Aeronautical propulsion system featuring a fan with improved mechanical strength

Abstract

The present description concerns an aeronautical propulsion system (1) comprising a fan disk (13) and a plurality of blades (14) attached and fixed in recesses (29) of the disk (13), each recess (29) having a minimum throat area (Sc) defined by the following formula: where: D is the diameter of the fan rotor (9); R is a hub-to-head ratio of the fan rotor (9); ω is a rotational speed of the fan rotor (9); n is the number of blades (14) in the fan rotor (9); s / c is a pitch relative to the tip (21) of the fan blades (14); and A = 1.26; B1 = 101.5 mm².tr².min-2; B2 = 75.1 mm².rpm.min⁻² and E = 57,970 mm⁴.rpm.min⁻². Figure for the abbreviation: Fig. 1

Description

Title of the invention: Aeronautical propulsion system having a fan with improved mechanical strength Technical field

[0001] The present application relates generally to the field of propulsion systems, and more particularly to aeronautical propulsion systems having a high, or even very high, dilution ratio. 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 comprise a low-pressure compressor and a high-pressure compressor, a combustion chamber and a turbine section which may comprise 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 appropriate, the low-pressure compressor are driven in rotation by the low-pressure turbine via a low-pressure shaft.

[0003] Technological research efforts have already made it possible to significantly improve the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, 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.

[0004] Thus, in order to improve the propulsive efficiency of the propulsion system and to reduce its specific consumption as well as the noise emitted by the fan section, propulsion systems have been proposed having a high BPR (bypass ratio in English, corresponding to the ratio between the flow rate of the secondary air flow and the flow rate of the primary air flow). To achieve such dilution ratios, the fan section can be decoupled from the low-pressure turbine, thus making it possible to independently optimize their respective rotational speeds. Generally, the 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 rotor of the fan section is then driven by the low-pressure shaft via the reduction mechanism at a rotational speed lower than that of the low-pressure shaft.

[0005] This decoupling makes it possible to reduce the rotation speed of the fan rotor. By Elsewhere, in order to reduce the fan rotor pressure ratio and improve the bypass ratio of the propulsion system, it has been proposed to increase the fan rotor diameter (and thus reduce the peripheral speed of the fan blades). The consequence is a reduction in centrifugal forces in the fan rotor. However, increasing the fan diameter has a negative impact on the mass and drag of the fan section, and by extension the propulsion system, and increases the stresses applied by the fan blades to the fan disc. However, the failure of the attachment of a fan blade to the fan disc causes a cascade of failures of the adjacent blades, until the fan disc fails. EXPOSED

[0006] An aim of the present application is to optimize the performance of the propulsion system in terms of specific consumption, mass and drag, while guaranteeing the mechanical strength of the fan rotor.

[0007] To this end, according to a first aspect, an aeronautical propulsion system is proposed comprising: - a drive shaft movable in rotation around an axis of rotation; - a fan shaft; - a fan section comprising a fan rotor rotated by the fan shaft, the fan rotor comprising a disc and a plurality of blades attached and fixed in cells of the disc; - a reduction mechanism coupling the drive shaft and the fan shaft in order to drive the fan shaft at a rotational speed lower than the rotational speed of the drive shaft; in which each cell has a neck section defined by the following formula: 4 „ (DMLR2jx^AH*10'i2+E A*-—i—----< Se < A*-—2------ where: Sc is the neck section and corresponds to a minimum section of a neck of a disc alveolus, measured in a plane extending circumferentially relative to the axis of rotation, in square millimeters; D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades, and is expressed in millimeters; R is a fan rotor hub-to-head ratio; co is a rotational speed of the fan rotor, when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level, is expressed in revolutions per minute; n is the number of blades in the fan rotor; s / c is a relative pitch at the fan blade tip level and is equal to the ratio of an inter-blade pitch to an axial chord of a fan blade; and A = 1.26; Bi = 101.5 mm2.tr2.min2; B2 = 75.1 mm2.tr2.min2 and E = 57,970 mm4.tr2.min -2

