Optimized fan section of an aeronautical propulsion system

Abstract

The invention relates to an aeronautical propulsion system comprising a fan section comprising: - a casing, - a rotor comprising a blade (14) extending over a blade height (47) from a blade root (37) to a blade tip (21), the blade (14) having: - an axial chord at the blade root, - an axial chord at the blade tip, and - a mean axial chord, - an axial chord parameter corresponding to a ratio of a product of the mean axial chord times the axial chord at the blade root to a product of the blade height (47) times the axial chord at the blade tip, and - a speed parameter corresponding to a difference of a product of a peripheral speed at the blade tip expressed in m / s times the value 3.263x10-3 m / s and the value 7.741x10-1, the blade (14) being configured such that the axial chord parameter is greater than or equal to the speed parameter.

Description

[0001] DESCRIPTION

[0002] OPTIMIZED FAN SECTION OF AN AERONAUTICAL PROPULSION SYSTEM

[0003] FIELD OF INVENTION

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

[0005] STATE OF THE ART

[0006] An aeronautical propulsion system generally comprises, from upstream to downstream in the direction of gas flow, a blower 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 a high pressure turbine and a low pressure turbine.

[0007] When the propulsion system is in operation, the high-pressure compressor is driven in rotation by the high-pressure turbine via a high-pressure shaft. The blower and, if equipped, the low-pressure compressor are driven in rotation by the low-pressure turbine via a low-pressure shaft.

[0008] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by different countries. In particular, an ambitious standard applies to both new types of aircraft and those already in service, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively working for several years now to contribute to the fight against climate change.

[0009] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into account the factors impacting all phases of design and development in order 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.

[0010] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes that minimize greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0011] This sustained research and development work focuses on new generations of aircraft engines, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and, as essential complements to technological progress, aviation biofuels.

[0012] One of the objectives of technological research is therefore to improve the environmental performance of aircraft. This is why, in all phases of design and development, relevant factors are taken into account to obtain aeronautical components and products that are less energy-intensive, more environmentally friendly, and whose integration and use in civil aviation have moderate environmental consequences, all with the aim of improving the energy efficiency of aircraft.

[0013] 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 (also known as the bypass ratio, abbreviated as BPR) have been proposed. The bypass ratio is the ratio of the secondary airflow rate to the primary airflow rate. To achieve high bypass ratios, the fan section can be decoupled from the low-pressure turbine, thereby allowing their respective rotational speeds to be optimized independently. Generally, this decoupling is achieved using a reduction mechanism located 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. Improving the system's propulsive efficiency can also be achieved by optimizing the fan section size. Indeed, due to its large diameter (designed in particular to achieve high bypass ratios and low fan pressure ratios), the fan section represents a significant portion of the propulsion system in terms of mass and therefore specific fuel consumption. Simultaneously, the fan section generates a very large share of the propulsion system's thrust.

[0014] SUMMARY OF THE INVENTION

[0015] One objective of this application is to optimize the fan section of the propulsion system, particularly to reduce its mass. To this end, the invention is the result of technological research aimed at significantly improving aircraft performance and, in this respect, contributes to reducing the environmental impact of aircraft.

[0016] This objective is achieved within the framework of this application through a fan section of an aeronautical propulsion system, the fan section being enclosed and comprising:

[0017] - a blower housing,

[0018] - an inner platform and an inner face of the blower housing defining between them a gas flow channel,

[0019] - a blower rotor comprising a blade extending in the duct over a blade height from a blade root located on the inner platform to a blade tip located on the inner face, the blade comprising a leading edge and a trailing edge, the blade having:

[0020] - an axial chord at the base of the blade extending along an axis of the fan from the leading edge to the trailing edge,

[0021] - an axial chord at the apex of the blade extending along the axis from the leading edge to the trailing edge, and

[0022] - a mean axial chord equal to the average between the axial chord at the blade root and the axial chord at the blade tip, in which the blade has an axial chord parameter corresponding to a ratio of a product of the mean axial chord by the axial chord at the blade root to a product of the blade height by the axial chord at the blade tip and a velocity parameter corresponding to a difference of a product of a peripheral velocity at the blade tip expressed in m / s by the value 3.263x10' 3 s / m and the value 7.741x10' 1 , the peripheral speed being measured in maximum permissible regime, the blade being configured so that the axial chord parameter is greater than or equal to the speed parameter.

