Optimisation of the high-pressure core of a high-bypass turbofan engine

Abstract

The invention relates to an aeronautical propulsion system (1), which comprises high-pressure bearings (12a, 12b) comprising a front bearing (12a) extending upstream of a combustion chamber of the propulsion system (1) and a rear bearing (12b) extending downstream of the combustion chamber, wherein a bypass ratio of the propulsion system (1) is greater than or equal to 8, a pressure ratio of the fan section is less than or equal to 1.45, and the high-pressure compressor (5) is dimensioned so as to comply with the following formula, wherein: XN is the maximum speed of the high-pressure shaft (10) (in rpm); Ri is a mean inner radius of the high-pressure compressor (5) (in m); and L is an axial distance between a centre of gravity (G12a) of the front bearing (12a) and a centre of gravity (G12b) of the rear bearing (12b) (in m).

Description

[0001] DESCRIPTION

[0002] Optimization of the high-pressure body of a high-bypass engine

[0003] TECHNICAL FIELD

[0004] This application relates generally to the field of propulsion systems, and more particularly to aeronautical propulsion systems comprising a shrouded or unshrouded fan and exhibiting a high, or even very high, bypass ratio.

[0005] STATE OF THE ART

[0006] 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.

[0007] 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 to obtain less energy-intensive and more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences, with the aim of improving the energy efficiency of aircraft.

[0008] Thus, in order to improve the propulsion efficiency of the propulsion system and reduce its specific fuel consumption as well as the noise emitted by the fan section, propulsion systems with a high bypass ratio (BPR, the ratio between the secondary airflow rate and the primary airflow rate) have been proposed. To achieve such bypass ratios, the fan section can be decoupled from the low-pressure turbine, thereby allowing their respective rotational speeds to be optimized independently. Generally, decoupling is achieved using a reduction mechanism placed between the upstream end of the low-pressure shaft and a rotor of the fan section. The fan section rotor is then driven by the low-pressure shaft via the reduction mechanism at a rotational speed lower than that of the low-pressure shaft.

[0009] Improving the propulsive efficiency of the fan section also involves reducing its pressure ratio. However, since the propulsion system has a high dilution rate, this requires reducing the flow rate in the high-pressure section, which generally translates into a reduction in its radial dimensions and an increase in its rotational speed.

[0010] However, the length of the high-pressure body varies at a lower ratio than its diameter, primarily because the axial clearances between the moving blades and the stators do not vary proportionally with the diameter of the high-pressure body. Consequently, the dynamic behavior of the high-pressure body is altered, particularly the positioning of its natural modes, which may fall within the stabilized operating range of the propulsion system. The dynamic state of the high-pressure body must be controlled to ensure its proper operation.

[0011] EXPOSED

[0012] One aim of this application is to optimize the propulsion system in order to increase its efficiency while controlling the dynamic behavior of the high-pressure shaft line.

[0013] For this purpose, an aeronautical propulsion system is proposed comprising:

[0014] - a blower section comprising a blower rotor connected to a blower shaft; - a first turbine configured to drive the blower rotor via a first shaft around an axis of rotation;

[0015] - a reduction mechanism coupling the first shaft and the blower shaft to drive the blower shaft at a rotational speed lower than that of the first shaft; - a second turbine configured to drive a compressor via a second shaft around the axis of rotation; and

[0016] - bearings configured to center the second shaft relative to the axis of rotation, the bearings comprising a front bearing extending upstream of a combustion chamber of the propulsion system and a rear bearing extending downstream of the combustion chamber,

[0017] in which a propulsion system dilution ratio is greater than or equal to 8, a blower section pressure ratio is less than or equal to 1.45, and the compressor is sized to comply with the following formula:

[0018] Ri /

[0019] / r

[0020] 3700 < - — 2

[0021] X 10 9 < 4900

[0022] XN

[0023] where: XN is the limiting speed of the second shaft, in revolutions per minute;

[0024] Ri is the average internal radius of the compressor, in meters; and

[0025] L is an axial distance between a center of gravity of the front bearing and a center of gravity of the rear bearing, in meters.