[0008] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following, taken individually or in combination: - the fan rotor hub-to-head ratio is between 0.22 and 0.32; - the fan rotor is sized so that a maximum moment that can be applied by a fan blade on the disc before disc rupture complies with the following formula: F}x^^ xR2-G <M<F^x^fi xR2-G 1 OH3 - 6 / 1- where: M is the maximum moment in Nm and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level; and Fj = 411 N; F2 = 555 N and G = 1203 Nm - the diameter of the fan rotor is between 80 inches and 185 inches inclusive, for example between 85 inches and 120 inches inclusive, for example of the order of 90 inches; - a reduction ratio of the reduction mechanism is greater than or equal to 2.5 and less than or equal to 11, for example greater than or equal to 2.7 and less than or equal to 3.5; - a dilution ratio of the propulsion system is greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive; - a peripheral speed at the tip of the fan rotor blades, when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level, is between 260 m / s and 400 m / s; - the fan blades are fixed in rotation relative to the disc - the propulsion system further comprises a drive turbine and a compressor connected directly by the drive shaft - the drive turbine comprises at least 3 and at most 5 stages; - the compressor has at least 2 and at most 4 stages - the propulsion system further comprises a high-pressure turbine and a high-pressure compressor connected via a high-pressure shaft, the high-pressure shaft rotating faster than the drive shaft, the high-pressure turbine being two-stage; and / or - the high pressure compressor has at least 8 and at most 11 stages.

[0009] According to a second aspect, there is provided an aircraft comprising at least one propulsion system according to the first aspect fixed to the aircraft by means of a mast.

[0010] According to a third aspect, there is provided a method of dimensioning or manufacturing a propulsion system comprising a reduction mechanism coupling a drive shaft and a fan rotor to drive the fan rotor at a speed lower than a speed of the drive shaft, in which the fan rotor is dimensioned so that each cell has a neck section defined by the following formula: where: Sc is the neck section and corresponds to a minimum section of a neck of a disc alveolus, measured in a plane extending circumferentially relative to the axis of rotation, in square millimeters (mm2); D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between a tip and a leading edge of the fan rotor blades, and is expressed in millimeters (mm); R is a fan rotor hub-to-head ratio; co is a rotational speed of the fan rotor, when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level, is expressed in revolutions per minute; n is the number of blades in the fan rotor; and s / c is a relative pitch at the fan blade tip level and is equal to the ratio of an inter-blade pitch to an axial chord of a fan blade; and A = 1.26; Bi = 101.5 mm2.tr2.min2; B2 = 75.1 mm2.tr2.min2 and E = 57,970 mm4.tr2.min2.

[0011] Optionally, the fan rotor is dimensioned so that a maximum moment that can be applied by a fan blade to the disc before disc rupture complies with the following formula: F x <F2x^K xR2-G 1 <)B OB' where M is the maximum moment in Nm and is measured when the propulsion system is stationary in take-off mode in a standard atmosphere and at sea level; and Fj = 411 N; F2 = 555 N and G = 1,203 Nm DESCRIPTION OF FIGURES

[0012] Other characteristics, aims and advantages will emerge from the following description, which is purely illustrative and non-limiting, and which must be read in conjunction with the appended drawings in which:

[0013] [Fig.l] is a schematic, partial and sectional view of an example of a propulsion system according to a first embodiment;

[0014] [Fig. 2] is a partial perspective view of a fan rotor of a propulsion system according to one embodiment;

[0015] [Fig. 3] is a schematic sectional view of an example of a reduction mechanism according to a first variant;

[0016] [Fig.4] is a schematic sectional view of an example of a reduction mechanism according to a second variant;

[0017] [Fig.5] is an example of an aircraft that may include at least one propulsion system according to one embodiment;

[0018] [Fig. 6] is a flowchart illustrating exemplary steps of a sizing or manufacturing method according to one embodiment.

