which the compressed air is mixed with fuel and the mixture is burned. The resulting hot combustion products are then expanded through the high-pressure turbine and the low-pressure turbine 17, 19 before being discharged through the nozzle 20, thereby driving the high-pressure turbine and the low-pressure turbine to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 through a suitable interconnecting shaft 27 or a second core shaft. The fan 23 usually provides most of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0283] Figure 2 An exemplary arrangement of a geared fan gas turbine engine 10 is shown in FIG. Figure 1 ) drives a spindle 26 which is coupled to a sun wheel or sun gear 28 of an epicyclic gear arrangement 30. Radially outwardly of and intermeshing with the sun gear 28 are a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously about the sun gear 28 while causing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the fan 23 via a connecting rod 36 so as to drive the fan in rotation about the engine axis 9. Radially outwardly of and intermeshing with the planet gears 32 is a ring gear or annular gear 38 which is coupled to the fixed structure 24 via a connecting rod 40.

[0284] It should be noted that the terms "low-pressure turbine" and "low-pressure compressor" used in this article may refer to the lowest pressure turbine stage and the lowest pressure compressor stage (i.e., excluding the fan 23), respectively, and / or the turbine stage and the compressor stage connected together by the interconnecting shaft with the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft driving the fan 23). In some documents, the "low-pressure turbine" and "low-pressure compressor" mentioned in this article may be alternatively referred to as "intermediate-pressure turbine" and "intermediate-pressure compressor". In the case of using such alternative naming, the fan 23 may be referred to as the first or lowest pressure compression stage.

[0285] exist Figure 3 The epicyclic gearbox 30 is shown in more detail by way of example in FIG. Each of the sun gear 28, the planetary gears 32, and the ring gear 38 includes teeth around its periphery for intermeshing with the other gears. However, for clarity, Figure 3 Only an exemplary portion of the teeth are shown. Four planetary gears 32 are shown, but it will be apparent to those skilled in the art that more or fewer planetary gears 32 may be provided within the scope of the claimed invention. Practical applications of the planetary epicyclic gearbox 30 typically include at least three planetary gears 32.

[0286] exist Figure 2 and Figure 3 The epicyclic gearbox 30 shown by way of example in FIG. 1 is a planetary type, in which the planet carrier 34 is coupled to the output shaft via connecting rods 36, in which the ring gear 38 is fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of another example, the epicyclic gearbox 30 may be a star arrangement in which the planet carrier 34 remains fixed, allowing the ring gear (or ring gear) 38 to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. By way of another alternative example, the gearbox 30 may be a differential gearbox in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0287] It should be understood that Figure 2 and Figure 3 The arrangement shown in is exemplary only, and various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement may be used to position the gearbox 30 in the engine 10 and / or for connecting the gearbox 30 to the engine 10. By way of another example, the connections (such as the gearbox 30) between the gearbox 30 and other components of the engine 10 (such as the input shaft (spindle) 26, the output shaft, and the fixed structure 24) may be such that the gearbox 30 is located within the engine 10 and / or is connected to the engine 10. Figure 2The connecting rods 36, 40 in the example may have any desired degree of stiffness or flexibility. By way of another example, any suitable arrangement of bearings between rotating and stationary components of the engine (e.g., between the input and output shafts from the gearbox and a stationary structure such as a gearbox housing) may be used, and the present disclosure is not limited to Figure 2 For example, where the gearbox 30 has a star arrangement (as described above), the skilled person will readily appreciate that the arrangement of the output and support links and bearing locations is typically different than Figure 2 The arrangement structure is shown by way of example in FIG.

[0288] Thus, the present disclosure extends to gas turbine engines having any arrangement in gearbox type (eg sun or planetary), support structure, input and output shaft arrangement and bearing locations.

[0289] Optionally, the gearbox may drive additional and / or alternative components (eg, an intermediate pressure compressor and / or a booster compressor).

[0290] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have alternative numbers of compressors and / or turbines and / or alternative numbers of interconnecting shafts. By way of further example, Figure 1 The gas turbine engine shown in has a splitter nozzle 18, 20, which means that the flow passing through the bypass duct 22 has its own nozzle 18, which is separated from the core engine nozzle 20 and radially outside the core engine nozzle. However, this is not restrictive, and any aspect of the present disclosure may also be applied to an engine in which the flow passing through the bypass duct 22 and the flow passing through the core 11 are mixed or combined before (or upstream) a single nozzle that may be referred to as a mixed flow nozzle. One or two nozzles (whether mixing or splitting) may have a fixed or variable area. Although the examples described relate to turbofan engines, the present disclosure may be applied to, for example, any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or, for example, a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0291] The geometry of the gas turbine engine 10 and its components is defined by conventional shafting, including an axial direction (aligned with the axis of rotation 9), a radial direction (in Figure 1 from bottom to top) and circumferential direction (perpendicular to Figure 1 view). The axial, radial, and circumferential directions are perpendicular to each other.

[0292] In the depicted embodiment, the engine 10 is a geared gas turbine engine having a gearbox 30 .

[0293] The fan 23 is attached to and driven by the low-pressure turbine 19 via a spindle 26 and an epicyclic gearbox 30. The spindle 26 is supported by three bearings 26a, 26b, 26c, including a front bearing 26a and two rear bearings 26b, 26c. In an alternative embodiment, more bearings 26a-c may be provided. The front bearing 26a may be referred to as the first bearing, the frontmost bearing 26b of the two rear bearings 26b, 26c may be referred to as the second bearing, and the last bearing 26c of the two rear bearings 26b, 26c may be referred to as the third bearing.

[0294] The front bearing 26a is a locating bearing 26a for the spindle 26; that is, a bearing on the spindle that is arranged to limit axial movement as well as radial movement of the spindle 26 and thereby axially locate the spindle. The front bearing 26a is mounted on a fixed structure to axially locate the spindle 26. The skilled person will appreciate that having multiple locating bearings on a single shaft can cause the shaft to bend or otherwise deform detrimentally when it is elongated in use, and therefore it is generally advantageous to use a single locating bearing per shaft.

[0295] In the depicted embodiment, the front bearing 26a is the spindle bearing 26a closest to the front of the engine 10. In alternative embodiments, there may be one or more other bearings, such as roller bearings, on the spindle 26 forward of the front bearing 26a, however any such bearings are not locating bearings, i.e., while such bearings may help to radially locate the spindle 26, the spindle may move axially relative to these bearings.

[0296] In the depicted embodiment, the front bearing 26a is the spindle bearing 26a closest to and behind the gearbox 30. In the depicted embodiment, the front bearing 26a is axially flush with or close to the outlet of the compressor 14. In alternative embodiments, the front bearing 26a may be located forward of the gearbox 30, however, the proximity of the front bearing 26a to the roller bearings of the fan 23 may increase complexity.

[0297] The front bearing 26a is mounted on the fixed structure.