[0026] Some preferred but not limiting characteristics of the propulsion system according to the first aspect are the following, taken individually or in combination:

[0027] - each bearing is associated with a bearing mode damper;

[0028] - bearing mode dampers include damping by pressurized fluid film;

[0029] - bearing mode dampers include a deformable cage mounted between a ring of each bearing and a stator part of the propulsion system;

[0030] - the speed limit of the second shaft also complies with the following formula:

[0031] 2370 3000

[0032]

[0033] where: Dg is a diameter of a rotor of the blower section, measured in a plane normal to the axis of rotation at the level of an intersection between a vertex and a leading edge of the blower rotor blades, and is expressed in meters;

[0034] BPR is a propulsion system dilution ratio and is measured when propulsion system 1 is stationary, uninstalled, in takeoff regime;

[0035] y is the adiabatic coefficient of air; and

[0036] OPR is an overall compression ratio of the propulsion system and is measured when the propulsion system 1 is stationary, uninstalled, in takeoff regime;

[0037] - the compressor has a hub-to-head ratio between 0.45 and 0.55;

[0038] - the compressor comprises at least eight stages and at most eleven stages;

[0039] - an average radius of an additional turbine bore is at most equal to 75 mm;

[0040] - a reduction rate of the reduction mechanism is greater than or equal to 2.5 and less than or equal to 11; and / or

[0041] - a high-pressure compressor pressure ratio is between 15 and 27.

[0042] According to a second aspect, it is proposed an aircraft comprising at least one propulsion system according to the first aspect fixed to the aircraft by means of a mast.

[0043] According to a third aspect, a method is proposed for dimensioning a propulsion system comprising a reduction mechanism coupling a first shaft and a blower rotor to drive the blower rotor at a speed lower than the speed of the first shaft, a second shaft coupling a turbine to a compressor to drive the compressor at a speed higher than the speed of the first shaft, and bearings configured to center the second shaft with respect to the axis of rotation, the bearings comprising a front bearing extending upstream of a combustion chamber of the propulsion system and a rear bearing extending downstream of the combustion chamber,

[0044] the compressor being sized to comply with the following formula:

[0045] Ri /

[0046] / r

[0047] 3700 < - — 2

[0048] 10 9 < 4900

[0049] XN

[0050] where: XN is the limiting speed of the second shaft, in revolutions per minute;

[0051] Ri is the average internal radius of the compressor, in meters; and

[0052] L is an axial distance between a center of gravity of the front bearing and a center of gravity of the rear bearing, in meters.

[0053] The speed limit of the second shaft can also be determined by the following formula:

[0054] 2370 3000

[0055]

[0056] where: Dg is a diameter of a rotor of the blower section, measured in a plane normal to the axis of rotation at the level of an intersection between a vertex and a leading edge of the blower rotor blades, and is expressed in meters;

[0057] BPR is a dilution rate of the propulsion system and is measured when the propulsion system 1 is stationary, uninstalled, in takeoff regime; y is the adiabatic coefficient of air; and

[0058] OPR is an overall compression ratio of the propulsion system and is measured when propulsion system 1 is stationary, uninstalled, in takeoff regime.

[0059] According to a fourth aspect, a manufacturing process for an aeronautical propulsion system is proposed, comprising the following steps:

[0060] - to dimension the aeronautical propulsion system in accordance with the process according to the third aspect; and

[0061] - to manufacture the aeronautical propulsion system.

[0062] DESCRIPTION OF THE FIGURES

[0063] Other features, purposes, and advantages will become apparent from the following description, which is purely illustrative and not exhaustive, and should be read in conjunction with the attached drawings on which:

[0064] Figure 1 is a schematic, partial and cross-sectional view of an example of a propulsion system conforming to a first embodiment, in which the blower section is faired;

[0065] Figure 2 is a schematic, partial, cross-sectional view of an example of a propulsion system conforming to a first embodiment, in which the blower section is unfaired;

[0066] Figure 3 is a schematic cross-sectional view of an example of a reduction mechanism according to a first variant;

[0067] Figure 4 is a schematic cross-sectional view of an example of an epicycloidal reduction mechanism according to a second variant;

[0068] Figure 5 is an example of an aircraft that may include at least one propulsion system conforming to the first or second embodiment;

[0069] Figure 6 is a flowchart illustrating examples of steps in a sizing or manufacturing process conforming to a form of realization.

[0070] Across all figures, similar elements bear identical references.

[0071] DETAILED DESCRIPTION

[0072] A propulsion system 1 has a principal direction extending along a longitudinal axis X and includes, 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).

[0073] The compressor section 4, 5 comprises a series of stages, each including a rotating impeller (rotor) 4a, 5a in front of a stationary impeller (stator) 4b, 5b. The turbine section 7, 8 also comprises a series of stages, each including a stationary impeller (stator) 7b, 8b behind which a rotating impeller (rotor) 7a, 8a rotates. In this application, the axial direction corresponds to the direction of the longitudinal axis X, corresponding to the rotation of the gas generator shafts, 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), respectively, are used with reference to a radial direction such that the internal part or face of an element is closer to the X axis than the external part or face of the same element.

[0074] In operation, an airflow F entering the propulsion system 1 is divided between a primary airflow F1 and a secondary airflow F2, which flow from upstream to downstream in the propulsion system 1.

[0075] The secondary airflow F2 (also called the "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.

[0076] The primary airflow F1 flows in a primary channel 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 F1 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.

