Optimization of fan behavior in an aeronautical propulsion system

Abstract

An aeronautical propulsion system (1) comprising: - a movable element (35) rotating about an axis (X) of the assembly (11a, 11b, 11c); and - a mode damper (25) having an axial length and a maximum radial clearance such that a surface viscosity of the mode damper (25), which corresponds to the ratio of the square of the axial length to the product of the cube of the maximum radial clearance multiplied by twice the value of Pi, is between 1.0 x 10⁴ m⁻¹ and 1.15 x 10⁶ m⁻¹: where: L is the axial length of the mode damper (25) and is expressed in meters (m); and C is the maximum radial clearance of the mode damper (25) and is expressed in meters (m). Figure 1

Description

Title of the invention: Optimization of the behavior of the fan in an aeronautical propulsion system. Technical field

[0001] The present application relates generally to the field of propulsion systems, and more particularly to mode dampers comprising a pressurized fluid (“squeeze film” in English) which can be mounted between the outer ring of a bearing or a reduction mechanism with epicyclic or planetary gear train and a rigid support fixedly attached to a stator part of the propulsion system. STATE OF THE ART

[0002] A propulsion system generally comprises, from upstream to downstream in the direction of gas flow, a fan section, a compressor section which may include a low-pressure compressor and a high-pressure compressor, a combustion chamber, and a turbine section which may include, in particular, a high-pressure turbine and a low-pressure turbine. The high-pressure compressor is driven in rotation by the high-pressure turbine via a high-pressure shaft. The fan and, where applicable, the low-pressure compressor are driven in rotation by the low-pressure turbine via a low-pressure shaft.

[0003] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into account the impactful factors in all phases of design and development in order to obtain aeronautical components and products that are less energy-intensive, more environmentally friendly, and whose integration and use in civil aviation have moderate environmental consequences, with the aim of improving the energy efficiency of aircraft.

[0004] In order to minimize mechanical loads, vibrations, and displacements between a moving part, such as a rotor, and a stator part of the propulsion system, it has been proposed to interpose a mode damper (or "squeeze-film") and, where appropriate, a flexible cage (also known as a squirrel cage) between the moving part and the stator part. For this purpose, a mode damper comprises an inner ring and an outer ring, which are fixed relative to the stator part by extending radially outward from the moving part, and a pressurized fluid that is confined between the inner and outer rings.

[0005] The performance and noise requirements of current propulsion systems are increasingly high. To improve system performance In propulsion systems, it is possible, for example, to reduce the dynamic play in the engine. To achieve this, it might be desirable to increase the radius of the mode dampers (the damping capacity of the mode dampers depends on their radius). However, the available space between the rotors and their respective drive shafts is increasingly reduced to improve the efficiency of the propulsion system and reduce its specific fuel consumption.

[0006] In the more specific case of short- and medium-range engines, the current trend is to increase the bypass ratio in order to reduce specific fuel consumption and improve the efficiency of the propulsion system. To achieve this, the fan is connected to the shaft that drives it using a reduction mechanism, which allows for independent optimization of their respective rotational speeds. The consequence is that the drive shaft rotates at higher speeds, which increases the loads transmitted from the rotor to the stators. However, mechanical integration constraints lead to large-diameter bearings, particularly in the vicinity of the reduction mechanism. Consequently, the dampers mounted above these bearings also end up with an unusually large diameter.However, the properties of mode dampers at these diameters are poorly understood, necessitating costly test campaigns to measure them and calibrate predictive models. Indeed, when the inertial and damping forces generated by the fluid in the mode damper are too high and the bearing rotational speed is high, the mode damper behaves like a rigid bearing, thus negating the beneficial effect of its damping. This is referred to as "locking" of the mode damper.

[0007] The increased rotational speed of the low-pressure shaft, made possible by the reduction mechanism, also leads to the appearance of a greater number of overall vibration modes within the operating range, unlike a conventional direct-drive architecture. However, the mode damper is generally optimized for only one particular overall vibration mode. It can therefore have detrimental effects on the other modes, leading to an increase in dynamic loads, dynamic backlash consumption, and vibrations in the other modes, instead of a reduction (due to a "lock-in" of the mode damper).

[0008] Increasing the rotational speeds of the low-pressure shaft can also lead to the presence of a bending mode of the low-pressure shaft within the operating range (supercritical low-pressure shaft), which can lead to excessive deformations of the low-pressure shaft, or even contact between the low-pressure shaft and the high-pressure shaft when passing through the low-pressure shaft mode. pressure and / or non-synchronous vibrations after passing through this mode, which can damage the propulsion system.

[0009] It is therefore necessary to optimize the mode damper so as to obtain a sufficient damping effort to fall below the HCF (High Cycle Fatigue) loading objectives (for vibration fatigue, for which structures are dimensioned for an infinite lifespan, typically corresponding to 107 cycles) and dynamic clearance consumption, and below the vibration and noise cabinet thresholds imposed by aircraft manufacturers, without risking blocking the mode damper. EXPOSED

[0010] One aim of the present application is to optimize the performance of the aeronautical propulsion system, in particular of propulsion systems with high bypass ratios, in particular in terms of vibration damping, dynamic clearance consumption and polycyclic loading, taking into account the constraints of integration in the propulsion system.

[0011] To this end, according to a first aspect, an aeronautical propulsion system is proposed comprising: - a movable element rotating around an axis of the assembly; and - a mode damper comprising an inner ring, an outer ring and a fluid confined between the inner ring and the outer ring, the inner ring and the outer ring being coaxial with the moving element; the inner ring extending radially between the moving element and the outer ring; and the mode damper having an axial length and a maximum radial clearance such that a surface viscosity of the mode damper, which corresponds to the ratio between the square of the axial length and the product between the cube of the maximum radial clearance multiplied by twice the value of Pi, is between 1.0 x 10⁴ m¹ and 1.15 x 10⁶ m¹: 1.0 x 10⁴ < surface viscosity = -3 < 1.15 x 10⁶ m¹ L is the axial length of the mode damper and is expressed in meters; and C is the maximum radial clearance of the mode damper and is expressed in meters.

[0012] Some preferred but non-limiting features of the assembly according to the first aspect are the following, taken individually or in combination: - the surface viscosity of the mode damper is less than or equal to 7.1 x 10⁵ m². - the mode damper also has a surface clearance, which corresponds to the ratio between the maximum radial clearance and the product between twice the value of Pi multiplied by the radius of implantation of the mode damper and the axial length of the mode damper, is between 2.5 x 102 m1 and 1.0 x 10 1 m 1: 2.5 x 10'2 < surface clearance = 1.0 x 10 loû: R is the radius of implantation of The fashion damper is expressed in meters - the surface clearance of the mode damper is greater than or equal to 3.7 x 102 m1; - the moving element includes a moving part of a bearing of the propulsion system; - the moving element includes a moving part of a propulsion system reduction mechanism; and / or - the mode damper further includes at least one sealing segment, preferably two sealing segments, axially delimiting the fluid between the inner ring and the outer ring.

[0013] According to a second aspect, an aeronautical propulsion system is proposed comprising an assembly according to the first aspect and a rotor coaxial with the moving element.