[0298] The rear bearings 26b, 26c are the lower two spindle bearings, located behind the front bearing 26a. In the depicted embodiment, the rear bearings 26b, 26c are the spindle bearings closest to the rear of the engine 10. In alternative embodiments, additional bearings may be located behind these bearings.

[0299] exist Figure 4In the embodiment shown, both rear bearings 26b, 26c are mounted on a tail bearing housing 29. The tail bearing housing 29 is a structure that is arranged not to rotate relative to a fixed structure and supports the bearings 26b, 26c of the spindle 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.

[0300] exist Figure 4 In the embodiment shown, the last bearing 26c of the two rear bearings 26b, 26c is positioned axially flush with or close to the outlet of the low-pressure turbine 19. More specifically, in the described embodiment, the last bearing 26c is at least substantially axially flush with the last / lowest-pressure rotor of the low-pressure turbine 19.

[0301] In various alternative embodiments, such as Figures 5 to 7 In the embodiment shown, one or both of the two rear bearings 26 b , 26 c may be arranged axially downstream of the leading edge of the lowest pressure set of blades 19 d of the low pressure turbine 19 (at the root of the blades).

[0302] In implementing schemes such as Figure 1 and Figure 4 In the illustrated embodiment, the low-pressure turbine 19 may have three stages; that is, three groups of rotor blades 19a, 19b, 19c. Each group of rotor blades has a corresponding axial position along the engine axis 9 and is offset from the axial positions of the other groups. The most upstream or axially forward group of blades is the highest pressure group of blades of the low-pressure turbine 19 and may be referred to as the first group of blades. The most downstream or axially last group of blades is the lowest pressure group of blades of the low-pressure turbine 19 and may be referred to as the last group of blades, or in these embodiments, the third group of blades. The middle group of blades 19b of the three groups may be referred to as the second group of blades.

[0303] In implementing schemes such as Figure 5 and Figure 6 In the illustrated embodiment, the low-pressure turbine 19 may have four stages; that is, four groups of rotor blades 19a, 19b, 19c, 19d. The most upstream or axially forward group of blades is the highest pressure group of blades of the low-pressure turbine 19 and may be referred to as the first group of blades 19a. The most downstream or axially last group of blades 19d is the lowest pressure group of blades of the low-pressure turbine 19 and may be referred to as the last group of blades, or in these embodiments, the fourth group of blades. The middle groups of blades 19b, 19c of the four groups may be referred to as the second group of blades and the third group of blades, respectively.

[0304] In alternative embodiments, turbine 19 may have fewer than three or more than four sets of blades, such as two or five sets of blades.

[0305] The length of the low-pressure turbine 19 may be defined as the distance between the leading edge of blade 19a in the first set of blades of the low-pressure turbine and the trailing edge of blade 19c / 19d in the last set of blades of the low-pressure turbine. The casing of the low-pressure turbine 19 may extend beyond the span between the first (highest pressure) blade and the last (lowest pressure) blade.

[0306] In various embodiments where the two rear bearings 26b, 26c are axially disposed downstream of the leading edge of the lowest pressure blade 19c, 19d of the low pressure turbine 19 (at the root of the blade), the length ratio (described below) and / or the spindle operating speed may or may not be within the ranges described in detail elsewhere herein. In such embodiments, the two rear bearings 26b, 26c may be axially disposed downstream of the centerline of the disk supporting the lowest pressure turbine blade 19d of the low pressure turbine 19, such as Fig.12 In such embodiments, the length L of the mandrel may be in the range of 1800 mm to 2900 mm, optionally 2300 mm to 2800 mm, and also optionally 2400 mm to 2750 mm.

[0307] In various embodiments, the low pressure turbine 19 may be a four-stage turbine 19 having four sets (19a-19d) of rotor blades, such as described above with respect to Figure 5 and Figure 6 As described. The frontmost group of rotor blades 19a may be defined as the first group, and the last group 19d may be defined as the fourth group. The two rear bearings 26b, 26c may be axially disposed downstream of the trailing edge of the low-pressure turbine blades of the third group 19c (at their root). In such an embodiment, one or both of the two rear bearings 26b, 26c may be axially disposed downstream of the leading edge of the low-pressure turbine blades 19d (the fourth group of blades) of the low-pressure turbine 19 (at the root of the blade). The pressure at both ends of the low-pressure turbine 19 is reduced—therefore the first group of blades 19a may be described as the highest pressure group of blades of the low-pressure turbine 19, and the fourth group of blades may be described as the lowest pressure group of blades.

[0308] In various such embodiments, the length ratio and / or the mandrel operating speed may or may not be within the ranges described in detail elsewhere herein. In such embodiments, the length L of the mandrel may be within the range of 1800 mm to 2900 mm, optionally 2300 mm to 2800 mm, and also optionally 2400 mm to 2750 mm.

[0309] In the low-pressure turbine 19, the turbine blades 19a-19d within each group are designed to be identical (within manufacturing tolerances). The blades between different groups may differ, for example, in size and / or shape. Each blade within group 19d has a mass m and a turbine blade radius r at mid-height. Generally speaking, heavier and larger blades are located in the group toward the rear of the turbine. The turbine blades each have a mid-height position at which the mid-height radius is measured. The mid-height position is the mid-position between the radially innermost point and the radially outermost point on the leading edge of the blade. The turbine blade radius at mid-height is measured in the radial direction between the axial centerline 9 of the engine 10 and the mid-height position.

[0310] Each blade 19a-19d also has a maximum rated angular velocity ω, which may also be referred to as a maximum take-off (MTO) speed. The MTO speed may be the maximum angular velocity at which the shaft is designed to rotate. ω may be between 5000 rpm and 9000 rpm, optionally in the range of 5000 rpm to 8000 rpm or 5000 rpm to 7000 rpm, and optionally about 5500 rpm, 6000 rpm (i.e., about 630 rad / sec), or 6500 rpm.

[0311] The value Y can be defined as follows:

[0312] Y=mrω 2

[0313] Y may have the unit kg.m.rad 2 .s -2 In some embodiments, the value Y of the lowest pressure group blades 19c (for a three-stage turbine) and 19d (for a four-stage turbine) of the low-pressure turbine 19 is between 45,000 and 100,000 kg.m.rad 2 .s -2 In the range of 50000 to 100000 kg.m.rad 2 .s -2 In the range of 55000 to 100000 kg.m.rad 2 .s -2 In the range of, and optionally also in the range of 60000 to 100000 kg.m.rad 2 .s -2 In such embodiments, the mass m of each blade in the last group may be in the range of 0.2 kg to 0.6 kg, and optionally may be about 0.4 kg. The radius r of each blade in the last group at mid-height may be in the range of 400 mm to 600 mm, and optionally may be about 500 mm (0.5 m).

[0314] The value Y can be considered to provide the centripetal force (F) acting on the blade when rotating at MTO speed. c )'s magnitude:

[0315]

[0316] where a c is the centripetal acceleration and v is the linear velocity, which is equal to ωr.