[0077] In a twin-spool 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 casing is greater than the rotational speed of the low-pressure casing. In a three-casing 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.

[0078] 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.

[0079] The fan section 2 includes at least the fan rotor 9, which is driven in rotation relative to a stator portion of the propulsion system 1 by the turbine section 7, 8. Each fan 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 12 or have variable pitch. In this case, the root of the blades 14 of each rotor 9 is pivotally mounted 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 a pitch-changing mechanism 15. The pitch-changing mechanism 15 is shown in dashed lines in Figure 1 to indicate that this feature is optional.

[0080] The fan section 2 may further include a fan stator 16, or rectifier, which comprises blades 17 mounted on a hub 18 of the fan stator 16 and whose function is to rectify the secondary airflow F2 exiting the fan rotor 9. The fan stator blades 17 may be fixed relative to the hub 18 or have 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.

[0081] The diameter Dg of the fan rotor can then be between 80 inches (203.2 cm) and 185 inches (469.9 cm) inclusive. When the fan rotor 9 is ducted, the diameter Dg is, for example, between 85 inches (215.9 cm) and 120 inches (304.8 cm) inclusive, for example, around 90 inches (228.6 cm), which allows for conventional integration of the propulsion system 1, particularly under the wing of an aircraft. When the fan rotor 9 is unducted, the diameter Dg is, for example, greater than or equal to 100 inches (254 cm), for example, between 120 inches (304.8 cm) and 156 inches (396.2 cm). The diameter of the fan rotor 9 is measured here in a plane normal to the X-axis of rotation at the intersection between a vertex 21 and a leading edge 22 of the fan rotor blades 14, and is expressed in meters. Note that since Figures 1 and 2 are partial views, the diameter Dg is only partially visible.

[0082] The fan rotor 9 also includes at least twelve blades 14 and at most twenty-four blades 14, for example at least sixteen blades 14 and at most twenty-two 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.

[0083] To improve the propulsive efficiency of propulsion system 1 and reduce its specific fuel consumption and the noise emitted by fan section 2, propulsion system 1 has a high bypass ratio. A high bypass ratio is defined here as 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 F1 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, 3). eedition) and at sea level (so-called SLS conditions, for Seal Level Standard). It should be noted that, in this application, unless otherwise specified, the parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined under these conditions. "Not installed" here means 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 shut down for a sufficient period for the components of the propulsion system to reach ambient temperature.

[0084] The blower 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 blower rotor 9, in order to independently optimize their respective rotational speeds. In this case, the propulsion system 1 further includes an additional shaft, referred to as the blower shaft 20. The low-pressure shaft 11 connects the low-pressure turbine 8 to an inlet of the reduction mechanism 19, while the blower shaft 20 connects the outlet of the reduction mechanism 19 to the blower rotor 9. The blower rotor 9 is thus driven by the low-pressure shaft 11 via the reduction mechanism 19 and the blower shaft 20 at a rotational speed lower than the rotational speed of the low-pressure turbine 8.

[0085] This decoupling allows for a reduction in the rotational speed and pressure ratio of the fan rotor 9 and an increase in the power extracted by the low-pressure turbine 8. Indeed, the overall efficiency of the 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, for example less than or equal to 1.50, for example between 1.05 and 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).

[0086] The propulsion system 1 is configured to provide a thrust between 18,000 lbf (80,068 N) and 70,000 lbf (311,375.51 N), preferably less than or equal to 51,000 lbf (222,411 N), for example, between 20,000 lbf (88,964 N) and 35,000 lbf (155,688 N). The fan section 2 may be shrouded or unshrouded. In the case of a shrouded fan section 2, the fan section 2 includes a fan casing 12, and the fan rotor 9 is housed within the fan casing 12.

[0087] A shrouded fan section 2 comprises a fan rotor 9 extending upstream of a fan stator. The fan stator blades are then generally called outlet guide vanes (OGVs) 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.

[0088] In an unshrouded fan section 2, the fan section 2 is not enclosed by a fan casing. Because the fan section 2 is unshrouded, the blades 14 of the fan rotor 9 have variable pitch. Propulsion systems comprising at least one unshrouded fan rotor 9 are known as "open rotor" or "unducted fan." The propulsion system 1 may comprise two unshrouded, 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 fan rotor(s) 9 may be positioned at the rear of the primary casing 3 to be of the pusher type or at the front of the primary casing 3 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.

[0089] 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.