[0019] Throughout the figures, similar elements bear identical references. DETAILED DESCRIPTION

[0020] A propulsion system 1 has a main direction extending along a longitudinal axis X and comprises, from upstream to downstream in the direction of flow of the gases in the propulsion system 1 when it is in operation, a fan 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 is here an aeronautical propulsion system 1 configured to be fixed to an aircraft 100 via a pylon (or mast).

[0021] The compressor section 4, 5 comprises a succession of stages each comprising a wheel of moving blades (rotor) 4a, 5a rotating in front of a wheel of fixed blades (stator) 4b, 5b. The turbine section 7, 8 also comprises a succession of stages each comprising a wheel of fixed blades (stator) 7b, 8b behind which a wheel of moving blades (rotor) 7a, 8a rotates.

[0022] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, in correspondence with the rotation of the shafts of the gas generator, and a radial direction is a direction perpendicular to this axis X and passing through it. Furthermore, the circumferential (or lateral, or even tangential) direction corresponds to a direction perpendicular to the longitudinal axis X and not passing through it. Unless otherwise specified, internal (respectively, interior) and external (respectively, exterior), respectively, are used with reference to a radial direction so that the internal part or face of an element is closer to the axis X than the external part or face of the same element.

[0023] In operation, an air flow F entering the propulsion system 1 is divided between a primary air flow F1 and a secondary air flow F2, which circulate from upstream to downstream in the propulsion system 1.

[0024] The secondary air flow F2 (also called “bypass air flow”) flows around the primary body 3. The secondary air flow F2 allows the periphery of the primary body 3 to be cooled and is used to generate the majority of the thrust provided by the propulsion system 1.

[0025] The primary air flow F1 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 oxidant, and the turbine section 7, 8. The passage of the primary air flow F1 through the turbine section 7, 8 receiving energy from the combustion chamber 6 causes rotation of the rotor of the turbine section 7, 8, which in turn drives rotation of the rotor of the compressor section 4, 5 as well as a rotor part 9 of the fan section 2.

[0026] In a twin-body propulsion system 1, the compressor section 4, 5 may comprise a low-pressure compressor 4 and a high-pressure compressor 5. The turbine section 7, 8 may comprise a high-pressure turbine 8 and a low-pressure turbine 7. The rotor of the high-pressure compressor 5 is rotated by the rotor of the high-pressure turbine 8 via a high-pressure shaft 10. The rotor of the low-pressure compressor 4 and the rotor portion 9 of the fan section 2 are rotated by the rotor of the low-pressure turbine 7 via a low-pressure shaft 11. Thus, the primary body 3 comprises a high-pressure body comprising the high-pressure compressor 5, the high-pressure turbine 8 and the high-pressure shaft 10, and a low-pressure body comprising the fan section 2, the low-pressure compressor 4, the low-pressure turbine 7 and the low-pressure shaft 11.The rotational speed of the high-pressure spool is greater than the rotational speed of the low-pressure spool. In a triple-spool propulsion system 1, the turbine section 7, 8 further comprises an intermediate turbine, positioned between the high-pressure turbine 8 and the low-pressure turbine 7 and configured to drive the rotor of the low-pressure compressor 4 via an intermediate shaft. The fan 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.

[0027] The low pressure shaft 11 is generally housed, over a section of its length, in 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, i.e. driven in the same direction around the longitudinal axis X. Alternatively, the low pressure shaft 11 and the high pressure shaft are counter-rotating, i.e. driven in opposite directions around the longitudinal axis X. Where appropriate, 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.

[0028] The fan section 2 comprises at least the fan rotor 9 capable of being driven in rotation relative to a fan casing 12 by the turbine section 7, 8. Each fan rotor 9 comprises a disc 13 and blades 14 extending radially from the disc 13. The blades 14 of each rotor 9 are attached and fixed in cells 29 formed in the disc 13 and are fixed relative to the disc 13 (fixed setting). Each blade 14 has a blade root 30 configured to be attached and fixed in an associated cell 29 of the disk 13, an aerodynamic blade 31 capable of extending into the air flow F and a stilt which connects the blade root 30 and the blade 31. The stilt can in particular extend between the outlet of the cell 29 and an aerodynamic platform 32 of the fan blade 14.