[0014] Some preferred but not limiting features of the system according to the first aspect are the following, taken individually or in combination: - the rotor includes a shaft of the propulsion system and the moving element includes a bearing which supports the rotor; - the rotor includes a drive shaft, the propulsion system further comprising a blower section including a blower shaft (20); a drive turbine driven in rotation by the rotor around the axis; and a reduction mechanism coupling the rotor and the blower shaft in order to drive the blower shaft at a rotational speed lower than the rotational speed of the rotor, the moving element comprising a moving part of the reduction mechanism; - the drive shaft is supported by at least one front bearing and one or two rear bearings, the moving element corresponding to the front bearing and, where applicable, to one of the rear bearings; and / or - a dilution ratio of the propulsion system is greater than or equal to 10, for example between 10 and 80 inclusive.

[0015] According to a third aspect, an aircraft is proposed comprising at least one aeronautical propulsion system according to the second aspect fixed to the aircraft by means of a mast.

[0016] According to a fourth aspect, a method for manufacturing an assembly of an aeronautical propulsion system is proposed, comprising the following steps: - determine an axial length and a maximum radial clearance of a mode damper such that a surface viscosity of the mode damper, which corresponds to the ratio between the square of the axial length and the product between the maximum radial clearance cubed multiplied by twice the value of Pi, is between 1.0 x 104 m 1 and 1.15 x 106 m1: 104 < surface viscosity = < 1.15 x 106 or ' L is the axial length of the mode damper and is expressed in meters; and C is the maximum radial clearance of the mode damper and is expressed in meters; - manufacture an inner ring and an outer ring of the mode damper according to the axial length and maximum radial play thus determined; - assemble the inner ring and the outer ring with a moving element; and - inject a fluid between the inner ring and the outer ring.

[0017] Some preferred but non-limiting features of the manufacturing process according to the fourth aspect are the following, taken individually or in combination: - the axial length and maximum radial play are determined so that the surface viscosity is greater than or equal to 3.7 x 102 m 1; - the axial length, the maximum radial clearance and an installation radius can be determined so that a surface clearance of the mode damper is between 2.5 x 10A-2 m 1 and 1.0 x 10A-1 m1, where the surface clearance is defined as follows: 2.5 x 10² < surface clearance = 1.0 x 10²: R is a radius of implantation of The mode damper is expressed in meters; and / or - the axial length, the maximum radial clearance and the implantation radius are determined so that the surface clearance is greater than or equal to 3.7 x 102 m'. DESCRIPTION OF THE FIGURES

[0018] Other features, purposes and advantages will become apparent from the following description, which is purely illustrative and not limiting, and which should be read in conjunction with the accompanying drawings on which:

[0019] Fig. 1 is a schematic, partial, cross-sectional view of an example of a propulsion system according to a first embodiment, in which the blower section is faired;

[0020] Fig. 2 is a schematic, partial, cross-sectional view of an example of a propulsion system according to a second embodiment, in which the blower section is unfaired;

[0021] Fig. 3 is an enlarged schematic cross-sectional view of part of an example assembly for a propulsion system according to a first variant;

[0022] The [Fig.4] is an enlarged schematic cross-sectional view of part of an example assembly for a propulsion system according to a second variant;

[0023] The [Fig.5] is an example of an aircraft that may include at least one propulsion system conforming to the first or second embodiment;

[0024] Figure 6 is a flowchart illustrating examples of steps in a manufacturing process for an assembly for a propulsion system; and

[0025] Fig. 7 is a graph representing the radial displacement of a moving element (bearing) of a bearing (in millimeters) as a function of the rotational speed of the shaft on which it is mounted (in rpm), in the presence of an imbalance present on the shaft.

[0026] Throughout the figures, similar elements bear identical references. DETAILED DESCRIPTION

[0027] A propulsion system 1 has a principal direction extending along a longitudinal axis X and comprises, from upstream to downstream in the direction of the gas flow in the propulsion system 1 when in operation, a blower section 2 and a primary body 3, often called a "gas generator", comprising a compressor section 4, 5, a combustion chamber 6 and a turbine section 7, 8. The propulsion system 1 here is an aeronautical propulsion system 1 configured to be fixed on an aircraft 100 by means of a pylon (or mast).

[0028] The compressor section 4, 5 comprises a series of stages, each including a rotating blade wheel (rotor) 4a, 5a in front of a fixed blade wheel (stator) 4b, 5b. The turbine section 7, 8 also comprises a series of stages, each including a fixed blade wheel (stator) 7b, 8b behind which a rotating blade wheel (rotor) 7a, 8a rotates.

[0029] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, corresponding to the rotation of the shafts of the gas generator, and a radial direction is a direction perpendicular to and passing through this axis X. Furthermore, the circumferential (or lateral, or tangential) direction corresponds to a direction perpendicular to and not passing through the longitudinal axis X. Unless otherwise specified, internal (respectively, inside) and external (respectively, outside) are used with reference to a radial direction such that the inner part or face of an element is closer to the axis X than the outer part or face of the same element.

[0030] In operation, an airflow F entering the propulsion system 1 is divided between a primary airflow Fl and a secondary airflow F2, which flow from upstream to downstream in the propulsion system 1.

[0031] The secondary airflow F2 (also called "bypass airflow") flows around the primary body 3. The secondary airflow F2 cools the periphery of the primary body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0032] The primary airflow Fl flows in a primary vein inside the primary body 3, passing successively through the compressor section 4, 5, the combustion chamber 6 where it is mixed with fuel to serve as an oxidizer, and the turbine section 7, 8. The passage of the primary airflow Fl through the turbine section 7, 8 receiving energy from the combustion chamber 6 causes a rotation of the rotor of the turbine section 7, 8, which in turn drives the rotation of the rotor of the compressor section 4, 5 as well as a rotor part 9 of the blower section 2.

[0033] In a twin-body propulsion system 1, the compressor section 4, 5 may include a low-pressure compressor 4 and a high-pressure compressor 5. The turbine section 7, 8 may include a high-pressure turbine 7 and a low-pressure turbine 8. The rotor of the high-pressure compressor 5 is driven in rotation by the rotor of the high-pressure turbine 7 via a high-pressure shaft 10. The rotor of the low-pressure compressor 4 and the rotor portion 9 of the blower section 2 are driven in rotation by the rotor of the low-pressure turbine 8 via a low-pressure shaft 11. Thus, the primary body 3 comprises a high-pressure body including the high-pressure compressor 5, the high-pressure turbine 7 and the high-pressure shaft 10, and a low-pressure body including the blower section 2, the low-pressure compressor 4, the low-pressure turbine 8 and the low-pressure shaft 11.The rotational speed of the high-pressure body is greater than the rotational speed of the low-pressure body. In a three-body propulsion system 1, the turbine section 7, 8 further includes an intermediate turbine, positioned between the high-pressure turbine 7 and the low-pressure turbine 8 and configured to drive the rotor of the low-pressure compressor 4 via an intermediate shaft. The blower rotor 9 and the rotor of the high-pressure compressor 5 remain driven by the low-pressure shaft 11 and the high-pressure shaft 10, respectively.