[0090] The reduction mechanism 19 may include, for example, an epicyclic gear train reduction mechanism, for example of the "epicyclic" or "planetary" type according to the terminology sometimes encountered by those skilled in the art, single-stage or two-stage.According to a first variant, the reduction mechanism 19 can be of the planetary type (“star” in English) (Figure 3) and comprise a sun pinion 19a (input of the reduction mechanism 19), centered on an X axis of rotation of the reduction mechanism 19 (generally coincident with the longitudinal axis X) and configured to be driven in rotation by the low pressure shaft 11, a ring 19b (output of the reduction mechanism 19) coaxial with the sun pinion 19a and configured to drive in rotation the blower shaft 20 around the X axis of rotation, and a series of satellites 19c distributed circumferentially around the X axis of rotation between the sun pinion 19a and the ring 19b, each satellite 19c being internally meshed with the sun pinion 19a and externally with the ring 19b.The series of satellites 19c is mounted on a satellite carrier 19d which is fixed relative to a stator part 19e of the propulsion system 1, for example relative to a housing of the compressor section 4, 5. According to a second variant, the reduction mechanism 19 can be of epicycloidal type (“planetary” in English) (Figure 4), in which case the ring 19b is fixedly mounted on the stator part 19e of the propulsion system 1 and the blower shaft 20 is driven in rotation by the satellite carrier 19d (which is therefore mobile in rotation relative to a stator part 19e of the propulsion system 1, for example relative to a housing of the compressor section 4, 5).

[0091] Regardless of the configuration of the reduction mechanism 19, the diameter of the ring gear 19b and the planet carrier 19d are greater than the diameter of the solar pinion 19a, so that the rotational speed of the blower rotor 9 is less than the rotational speed of the low pressure shaft 11.

[0092] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11. In the case of a propulsion system 1 comprising a shrouded fan rotor 9, the reduction ratio can be greater than or equal to 2.7 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 between 9.0 and 11.0.

[0093] The high-pressure shaft 10 is supported by at least one front bearing 12a and one rear bearing 12b positioned respectively near an upstream end and a downstream end of the high-pressure shaft 10. The front bearing 10a can be mounted on the intermediate housing 23. The rear bearing 12b can be mounted on the inter-turbine housing 24. If necessary, the high-pressure shaft 10 can be supported by a third bearing (not shown), generally located downstream of the shaft 10, near the rear bearing 12b.

[0094] The high-pressure compressor 5 is preferably an axial compressor, that is to say, it does not include a centrifugal stage.

[0095] The movable blade wheels 5a of the high-pressure compressor 5 and the movable blade wheels 7a of the high-pressure turbine 7 are connected to each other and to the bearings by ferrules 10b which, together with the high-pressure shaft 10, form the high-pressure shaft line 10, 10b. This high-pressure shaft line therefore extends from the front high-pressure bearing to the rear high-pressure bearing.

[0096] The low-pressure shaft 11 is supported by three or four bearings 11a, 11b and / or 11c, in order to better control the deformation modes of the low-pressure shaft 1 and possibly shift the deformation modes of the low-pressure shaft 11 in the transient regime of the propulsion system 1, with safety margins compared to the stabilized regimes.

[0097] The Applicant observed that the position of the bearings 11a-11c of the low-pressure shaft 11 could influence the deformation modes of the low-pressure shaft 11. The low-pressure shaft can thus comprise one or two front bearings 11a, extending upstream of the combustion chamber 6, and two rear bearings 11b, 11c, extending downstream of the combustion chamber 6. A first front bearing 11a can be mounted on the low-pressure shaft 11 and on an inlet housing 26 of the propulsion system 1, extending 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 an inter-compressor housing 23 (or intermediate housing) of the propulsion system 1, i.e., between the low pressure compressor 4 and 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 rear bearing 12b furthest downstream of the high-pressure shaft 10. The second rear bearing 11c can be mounted on the exhaust housing 27. If necessary, the first and second rear bearings 11b and 11c can be mounted on the same cylindrical ferrule, which is itself fixed to the exhaust housing 27.

[0098] The inlet casing 26, the intermediate casing 23, the exhaust casing 27 and, where applicable, the inter-turbine casing 24 form structural casings of the propulsion system 1 through which the forces generated by the propulsion system 1 pass.

[0099] Decoupling the low-pressure shaft 11 and the fan rotor 9 by means of the reduction mechanism 19 makes it possible to obtain an efficient propulsion system 1 with a fan pressure ratio of less than 1.45 and a high bypass ratio. Consequently, the power required at constant thrust decreases. The amount of energy needed to drive the fan rotor 9 is therefore reduced, so that the inlet flow rate of the high-pressure compressor 5, and thus the inlet cross-section of the high-pressure compressor 5, can be reduced. This results in a decrease in the overall size of the high-pressure unit, and in particular the radial size of the high-pressure compressor (maximum external diameter of its blades). Such a reduction in the size of the high-pressure unit is generally desirable, as it allows both the reduction of drag and the weight of the propulsion system 1 to decrease, and therefore further improves its specific fuel consumption.