[0029] Each cell 29 extends between two teeth 33 of the disc 13 and forms a cavity intended to receive the blade root 30 which is open on an external face of the disc. A cell 29 thus has a bottom which extends opposite an internal face of the blade root 30 and two lateral flanks which extend radially from the bottom, on either side of the intrados and extrados faces of the blade root 30, to the external face of the disc. The flanks converge near the external face of the disc to form a neck. The section of the cavity at the neck 34 (i.e. in a plane circumferential to the axis of rotation X which intersects the external face of the disk) is less than its section at mid-height of the flanks (i.e. in a plane circumferential to the axis of rotation X which intersects the flanks at mid-height of the cavity) in order to radially retain the blade root 30 in the cell 29.The blade 14 then extends through the opening so that the blade 31 and Péchasse are located outside the cell 29 while the root 30 is housed in the cell 29.

[0030] Each fan blade 14 has a leading edge 14a and a trailing edge 14b. The leading edge 14a is configured to extend opposite the flow of gases entering the fan rotor 9. It corresponds to the front part of an aerodynamic profile which faces the air flow and which divides the air flow into a pressure flow and an extrados flow. The trailing edge 28b corresponds to the rear part of the aerodynamic profile, where the pressure and extrados flows meet. The fan blade 14 further has an axial chord, which corresponds to the straight line segment parallel to the axis X of rotation which connects a point of intersection between the trailing edge 14b and the tip 21 and the leading edge 14a of the blade.

[0031] The fan section 2 may further comprise 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 straighten the secondary air flow F2 which flows out of the fan rotor 9. The blades 17 of the fan stator 18 are for example fixed relative to the hub 18.

[0032] In order to improve the propulsive efficiency of the propulsion system 1 and to reduce its specific consumption as well as the noise emitted by the fan section 2, the propulsion system 1 has a high bypass ratio. By high bypass ratio, here is meant a bypass 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 air flow F2 and the mass flow rate of the primary air flow F1 are measured when the 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 will be noted that, in the present application, the parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined under these conditions.By "not installed" it will be understood here that the measurements are carried out when the propulsion system 1 is in a test bench (and not installed on an aircraft 100), the measurements then being simpler to carry out. The distances (length, radius, diameter, etc.) are on the other hand measured at ambient temperature (approximately 20°C) when the propulsion system 1 is cold, that is to say when the propulsion system 1 has been stopped for a sufficient period for the parts of the propulsion system 1 to be at ambient temperature, it being understood that these dimensions vary little compared to the conditions in which the propulsion system 1 is in take-off mode.

[0033] The fan rotor 9 is decoupled from the low-pressure shaft 11 using a reduction mechanism 19, placed between an upstream end of the low-pressure shaft 11 and the fan rotor 9, in order to independently optimize their respective rotational speed. In this case, the propulsion system 1 further comprises an additional shaft, called 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 therefore 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.

[0034] This decoupling makes it possible to reduce the rotational speed and the pressure ratio of the fan rotor 9 and to increase the power extracted by the low-pressure turbine 7. Indeed, the overall efficiency of the propulsion systems is conditioned to the first order by the propulsive efficiency, which is favorably influenced by a minimization of the variation in kinetic energy of the air when passing through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow rate generating the propulsive force is constituted by the secondary air flow F2 of the propulsion system 1, the kinetic energy of the secondary air flow F2 being mainly affected by the compression that the secondary air flow F2 undergoes when passing through the fan section 2. The propulsive 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 propulsive efficiency. In order to optimize the propulsive efficiency of the propulsion system 1, the pressure ratio of the fan, which corresponds to the ratio between the average pressure at the outlet of the fan stator 17 (or, in the absence of a stator, of the fan rotor 9) and the average pressure at the inlet of the fan 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 path at the inlet of the fan rotor 9 to the tip 21 of the fan blade 14).