[0034] The low-pressure shaft 11 is generally housed, along a portion of its length, within the high-pressure shaft 10 and is coaxial with the high-pressure shaft 10. The low-pressure shaft 11 and the high-pressure shaft 10 may be co-rotating, that is, driven in the same direction around the longitudinal axis X. Alternatively, the low-pressure shaft 11 and the high-pressure shaft may be counter-rotating, that is, driven in opposite directions around the longitudinal axis X. If applicable, the intermediate shaft is housed between the high-pressure shaft 10 and the low-pressure shaft 11. The intermediate shaft and the low-pressure shaft 11 may be co-rotating or counter-rotating.

[0035] The blower section 2 comprises at least the blower rotor 9 adapted to be driven in rotation relative to a stator portion of the propulsion system 1 by the turbine section 7, 8. Each blower rotor 9 comprises a hub 13 and blades 14 extending radially from the hub 13. The blades 14 of each rotor 9 may be fixed relative to the hub 13 or have a variable pitch. In this case, the base of the blades 14 of each rotor 9 is mounted pivoting about a pitch axis and is connected to a pitch-changing mechanism 15 mounted in the propulsion system 1, the pitch being adjusted according to the flight phases by the pitch-changing mechanism 15. The pitch-changing mechanism 15 is illustrated in dashed lines on [Fig. 1] to show that this feature is optional.

[0036] The fan section 2 may further include a fan stator 16, or rectifier, which comprises blades 17 mounted on a hub of the fan stator 16 and whose function is to rectify the secondary airflow F2 exiting the fan rotor 9. The blades 17 of the fan stator 18 may be fixed relative to the hub or have a variable pitch. Similar to the rotor blades 14, the base of the stator blades 17 is pivotally mounted about a pitch axis X and is connected to a pitch-changing mechanism 15a, which is generally separate from that of the fan rotor 9, the pitch being adjusted according to the flight phases by the pitch-changing mechanism.

[0037] The fan rotor 9 also comprises at least twelve blades 14 and at most twenty-four blades 14, for example at least fourteen blades 14 and at most twenty-two blades 14, for example sixteen blades 14. The number of blades 16 in the fan stator 17 depends on the acoustic criteria defined for the propulsion system 1 and is at least equal to the number of blades 14.

[0038] In order to improve the propulsive efficiency of propulsion system 1 and reduce its specific fuel consumption as well as the noise emitted by the fan section 2, propulsion system 1 has a high bypass ratio. A high bypass ratio is understood here to be a ratio greater than or equal to 10, for example, between 10 and 80 inclusive. To calculate the bypass ratio, the mass flow rate of the secondary airflow F2 and the mass flow rate of the primary airflow Fl are measured when propulsion system 1 is stationary, uninstalled, in takeoff mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) Manual, Doc 7488 / 3, 3rd edition) and at sea level. It should be noted that, in this application, the parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined under these conditions.By "not installed," we mean here that the measurements are taken when the propulsion system 1 is in a test bench (and not installed on an aircraft 100), as the measurements are then simpler to perform. The distances (length, radius, diameter) are, however, measured at ambient temperature (approximately 20°C) when the propulsion system 1 is cold, that is, when the propulsion system has been stopped for a sufficient period for the components of the propulsion system to reach ambient temperature.

[0039] The fan rotor 9 is decoupled from the low-pressure shaft 11 by means of a reduction mechanism 19, positioned between an upstream end of the low-pressure shaft 11 and the fan rotor 9, in order to independently optimize their respective rotational speeds. In this case, the propulsion system 1 further comprises an additional shaft, referred to as the fan shaft 20. The low-pressure shaft 11 connects the low-pressure turbine 7 to an inlet of the reduction mechanism 19, while the fan shaft 20 connects the outlet of the reduction mechanism 19 to the fan rotor 9. The fan rotor 9 is thus driven by the low-pressure shaft 11 via the reduction mechanism 19 and the fan shaft 20 at a rotational speed lower than the rotational speed of the low-pressure turbine 7.

[0040] This decoupling makes it possible to reduce the rotational speed and pressure ratio of the fan rotor 9 and to increase the power extracted by the low-pressure turbine 7. Indeed, the overall efficiency of propulsion systems is primarily determined by the propulsion efficiency, which is favorably influenced by minimizing the variation in the kinetic energy of the air as it passes through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow generating the propulsive force consists of the secondary airflow F2 of the propulsion system 1, the kinetic energy of the secondary airflow F2 being mainly affected by the compression that the secondary airflow F2 undergoes as it passes through the fan section 2. The propulsion efficiency and the pressure ratio of the fan section 2 are therefore linked: the lower the pressure ratio of the fan section 2, the better the propulsion efficiency.In order to optimize the propulsive efficiency of the propulsion system 1, the blower pressure ratio, which corresponds to the ratio between the average pressure at the outlet of the blower stator 17 (or, in the absence of a stator, of the blower rotor 9) and the average pressure at the inlet of the blower rotor 9, is less than or equal to 1.70, preferably less than or equal to 1.50, for example greater than or equal to between 1.05 and less than 1.45. The average pressures are measured here over the height of the blade 14 (from the surface which radially delimits the flow duct inside the inlet of the blower rotor 9 to the apex 21 of the blower blade 14).

[0041] The propulsion system 1 is configured to provide a thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,241 N), preferably between 20,000 Ibf (88,964 N) and 35,000 Ibf (15,568 N).

[0042] The blower section 2 may be shrouded or unshrouded. In the case of a shrouded blower section 2, the blower section 2 comprises a blower housing 12 and the blower rotor 9 is housed in the blower housing 12.

[0043] A shrouded blower section 2 comprises a blower rotor 9 extending upstream of a blower stator. The blower stator blades are then generally referred to as outlet guide vanes ("Outlet Guide Vane" or "OGV" in (English) and have a fixed pitch relative to the fan stator hub. Furthermore, the bypass ratio of the propulsion system 1 is, for example, greater than or equal to 10, for example, between 10 and 35 inclusive, for example, between 10 and 18 inclusive. The peripheral velocity at the tip 21 of the fan rotor blades 9 can also be between 260 m / s and 400 m / s. The fan rotor blades 14 can be fixed or have a variable pitch. The fan pressure ratio can then be between 1.20 and 1.45.

[0044] In an unfaired fan section 2, the fan section 2 (which may also be referred to as the propeller) is not enclosed by a fan casing. Since the fan section 2 is unfaired, the blades 14 of the fan rotor 9 have variable pitch. Propulsion systems comprising at least one unfaired fan rotor 9 are known as "open rotors" or "unducted fans." The propulsion system 1 may comprise two unfaired, counter-rotating fan rotors 9. Such a propulsion system 1 is known by the acronym CROR for "Contra-Rotating Open Rotor" or UDF for "Unducted Double Fan." The blower rotor(s) 9 can be placed at the rear of the primary body 3 so as to be of the pusher type or at the front of the primary body 3 so as to be of the tractor type.Alternatively, the propulsion system 1 may comprise a single unducted fan rotor 9 and an unducted fan stator 16 (rectifier). Such a propulsion system 1 is known by the English acronym USF for "Unducted Single Fan". In the case of a USF-type propulsion system 1, the blades 17 of the rectifier 16 are fixed in rotation relative to the X-axis of rotation of the upstream fan rotor 9 and therefore do not experience centrifugal force. The blades 17 of the rectifier 16 also have variable pitch.