[0100] However, reducing the radial dimensions of the high-pressure core also alters its dynamic behavior, particularly the positioning and number of high-pressure shaft line modes 10, 10b within the high-pressure core's operating range. The dynamic state of the high-pressure shaft line in conventional engines includes at least one (usually two) natural mode of deformation within the operating range of the propulsion system 1, with the third natural mode being beyond the operating range of the propulsion system 1. However, reducing the radial dimensions of the high-pressure core directly reduces its stiffness. Therefore, a third natural mode of deformation may appear within the operating range of the high-pressure shaft line 10, 10b.However, reducing its radial footprint, and in particular lowering the flow path (bringing the X-axis closer), reduces the space available for positioning the bearings. To avoid this third mode of deformation, it is possible to limit the overall compression ratio of the propulsion system 1. However, this has a negative impact on engine efficiency. It is also possible to increase the hub-to-head ratio of the high-pressure compressor 5 (which corresponds to the ratio between an external radius of the high-pressure compressor 5 and an internal radius of the rotor (moving blades) of the high-pressure compressor 5, measured at the inlet of the high-pressure compressor 5 at 50% of the blade chord), by raising the flow path (i.e., by increasing the internal radius of the high-pressure compressor rotor and reducing the radial length of the high-pressure compressor 5 blades).However, this leads to a loss of efficiency in the high-pressure compressor.

[0101] In order to obtain a propulsion system 1 with high propulsive efficiency with a compact high-pressure body, while controlling the dynamic behavior of the high-pressure shaft 10, the high-pressure compressor 5 is dimensioned so that its dimensions and its limiting speed comply with the following formula (1):

[0102] 3700 <

[0103]

[0104] 10 9 < 4900

[0105] where: Ri is the average internal radius of the high-pressure compressor 5, in meters (m);

[0106] L is the axial distance between the front bearing 12a and the downstream bearing 12b of the high-pressure body (in meters (m); and

[0107] XN is the high-pressure body's limiting speed, in revolutions per minute (rpm).

[0108] The mean internal radius Ri corresponds to the arithmetic mean of the internal radii of the high pressure compressor 5, each internal radius being measured in a plane normal to the X axis at mid-distance between the leading edge 5c and the trailing edge 5d of the moving blades 5a of the rotor of the high pressure compressor 5 (i.e. at 50% of the chord at the base of the blade) at the level of the hub.

[0109] L is an axial distance measured between the center of gravity Gi2a of the front bearing 12a and the center of gravity Gi2b of the rear bearing 12b; that is, the distance between the orthogonal projection of the center of gravity Gi2a of the front bearing 12a onto the axis of rotation X and the orthogonal projection of the center of gravity Gi2b of the rear bearing 12b onto the axis of rotation X. It should be noted here that a bearing comprises at least one outer ring and at least one inner ring that are coaxial, as well as bearings (balls, rollers, etc.) mounted between the inner and outer rings and configured to allow relative movement of the inner ring with respect to the outer ring. The inner ring is mounted on the high-pressure shaft 10 and driven in rotation by this shaft 10; the outer ring is mounted on a stator portion (typically, the intermediate housing or the inter-turbine housing) of the propulsion system 1.The center of gravity Gi2a, Gi2b of a bearing 12a, 12b then corresponds to the center of gravity of the assembly formed by the inner ring, the outer ring, and the bearings. When the inner and / or outer ring is formed integrally and in a single piece with a support (flange, shell, housing, etc.) of the propulsion system 1, such that the ring cannot be distinguished from the support, the center of gravity is determined without taking the ring into account. Furthermore, when the bearing is double (duplex) and comprises two rows of bearings that can be housed in the same cage, the center of gravity of the bearing carrying the radial loads is chosen to determine the axial distance L (the other bearing generally carrying the axial loads).An example of a double bearing can, for example, include a row of balls (retaining axial forces) followed by a row of rollers (retaining radial forces): the axial distance L therefore starts from the center of gravity of the row of rollers.

[0110] The limiting speed XN corresponds to the absolute maximum speed that the high-pressure shaft 10, 10b, is likely to encounter during the entire flight (according to the European certification regulation EASA CS-E 740 (or according to the American certification regulation 14-CFR Part 33)). The limiting speed corresponds to the maximum rotational speed when the propulsion system is in good condition (and potentially at the end of its service life). It is therefore likely to be reached by the high-pressure shaft 12 under flight conditions. This limiting speed is part of the data declared in the engine certification (the "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).

[0111] The speed limit XN of the high-pressure shaft 10 can, for example, be between 15,000 revolutions per minute and 27,000 revolutions per minute.

[0112] The length L can for example be between 990 mm and 1600 mm for a high-pressure body comprising between ten and thirteen stages of compressor and turbine.

[0113] The average internal radius Ri can be between 140 mm and 210 mm.