[0317] The skilled person will appreciate that the higher centripetal force (F c ) may increase the risk of a blade-off event for that blade.

[0318] The secondary span S is the distance between the two last bearings and may be in the range of 250 mm to 350 mm, optionally in the range of 275 mm to 325 mm, and further optionally may be about 300 mm.

[0319] The first blade to bearing ratio can be defined as:

[0320]

[0321] The ratio of the first blade to the bearing can be 2.0×10 -6 to 7.5×10 -6kg-1 .rad -2 .s 2 in the range of, and optionally in the range of 3.0×10 -6 Up to 4.5×10 -6kg-1 .rad -2 .s 2 The value of this ratio can be lower than that of a conventional engine because the subspan S is smaller than the value Y. The skilled person will appreciate that increasing the angular velocity in flight can improve engine efficiency, and that the MTO speed provides a measure of the maximum rotational speed that can be obtained from the engine 10, and therefore provides a measure of the maximum force (indicated by Y). However, the inventors recognize that the engine 10 should not increase linearly in proportion to the increase in force (Y provides a measure of force), but rather the subspan length should be increased as little as possible so as to relatively reduce the engine length and weight, thereby allowing efficiency gains to be increased by avoiding additional weight, and so as to avoid the formation of unwanted vortex patterns within the subspan. Although conventional wisdom suggests that a larger subspan is needed to improve the reaction force from the low-pressure turbine 19, the inventors have found that the risk of introducing vortex patterns and the introduction of more length and weight offset the benefits of the force reaction force.

[0322] In an alternative or additional embodiment, the second blade to bearing ratio may be defined as:

[0323]

[0324] The value of the blade mass m multiplied by the blade radius r may be in the range of 180 kg.mm to 280 kg.mm. The secondary span S, i.e. the distance between the two last bearings, may be in the range of 250 mm to 350 mm. In various embodiments, the secondary span may be greater than or equal to any one of 250 mm, 255 mm, 260 mm, and 265 mm. In various embodiments, the secondary span may be less than or equal to any one of 350 mm, 345 mm, 340 mm, or 335 mm. The ratio of the second blade to the bearing may be in the range of 0.8 to 6.0 kg. -1 In the range of, and optionally in the range of 0.9 to 3.9 kg -1 In the range of, and optionally also in the range of 1.2 to 2.6 kg -1 In such embodiments, the gear ratio of the gearbox 30 may be greater than 3, and optionally in the range of 3.1 to 3.8.

[0325] In such embodiments, the operating speed of the engine 10 / spindle 26 may be in the range of 5400 rpm to 5700 rpm (i.e., approximately 565 to 597 rad / sec) at cruise, and optionally in the range of 5500 rpm to 5600 rpm. Additionally or alternatively, the operating speed of the spindle 26 may be in the range of 5800 rpm to 6200 rpm at MTO, and optionally in the range of 5900 rpm to 6100 rpm.

[0326] In such embodiments, the length L of the mandrel 26 may be in the range of 1800 mm to 2900 mm, optionally 2000 mm to 2900 mm, further optionally 2300 mm to 2800 mm, and further optionally 2400 mm to 2750 mm.

[0327] Specifically, in Figure 5 and Figure 6 In the embodiment shown, both the aft bearings 26b, 26c are mounted on an aft bearing housing 29. The aft bearing housing 29 is located axially to the rear of the low pressure turbine 19. Two bearing discs 29a, 29b extend from the aft bearing housing 29. Although the forwardmost bearing disc 29a of the bearing housing 29 of this embodiment is angled forward of the aft bearing housing 29, extending towards the casing of the low pressure turbine 19 and in this case extending within the casing, the forwardmost bearing 26b of the aft bearing, mounted on the forwardmost bearing disc 29a, is still located axially to the rear of the leading edge of the lowest pressure turbine blade of the low pressure turbine 19 (at the root of the blade). The last bearing 26c of the aft bearing, mounted on the last bearing disc 29b, is located axially to the rear of the low pressure turbine 19.

[0328] exist Figure 7In the illustrated embodiment, the aft bearing housing 29 has only one bearing disc 29c extending therefrom. The single bearing disc 29c is located adjacent to the Figure 5 and Figure 6 The foremost bearing disc 29a of the embodiment shown is in a similar position and has a similar shape thereto, but this time the single bearing disc supports the last bearing 26c in the rear bearing, thereby maintaining the bearing 26c axially rearward of the leading edge of the lowest pressure turbine blade of the low pressure turbine 19. The foremost bearing 26b in the rear bearing is mounted on the front bearing housing 31, rather than on the rear bearing housing 29. The front bearing housing 31 has a single bearing disc 31a, which is arranged to support the foremost bearing 26b in the rear bearing. In this embodiment, the foremost bearing 26b in the rear bearing is located in front of and adjacent to the low pressure turbine 19. In this embodiment, the foremost bearing 26b in the rear bearing remains axially flush with a portion of the low pressure turbine 19, and more specifically, remains axially flush with the leading edge of the highest pressure turbine blade of the low pressure turbine 19. In such an embodiment, the core shaft 26 may be shorter overall, with only one rear bearing 26c located behind the leading edge of the lowest pressure turbine blade of the low-pressure turbine 19 and positioned to be at least partially axially flush with the low-pressure turbine 19, rather than being strictly located behind the low-pressure turbine 19 (more specifically, behind the leading edge of the lowest pressure turbine blade of the low-pressure turbine 19) because the core shaft 26 may not extend far to the rear of the low-pressure turbine 19 in view of the secondary span S.

[0329] The length L of the spindle 26 may be defined between the front bearing 26a and the last rear bearing 26c, as shown in FIG. Figure 5 and 6 The length L of the mandrel 26 is marked (i.e., between the first bearing and the third bearing). This length L may be referred to as the main span of the mandrel 26, or the bearing-to-bearing length of the mandrel 26. In the described embodiment, the length L of the mandrel 26 is in the range of 1800 mm to 2900 mm, more specifically in the range of 2300 mm to 2800 mm, and even more specifically in the range of 2400 mm to 2750 mm.

[0330] The secondary span S of the spindle 26 may be defined as the distance between the rear bearings 26 b , 26 c (ie, between the second and third bearings on the spindle 26 ).

[0331] The secondary span S is equal to the length L of the spindle 26 minus the distance D between the front bearing 26a and the first rear bearing 26b (ie, minus the distance between the first bearing and the second bearing on the spindle 26).

[0332] The length ratio of the mandrel can be defined as:

[0333]

[0334] In various embodiments, the bearings 26a-26c are arranged so that the length ratio of the secondary span to the mandrel length is in the range of 0.08 to 0.14, and optionally in the range of 0.08 to 0.13. In various embodiments, the mandrel 26 can have any suitable length ratio, such as in the range of 0.09 to 0.13, or in the range of 0.10 to 0.12, or such as about or at least 0.08, 0.09, 0.10, 0.11, 0.12, 0.13, or 0.14. The mandrel length ratio can be, for example, between any two of the values ​​in the previous sentence.