[0045] Removing the fairing around the fan section 2 significantly increases the bypass ratio without the propulsion system 1 being negatively impacted by the mass of the housings or nacelles intended to surround the fan section 2. The bypass ratio of the propulsion system 1, including an unfaired fan section 2, is thus greater than or equal to 40, for example, between 40 and 80 inclusive. Furthermore, the peripheral velocity at the tip 21 of the fan blades 14 of the fan rotor(s) 9 can be between 210 m / s and 260 m / s. The fan pressure ratio can then be, for example, between 1.05 and 1.20.

[0046] The reduction mechanism 19 may comprise an epicyclic or planetary reduction mechanism, single-stage or two-stage. For example, the reduction mechanism 19 may be of the planetary type ('star' in English) and comprise a sun pinion (input of the reduction mechanism 19), centered on an axis X of rotation of the reduction mechanism 19 (generally coinciding with the longitudinal axis X) and configured to be driven in rotation by the low-pressure shaft 11, a ring gear (output of the reduction mechanism 19) coaxial with the solar pinion and configured to drive the blower shaft 20 in rotation around the X-axis of rotation, and a series of planet gears distributed circumferentially around the X-axis of rotation between the solar pinion and the ring gear, each planet gear being internally meshed with the solar pinion and externally with the ring gear. The series of planet gears is mounted on a planet carrier which is fixed relative to a stator portion of the propulsion system 1, for example, relative to a housing of the compressor section 4, 5. Alternatively, the reduction mechanism 19 can be epicyclic (or "planetary"), in which case the ring gear is fixedly mounted on the stator portion of the propulsion system 1 and the blower shaft 20 is driven in rotation by the planet carrier.

[0047] Regardless of the configuration of the reduction mechanism 19, the diameter of the ring gear and the satellite carrier are greater than the diameter of the solar pinion, so that the rotational speed of the blower rotor 9 is less than the rotational speed of the low pressure shaft 11.

[0048] In the case of a propulsion system 1 comprising a shrouded fan rotor 9, the reduction ratio can be greater than 2.5 and less than or equal to 6.0, typically around 3.0. In the case of a propulsion system 1 comprising an unshrouded fan rotor, the reduction ratio can be greater than 7.0 and less than or equal to 11.0, typically around 9.0.

[0049] The low-pressure shaft 11 is supported by three or four bearings 1a, 11b and / or 1le, in order to control the deformation modes of the low-pressure shaft 1. The low-pressure shaft 11 can thus comprise one or two front bearings 1a, which extend upstream of the combustion chamber 6, and two rear bearings 11b, 1le, which extend downstream of the combustion chamber 6. The bearings 1a, 11b, 1le comprise a first ring mounted on the low-pressure shaft 11 and a second ring mounted on a stator portion of the propulsion system 1, typically on a housing of the propulsion system 1 through which forces are transmitted in the propulsion system 1. Thus, a first front bearing 1a can be mounted on the low-pressure shaft 11 and on an inlet housing 26 of the propulsion system 1, which extends between the blower rotor 9 and the low pressure compressor 4.If necessary, a second front bearing can be mounted on the low-pressure shaft 11 and on the inter-compressor housing 23 (or intermediate housing) or on the inlet housing 26 of the propulsion system 1, i.e., between the low-pressure compressor 4 and the high-pressure compressor 5. A first rear bearing 11b can be mounted on the low-pressure shaft 11 and on the inter-turbine housing 24 (i.e., on the housing extending between the high-pressure turbine 7 and the low-pressure turbine 8), upstream of the low-pressure turbine 8. Alternatively, the first rear bearing 11b can be mounted. on the exhaust housing 27, which extends immediately downstream of the low-pressure turbine 8. The first rear bearing 11b extends downstream of the most downstream bearing 12b of the high-pressure shaft 10. The second rear bearing 11e can be mounted on the exhaust housing 27 or, if integration permits, on the inter-turbine housing. If applicable, the first and second rear bearings 11b, 11e can be mounted on the same cylindrical ferrule, which is itself fixed to the exhaust housing 27.

[0050] In one embodiment, the low-pressure shaft 11 is supported by a front bearing 11a and two rear bearings 11b, 11e.

[0051] All or part of the bearings, in particular the lla-llc bearings of the low-pressure shaft 11, may also include a flexible cage, mounted between the inner ring and a rigid support fixedly attached to the stator part of the propulsion system 1 in order to control the stiffness of the bearings and to fine-tune the position of the first mode of deformation. For this purpose, the cage comprises a generally cylindrical wall mounted between the rigid support and the second ring of the bearing, and radially deformable columns to allow radial displacement of the generally cylindrical wall, and therefore of the bearing, relative to the rigid support. Examples of bearings with a flexible cage are described in documents WO 2021 / 001610 and WO 2022 / 195198 filed on behalf of the Applicant.

[0052] The twin-body propulsion system 1 may in particular include a single- or two-stage high-pressure turbine 7, a high-pressure compressor 5 comprising at least eight stages and at most eleven stages, a low-pressure turbine 8 comprising at least three stages and at most five stages and a low-pressure compressor 4 comprising at least two stages and at most four stages.

[0053] The overall compression ratio of the propulsion system 1, which corresponds to the pressure ratio between the pressure at the outlet of the high-pressure compressor 5 and the pressure at the inlet of the blower rotor 9 (measured at the base of the blower rotor 9), shall be greater than or equal to 40 and less than or equal to 70, preferably greater than or equal to 44 and less than or equal to 55.

[0054] The propulsion system 1 further comprises an assembly including at least one mode damper 25 (known in English as a "squeeze film") capable of sufficiently damping the vibrations generated by a moving element 35 of the propulsion system 1 while ensuring that it can be integrated under the flow stream. The mode damper 25 can, in particular, be mounted between a fixed part of a bearing of the propulsion system 1 and a stator part of the propulsion system 1 (e.g., inlet housing, intermediate housing, turbine housing or exhaust housing), or between a fixed part of the reduction mechanism 19 (ring support or planet carrier, depending on the configuration of the reduction mechanism 19) and a stator part of the propulsion system 1 (e.g., inlet housing or intermediate housing). The propulsion system 1 may include a mode damper 25 on each bearing supporting the low-pressure shaft 11 and / or on each bearing supporting the blower shaft 20. Preferably, the propulsion system 1 includes a mode damper 25 on the upstream bearing 1a of the low-pressure shaft 11 (or at least one of the upstream bearings when the low-pressure shaft 11 comprises several) insofar as this upstream bearing has a larger radius due to the presence of the reduction mechanism. Optionally, the propulsion system 1 may further include a mode damper on the downstream bearing 11b, 1le of the low-pressure shaft 11 (or at least one of the downstream bearings 11b, 1le when the low-pressure shaft 11 comprises several).

[0055] The mode damper 25 includes an inner ring 30 and an outer ring 32, which together define a damping chamber 34, and a fluid 29 confined within the damping chamber 34. The fluid 29 is pressurized within the damping chamber 34 and is configured to dampen the mode change (viscous damping) of the stationary part on which the mode damper 25 is mounted. The inner ring 30 and the outer ring 32 are annular and coaxial with the moving element 35 (the bearing housing, a moving part (satellites or ring gear) of the reduction mechanism 19). The inner ring 30 therefore extends radially between the moving element and the outer ring 32.