[0023] Such a blower section is advantageously and optionally complemented by the following various characteristics, taken alone or in combination: the velocity parameter is a first velocity parameter, the blade having a second velocity parameter corresponding to a sum of a product of the peripheral velocity at the blade tip expressed in m / s by the value 3.263x10 3 s / m and the value 9.2714x10 1 , the axial chord parameter is less than or equal to the second velocity parameter;

[0024] - a diameter of the blower rotor is greater than or equal to 177.8 cm and less than or equal to 304.8 cm, in particular greater than or equal to 213.36 cm and less than or equal to 266.7 cm and preferably greater than or equal to 215.9 cm and less than or equal to 266.7 cm; the blade comprises a carbon fibre composite material; the blade being a first blade of a plurality of blades included in the blower rotor, a number of the plurality of blades being greater than or equal to 16 and less than or equal to 24; the blower rotor has a strength greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.2, where the strength is equal to a ratio between a straight chord at the blade tip extending from the leading edge to the trailing edge along the blade tip and an inter-blade pitch at the blade tip;and the fan rotor exhibits a strength strictly less than 1.0, where the strength is equal to a ratio between a direct chord at the blade tip extending from the leading edge to the trailing edge along the blade tip and an inter-blade pitch at the blade tip. The presentation also relates to an aeronautical propulsion system comprising:;

[0025] - a movable drive shaft rotating around an axis;

[0026] - a blower shaft;

[0027] - a blower section such as has been presented so far, the blower rotor being driven in rotation by the blower shaft; and

[0028] - a reduction mechanism coupling the drive shaft and the blower shaft in order to drive the blower shaft at a rotational speed lower than the rotational speed of the drive shaft.

[0029] Such a system is advantageously and optionally complemented by the following features: the system is configured to provide a thrust between 80,068 N and 331,374 N, preferably between 88,964 N and 155,688 N; 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;and a nacelle surrounding the fan section, the nacelle comprising an upstream axial end with respect to a direction of gas flow in the propulsion system, the upstream axial end being situated vertically above the axis, the nacelle comprising a lower upstream axial end situated vertically below the axis, an average distance between a first distance along the axis separating the upper upstream axial end and an intersection between the inner face of the casing above the axis and the leading edge of the blade when the blade is above the axis and a second distance along the axis separating the lower upstream axial end and an intersection between the inner face of the casing below the axis and the leading edge of the blade when the blade is below the axis, a ratio of the average distance to a diameter of the fan rotor being less than or equal to 0.4.

[0030] The presentation also covers a method for dimensioning a fan section of an aeronautical propulsion system, the fan section being enclosed and comprising:

[0031] - a blower casing, - an inner platform and an inner face of the blower casing defining a gas flow path between them,

[0032] - a blower rotor comprising a blade extending in the duct over a blade height from a blade root located on the inner platform to a blade tip located on the inner face, the blade comprising a leading edge and a trailing edge, the blade having:

[0033] - an axial chord at the base of the blade extending along an axis of the fan from the leading edge to the trailing edge,

[0034] - an axial chord at the apex of the blade extending along the axis from the leading edge to the trailing edge, and

[0035] - a mean axial chord equal to the average between the axial chord at the blade root and the axial chord at the blade tip, in which the blade has an axial chord parameter corresponding to a ratio of a product of the mean axial chord by the axial chord at the blade root to a product of the blade height by the axial chord at the blade tip and a velocity parameter corresponding to a difference of a product of a peripheral velocity at the blade tip expressed in m / s by the value 3.263x10' 3 s / m and the value 7.741x10 1 , the peripheral speed being measured in maximum permissible regime, the process including a step of dimensioning the blower section so that the axial chord parameter is greater than or equal to the speed parameter.

[0036] Such a process is advantageously and optionally complemented by the following features: the fan section is further dimensioned so that the thrust of the aeronautical propulsion system is between 80,068 N and 331,374 N, preferably between 88,964 N and 155,688 N; and the fan section is further dimensioned so that the bypass ratio of the propulsion system is greater than or equal to 10, for example, between 10 and 35 inclusive, or for example, between 10 and 18 inclusive. Finally, the presentation relates to a method for manufacturing a fan section of an aeronautical propulsion system, comprising the dimensioning of the fan section as described so far and the manufacturing of the section.