[0035] The propulsion system 1 is configured to provide a thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,2411 N), for example between 20,000 Ibf (88,964 N) and 35,000 Ibf (15,5688 N).

[0036] The fan section 2 comprises a fan casing 12, the fan rotor 9 being housed in the fan casing 12. The fan section 2 comprises in particular a fan rotor 9 extending upstream of a fan stator. The vanes 16 of the fan stator 17 are then generally called outlet vanes (“Outlet Guide Vane” or “OGV” in English) and have a fixed setting relative to the hub of the fan stator. 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 speed at the tip 21 of the vanes of the fan rotor 9 may also be between 260 m / s and 400 m / s. The blades 14 of the fan rotor 9 may be fixed or have variable pitch. The fan pressure ratio may then be between 1.20 and 1.45.

[0037] The reduction mechanism 19 may comprise, for example, a reduction mechanism 19 with an epicyclic gear train, 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 may be of the planetary (“star” in English) type ([Fig.3]) and comprise a sun gear 19a (input of the reduction mechanism 19), centered on an axis X of rotation of the reduction mechanism 19 (generally confused with the longitudinal axis X) and configured to be driven in rotation by the low pressure shaft 11, a crown 19b (output of the reduction mechanism 19) coaxial with the sun gear 19a and configured to drive in rotation the fan shaft 20 around the axis X of rotation, and a series of satellites 19c distributed circumferentially around the axis X of rotation between the sun gear 19a and the crown 19b, each satellite 19c being meshed internally with the sun gear 19a and externally with the crown. 19b. The series of satellites 19c is mounted on a planet carrier 19d which is fixed relative to a stator part 19e of the propulsion system 1, for example relative to a casing of the compressor section 4, 5. According to a second variant, the reduction mechanism 19 may be of the epicyclic type (“planetary” in English) ([Fig. 4]), in which case the ring gear 19b is fixedly mounted on the stator part 19e of the propulsion system 1 and the fan shaft 20 is driven in rotation by the planet carrier 19d which is therefore movable in rotation relative to a stator part 19e of the propulsion system 1, for example relative to a casing of the compressor section 4, 5).

[0038] Whatever the configuration of the reduction mechanism 19, the diameter of the crown 19b and of the planet carrier 19d are greater than the diameter of the sun gear 19a, so that the rotation speed of the fan rotor 9 is lower than the rotation speed of the low pressure shaft 11.

[0039] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11, for example greater than or equal to 2.7 and less than or equal to 3.5, typically around 3.0.

[0040] The limit speed (redline speed) of the low pressure shaft 11, which corresponds to the absolute maximum speed likely to be encountered by the low pressure shaft 11 during the entire flight, is between 8500 rpm and 12000 rpm, for example between 9000 rpm and 11000 rpm. The limit speed corresponds to the absolute maximum speed likely to be encountered by the low pressure shaft 11 during the entire flight (according to the European certification regulation EASA CS-E 740 (or according to the American certification regulation 14-CFR Part 33.87)). The limit speed corresponds to the maximum rotation speed when the propulsion system is healthy. It is therefore likely to be reached by the low pressure shaft 11 in flight conditions. This limit speed is part of the data declared in the engine certification (type certificate data sheet).Indeed, this rotation speed is usually used as a reference speed for the dimensioning of propulsion systems 1 and in certain certification tests (such as blade loss or rotor integrity tests).

[0041] 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 fan rotor 9 (measured at the level of the foot 30 of the fan rotor 9), may be greater than or equal to 40 and less than or equal to 70, for example greater than or equal to 44 and less than or equal to 55.