[0056] The fluid 29 may include any type of viscous fluid 29, for example oil or fuel.

[0057] The supply parameters of the mode damper 25 can be conventional. For example, the supply pressure of the fluid 29 can be between 0.5 bar (50 kPa) and 30 bar (3000 kPa), preferably between 4 bar (400 kPa) and 18 bar (1800 kPa); the temperature of the fluid 29 can be between 10°C and 300°C, preferably between 50°C and 200°C; the supply flow rate of the fluid 29, which depends on the supply pressure and the sealing system of the mode damper 25, can be between ten liters per hour (L / h) and several hundred liters per hour.

[0058] The damping chamber 34 is radially delimited internally by the inner ring 30 and radially externally by the outer ring 32. The inner surface 33 of the outer ring 32 and the outer surface 31 of the inner ring 30 which delimit the damping chamber 34 may be smooth, in which case the inner ring 30 and the outer ring 32 are spaced radially apart from each other so as to define the damping chamber 34. Alternatively, a recess may be formed in at least one of the inner surface 33 of the outer ring 32 and the outer surface 31 of the inner ring 30, the damping chamber 34 then being delimited by the walls of the recess.

[0059] When the mode damper 25 is mounted on a bearing, the inner ring 30 of the mode damper 25 corresponds to the second ring of the bearing (or is integral with the second ring of the bearing). Similarly, when the mode damper 25 is mounted on a reduction mechanism 19, the inner ring 30 is integral with the crown support or the planet carrier, depending on the configuration of the reduction mechanism 19.

[0060] The inner ring 30 is fixed relative to the outer ring 32. For example, the assembly further includes an anti-rotation device configured to lock the inner ring 30 relative to the outer ring 32. The anti-rotation device may, for example, include a radial spacer mounted in notches formed in the inner ring 30 and in the stator part of the propulsion system 1 (or, alternatively, in the outer ring 32 of the mode damper 25) so as to circumferentially lock the inner ring 30 relative to the outer ring 32.

[0061] The damping chamber 34 is axially delimited by axial terminals 34a, 34b configured to prevent the fluid 29 from leaking during the operation of the propulsion system 1. The axial terminals 34a, 34b may correspond to the upstream and downstream axial limits of the damping chamber 34, typically the upstream and downstream faces of the recess formed in the outer ring 32 and / or the inner ring 30. Alternatively, at least one of the axial terminals 34a, 34b may be formed by a sealing segment mounted between the inner ring 30 and the outer ring 32 so as to axially delimit the damping chamber 34. Preferably, the mode damper 25 comprises two sealing segments, each forming one of the axial terminals 34a, 34b.

[0062] Each sealing segment 34a, 34b can be made of a metallic material, a composite material, a ceramic or an elastomeric material, preferably a metallic or elastomeric material.

[0063] In one embodiment, the mode damper 25 is circumferentially continuous around the X axis, that is to say that the damping chamber 34 is devoid of separating means forming independent damping chamber sectors 34.

[0064] Optionally, the mode damper 25 includes a fluid supply groove 36 29 configured to distribute the fluid 29 circumferentially in the damping chamber 34. The supply groove 36 can, for example, be formed in the external surface 31 of the inner ring 30 and extend circumferentially and continuously over the entire circumference of the inner ring 30.

[0065] In order to improve the performance of the propulsion system 1 and reduce the noise generated by the propulsion system 1, the axial length and maximum radial clearance of the mode damper 25 are configured so that the surface viscosity of The mode 25 damper, which corresponds to the ratio between the square of the axial length and the product between the maximum cubed radial clearance and twice the value of Pi, is between 1.0 x 104 m1 and 1.15 x 106 m1:

[0066] 1.0 x 10⁴ £ surface viscosity = ^4 S 1.15 x 10⁶ : L is the length axial length of the mode 25 damper and is expressed in meters (m); and C is the maximum radial clearance of the mode 25 damper and is expressed in meters (m).

[0067] The axial length L of the mode 25 damper corresponds to the distance along which the oil film operates (viscous damping). When the mode 25 damper lacks a feed groove ([Fig. 3]), the axial length L therefore corresponds to the axial distance between the axial terminals 34a, 34b of the damping chamber 34, and, where applicable, between the sealing segments of the damping chamber 34. When the mode 25 damper includes a feed groove 36 ([Fig. 4]), the axial length does not include the length of the feed groove 34; it therefore corresponds to the sum of the axial distance L1 between the upstream axial terminal 34a and the upstream limit of the groove and the axial distance L2 between the downstream limit of the groove and the downstream axial terminal 34b.

[0068] The maximum radial play of the mode damper 25 corresponds to the maximum radial distance, outside the feed groove 36, between the inner surface 33 of the outer ring 32 and the outer surface 31 of the inner ring 30. The maximum radial play therefore corresponds to the maximum thickness of the part of the oil film which plays a role in the damping of the modes.

[0069] When the moving element 35 rotates at a given rotational speed co, the inner ring 30 is subjected to a precessional motion (i.e., an orbit without relative rotation of the inner ring 30 around its axis) of the same frequency, resulting from the presence of an (inevitable) imbalance on the moving element 35, with a relative eccentricity e with respect to the outer ring 32. The damping generated by the mode damper 25 then depends on the tangential force generated by the compression of the fluid film 29 between the inner and outer rings. The tangential force Ft can be estimated by the following formula:

[0070] F7=-Crxexw

[0071] where: Cv is the viscosity of the mode 25 damper.

[0072] Viscosity is a drag coefficient that reflects the ability of the mode 25 damper to dissipate energy. This viscosity can be estimated, in particular, using the following formula:

[0073] . ~ / L)3 Cv - Ak h'aRX'q )

[0074] where: p depends on the properties of the fluid 29 of the mode damper 25;

[0075] R is the implantation radius of the mode damper 25, that is to say the maximum radius between the X axis and the external surface 31 of the inner ring 30, the implantation radius being measured in a plane normal to the X axis and expressed in meters (m); and

[0076] A is a constant.

[0077] The surface viscosity defined above is therefore a parameter allowing us to approximate the viscosity of the mode 25 damper by taking into account the surface area of ​​the mode 25 damper:

[0078] }, (2) surface viscosity -

[0079] R, L and C being distances, these parameters are determined when the propulsion system 1 (and therefore the assembly including the mode damper 25) is cold, as specified above.

[0080] When the mode damper 25 is dimensioned to comply with formula (1), the damping provided by the mode damper 25 is improved while taking into account the mechanical integration constraints of the assembly including the damper. In particular, increasing the surface viscosity makes it possible to increase the damping of the moving element 35. However, this increase does not need to be achieved at the expense of the available space in the propulsion system 1.

[0081] The present description applies mainly to propulsion systems in which at least one of the rotors (low pressure shaft 11, or high pressure shaft 10, rotating part of the reduction mechanism 19, etc.) has several natural modes of deformation included in the operating range of the propulsion system 1. In general, this will more particularly be the low pressure shaft 11 (in the case of direct drive propulsion systems or systems including a reduction mechanism 19).