[0037] PRESENTATION OF THE DRAWINGS

[0038] Other features and advantages will become clearer from the following description, which is purely illustrative and not exhaustive, and should be read in conjunction with the attached figures, including:

[0039] - Figure 1 schematically represents an aircraft including propulsion systems,

[0040] - Figure 2 schematically represents, in partial view and in cross-section, an example of a propulsion system in which the fan section is enclosed,

[0041] - Figure 3 schematically represents a first example of a planetary reduction mechanism,

[0042] - Figure 4 schematically represents a first example of an epicycloidal reduction mechanism,

[0043] - Figure 5 schematically represents, in partial view and in cross-section, an example of a blower section,

[0044] - Figure 6 schematically represents a cross-section of two blades in a blower section,

[0045] - Figure 7 schematically represents, in partial view and cavalier perspective, an example of a blower section and

[0046] - Figure 8 schematically represents, in partial view and in section, an example of a blower section.

[0047] DETAILED DESCRIPTION OF A METHOD OF IMPLEMENTATION

[0048] In the example illustrated in Figure 1, the aircraft is an airplane 100 comprising a fuselage 101 and two wings 102. In this example, the aircraft includes two propulsion systems 1, each propulsion system 1 being attached to a respective wing 102 of the airplane 100 by means of a pylon. In another embodiment, the aircraft could include one or more propulsion systems attached to the fuselage 101.

[0049] Figure 2 schematically represents, in partial view and in cross-section, a first example of a propulsion system 1.

[0050] In this example, propulsion system 1 is a twin-shaft, shrouded-fan gas turbine engine.

[0051] In Figure 2, the propulsion system 1 has a principal direction extending along a longitudinal axis X. The propulsion system 1 comprises a fan section 2 and a primary body 3, often called the "gas generator". The X direction also defines the principal direction of gas flow through the propulsion system 1.

[0052] The blower section 2 includes a blower 22 and a blower housing 12. The blower 22 includes a blower rotor 9. The blower housing 12 surrounds the blower rotor 9. The blower rotor 9 is mounted to rotate relative to the blower housing 12.

[0053] Fan section 2 is enclosed. For example, the propulsion system 1 may include a nacelle 28 which surrounds fan section 2.

[0054] Referring to Figure 5, the blower section 2 comprises an inner platform 34 and an inner face 35 of the blower housing 12, defining between them a gas flow channel 36. The flow channel 36 therefore extends from the inner platform 34 to the inner face 35.

[0055] The term "platform" here refers to any element of the propulsion system from which a fan blade can be mounted. The platform may be, in particular, a hub or a housing that surrounds the X-axis of the propulsion system. The inner face 35 of the fan housing is also referred to as the "outer platform 35" in the following text. The inner face 35 of the fan housing surrounds the inner platform 34, such that the inner platform 34 lies radially between the X-axis of the propulsion system and the inner face 35 of the fan housing. The inner platform 34 and the inner face 35 of the fan housing each define a flow-stream side surface.

[0056] These surfaces can be at varying distances from the X-axis of the propulsion system 1, as illustrated in Figure 5. The inner surface defined by the inner platform 34 can move away from the X-axis of the propulsion system 1 when traversing this inner surface in the direction of gas flow. The outer surface defined by the outer platform 35 can move towards the X-axis of the propulsion system 1 when traversing this outer surface in the direction of gas flow.

[0057] The outer surface defined by the outer platform 35 and the inner surface defined by the inner platform 34 can alternatively be cylindrical surfaces, each at a constant radial distance to the X axis of the propulsion system 1.

[0058] The blower rotor 9 includes at least one blade 14 extending in the channel 36. The blade 14 may include a carbon fiber composite material.

[0059] The fan blade 14 has a leading edge 14a and a trailing edge 14b (see, for example, Figure 5). The leading edge 14a is configured to extend in relation to the flow of gases entering the fan rotor 9. It corresponds to the forward part of an airfoil that faces the airflow and divides the airflow into an intrados (lower surface) flow and an extrados (upper surface) flow. The trailing edge 14b, on the other hand, corresponds to the rear part of the airfoil, where the intrados and extrados flows meet.Note here that, when the blade 14 includes a leading edge shield and / or trailing edge shield, the leading edge 14a (respectively the trailing edge 14b) of the blade 14 corresponds to the front part of the shield profile which reconstitutes the leading edge (respectively the rear part of the shield profile which reconstitutes the trailing edge 14b) and whose function is to divide the flow into an intrados flow and an extrados flow (respectively to join the flows).