[0042] In order to optimize the performance of the propulsion system 1 in terms of specific consumption, mass and drag, while guaranteeing the mechanical strength of the fan rotor 9, each cell 29 has a neck section defined by the following formula (1): (D4^i-^)x^li)^i2+E (1) A.--------_____------s Scs A-->—------ where: Sc is the neck section and corresponds to a minimum section of a neck 34 of a cell 29 of the disc 13, measured in a plane extending circumferentially relative to the axis of rotation X, in square millimeters (mm2); D is the diameter of the fan rotor 9, measured in a plane normal to the axis of rotation X at an intersection between the tip 21 and the leading edge 22 of the blades 14 of the fan rotor 9, and is expressed in millimeters (mm); R is a hub-to-head ratio of the fan rotor 9. It corresponds to the ratio between the inner radius Ri and the outer radius Re of the fan rotor 9. The inner radius Ri is equal to the distance between the axis of rotation X and the point of intersection between the leading edge 22 and the surface which radially delimits on the inside the flow path at the inlet of the fan rotor 9 (and corresponds to the point of connection of the leading edge 22 with the aerodynamic surface of a platform 32 of the fan rotor 9). The outer radius Re is equal to the distance between the axis of rotation X and the point of intersection between the leading edge 22 and the tip 21 of the fan blades (and corresponds to half the fan diameter D); co is a rotational speed of the fan rotor 9, when the propulsion system 1 is stationary in take-off mode in a standard atmosphere and at sea level, is expressed in revolutions per minute;

[0043] n is the number of blades 14 in the fan rotor 9; s / c is a pitch relative to the tip 21 of the fan blades 14, which is equal to the ratio between an inter-blade pitch of the fan blades 14 and the axial chord of the fan blades 14 and is dimensionless, knowing that the inter-blade pitch is equal to ^2; and A = 1.26; Bi = 101.5 mm2.tr2.min2; B2 = 75.1 mm2.tr2.min2 and E = 57,970 mm4.tr2.min2.

[0044] When the neck section Sc complies with formula (1), the neck section 34 of the cells 29 (and therefore the distance between two adjacent teeth 33 of the disk 13) is limited, which makes it possible to guarantee the mechanical strength of the fan blades 14 throughout the flight, even in the event of blade loss (and therefore significant unbalance on the fan shaft 20), while allowing the optimization of the hub-head ratio of the fan rotor 9 and therefore the performance of the propulsion system 1, at iso-diameter D of the fan rotor 9. Note that the optimization of the hub-head ratio also makes it possible to reduce the specific consumption of the propulsion system 1. Thus, a fan disk whose neck section complies with formula (1) presents a lower risk of rupture of the disk 13 in the event of blade loss (in particular at the teeth 33 of the disk 13) while guaranteeing the efficiency of the blower 2.

[0045] In particular, at iso-diameter of the fan rotor 9, the reduction in the section of col Sc in order to comply with formula (1) makes it possible to reduce the internal radius Ri of the fan rotor 9. However, the reduction of the internal radius Ri, at fan iso-diameter, makes it possible to reduce the hub-to-head ratio of the fan rotor 9, knowing that the lower the hub-to-head ratio, the more efficient the fan rotor 9 is and therefore the more efficient the propulsion system 1 is. For example, the hub-to-head ratio of the fan rotor 9 is between 0.22 and 0.32, for example between 0.22 and 0.30, in particular when the blades 14 are fixed pitch since it is then not necessary to integrate a pitch change mechanism within the fan disc 13.

[0046] In order to obtain both a neck section Sc complying with formula (1), a low hub-head ratio R and a fan disc 13 which is sufficiently strong from a mechanical point of view, the fan disc 13 may in particular be made from materials having a mechanical strength greater than the mechanical strength of the titanium alloys usually used. The fan disc 13 may for example be made from a nickel-based superalloy such as Inconel^ 718.

[0047] For a section Sc which complies with formula (1), the diameter D of the fan rotor can then be between 80 inches (2,032 mm) and 185 inches (4,699 mm) inclusive, for example between 85 inches (2,159 mm) and 120 inches (3,048 mm) inclusive, for example of the order of 90 inches (2,286 mm), which makes it possible to integrate the propulsion system 1 in a conventional manner, in particular under the wing of an aircraft 100.