[0082] The Applicant observed that optimizing the damping of the first deformation mode could result in an increase in the stiffness of the mode damper 25 at high frequencies in the assembly comprising the mode damper 25 and the moving element 35, and in bypassing the flexible cage, as the forces applied by the moving element 35 would then be transmitted directly to the stator part of the propulsion system 1 by the mode damper 25. Dimensioning the mode damper 25 to comply with formula (1), on the contrary, optimizes the damping provided by the mode damper 25 for the first three deformation modes and avoids the risk of locking. In particular, a mode damper 25 with a surface viscosity greater than or equal to 1.0 x 10⁴ and less than or equal to 1.15 x 10⁶ has the following effect: - to reduce the amplitude of the first mode of deformation of the moving element 35 without changing its frequency; and - to increase the frequency of the second and third modes of deformation and to reduce their amplitude.

[0083] Preferably, the surface viscosity of the mode damper 25 is less than or equal to 7.1 x 105. Indeed, beyond this value, the amplitude of the second and third modes of deformation of the moving element 35 increases again (as does the frequency of the second and third modes of deformation).

[0084] In order to take into account further the integration constraints in the optimization of the damping, the mode 25 damper has a surface clearance, which corresponds to the ratio between the maximum radial clearance C (between the external surface 31 of the inner ring 30 and the internal surface 33 of the outer ring 32) and the product between twice the value of Pi multiplied by the mounting radius R of the mode 25 damper and the axial length L of the mode 25 damper, between 2.5 x 102 m1 and 1.0 x 10 1 m1:

[0085] 2.5 x 10'2 < surface clearance = 1.0 x 10"1.

[0086] The surface clearance parameter takes into account in particular the implantation radius of the mode damper 25, and therefore the available space in the propulsion system 1. Indeed, the smaller the implantation radius R, the larger the surface clearance but the less effective the damping. A mode 25 damper exhibiting both a surface viscosity between 10A4 m1 and 1.15 x 10A6 m1 and a surface clearance between 2.5*102 m1 and 1.0*101 m1 is therefore capable of sufficiently reducing the vibration response of the moving element 35 around which it is mounted, while respecting strict size criteria, thanks to the adaptation of the axial length and radial clearance of the mode 25 damper. Thus, the dimensioning of the mode 25 damper(s) is adapted according to their position in the propulsion system 1, taking into account, in particular, their mounting radius R.

[0087] In a conventional engine, given the values ​​usually selected for the axial length L and the maximum radial clearance C of the mode damper, the surface viscosity is greater than 115 x 106, which leads to the mode damper being locked. Conversely, by selecting values ​​for the axial length L and the maximum radial clearance C such that the surface viscosity of the mode damper 25 remains less than ]J 5 x IfA, the mode damper 25 is no longer locked and regains a classic damper behavior. In order to optimize the behavior of the mode 25 damper, increasing the surface viscosity (without exceeding the limit of 115x 106) then makes it possible to reduce the response on the modes of deformation.

[0088] Preferably, the surface clearance of the mode 25 damper is greater than or equal to 3.7 x 102 m1 in order to optimize the integration of the mode 25 damper. Advantageously, when the viscosity is high, a surface clearance greater than or equal to 3.7 x 102 m1 has the effect of moderating the damping of the mode 25 damper and thus limiting its stiffness in the high frequencies.

[0089] The present description is particularly applicable to propulsion systems comprising a reduction mechanism 19 and a high-speed low-pressure shaft 11. The term "high-speed low-pressure shaft 11" here refers to a low-pressure shaft 11 whose redline speed, which corresponds to the absolute maximum speed that the low-pressure shaft 11 is likely to encounter during the entire flight (according to European certification regulation EASA CS-E 740 (or according to US certification regulation 14-CFR Part 33.87)), is greater than or equal to 8000 revolutions per minute. For example, the redline speed may be between 8500 and 12000 revolutions per minute, preferably between 9000 and 11000 revolutions per minute. The redline speed corresponds to the maximum rotational speed when the propulsion system is in good working order (and potentially at the end of its service life).It is therefore likely to be reached by the low-pressure shaft 11 under flight conditions. This speed limit is part of the data declared in the engine certification ("type certificate data sheet"). Indeed, this rotational speed is usually used as a reference speed for the sizing of propulsion systems 1 and in certain certification tests (such as blade loss or rotor integrity tests).

[0090] A first example of a propulsion system 1 in which the present description finds application is a USF type engine 1 comprising a single unshod fan rotor 9 and an unshod fan stator 16, which are fixed in rotation with respect to the X-axis of the fan rotor 9. The fan rotor 9 has a diameter D9 greater than 132 inches (3.3528 meters (m)) and is driven via a reduction mechanism by a drive turbine whose limiting speed is between 7,000 and 14,000 revolutions per minute (rpm). The low-pressure shaft is supported by three bearings, a front bearing mounted on the inlet casing and two rear bearings mounted on the exhaust casing. The dilution ratio of this propulsion system 1 is greater than or equal to 40, for example between 40 and 80 inclusive, and its overall compression ratio is greater than 40. This propulsion system 1 is also configured to provide a thrust of between 50,000 and 350,000 N..

[0091] A second example of a propulsion system 1 in which the present description finds application is an engine comprising a shrouded fan section. The rotor of The fan 9 has a diameter D9 between 70 and 120 inches (between 1.778 and 3.048 meters (m)) and is driven via a reduction mechanism by a drive turbine with a maximum speed between 7,000 and 14,000 revolutions per minute (rpm). The low-pressure shaft is supported by three bearings: a front bearing mounted on the intermediate housing and two rear bearings mounted on the exhaust housing. The bypass ratio of this propulsion system 1 is greater than 10, less than or equal to 35, preferably less than or equal to 18, and its overall compression ratio is greater than 40. This propulsion system 1 is further configured to provide a thrust between 50,000 and 350,000 N.

[0092] In these two examples of propulsion systems 1, the mode damper 25 can be mounted on the reduction mechanism 19 and / or at least one bearing of the low pressure shaft 11. For example, the mode damper 25 can be mounted on the front bearing 1la and on one of the downstream bearings 11b, 1le which are mounted on the exhaust housing, these two bearings 11b, 1le being for example roller or ball bearings.

[0093] The invention applies however to any type of bearing capable of comprising a mode 25 damper, and in particular to roller bearings, ball bearings or tapered roller bearings.

[0094] It should be noted here that, regardless of the configuration of the propulsion system 1, the diameter D9 of the fan rotor 9 is measured in a plane normal to the axis of rotation of the fan rotor, at the intersection between a vertex 21 and a leading edge 22 of the blades 14 of the fan rotor 9, and is expressed in meters (m). Note that since Figures 1 and 2 are partial views, the diameter D9 is only partially visible.

[0095] It should be noted that these two examples of propulsion systems present severe mechanical integration constraints (related in particular to the dimensions of the high-pressure body, which is increasingly compact, and to the integration of the reduction mechanism, leading to bearings with a large radius due to the axial proximity of the reduction mechanism) leading to constrained bearing environments (bearings supporting shafts radially distant from the X axis).