[0114] When the high-pressure compressor 5 is sized to comply with formula (1), and in particular such that the ratio (Ri / L 2 / XN) x 10 9remaining below 4900, the dynamic behavior of the high-pressure rotor body is controlled, while still allowing for an efficient propulsion system 1, and in particular one with a high bypass ratio and a compact high-pressure compressor 5. The high-pressure shaft line 10, 10b can still include three modes of deformation within its operating range: however, the positioning of the 3 ème The mode is acceptable in that it is beyond a stabilized regime and the amplitude of the deformation modes can be sufficiently controlled by a person skilled in the art to ensure proper functioning of the propulsion system 1. In particular, a person skilled in the art can implement at least one of the following means to compensate for these deformation modes:

[0115] - Addition, on all or part of the bearings 12a, 12b of the high-pressure shaft 10, of a mode damper 25 comprising a pressurized oil film configured to dampen the mode change (viscous damping). This type of damping is generally known by the English term "squeeze film". The mode damper 25 is specifically configured to dampen the vibrations generated by the high-pressure shaft line 10, 10b. Taking into account the mean inside radius Ri of the high-pressure compressor 5, formula (1) specifically guarantees the possibility of integrating one or more mode damper 25, despite a lowered flow path. The mode damper 25 can, in particular, be mounted between a fixed part of the bearing 12a, 12b of the high-pressure shaft 10 and a stator part of the propulsion system 1 (intermediate casing 23 or inter-turbine casing 24, for example).Preferably, at least one mode 25 damper is placed on the front bearing 12a of the high-pressure shaft 10. When the high-pressure shaft 10 has a double bearing, the mode 25 damper is preferably placed on the bearing that carries the radial loads. When the double bearing is a double angular contact ball bearing, a bearing damper is preferably placed on each row of balls in the bearing.

[0116] - Addition of a flexible cage to each mode 25 bearing damper. Such a flexible cage can be mounted between the outer ring of the bearing 12a, 12b and a rigid support fixedly attached to a stator section (inter-compressor housing 23, inter-turbine housing 24, etc.) of the propulsion system 1 in order to control the stiffness of the bearings 12a, 12b and to fine-tune the position and amplitude of the deformation modes. The cage comprises, for this purpose, a generally cylindrical wall mounted between the rigid support and the outer ring of the bearing, and radially deformable columns to allow radial displacement of the generally cylindrical wall, and therefore of the bearing 12a, 12b, relative to the rigid support. Examples of bearings with a flexible cage and of a pressurized fluid film damper comprising a flexible cage are described in documents WO 2021 / 001610 and WO 2022 / 195198 on behalf of the Applicant.Damping the response of the high-pressure shaft line 10, 10b using mode dampers associated with flexible cages thus limits the resonance of the low-pressure shaft when its frequency reaches the first mode of deformation. This involves adapting the technology of the high-pressure turbine blades 7, for example, their cooling flow rate. Indeed, complying with formula (2) implies increasing the rotational speed of the high-pressure shaft 10.

[0117] - Modification of the shape of the moving blades 7a of the high-pressure turbine 7 in order to reduce the ratio between the chord at the tip and the chord at the root of the blade (blade distal tapering). For example, the ratio can be between 0.85 and 1.0. In particular, reducing this ratio makes it possible to increase the maximum possible rotational speed of the high-pressure shaft line 10, 10b).

[0118] - stiffen the high pressure shaft line 10, 10b by increasing the radius of the ferrules 10b between the moving blades 5a of the high pressure compressor 5 or the thickness of the high pressure shaft 10 between the high pressure compressor 5 and the high pressure turbine 7.

[0119] It should be noted that, when the ratio (Ri / L 2 / XN x 10 9If the pressure is less than 3700, the overall dimensions of the high-pressure compressor 5 become too small. In particular, the flow path is too low, making it impossible to integrate the bearings 12a, 12b, and the first-stage disc of the high-pressure compressor 5 without altering the diameter of the low-pressure shaft 11 (which affects its dynamic position), since this shaft is located within the high-pressure shaft 10. It is then necessary to modify the dynamic behavior of the high-pressure shaft line 10, 10b, by limiting the overall compression ratio of the propulsion system 1, increasing the hub-to-head ratio of the high-pressure compressor 5, or reducing the engine bypass ratio. However, this results in a loss of efficiency for both the high-pressure compressor 5 and the engine.

[0120] Furthermore, in one embodiment, the limiting speed of the high-pressure shaft 10 can be dimensioned to comply with the following formula:

[0121] < <

[0122]

[0123] where: Dg is the diameter of the blower rotor 9 and is expressed in meters;

[0124] BPR is the dilution rate of propulsion system 1; y is the adiabatic coefficient of air; and

[0125] OPR is the overall compression ratio of the propulsion system 1, i.e. the ratio between the pressure at the outlet of the high-pressure compressor 5 and the pressure at the inlet of the blower section 2.