[0048] In order to further improve the aerodynamic performance of the propulsion system, the fan rotor 9 is dimensioned so that the maximum moment M that can be applied by a fan blade 14 to the tooth 33 of the adjacent disc 13 before the tooth 33 breaks complies with the following formula: F x x.R2-G <m<fkr2-g^ where: M is the maximum moment, in Nm, and is measured when the propulsion system 1 is stationary in take-off mode in a standard atmosphere and at sea level; and F! = 411N; F2 = 555 N and G = 1,203 Nm

[0049] When the fan rotor 9 is dimensioned so as to mechanically resist a maximum moment M as defined in formula (1), it is possible to reduce the hub-head ratio of the fan rotor 9, to iso-diameter D of the fan rotor 9, while guaranteeing the mechanical strength of the fan blades 14 throughout the flight. The centrifugal forces applied to the fan disk 13, in particular the force applied by the blade root 30 on the teeth 33 of the disk 13, are certainly greater: however, the use of the materials described above makes it possible to mechanically resist these centrifugal forces, in particular in the event of blade loss.

[0050] The choice of the number of blades 14 in the fan rotor 9 also has an impact on the maximum moment that can be applied by a blade 14; in particular, with the same hub-to-head ratio, reducing the number of blades 14 makes it possible to increase the maximum moment. Alternatively, the number of blades 14 may be selected to allow the hub-to-head ratio to be reduced (and therefore the efficiency of the fan rotor 9 to be improved) without impacting the maximum moment M that can be applied by a fan blade 14. For example, the fan rotor 9 may comprise at least 12 fan blades 14 and at most 24 fan blades 14, for example at least 16 fan blades 14 and at most 22 fan blades 14.

[0051] 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.

[0052] A double-body propulsion system 1 whose neck section Sc complies with formula (1) and, where appropriate, dimensioned so that the maximum moment complies with formula (2), may in particular comprise a two-stage high-pressure turbine 8, a high-pressure compressor 5 comprising at least eight stages and at most eleven stages, a low-pressure turbine 7 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.

Claims

Claims

1. Aeronautical propulsion system (1) comprising: - a drive shaft (11) movable in rotation around an axis of rotation (X); - a fan shaft (20); - a fan section (2) comprising a fan rotor (9) driven in rotation by the fan shaft (20), the fan rotor (9) comprising a disc (13) and a plurality of blades (14) attached and fixed in cells (29) of the disc (13); - a reduction mechanism (19) coupling the drive shaft (11) and the fan shaft (20) in order to drive the fan shaft (20) at a rotational speed lower than the rotational speed of the drive shaft (11); in which each cell (29) has a neck section defined by the following formula: where: Sc is the neck section and corresponds to a minimum section of a neck (34) of a cell (29) of the disc (13), measured in a plane extending circumferentially relative to the axis of rotation (X), in square millimeters (mm2); D is the diameter of the fan rotor (9), measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in millimeters (mm); R is a fan rotor hub-to-head ratio (9); co is a rotational speed of the fan rotor (9), when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level, is expressed in revolutions per minute (rpm); n is the number of blades (14) in the fan rotor (9); s / c is a relative pitch at the tip (21) of the fan blades (14) and is equal to the ratio between an inter-blade pitch and an axial chord of a fan blade (14); and A = 1.26; Bi = 101.5 mm2.tr2.min2; B2 = 75.1 mm2.tr2.min2 and E = 57,970 mm4.tr2.minA

2. A propulsion system (1) according to claim 1, wherein the ratio fan rotor head hub (9) is between 0.22 and 0.

32.

3. Propulsion system (1) according to one of claims 1 and 2, wherein the fan rotor (9) is dimensioned so that a maximum moment that can be applied by a fan blade (14) on the disc (13) before rupture of the disc (13) complies with the following formula: F XR2-G <M<FxR2-G * 6«-' on3 où : M est le moment maximal en Newton mètres (N.m) et est mesuré lorsque le système propulsif (1) est stationnaire en régime de décollage dans une atmosphère standard et au niveau de la mer; et Fi = 411 Newton (N) ; F2 = 555 Newton (N) et G = 1 203 Newton mètres (N.m).