[0096] Examples of dimensioning mode 25 shock absorbers adapted in particular for assemblies 1 to 7 comprising a bearing supporting the low-pressure shaft 11 of a propulsion system 1, are as follows: No. R(m) L(m) C (mm) Surface viscosity (m1) Surface clearance (m') 1 0.1 20 0.7 8.25 x 10A4 2.48 x 10A-2 2 0.2 12 0.7 5.35 x 10A4 3.71 x 10A-2 3 0.2 20 0.7 1.32 x 10A5 1.98 x 10A-2 4 0.3 12 0.7 5.35 x 10A4 2.48 x 10A-2 5 0.3 12 0.7 4.01 x 10A4 1.86 x 10A-2 6 0.4 10 0.9 1.16 x 10A4 1.91 x 10A-2 7 0.4 12 0.9 4.31 x 10A4 4.09 x 10A-2

[0097] These examples of mode damper 25 all make it possible to reduce the vibration response of the moving element 35 (bearing or reduction mechanism 19) around which they are mounted and can in particular be implemented in propulsion systems 1 conforming to the first and second examples of propulsion systems described above.

[0098] In these examples, the mode 25 dampers Nos. 1 to 7 exhibit high surface viscosity and low surface clearance, which optimizes the damping of bearing deformation modes, particularly in propulsion systems such as the examples described above. Their low surface clearance also helps to limit cavitation effects. The mode 25 damper No. 2 is particularly optimized in terms of size (small mounting radius and axial length) and damping (high surface viscosity and low surface clearance) and can, for example, be mounted on a low-pressure shaft bearing 11 due to its small mounting radius R and short axial length L. The mode damper No. 3 is also optimized in terms of radial size due to its small mounting radius and damping (high surface viscosity).

[0099] Mode 25 dampers No. 4 to 7 will be particularly advantageous for damping located above the reduction mechanism 19.

[0100] Figure 7 illustrates the influence of the surface viscosity of a mode 25 damper on the displacement of the front bearing 1 of the second example of a propulsion system 1 as a function of the rotational speed of the low-pressure shaft 11 of this propulsion system 1. In this figure: - the Cv curve— corresponds to a mode 25 damper with a surface viscosity much lower than 104 m1; - the Cv- curve corresponds to a mode 25 damper with a surface viscosity greater than or equal to 104 m1; - the Cv curve corresponds to a mode 25 damper with a surface viscosity between 104 m1 and 7.1 x 105 m 1; - the Cv+ curve corresponds to a mode 25 damper with a surface viscosity greater than 7.1 x 105 m 1 and less than 1.15 x 106 m 1; - the Cv++ curve corresponds to a mode 25 damper with a surface viscosity greater than 1.15 x 10⁶ m²; and - the Cv+++ curve corresponds to a mode 25 damper with a surface viscosity even greater than Cv++.

[0101] As can be seen from this figure, the amplitude of the three deformation modes of the bearing comprising a mode damper with a surface viscosity between 1.0 x 10⁴ and 1.15 x 10⁶ m² is reduced. Such a mode damper is therefore capable of damping vibrations and reducing the HCF loads of such a bearing throughout the entire operating range of the propulsion system. In contrast, the amplitude of the second deformation mode of the bearing comprising a mode damper with a surface viscosity less than 1.0 x 10⁴ m² is very high (curves C— and C-), as is the amplitude of the third deformation mode of the bearing comprising a mode damper with a surface viscosity greater than 1.15 x 10⁶ m². Manufacturing process

[0102] An assembly for a propulsion system conforming to this exposition can be manufactured in accordance with the following steps, after defining the damping requirement of the mode 25 damper (surface viscosity) and specifying the geometry of the mode 25 damper (in particular its axial length L and its maximum radial clearance C as a function of the defined surface viscosity).

[0103] A mode 25 damper comprising an inner ring 30 and an outer ring 32 is dimensioned so that the mode 25 damper has the axial length and maximum radial clearance such that the surface viscosity of the mode 25 damper is between 1.0 x 104 and 7.1 x 105 m1.

[0104] The mode damper 25 is then manufactured and assembled with a movable element 35, for example a bearing, so that the inner ring 30 and the outer ring 32 are coaxial with the bearing and the first ring (which are movable around the X axis) of the bearing and that the inner ring 30 extends radially between the bearing and the outer ring 32.

[0105] A fluid 29 is then injected under pressure into the damping chamber 34 of the mode damper 25, between the inner ring 30 and the outer ring 32.

[0106] Comparative example:

[0107] The engine 1 is a twin-body propulsion system comprising a shrouded blower section 2 driven in rotation by means of a gear reduction mechanism corresponding to the current technical standard (at the date of filing of this application) which we seek to improve.

[0108] The engine 2 is a twin-body propulsion system 1 comprising a shrouded fan section 2 conforming to the second example described above.

[0109] The low-pressure shaft of motors 1 and 2 are each supported by a front ball bearing comprising a mode 25 damper conforming to formulas (1) and (3) and two rear roller bearings, the most downstream roller bearing also including a mode 25 damper conforming to formulas (1) and (3). The reduction mechanism and the second downstream bearing do not include a mode damper. Dimensioning parameter (SLS unless otherwise specified) Motor 1 (reference motor) Motor 2 (as disclosed in this application) Mode damper mounting radius of front low-pressure shaft bearing (m) N / A 115.6 mm Axial length of mode damper of front low-pressure shaft bearing (m) N / A 25 mm Maximum radial clearance of mode damper of front low-pressure shaft bearing (m) N / A 0.5 mm Surface viscosity of the mode damper of the front bearing (m1) N / A 1.59 x 104 m 1 Surface clearance of the mode damper of the front bearing (m ■') N / A 2.75 x 102 m 1 Axial length of the mode damper of the first rear bearing of the low pressure shaft (m) N / AN / A Maximum radial clearance of the mode damper of the first rear bearing of the low pressure shaft (m) N / AN / A Surface viscosity of the mode damper of the first rear bearing (m1) N / AN / A Setting radius of the mode damper of the second N / A 95.1 mm. Low-pressure shaft rear bearing (m) Axial length of the mode damper of the second low-pressure shaft rear bearing (m) N / A 25 mm Maximum radial clearance of the mode damper of the second low-pressure shaft rear bearing (m) N / A 0.5 mm Surface viscosity of the mode damper of the second rear bearing (m1) N / A 1.59 x 10⁴m¹ Surface clearance of the mode damper of the second rear bearing (m1) N / A 3.35 x 10²m¹ Low-pressure shaft limit speed (Ni) (redline) 7,917 rpm 10,143 rpm Frequency of the first low-pressure shaft deformation mode 158 Hz 155 Hz Propulsion system thrust 147,000 N 148,000 N Diameter 2235 mm 2235 mm Number of blower blades 18 18 Blower pressure ratio 1.39 1.39 Dilution ratio 13 13 Peripheral speed of blower blades 334 m / s 336 m / s Blower rotor hub-to-head ratio 0.3 0.3 Blower rotor redline speed 2990 rpm 3007 rpm High-pressure shaft (redline) speed limit: 19,500 rpm 23,100 rpm Low-pressure compressor compression ratio: 2.51 2.51 Overall compression ratio: 40 50 Reduction ratio: 2.65 3.73 Number of low-pressure compressor stages: 3 3 Number of high-pressure compressor stages (axial and concentric): 9 10 Number of high-pressure turbine stages: 2 2 Number of low-pressure turbine stages: 4 4 High-pressure turbine (redline) inlet temperature: 1810 K 1930 K Low-pressure turbine (redline) inlet temperature: 1317 K 1370 K Low-pressure shaft diameter under high-pressure turbine: 62 mm 45 mm High-pressure turbine disc bore radius: 70 mm 53 mm

[0110] The engine 1 has a subcritical shaft dynamic situation, i.e., the first bending mode of the low-pressure shaft is outside the operating range with a sufficient margin. This characteristic limits the engine's performance in terms of the overall compression ratio / temperature of the high-pressure turbine and in terms of the mechanical load (stress at the base of the blade) of the low-pressure turbine.