[0126] When the limiting speed XN complies with formula (2), the limiting speed of the high-pressure shaft 10 is sufficiently high to obtain an efficient high-pressure compressor 5, despite its small size. Typically, the hub-to-head ratio of the high-pressure compressor can be between 0.49 and 0.53. Furthermore, the mean bore radius Rm_a of the high-pressure turbine, which corresponds to the arithmetic mean of the bore radii of the high-pressure turbine rotor 7, each bore radius being measured in a plane normal to the X-axis at the midpoint between the leading and trailing edges of the moving blades 7a of the high-pressure turbine rotor, is between 45 and 75 mm.

[0127] Furthermore, when the limit speed XN remains below the upper bound defined in formula (2), the fourth mode remains outside the operating range of the high-pressure shaft line 10, 10b.

[0128] The propulsion system 1 also respects formula (1), its modes of deformation are also controlled.

[0129] Comparative examples:

[0130] Engine 1 is a twin-body propulsion system comprising a shrouded fan section 2 corresponding to the current technical standard (at the date of filing of this application) which we seek to improve.

[0131] The engines 2, 3 and 4 are 1 twin-body propulsion systems comprising a shrouded fan section 2 conforming to the teaching of this application, the high-pressure shaft 10 of which is dimensioned so as to comply with formulas (1) and (2) defined above.

[0132] Engine 1 and engines 2, 3, and 4 all have the same number of stages in the low-pressure section (identical fan section, low-pressure compressor, and low-pressure turbine), the same bypass ratio and pressure ratio in the fan section, the same type of stages (only axial) in the high-pressure compressor, and a high-pressure shaft supported by exactly one ball bearing upstream of the combustion chamber and one roller bearing downstream of the combustion chamber. Furthermore, the upstream bearing of each engine includes pressurized fluid film damping and a flexible cage.

[0133]

[0134]

[0135]

[0136]

[0137]

[0138]

[0139]

[0140] The bypass ratio and pressure ratio of the blower in engine 2 are identical in all four engines. However, the overall pressure ratio and the high-pressure turbine inlet temperature of the engines conforming to one embodiment are higher, and their radial dimensions are smaller—this is particularly evident in the smaller bore radius of the high-pressure turbine (59 / 56 / 57 mm versus 66 mm in engine 1), the hub-to-head ratio, and the average internal radius of the high-pressure compressor (167 / 157 / 175 mm versus 185 mm for the average internal radius and 0.5 / 0.47 / 0.52 versus 0.55 for the hub-to-head ratio). The engines conforming to one embodiment are therefore more efficient than engine 1 (since their high-pressure casing is more efficient), while also having a smaller radial footprint.It should be noted, however, that the number of stages in the high-pressure compressor 5 has been increased in engines conforming to an embodiment in order to increase their compression ratio.

[0141] It is therefore clear from the comparative table that engine 1 is less efficient than the engines conforming to one embodiment. This is due to the fact that its high-pressure compressor does not comply with formula (1), since its Ri / L ratio 2 / XN x10 9is greater than 4900. The dynamic behavior of its high-pressure shaft therefore had to be controlled by making compromises in terms of efficiency (reduction of the overall compression ratio and the temperature at the inlet of the high-pressure turbine, increase of the hub-to-head ratio of the high-pressure turbine) and size (higher flow channel due to the larger average internal radius of the high-pressure compressor and bore of the high-pressure turbine). Furthermore, the limiting speed of the high-pressure shaft of the engines conforming to one embodiment is higher than that of engine 1, which further improves the performance of the high-pressure compressor, despite its small size.

[0142] In order to control the deformation modes of the high-pressure shaft of engines 2, 3 and 4, the following modifications were made compared to engine 1: a bearing mode damper was integrated into the downstream bearing, the position of the interstage high-pressure compressor ferrules was optimized for stiffness.

[0143] Finally, the motor 2, 3 and 4 also respecting formula (2), the mechanical load at the attachment of the blade disc 5a of the high pressure compressor is increased, and therefore the mass of the high pressure compressor 5 is reduced (via the decrease in the external radius of the high pressure compressor 5).