[0060] The blade 14 extends in the channel 36 over a blade height 47 from a blade root 37 located on the inner platform 34 to a blade tip 21 located on the outer platform 35. The blade root is located in the gas flow channel 36. The blade further includes, beneath the platform, a blade base which secures the blade to the hub. The blade height is measured along a radial axis R orthogonal to the X-axis of the propulsion system 1 and passing through the X-axis of the propulsion system 1. Preferably, and with reference to Figure 5, the blade height 47 is measured between an intersection 43 between the inner platform 34 and the leading edge 14a and an intersection 45 between the outer platform 35 and the leading edge 14a.The fan blade 14 also has a direct chord at the blade tip 39 extending from the leading edge 14a to the trailing edge 14b along the blade tip 21 and a direct chord at the blade root 38 extending from the leading edge 14a to the trailing edge 14b along the blade root 37. The direct chord, or normal chord, refers to the distance of the segment between the leading edge and the trailing edge without the segment being projected. The direct chord at the blade tip extends from an upstream point, which is at the intersection between the leading edge 14a and the inner face 35 of the fan casing, to a downstream point, which is at the intersection between the trailing edge 14b and the inner face 35 of the fan casing. The direct rope at the base of the wing extends from an upstream point which is at the intersection between the leading edge 14a and the inner platform 34 and a downstream point which is at the intersection between the trailing edge 14b and the inner platform 34.

[0061] The fan blade 14 also has an axial chord at the blade apex extending along the X-axis from the leading edge 14a to the trailing edge 14b and an axial chord at the blade root extending along the X-axis from the leading edge 14a to the trailing edge 14b. The term axial chord at the blade apex (respectively at the blade root) refers to the length of the projection along the X-axis of the direct chord at the blade apex (respectively at the blade root).

[0062] We can define an average axial chord equal to the average between the axial chord at the base of the blade and the axial chord at the top of the blade.

[0063] D is the diameter of the fan rotor, measured in a plane normal to the X-axis of the propulsion system 1 from the X-axis to the intersection point 45 between the blade tip 21 and the leading edge 14a of the blade. The diameter D is expressed in meters. The diameter D of the fan rotor 9 can be between 70 inches (177.8 cm) and 120 inches (304.8 cm) inclusive, specifically between 84 inches (213.36 cm) and 105 inches (266.7 cm) inclusive, and preferably between 85 inches (215.9 cm) and 105 inches (266.7 cm) inclusive. For example, the diameter D of the fan rotor 9 is approximately 90 inches (228.6 cm). These different values ​​of diameter D allow the propulsion system 1 to be integrated in a conventional manner, in particular under a wing of aircraft 100.

[0064] Blade 14 has an axial chord parameter corresponding to a ratio of a product of the average axial chord by the axial chord at the base of the blade to a product of the blade height 47 by the axial chord at the top of the blade.

[0065] The axial chord parameter can be denoted Pc. It obeys the relation Pc = — — in which C xm denotes the mean axial chord, C x0 denotes the axial chord, h denotes the blade height 47 and C xl00 designates the axial chord at the apex of the blade.

[0066] Blade 14 has a velocity parameter corresponding to the difference between the product of a peripheral velocity at the blade tip, expressed in m / s, and the value 3.263x10 3 s / m and the value 7.741x10 1The peripheral speed is measured at the maximum permissible speed. This speed corresponds to the maximum speed of the blower during normal operation. This term is also known by its English name, 'Maximum Permissible Rotational Speed'.

[0067] The speed parameter can be denoted as Pv.

[0068] It respects the relationship Pv = 3.263 x 10 -3 x V p - 7.741 x 10 -1 which can also be written as Pv = 0.003263 x V p - 0.7741 and in which V p denotes peripheral speed.

[0069] The peripheral velocity is the velocity of a point on the blade tip 21, for example the intersection point 45, in the circumferential direction, i.e. along the circumferential axis 0. The circumferential axis 0 is an axis passing through the point on the blade tip 21, the circumferential axis 0 being orthogonal to the X axis of the propulsion system and to the radial axis R when the radial axis R passes through the point on the blade tip 21. The peripheral velocity at the blade tip can be greater than or equal to 260 m / s and less than or equal to 400 m / s, for example greater than or equal to 270 m / s and less than or equal to 380 m / s.

[0070] In order to optimize the performance of propulsion system 1, the axial chord parameter is greater than or equal to the velocity parameter.