[0111] The motor 2 has a supercritical shaft dynamic situation, that is to say that the first bending mode of the low pressure shaft 11 is in the operating range of the low pressure shaft 11. This dynamic situation is however permitted by the placement of mode dampers 25 on the upstream bearing 1 and the second downstream bearing of the low pressure shaft 11 according to formulas (1) and (3). The ability to access this modal positioning of the low-pressure shaft 11 thus makes it possible to increase the compression ratio of the high-pressure compressor (and therefore the overall compression ratio) and the temperature of the high-pressure turbine, as well as the mechanical load (stress at the tip of the blade) of the low-pressure turbine. This improves the performance of the propulsion system 1 by enhancing its thermal efficiency and reducing its mass.

Claims

Demands

1. Assembly (lia, 11b, 11c) of an aeronautical propulsion system (1) comprising: - a movable element (35) rotating about an axis (X) of the assembly (lia, 11b, 11c); and - a mode damper (25) comprising an inner ring (30), an outer ring (32) and a fluid (29) confined between the inner ring (30) and the outer ring (32), the inner ring (30) and the outer ring (32) being coaxial with the movable element (35); the inner ring (30) extending radially between the movable element (35) and the outer ring (32);and the mode damper (25) having an axial length and a maximum radial clearance such that a surface viscosity of the mode damper (25), which corresponds to the ratio between the square of the axial length and the product between the maximum radial clearance cubed multiplied by twice the value of Pi, is between 1.0 x 104 m 1 and 1.15 x 106 m1 : 1.0 x K)4 < surface viscosity - “3 <1.15x106 or ' L is the axial length of the mode damper (25) and is expressed in meters (m); and C is the maximum radial clearance of the mode damper (25) and is expressed in meters (m).;

2. Assembly (11a, 11b, 11c) according to claim 1, wherein the surface viscosity of the mode damper (25) is less than or equal to 7.1 x 10⁵ m

3. Assembly (11a, 11b, 11c) according to any one of claims 1 and 2, wherein the mode damper (25) further has a surface clearance, which corresponds to the ratio between the maximum radial clearance and the product between twice the value of Pi multiplied by the implantation radius (R) of the mode damper (25) and the axial length (L) of the mode damper (25), is between 2.5 x 10² m₁ and 1.0 x 10¹ m₁: 2.5 x 10² < surface clearance = < 1.0 x 10¹ where: R is the implantation radius of the mode damper (25) and is expressed in meters (m).

4. Assembly (lia, 11b, 1 le) according to claim 3, wherein the surface clearance of the mode damper (25) is greater than or equal to 3.7 x 102 m1.

5. Assembly (lia, 11b, 11c) according to any one of claims 1 to 4, wherein the moving element (35) comprises a moving part of a bearing (lia, 11b, 11c) of the propulsion system (1).

6. Assembly (11a, 11b, 11c) according to any one of claims 1 to 5, wherein the moving element (35) comprises a moving part of a reduction mechanism (19) of the propulsion system (1).

7. Assembly (11a, 11b, 11c) according to any one of claims 1 to 6, wherein the mode damper (25) further comprises at least one sealing segment (34a, 34b), preferably two sealing segments (34a, 34b), axially delimiting the fluid (29) between the inner ring (30) and the outer ring (32).

8. Aeronautical propulsion system comprising an assembly (lia, 11b, 1 le) according to any one of claims 1 to 7 and a rotor coaxial with the moving element (35).

9. System according to claim 8, wherein the rotor includes a shaft (11) of the propulsion system and the moving element (35) includes a bearing (11a, 1b, 11e) which supports the rotor.

10. Aeronautical propulsion system (1) according to claim 8, wherein the rotor comprises a drive shaft (11), the propulsion system further comprising: - a fan section (2) comprising a fan shaft (20); - a drive turbine (8) driven in rotation by the rotor (11) about the axis (X); and - a reduction mechanism (19) coupling the rotor (11) and the fan shaft (9) in order to drive the fan shaft (20) at a rotational speed lower than the rotational speed of the rotor (11), the moving element (35) comprising a moving part of the reduction mechanism (19).

11. Aeronautical propulsion system (1) according to claim 10, wherein the drive shaft (11) is supported by at least one front bearing (1a) and one or two rear bearings (11b), the moving element corresponding to the front bearing (1a) and where appropriate to one of the rear bearings (11b).

12. Aeronautical propulsion system (1) according to any one of claims 10 and 11, wherein a dilution ratio of the propulsion system (1) is greater than or equal to 10, for example between 10 and 80 inclusive.

13. Aircraft (100) comprising at least one aeronautical propulsion system (1) according to any one of claims 7 to 12 fixed to the aircraft by means of a mast.

14. A method for manufacturing an assembly (11a, 11b, 11c) of an aeronautical propulsion system (1) comprising the following steps: - determining an axial length and a maximum radial clearance of a mode damper (25) such that a surface viscosity of the mode damper (25), which corresponds to the ratio of the squared axial length to the product of the cubed maximum radial clearance multiplied by twice the value of Pi, is between 1.0 x 10⁴ m and 1.15 x 10⁶ m: 10⁴ < surface viscosity = 3 < 1.15 x 10⁶ where L is the axial length of the mode damper (25) and is expressed in meters (m); and C is the maximum radial clearance of the mode damper (25) and is expressed in meters (m); - manufacturing an inner ring (30) and an outer ring (32) of the mode damper (25) depending on the axial length and maximum radial clearance thus determined; - assemble the inner ring (30) and the outer ring (23) with a moving element (35);and - inject a fluid (29) between the inner ring (30) and the outer ring (32).;

15. A manufacturing method according to claim 14, wherein the axial length and maximum radial clearance are determined such that the surface viscosity is greater than or equal to 3.7 x 10² m1.

16. A manufacturing method according to any one of claims 14 and 15, wherein the axial length, maximum radial clearance and an implantation radius are further determined such that a surface clearance of the mode damper (25) is between 2.5 x 10A-2 m1 and 1.0 x 10A-lm', where the surface clearance is defined as follows: 29 2.5 x la2 < surface play = < 1.0 x 10'loû : R is a radius of implantation of the mode damper (25) and is expressed in meters (m).

17. A manufacturing method according to claim 16, wherein the axial length, maximum radial clearance and implantation radius are determined so that the surface clearance is greater than or equal to 3.7 x 102 m1.

Images