4. Propulsion system (1) according to one of claims 1 to 3, in which the diameter of the fan rotor (9) is between 80 inches (2,032 mm) and 185 inches (4,699 mm) inclusive, for example between 85 inches (2,159 mm) and 120 inches (3,048 mm) inclusive, for example of the order of 90 inches (2,286 mm).

5. Propulsion system (1) according to one of claims 1 to 4, wherein a reduction ratio of the reduction mechanism (19) is greater than or equal to 2.5 and less than or equal to 11, for example greater than or equal to 2.7 and less than or equal to 3.

5.

6. Propulsion system (1) according to one of claims 1 to 5, in which a dilution ratio of the propulsion system (1) is greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive.

7. Propulsion system (1) according to one of claims 1 to 6, in which a peripheral speed at the tip (21) of the blades (14) of the fan rotor (9), when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level, is between 260 m / s and 400 m / s.

8. Propulsion system (1) according to one of claims 1 to 7, in which the fan blades (14) are fixed in rotation relative to the disc (13).

9. Propulsion system (1) according to one of claims 1 to 8, further comprising a drive turbine (8) and a compressor (4) directly connected by the drive shaft (11).

10. Propulsion system (1) according to claim 9, wherein the drive turbine (8) comprises at least 3 and at most 5 stages.

11. Propulsion system (1) according to one of claims 9 and 10, in which the compressor (4) comprises at least two and at most four stages.

12. Propulsion system (1) according to one of claims 1 to 11, further comprising a high pressure turbine (7) and a high pressure compressor (5) connected via a high pressure shaft (10), the high pressure shaft (10) rotating faster than the drive shaft (11), the high pressure turbine (7) being two-stage.

13. A propulsion system (1) according to claim 12, wherein the high pressure compressor (5) comprises at least eight and at most eleven stages.

14. Aircraft (100) comprising at least one propulsion system (1) according to one of claims 1 to 13 fixed to the aircraft via a mast.

15. A method of dimensioning a propulsion system (1) comprising a reduction mechanism (19) coupling a drive shaft (11) and a fan rotor (9) for driving the fan rotor (9) at a speed lower than a speed of the drive shaft (11), wherein the fan rotor (9) is dimensioned such that each cell (29) has a neck section defined by the following formula: 4 \ V / L Ç \ ) Ini 1 <- >______________ Bj — B? where: Sc is the neck section and corresponds to a minimum section of a neck (34) of a cell (29) of the disk (13), measured in a plane extending circumferentially relative to the axis of rotation (X), in square millimeters (mm2);D is the diameter of the fan rotor (9), measured in a plane normal to the axis of rotation (X) at an intersection between a tip (21) and a leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in millimeters (mm); R is a hub-to-head ratio of the fan rotor (9); co is a rotational speed of the fan rotor (9), when the propulsion system (1) is stationary in takeoff mode in a standard atmosphere and at sea level, is expressed in revolutions per minute (rpm); n is the number of blades (14) in the fan rotor (9); and s / c is a relative pitch at the tip (21) of the fan blades (14) and is equal to the ratio of an inter-blade pitch to an axial chord of a fan blade (14); and; A = 1.26; Bi = 101.5 mm2.tr2.min2; B2 = 75.1 mm2.tr2.min2 and E = 57,970 mm4.tr2.min2.

16. A sizing method according to claim 15, wherein the fan rotor (9) is sized such that a maximum moment that can be applied by a fan blade (14) to the disc (13) before rupture of the disc (13) complies with the following formula: xR 2 -G <m<f2x^f xr^-gwhere M is the maximum moment in Newton meters (Nm) and is measured when the propulsion system (1) is stationary in take-off mode in a standard atmosphere and at sea level; and Fi = 411 Newton (N); F2 = 555 Newton (N) and G = 1,203 Newton meters (Nm).

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