[0071] In other words, Pc > Pv, a relationship that can also be written in the form

[0072] The axial chord parameter is chosen here to be sufficiently large to significantly increase the margin with respect to the natural mode of vibration of the blade in bending. Bending is a deformation of the blade during which the blade tip moves closer to the blade root, the blade curving around a direction parallel to the drive axis, i.e., the X-axis of the propulsion system. The margin depends in particular on the difference between the blade's rotational speed at maximum permissible speed and the frequency of the natural mode of vibration in bending. When the axial chord parameter is chosen to be greater than or equal to the speed parameter, this allows the fan to be more stable with respect to the natural mode of vibration of the blade in bending under greater external disturbances. It becomes possible to operate the fan stably by reducing the ratio of the length between the shroud inlet and the fan to the fan diameter.This helps to reduce the mass of the engine and the drag of the engine via the drag of the fairing.

[0073] It should be noted that the velocity parameter Pv can be a first velocity parameter, denoted Pv1, and that blade 14 has a second velocity parameter, Pv2. The second velocity parameter corresponds to the sum of a product of the peripheral velocity at the blade tip, expressed in m / s, and the value 3.263 x 10⁻¹⁰⁸⁴. 3 s / m and the value 9.2714x10 1 , the peripheral speed being measured at maximum permissible operating speed.

[0074] The second speed parameter Pv2 = 3.263 x 10 -3 x V p + 9.2714 x 10 -1 which can also be written as Pv2 = 0.003263 x V p + 0.92714 and in which V p denotes peripheral speed.

[0075] Advantageously, the axial chord parameter is less than or equal to the second velocity parameter.

[0076] In other words, Pc < Pv2, a relationship that can also be written in the form 3.263 x 10⁻¹⁰ 3 x + 9.2714 x 10” 1

[0077] If the propulsion system 1 includes a nacelle 28, it is possible to choose particular geometric conditions in relation to the diameter D of the blower rotor 9 and the distance separating the blower rotor and the inlet of the nacelle.

[0078] The nacelle may have an upstream axial end, relative to the direction of gas flow in the propulsion system, located more or less far from the fan 22 depending on the angle around the X-axis. Referring to Figure 8, the nacelle 28 comprises an upper part 280 situated vertically above the X-axis and a lower part 281 situated vertically below the X-axis. The upper part 280 includes an upper upstream axial end 330, and the lower part 281 includes a lower upstream axial end 331. The upper upstream axial end 330 is located upstream of the lower upstream axial end 331, such that the upper part 280 extends forward beyond the lower part 281. The axial distance between the upper upstream axial end 330 and the fan 22 is greater than the axial distance between the lower upstream axial end 331. and blower 22.The nacelle 28 exhibits continuity around the X-axis, such that the axial distance between the upstream axial end for a certain angle around the X-axis and the blower 22 varies according to the angle between a maximum at noon, i.e., above the X-axis at the upper upstream axial end 330, and a minimum at 6 o'clock, i.e., below the X-axis at the lower upstream axial end 331. The nacelle is globally symmetrical with respect to a vertical plane passing through the X-axis. Figure 8 shows a cross-section in this vertical plane.

[0079] To characterize the distance between the blower inlet and the upstream end of the nacelle, we can choose the average distance L between: the distance 480 separating the upper upstream axial end along the X-axis

[0080] 330 and the intersection point 450 between the inner face 350 of the housing above the X-axis and the leading edge of the blade 140 when it is above the X-axis, and the distance 481 separating along the X-axis the lower upstream axial end

[0081] 331 and the intersection point 451 between the inner face 351 of the casing below the X axis and the leading edge of the blade 141 when it is below the X axis.

[0082] It is possible to choose a ratio of the distance between the blower inlet and the upstream end of the nacelle to the diameter D of the blower rotor 9 less than or equal to 0.4.

[0083] Advantageously, it can also be required that this ratio be greater than or equal to 0.2.

[0084] The fan rotor 9 may comprise a plurality of fan blades 14. Each of the fan blades 14 may be fixed relative to the fan hub 13 or have a variable pitch. In the latter case, each of the fan blades 14 is pivotally mounted relative to the fan hub 13 along a pitch axis and is connected to a pitch-changing mechanism (not shown) mounted in the propulsion system 1. The pitch-changing mechanism allows the pitch angle of the fan blades 14 to be adjusted according to the flight phases.