[0335] The inventors have recognized that the ratio of the distance between the two rear bearings 26 b , 26 c to the shaft length L may be the main parameter controlling the rotor dynamics of the spindle 26 .

[0336] The value of this length ratio of the engine 10 of various embodiments may be relatively low compared to previous engines.

[0337] As the engine 10 gets larger, the spindle 26 gets longer, however, increasing the distance S (sub-span) between the two rear bearings 26b-26c linearly with increasing shaft length L can lead to detrimental effects. Once the sub-span S reaches a certain length, there may be no benefit in extending the distance in terms of the effect on the portion of the spindle 26 between the first bearing 26a and the second 26b bearing and / or in terms of the reaction moment to the torque of the bearing plates 29a-29b, 31a. In fact, if this distance S increases linearly with the spindle length L, the sub-span may become long enough to cause the vortex pattern between the second bearing 26b and the third bearing (as described below with respect to Fig.11 Said) within the engine operating range, this may exacerbate the undesired movement.

[0338] Additionally or alternatively, a relatively larger tail bearing housing (TBH) 29 may be used to accommodate a wider subspan S between the bearings 26b, 26c, such as Figure 8 As shown. A larger TBH 29 may add undesirable weight and / or volume to the engine 10. Alternatively or additionally, the bearing plates 29a, 29b may be angled wider to accommodate a wider sub-span between the bearings 26b, 26c, such as Fig. 9 From a structural point of view, bearing discs 29a, 29b angled further away from the direction perpendicular to the engine axis 9 / parallel to the radius may be a suboptimal solution, for example, the moment of the bearing disc may not be as Figure 8 Alternatively or additionally, one of the bearings 29a, 29b may be located on a different bearing structure 31 than on the TBH 29.

[0339] In various embodiments, the bearing plates 29a, 29b, 31a (whether they are on the TBH 29 or on a separate bearing structure 31) can be oriented to be at least substantially perpendicular to the engine axis 9, for example, having an angle (θ) with the radius of the engine 10 within the range of: between 0° and 20°, and optionally between 0° and 15° (i.e., an angle with the engine axis 9 within the range of: between 90° and 70°, and optionally between 90° and 75°).

[0340] In such Fig. 9 and Fig.19 In the illustrated embodiment, the bearing discs 29a, 29b each extend inwardly from the remainder of the TBH 29 towards the engine axis 9 at an at least substantially constant angle (the bearing discs 29a, 29b may, for example, comprise a solid disc or a series of circumferentially spaced struts around an axis extending between the inner and outer rings). An angle θ may thus be defined between the discs 29a, 29b and the radial direction. 1 , Θ 2 In other embodiments, such as Fig. 20 As shown, the bearing plates 29a, 29b may include two or more portions extending at different angles (e.g. Fig. 20 In the example shown here, the frontmost bearing disc 29a has three parts and the rear bearing disc 29b has two parts). In such a case, the angle selected is the angle for the part with the longest radial extent, such as for Fig. 20 The two examples shown are labeled. In embodiments where there is no single portion having a greater radial extent than other portions, the average angle may instead be taken over the two or more portions having the largest radial extents.

[0341] In the described embodiment, the spindle 26 has an operating speed range with a lower limit of 1500 rpm (e.g., when idling on the ground) and an upper limit of 6200 rpm (e.g., at maximum takeoff MTO). Specifically, in various embodiments, the spindle cruising operating speed is in the range of 5400 rpm to 5700 rpm, and optionally in the range of 5500 rpm to 5600 rpm. At MTO, the operating speed of the spindle may be in the range of 5800 rpm to 6200 rpm, and optionally in the range of 5900 rpm to 6100 rpm. The cruising operating speed range of a particular aircraft is typically much lower than the maximum rated rotation rate (MTO speed) of the spindle 26 of the aircraft. During normal operation, the engine 10 may operate within the MTO operating speed range for a relatively short duration, such as five or ten minutes.

[0342] Figure 6The toughening effect on the spindle 26 obtained by arranging the bearings 26a-26c as described herein, and specifically using two rear bearings 26b, 26c, is shown. The positioning of the bearings is selected to control or help control the swirl pattern of the engine 10.

[0343] Specifically, the boundary conditions at the location of the second bearing 26b (the forward-most rear bearing) on ​​the mandrel 26 can be changed by extending the mandrel 26 beyond the second bearing 26b to the third bearing 26c (the rear-most bearing). This can change the boundary conditions at the second bearing 26b from a simply supported / pinned boundary condition to a cantilever boundary condition, thereby affecting the shape of the mandrel 26 when it is bent. Figure 6 As shown, this effect can be most clearly seen just to the left of the second bearing 26b, where the angle of the core shaft at the second bearing 26b in the upper two-bearing configuration (pin-to-pin boundary condition) is steeper than the angle in the lower three-bearing configuration (pin-to-cantilever boundary condition), in which the shaft 26 is closer to horizontal when entering the second bearing 26b.

[0344] In rotor dynamics, the critical speed is the theoretical angular velocity that excites the natural frequency of a rotating object such as a shaft (e.g., spindle 26). Higher frequency vortex modes (e.g., twice the natural / resonant frequency) may also be induced. The engine 10 of the described embodiment may be susceptible to vortex modes (resonant frequencies) within or near the engine operating range that may result in undesirable and potentially harmful movement of the spindle 26, such as Figure 6 This alignment of the swirl pattern with the engine operating range may be a result of the longer spindle 26 of the larger engine 10.

[0345] While the diameter and / or thickness of the mandrel 26 may be varied to increase stiffness and thereby push the swirl mode outside the engine operating envelope, the resulting increase in size and / or weight and / or impact effects on other components may remove or reduce the feasibility of this option. For example, the mandrel diameter may be constrained by other engine components located radially outboard of the mandrel 26 (e.g., in the depicted embodiment, the mandrel 26 is radially inward relative to the interconnecting shaft 27 connecting the second (high pressure) turbine 17 to the second (high pressure) compressor 15).

[0346] In the depicted embodiment, the low-pressure turbine 19 directly drives the low-pressure compressor 14 and indirectly drives the fan 23 via a reduction gearbox 30. The high-pressure system 15, 17, including the high-pressure compressor 15 and the high-pressure turbine 17 and the interconnecting shaft 27 between the two, is located between the low-pressure compressor 14 and the low-pressure turbine 19. Therefore, the core shaft 26 is longer than the interconnecting shaft 27 because it extends over the entire length of the high-pressure system 15, 17 and the additional length of the low-pressure system 14, 19.

[0347] The relatively longer length L of the spindle 26 reduces the natural frequency of the spindle compared to previous engine architectures so that the natural frequency is within the operating speed of the engine 10 because the natural frequency is inversely proportional to the main span (spindle length L).