Claims

DEMANDS 1. Aeronautical propulsion system (1) comprising: - a blower section comprising a blower rotor (9) connected to a blower shaft (20); - a first turbine (8) configured to drive the blower rotor (9) via a first shaft (11) around an axis of rotation (X); - a reduction mechanism (19) coupling the first shaft (11) and the blower shaft (20) in order to drive the blower shaft (20) at a rotational speed lower than the rotational speed of the first shaft (11); - a second turbine (7) configured to drive a compressor (5) via a second shaft (10) around the axis of rotation (X); and - bearings (12a, 12b) configured to center the second shaft (10) relative to the axis of rotation (X), the bearings (12a, 12b) comprising a front bearing (12a) extending upstream of a combustion chamber of the propulsion system (1) and a rear bearing (12b) extending downstream of the combustion chamber, in which a dilution ratio of the propulsion system (1) is greater than or equal to 8, a pressure ratio of the blower section is less than or equal to 1.45, and the compressor (5) is sized to comply with the following formula: Ri / / r 3700 < - — 2 X 10 9 < 4900 XN where: XN is the limiting speed of the second shaft (10), in revolutions per minute (rpm); Ri is an average internal radius of the compressor (5), in meters (m); and L is an axial distance between a center of gravity (Gi2a) of the front bearing (12a) and a center of gravity (Gi2t>) of the rear bearing (12b), in meters (m).

2. Propulsion system (1) according to claim 1, wherein each bearing (12a, 12b) is associated with a bearing mode damper (25).

3. Propulsion system according to claim 2, wherein the bearing mode dampers (25) comprise damping by pressurized fluid film.

4. Propulsion system one of claims 2 and 3, wherein the bearing mode dampers (25) comprise a deformable cage mounted between a ring of each bearing and a stator part (23, 24, 26, 27) of the propulsion system (1).

5. Propulsion system according to any one of claims 1 to 4, wherein the limiting speed of the second shaft (11) further complies with the following formula: 2370 3000 where: Dg is a diameter of a rotor of the blower section, measured in a plane normal to the axis (X) of rotation at the level of an intersection between a vertex (21) and a leading edge (22) of the blades (14) of the blower rotor (9), and is expressed in meters; BPR is a dilution rate of the propulsion system (1) and is measured when the propulsion system 1 is stationary, uninstalled, in takeoff regime; y is the adiabatic coefficient of air; and OPR is an overall compression ratio of the propulsion system (1) and is measured when the propulsion system 1 is stationary, uninstalled, in takeoff regime.

6. Propulsion system (1) according to any one of claims 1 to 5, wherein the compressor (5) has a hub-to-head ratio between 0.45 and 0.

55.

7. Propulsion system (1) according to any one of claims 1 to 6, wherein the compressor (5) comprises at least eight stages and at most eleven stages.

8. Propulsion system (1) according to any one of claims 1 to 7, wherein an average radius of an additional turbine bore (7) is at most equal to 75 mm.

9. Propulsion system (1) according to any one of claims 1 to 8, wherein a reduction ratio of the reduction mechanism is greater than or equal to 2.5 and less than or equal to 11.

10. Propulsion system (1) according to any one of claims 1 to 9, wherein a high-pressure compressor pressure ratio is between 15 and 27.

11. Aircraft comprising at least one propulsion system (1) according to any one of claims 1 to 10 fixed to the aircraft by means of a mast.

12. Method for dimensioning a propulsion system (1) comprising a reduction mechanism (19) coupling a first shaft (11) and a blower rotor (9) to drive the blower rotor (9) at a speed lower than the speed of the first shaft (11), a second shaft (10) coupling a turbine (7) to a compressor (5) to drive the compressor (5) at a speed higher than the speed of the first shaft (11), and bearings (12a, 12b) configured to center the second shaft (10) with respect to the axis of rotation (X), the bearings (12a, 12b) comprising a front bearing extending upstream of a combustion chamber of the propulsion system (1) and a rear bearing extending downstream of the combustion chamber, the compressor (5) being sized to comply with the following formula: Ri / / r 3700 < - — 2 10 9 < 4900 XN where: XN is the limiting speed of the second shaft (10), in revolutions per minute (rpm); Ri is an average internal radius of the compressor (5), in meters (m); and L is an axial distance between a center of gravity (Gi2a) of the front bearing (12a) and a center of gravity (Gi2t>) of the rear bearing (12b), in meters (m).

13. A dimensioning method according to claim 12, wherein the limiting speed of the second shaft (11) further complies with the following formula: 2370 3000 where: Dg is a diameter of a rotor of the blower section, measured in a plane normal to the axis (X) of rotation at the level of an intersection between a vertex (21) and a leading edge (22) of the blades (14) of the blower rotor (9), and is expressed in meters; BPR is a dilution rate of the propulsion system (1) and is measured when the propulsion system 1 is stationary, uninstalled, in takeoff regime; y is the adiabatic coefficient of air; and OPR is an overall compression ratio of the propulsion system (1) and is measured when the propulsion system 1 is stationary, uninstalled, in takeoff regime.

14. Method for manufacturing an aeronautical propulsion system comprising the following steps: - to dimension the aeronautical propulsion system in accordance with the method according to one of claims 12 and 13; and - to manufacture the aeronautical propulsion system.

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