[0348] Natural frequencies exist in various modes (which may be referred to as vortex modes), such as Fig.11 As shown. The first resonance (mode 1) has a total of two nodes (non-moving points), one node at each constrained end (bearings 26a, 26b on the shaft 26). The second resonance (mode 2) has one additional node in the center (three nodes total). The third resonant mode (mode 3) has four nodes, and the fourth resonant mode (mode 4) has five nodes, etc. The nodes are evenly spaced along the span of the shaft 26 between the constrained ends 26a, 26b. The amplitude decreases as the mode number increases, with the maximum amplitude of the mode 1 resonance being greater than the maximum amplitude of the mode 2 resonance, etc. Therefore, the first resonance can be the most destructive because it causes the largest radial displacement of the shaft 26. Therefore, it can be particularly advantageous to avoid operating speeds that may trigger the first resonance by moving the first resonance out of the operating speed range.

[0349] The natural frequency of a simply supported (pinned) beam is related to the length of the beam by:

[0350]

[0351] in:

[0352] n is the mode number (where mode 1 has nodes only at each pinned end, such as Figure 6 As shown, mode 2 has an additional node at the midpoint between the two ends, etc.);

[0353] f n is the frequency of the nth mode (resonant frequency), measured in Hz;

[0354] K n is a dimensionless factor that depends on the order n of the mode and also on the applicable boundary conditions, and can be derived for various boundary conditions as detailed in, for example, "Roark's Formulas for Stress and Strain", 7th edition, by Warren C. Young and Richard G. Budynas.

[0355] E is the elastic modulus of the beam (N / m 2 );

[0356] I is the area moment of inertia of the beam (m 4 );

[0357] g is the acceleration due to gravity (m / s 2 );

[0358] l is the length of the beam between the pinned ends (m); and

[0359] W is the load per unit length on the mandrel (N / m).

[0360] The node positions (according to the beam length l) and K of the first five modes n The values ​​of can be listed in Table 1. The data in Table 1 are for beams with pins at both ends (boundary conditions), so K n Equal to (nπ) 2 For example, for the second mode, n=2 and K n =(2π) 2 =39.5 (accurate to three significant figures).

[0361] Table 1 :

[0362]

[0363] The bending stiffness of the mandrel 26 is related to both the material properties and the area moment of inertia I (geometry). The area moment of inertia of a tubular structure such as the mandrel 26 is:

[0364]

[0365] where r 1 and r 2 are the inner and outer radii of the tube, respectively.

[0366] To increase the frequency, available options include using a stiffer material for the shaft 26, reducing the length L of the shaft 26, or increasing the diameter of the shaft. Diameter and shaft length are quadratic functions of frequency and therefore have the greatest impact on changes of a given magnitude. However, diameter is limited by the lack of available space within the engine 10.

[0367] Providing a second rear bearing 26c modifies the boundary conditions to simulate a cantilever, and as described above, this change in boundary conditions has a stiffening effect on the shaft 26, thereby increasing its natural frequency. The change in boundary conditions is reflected in K n , as shown in Table 2. Specifically, the beam is now pinned at only one end (simply supported) and fixed at the other end, so

[0368] K n =π 2 (n+1 / 4) 2

[0369] For the example of n=2, K n =((2.25)2 π 2 =50.0 (accurate to three significant figures).

[0370] Table 2

[0371]

[0372] Assume that except K n With all parameters except φ being constant, it can be seen that a simple change in boundary conditions can increase the first eigenfrequency by more than 50%.

[0373] Thus, the relative positioning of the spindle bearings 26a-c can be used to tune the stiffening effect to control the resonant frequency.

[0374] The ratio of the subspan to the turbine length can be defined as:

[0375]

[0376] In various embodiments, the ratio of the subspan to the turbine length may be equal to or less than 1.05, optionally within the range of 0.85 to 1.05, and further optionally within the range of 0.85 to 0.95.

[0377] In such embodiments, the length ratio and / or spindle operating speed may or may not be within the ranges described in detail elsewhere herein. Similarly, the two rear bearings 26b, 26c may or may not be located behind the respective blades of the turbine 19, as described in detail elsewhere herein. The skilled artisan will appreciate that a variety of different methods for controlling or changing vibration modes are disclosed herein, and these methods may be used alone or in any suitable combination.

[0378] In various additional or alternative embodiments, the engine 10 may have a core shaft 26 having a length (L) in the range of 1800 mm to 2900 mm or 2750 mm between the frontmost bearing 26a and the rearmost bearing 26c, and an axial spacing (secondary span S) in the range of 250 mm to 350 mm between the two rear bearings 26b, 26c, so that the core shaft 26 has only a natural frequency of two nodes (one node at each end of the span between the first bearing 26a and the second bearing 26b, as shown in Tables 1 and 2 above, the mode 1 frequency or the first resonance) that is not within the operating range of the engine 10.

[0379] In such embodiments:

[0380] The spindle 26 may have an operating speed range with a lower limit of 1500 rpm and an upper limit of 6200 rpm; and / or

[0381] • The fan 23 may have a fan diameter in the range of 330 cm to 380 cm (130-150), and the gearbox 30 may have a gear ratio between 3.1 and 3.8.

[0382] Therefore, the spindle length L and the distance S between the two rear bearings 26b, 26c and the operating speed can be selected so that the spindle 26 does not have a first resonance mode between the first bearing 26a and the second bearing 26b within the engine operating range. The fan diameter of the fan 23 of the engine 10 can be selected to be suitable for the expected operating speed, so the fan diameter can be selected so that the appropriate operating speed of the fan does not cause the first resonance mode of the spindle 26 between the first bearing 26a and the second bearing 26b within the engine operating range.

[0383] The skilled person will appreciate that the mode 1 resonance (first resonance) has only two nodes, whereas the mode 2 frequency has an additional central node and a smaller maximum amplitude A.

[0384] When designing 3000 the engine 10 prior to manufacturing, the designer has greater scope to adjust engine parameters. The design method 3000 may include selecting 3002 a position for the front bearing 26a and the frontmost bearing 26b of the rear bearings 26b, 26c, such as selecting the positions of the front bearing and the frontmost bearing based on the spacing between the front bearing and the frontmost bearing (e.g., in the range of 1450 mm to 2500 mm) and the positions of other engine components. The method 3000 may also include increasing (or decreasing) 3004 the length of the spindle 26 behind the frontmost bearing 26b of the rear bearings 26b, 26c to provide a suitable shaft length for the secondary span (i.e., the distance between the two rear bearings 26b, 26c) so that the secondary span is within a suitable range (e.g., in the range of 250 mm to 350 mm) so that the first resonance of the spindle 26 is not within the operating speed range of the spindle 26. The last of the two rear bearings (in the depicted embodiment, the third bearing) 26 c may be located at the rear end of the spindle 26 or adjacent the end of the spindle 26 .

[0385] In various such embodiments, the spindle 26 / engine 10 may have a cruising operating speed range of 5400 rpm to 5700 rpm, and optionally 5500 rpm to 5600 rpm.

[0386] In various such embodiments, the spindle 26 / engine 10 may have an MTO operating speed range of 5800 rpm to 6200 rpm, and optionally 5900 rpm to 6100 rpm.

[0387] In various such embodiments, the mandrel 26 may have a length L of 1800 rpm to 2900 mm, optionally 2300 rpm to 2800 mm, and also optionally 2400 rpm to 2750 mm.

[0388] In various embodiments, the bearing stiffness at the forwardmost bearing 26b of the two rear bearings 26b, 26c is controlled. Control of the stiffness may allow or facilitate management of vibration modes.

[0389] The front one of the two rear bearings 26b (also referred to as the second bearing) has a bearing stiffness in the range of 30 kN / mm to 100 kN / mm, and optionally about 50 kN / mm, by means of a spring bar (as described below).

[0390] The stiffness of the second bearing 26b is determined in part by engine conditions. The lower the excitation at this bearing 26b, the lower the stiffness provided, but the higher the excitation, the higher the stiffness and the higher the damping. The bearing stiffness is a variable parameter that depends on engine conditions and may vary within the ranges listed above, for example during normal operation of the aircraft.

[0391] The forward rear bearing 26b comprises an outer ring 51 surrounding the spindle 26; the ring 51 may be referred to as a race and may contain a ball bearing 52 or the like, and oil arranged to lubricate the race in use. An inner ring 53 (or race) radially inward of the outer ring 51 may be used to retain the ball bearing 52 in a channel formed between the races 51, 53.

[0392] The outer race 51 of the rear bearing 26b is mounted on the fixed support structure 24 of the engine by one or more bearing support structures 50, 55. In the embodiment described, there are two bearing support structures 50, 55. The bearing support structures 50, 55 can together form the bearing disc 29a, as shown by comparing Fig.12 and Fig.14 Schematically shown. The bearing support structures 50, 55 of the described embodiment are each connected to the same aft bearing housing 29. In the described embodiment, the aft bearing housing 29 forms part of the fixed support structure of the engine 10. The disk 29a (including the component support structures 50 and 55) extends from the aft bearing housing 29 to the bearing 26b. In the described embodiment, the first bearing support structure 50 is mounted on the fixed support structure 24 of the engine 10 at a first position 58a. In the described embodiment, the first position 24a is axially rearward of the bearing 26b. In the described embodiment, the first bearing support structure 50 provides an outer race 51; specifically, the outer race 51 is formed integrally with the first bearing support structure 50, as shown Fig.13In an alternative embodiment, the outer race 51 may be formed separately from the first bearing support structure 50 and mounted on the first bearing support structure.

[0393] In the depicted embodiment, the first bearing support structure 50 includes a plurality of connection members 57 spaced circumferentially about the engine axis 9 to connect the outer race 51 to the fixed support structure 58. The connection members 57 extend axially along a portion of the engine 10.

[0394] For example, in some embodiments, there may be twenty or forty evenly spaced connection members 57, in other embodiments, the number and / or spacing may vary. The one or more connection members may extend between the outer races of the bearings 26b, 26c and the bearing housings 29, 31. The bearing housings 29, 31 may be rigidly connected to the fixed support structure 58, thereby effectively becoming part of the support structure 58.

[0395] The or each connecting member 57 may comprise a metal rod or stay and may be referred to as a spring bar. The or each connecting member 57 may be arranged to provide some flexibility to the bearings 26b, 26c, thereby allowing some radial and / or axial movement (e.g. due to elongation) to be accommodated in use. Therefore, the flexibility of the first bearing support structure 50 may be referred to as the stiffness of the spring bar.

[0396] In the depicted embodiment, the second bearing support structure 55 is mounted at a second location 58b on the fixed support structure 58 of the engine 10. In the depicted embodiment, the first location 58a and the second location 58b on the fixed support structure 58 are each located on the aft bearing housing 29 portion of the fixed structure 58.

[0397] In the depicted embodiment, the second location 58b is axially rearward of the (forward-most rear) bearing 26b, but axially forward of the first location 58a on the fixed support structure 58, and radially outward from both the bearing 26b and the first location 58a. In this embodiment, the second bearing support structure 55 is connected between the radially outer surface of the first bearing support structure 50 and the fixed support structure 58, as shown in FIG. Fig.14 As shown, in other embodiments, the connection may be different. In the described embodiment, the second bearing support structure 55 (compared to the first bearing support structure 50) has a relatively high stiffness and therefore contributes negligibly to the flexibility of the bearing 26b. Its stiffness may be considered as an effective stiffness.

[0398] The squeeze film damper 56 is disposed between the first bearing support structure 50 and the second bearing support structure 55 and is located in the region of the outer race 51. In the depicted embodiment, a channel 56a defined by a raised lip is disposed around the periphery of the first bearing support structure 50 to accommodate the squeeze film damper 56 and to locate an O-ring seal (not shown; which is axially spaced apart to be one at each end of the channel 56a). In alternative embodiments, such a channel and lip may not be provided, and / or the channel and lip may instead be provided on the second bearing support structure 55 or as described above, and the squeeze film damper 56 may be (in some cases, completely) contained by an O-ring seal or the like. The film is in addition to the hydrodynamic oil layer between the bearing 52 and the spindle 26 (i.e., the ball bearing 52 and the like are typically lubricated by the oil layer, and the squeeze film damper 56 is a separate layer from the ball bearing 52).

[0399] The squeeze film damper 56 includes a film layer, typically an oil layer or another lubricant layer, located between the bearing 26b and the housing 30. The squeeze film damper 56 is arranged to soften the bearing support to increase the damping effect. The stiffness of the squeeze film damper 56 generally depends on the temperature and the shaft rotation rate. The technician will understand that any centrifugal force in the rotation of the bearing 26b can be damped by the squeeze film damper 56. The squeeze film damper 56 can provide some structural isolation of the support structure 58 from the core shaft 26, can reduce the amplitude of the rotor in response to imbalance, and can help to suppress rotor dynamic instability. This damping may be particularly useful if the blade group has an uneven mass distribution (for example, due to damage or manufacturing defects in use), or if an unexpected event during manufacturing, maintenance or operation causes the blade group to be out of radial alignment, causing the blade group to be slightly tilted relative to the engine axis 9.

[0400] The inventors have realised that the overall bearing stiffness can be thought of as having three components which can be thought of as springs arranged, for example, in parallel or in a combination of series and parallel depending on the arrangement of the components, namely:

[0401] The stiffness of the oil layer in the bearing (fluid stiffness of the squeeze film damper);

[0402] The stiffness of the bearing support 50 (which is usually most affected by the stiffness of the connecting member 57, which may be referred to as the stiffness of the spring bar); and

[0403] • The stiffness of the fixed support structure 29, 58 on which the bearings are mounted.

[0404] The fixed support structure 58 has a much greater stiffness than the other two contributors to the bearing stiffness and is therefore considered to be effectively stiff. Any flexibility in the fixed support structure 58 may only become apparent under extreme conditions such as a blade-off event. The stiffness of the oil layer varies significantly depending on the shaft rotation rate and temperature, but at and above cruising speeds, the stiffness of the oil is much greater than the stiffness of the spring rods of the bearings. Therefore, the stiffness of each bearing is considered according to the stiffness of its spring rods.

[0405] The stiffness of the spring bar of the bearing 26b is defined as the radial stiffness, ie the linear deflection δ measured along the radius of the engine 10, caused by the force F. This is Fig.15A and Fig. 15B The diagonal hatching shows that the fixed structure 58 is considered rigid / non-moving.

[0406] Fig.15A The bearing support structure 50, 56, 55 is shown (with the spindle 26, inner race 53 and ball bearing 52 excluded for clarity) with a portion of the radius r of the engine 10 marked with a dashed line. The marked radial line is located at the axial center point of the bearing race 51. A force F is shown applied in a radially outward direction (i.e., away from the axial center line) along the radius r. Fig. 15B The initial position of the first bearing structure 50 (the position before the force F is applied) is shown in dashed lines, and the final position of the first bearing structure 50 (the position during the application of the force F) is shown in solid lines. The skilled person will understand that the deflection shown is much larger than the deflection expected in normal operation, and the deflection shown is provided only for ease of understanding. In addition, during normal operation, after the force F is removed, the first bearing support structure 50 should return to its initial position. The displacement or deflection δ is then measured along the radius r at the axial center point of the bearing race 51. The two black dots show the position of the inner surface of the first support structure 50 before and during the application of the force F. The displacement δ is the distance between these points. The inner surface of the first bearing support structure 50 is selected only for ease of illustration, and the skilled person will understand that another point may be selected alternatively, such as the radial outer surface or radial center point of the first bearing support structure 50. The displacement reflects a combination of the compression of the squeeze film damper 56, the bending of the first bearing support structure 50 (particularly the spring rod 57), and any bending of the second bearing support structure 55. Therefore, bearing stiffness is a measure of the radial displacement caused by applying a radial force at the axial center point of bearing 26b.

[0407] In the described embodiment, a force F of 50 kN causes a deflection δ of 1 mm when applied to the forward most bearing 26b of the two rear bearings 26b, 26c, so the stiffness of that bearing 26b as defined herein is 50 kN / mm. The bearing stiffness may vary between the two rear bearings 26b, 26c and between embodiments.

[0408] Thus, in the depicted embodiment, the forwardmost of the rear bearings 26b has a bearing stiffness in the range of 30 kN / mm to 100 kN / mm.

[0409] The stiffness ratio at the forward-most rear bearing 26b (ie, the second bearing along the spindle in the described embodiment) may be defined as:

[0410]

[0411] In various embodiments, the stiffness ratio may be between 0.08 and 0.5 kN / mm 2 and optionally in the range of 0.09 to 0.40 kN / mm 2 Optionally, the stiffness ratio may be at least substantially equal to 0.25, 0.30 or 0.35 kN / mm 2 . Fig.10 A method 1000 is shown that may be performed and includes starting 1002 the engines 10 of the aircraft and reaching a cruise condition, and operating 1004 the aircraft in the cruise condition.

[0412] The engine 10 may have a cruising operating speed range of 5400 rpm to 5700 rpm and optionally 5500 rpm to 5600 rpm.

[0413] The engine 10 may be operated such that the first resonant mode does not exist between the first bearing 26a and the second bearing 26b under cruising conditions.

[0414] The engine 10 may be operated such that the first resonant mode between the first bearing 26a and the second bearing 26b is not anywhere within the engine operating range (including both MTO and cruise).

[0415] Unless otherwise stated, the lengths defined herein are for the corresponding parts when the engine is off (i.e., at zero speed / on the bench, at room temperature). These values ​​generally do not vary significantly within the operating range of the engine (e.g., the shaft length has only a few millimeters or less of elongation at operating temperature); therefore, the stiffness at cruise conditions of an aircraft equipped with the engine (these cruise conditions are defined elsewhere herein) can be the same as the stiffness when the engine is not in use. However, in the case where the length varies within the operating range of the engine, the values ​​defined herein should be understood to be the length when the engine is at room temperature and not moving.

[0416] In contrast, since the oil layer stiffness is speed dependent and contributes to the bearing stiffness, the bearing stiffness is defined under cruising conditions / with the shaft rotating at a suitable cruising speed.

[0417] Fig.17 It is shown how the bearing stiffness defined herein can be measured. Fig.17 A graph showing the displacement δ due to an applied load L (e.g., force, moment, or torque) applied to any component whose stiffness is to be measured. P At load levels above L, there is a nonlinear region where displacement is caused by movement of the component when loaded (or relative movement of individual parts of the component) rather than deformation of the component (e.g., movement within the gap between parts). For bearing 26b, the amount of displacement possible in this nonlinear region may be very small. Q At load levels of , the elastic limit of the component has been exceeded and the applied load no longer causes elastic deformation, but rather plastic deformation or component failure may occur. Between points P and Q, the applied load and the resulting displacement have a linear relationship. The stiffness defined herein can be determined by measuring the gradient of the linear region between points P and Q (where the stiffness is the inverse of the gradient). The gradient of the linear region can be found as large as possible by providing a larger displacement for measurement to increase the accuracy of the measurement. For example, by applying a displacement equal to or slightly greater than L P and equal to or slightly less than L Q The gradient can be found by applying a load. L can be estimated based on material properties before testing. P and L Q value in order to apply a suitable load.

[0418] It should be understood that the present invention is not limited to the above-described embodiments, and various modifications and improvements may be made without departing from the concepts described herein. Unless mutually exclusive, any feature may be used alone or in combination with any other feature, and the present disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A gas turbine engine (10) for an aircraft, comprising: An engine core (11), the engine core comprising a turbine (19), a compressor (14) and a spindle (26) connecting the turbine (19) to the compressor, wherein the compressor (14) is a lowest pressure compressor of the engine (10), and the turbine (19) is a lowest pressure turbine of the engine (10) and has lowest pressure set blades (19c, 19d), each blade of the lowest pressure set blades (19c, 19d) having a mass m, a radius r at mid-height of the blade, and a cruising angular velocity ω; a fan (23), the fan being located upstream of the engine core (11), the fan comprising a plurality of fan blades; and a gearbox (30) receiving input from the spindle (26) and outputting drive to the fan (23) so as to drive the fan at a lower rotational speed than the spindle (26), And wherein the engine core (11) further comprises three bearings (26a-c) arranged to support the spindle (26), the three bearings comprising a front bearing (26a) and two rear bearings (26b, 26c), wherein a secondary span S is defined as an axial distance between the two rear bearings (26b, 26c), and wherein a ratio of the first blade to the bearing is: With 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 value in the range of .

2. The gas turbine engine (10) according to claim 1, wherein the ratio of the first blade to the bearing is between 2.0×10 -6 to 6.5×10 -6 kg -1 .rad -2 .s 2 within the range.

3. A gas turbine engine (10) according to claim 1 or claim 2, wherein the ratio of the second blade to the bearing is: With 0.8 to 6.0kg -1 value in the range of .

4. The gas turbine engine (10) according to claim 3, wherein the ratio of the second blade to the bearing is between 0.8 and 5.0 kg -1 within the range.

5. The gas turbine engine (10) according to claim 1 or claim 2, wherein the secondary span S is in the range of 250 mm to 350 mm.

6. A gas turbine engine (10) according to claim 1 or claim 2, wherein m multiplied by r is in the range of 180 to 280 kg.mm.

7. The gas turbine engine (10) of claim 1 or claim 2, wherein the gearbox (30) has a gear ratio greater than 3.

8. The gas turbine engine (10) of claim 1 or claim 2, wherein the spindle (26) has an operating speed in the range of 1500 rpm to 6200 rpm.

9. The gas turbine engine (10) according to claim 1 or claim 2, wherein the length ratio of the subspan between the two rear bearings (26b, 26c) to the spindle length is equal to or less than 0.

14.

10. The gas turbine engine (10) according to claim 1 or claim 2, wherein: (i) the mass m of the blades in the lowest pressure group of blades (19c, 19d) is in the range of 0.2 to 0.6 kg; and / or (ii) The radius r of the blades in the lowest pressure group of blades (19c, 19d) at mid-height is in the range of 400 to 600 mm.

11. A gas turbine engine (10) according to claim 1 or claim 2, wherein: (i) the spindle (26) has a length L between the front bearing (26a) and the last rear bearing (26c) in the range of 1800 to 2900 mm; and / or (ii) the turbine (19) has a turbine length defined between the leading edge of its most upstream blade and the trailing edge of its most downstream blade, and wherein the ratio of the secondary span to the turbine length is: Equal to or less than 1.

05.

12. A gas turbine engine (10) according to claim 1 or claim 2, wherein the rear bearing (26b, 26c) is located axially flush with or behind: (i) the leading edge of the lowest pressure turbine blade (19c, 19d) of the turbine (19) at the root of the blade; and / or (ii) the trailing edge of a turbine blade of a third set of turbine blades (19c) from the front of the turbine (19) at the root of the blade, wherein the turbine (19) comprises four sets of turbine blades (19a-19d).

13. The gas turbine engine (10) according to claim 1 or claim 2, wherein: (i) the forwardmost of the rear bearings (26b) has a bearing stiffness in the range of 30 kN / mm to 100 kN / mm; and / or (ii) the ratio of the bearing stiffness at the frontmost rear bearing (26b) to the stiffness of the distance between the two rear bearings (26b, 26c) is between 0.08 and 0.5 kN / mm 2 within the range of Therein, the bearing stiffness is defined by the radial displacement caused by applying a radial force at the axial center point of the bearing.

14. A gas turbine engine (10) according to claim 1 or claim 2, further comprising a tail bearing housing (29), which is located rearward of the turbine (19) and comprises two bearing discs (29a, 29b), each bearing disc being arranged to support one of the two rear bearings (26b, 26c).

15. The gas turbine engine (10) of claim 2, wherein the first blade to bearing ratio is 3.0×10 -6 to 4.5×10 -6 kg -1 .rad -2 .s 2 within the range.

16. The gas turbine engine (10) of claim 4, wherein the ratio of the second blade to the bearing is between 0.8 and 4.0 kg -1 within the range.

17. The gas turbine engine (10) of claim 16, wherein the second blade to bearing ratio is between 0.8 and 3.0 kg -1 within the range.

18. The gas turbine engine (10) of claim 17, wherein the ratio of the second blade to the bearing is between 0.8 and 2.0 kg -1 within the range.

19. The gas turbine engine (10) of claim 7, wherein the gearbox (30) has a gear ratio in the range of 3.1 to 3.

8.

20. The gas turbine engine (10) of claim 7, wherein the fan (23) has a fan diameter in the range of 330 cm to 380 cm.

21. The gas turbine engine (10) of claim 8, wherein the operating speed of the spindle (26) is one or more of: (i) at cruising speed, within the range of 5400 rpm to 5700 rpm; and / or (ii) Under maximum takeoff conditions, within the range of 5800 rpm to 6200 rpm.

22. The gas turbine engine (10) of claim 21, wherein the operating speed of the spindle (26) is one or more of: (i) at cruising speed, within the range of 5500 rpm to 5600 rpm; and / or (ii) Under maximum takeoff conditions, within the range of 5900 rpm to 6100 rpm.

23. The gas turbine engine (10) of claim 9, wherein a length ratio of the subspan between the two rear bearings (26b, 26c) to the spindle length is in the range of 0.08 to 0.

13.

24. The gas turbine engine (10) according to claim 11, wherein the mandrel (26) has a length L between the front bearing (26a) and the last rear bearing (26c) in the range of 2300 to 2800 mm.

25. A gas turbine engine (10) according to claim 24, wherein the spindle (26) has a length L between the front bearing (26a) and the last rear bearing (26c) in the range of 2400 to 2750 mm.

26. The gas turbine engine (10) of claim 11, wherein a ratio of subspan to turbine length is in the range of 0.85 to 1.

05.

27. The gas turbine engine (10) of claim 26, wherein a ratio of subspan to turbine length is in the range of 0.85 to 0.

95.

28. The gas turbine engine (10) of claim 12, wherein the turbine (19) comprises a total of four sets of turbine blades (19a-19d).

29. The gas turbine engine (10) of claim 14, wherein the bearing disk is oriented at least substantially perpendicular to the engine axis (9).

30. A method (1000) of operating a gas turbine engine (10) for an aircraft, the gas turbine engine comprising: An engine core (11), the engine core comprising a turbine (19), a compressor (14) and a spindle (26) connecting the turbine to the compressor, and wherein the compressor (14) is a lowest pressure compressor of the engine (10), and the turbine (19) is a lowest pressure turbine of the engine (10) and has lowest pressure set blades (19c, 19d), each blade of the lowest pressure set blades (19c, 19d) having a mass m, a radius r at the mid-height of the blade and a cruising angular velocity ω, and wherein the engine core (11) further comprises three bearings (26a-c) arranged to support the spindle (26), the three bearings comprising a front bearing (26a) and two rear bearings (26b, 26c), wherein a subspan (S) is defined as the axial distance between the two rear bearings (26b, 26c); a fan (23), the fan being located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox (30) arranged to receive input from the spindle (26) and output drive to the fan (23) so as to drive the fan at a lower rotational speed than the spindle, The method comprises: The engine (10) is operated (1004) such that a ratio of the first blade to the bearing is: When cruising, it has a 2.0×10 -6 to 7.5×10 -6 kg -1 .rad -2 .s 2 value in the range of .

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