[0085] The blower rotor 9 preferably comprises at least fourteen blower blades 14 and at most twenty-four blower blades 14, even more preferably at least sixteen blower blades 14, or at least seventeen blades and at most twenty-four blower blades 1, or at most twenty-two blower blades 14. Figure 7 illustrates the case where the blower rotor 9 comprises 22 blades.

[0086] The fan section 2 may also include a fan stator 16 fixedly mounted on the fan housing 12. The fan stator 16 includes fixed blades 17, generally called outlet guide vanes (OGVs). This set of fixed blades straightens and regulates the airflow downstream of the fan rotor 9 to contribute to engine thrust. This set of fixed blades also acts as a noise reducer.

[0087] Alternatively, the outlet blades 17 could have variable pitch. If so, and similarly to the fan blades 14 of the fan rotor 9, the base of the outlet blades 17 is pivotally mounted about a pitch axis and is connected to a pitch-changing mechanism (not shown), the pitch being adjusted according to the flight phases by the pitch-changing mechanism.

[0088] The number of outlet blades 17 depends on the acoustic criteria defined for the propulsion system 1 and is at least equal to the number of blower blades 14.

[0089] The primary body 3 comprises a compressor section 29, a combustion chamber 6 and a turbine section 30.

[0090] Compressor section 29 includes a low pressure compressor 4 and a high pressure compressor 5.

[0091] The low-pressure compressor 4 includes a rotor 41 suitable for being driven in rotation relative to the housing 31 of the propulsion system 1 and a stator 42 fixedly mounted on the housing 31.

[0092] The rotor 41 of the low-pressure compressor 4 includes movable wheels 4a and the stator 42 of the low-pressure compressor 4 includes fixed wheels 4b. The movable wheels 4a are arranged alternately with the fixed wheels 4b, thus forming a succession of low-pressure compressor stages.

[0093] Similarly, the high-pressure compressor 5 includes a rotor 51 suitable for being driven in rotation relative to the housing 31 of the propulsion system 1 and a stator 52 fixedly mounted on the housing 31.

[0094] The rotor 51 of the high-pressure compressor 5 comprises rotating wheels 5a, and the stator 52 of the high-pressure compressor 5 comprises stationary wheels 5b. The rotating wheels 5a are arranged alternately with the stationary wheels 5b, thus forming a succession of high-pressure compressor stages. The turbine section 30 comprises a high-pressure turbine 7 and a low-pressure turbine 8.

[0095] The high-pressure turbine 7 comprises a rotor 71 suitable for being driven in rotation relative to the casing 31 of the propulsion system 1 and a stator 72 fixedly mounted on the casing 31.

[0096] The rotor 71 of the high-pressure turbine 7 comprises rotating wheels 7a and the stator 72 of the high-pressure turbine 7 comprises fixed wheels 7b. The rotating wheels 7a are arranged alternately with the fixed wheels 7b, thus forming a succession of high-pressure turbine stages.

[0097] Similarly, the low-pressure turbine 8 comprises a rotor 81 suitable for being driven in rotation relative to the casing 31 of the propulsion system 1 and a stator 82 fixedly mounted on the casing 31.

[0098] The rotor 81 of the low-pressure turbine 8 comprises moving wheels 8a and the stator 82 of the low-pressure turbine 8 comprises fixed wheels 8b. The moving wheels 8a are arranged alternately with the fixed wheels 8b, thus forming a succession of low-pressure turbine stages.

[0099] The propulsion system 1 includes a low-pressure shaft 11 connecting the rotor 41 of the low-pressure turbine 4 to the rotor 81 of the low-pressure compressor 8, the low-pressure shaft 11 being mounted rotatably relative to the housing 31 around the longitudinal axis X.

[0100] When the propulsion system 1 is in operation, the rotor 81 of the low-pressure turbine 8 drives the rotor 41 of the low-pressure compressor 4 through the low-pressure shaft 11.

[0101] The propulsion system 1 further comprises a blower shaft 20 and a reduction mechanism 19. The blower rotor 9 is fixedly mounted on the blower shaft 20. The reduction mechanism 19 has an inlet and an outlet. The inlet of the reduction mechanism 19 is connected to the low-pressure shaft 11, and the outlet of the reduction mechanism 19 is connected to the blower shaft 20. Thus, when the propulsion system 1 is in operation, the rotor 81 of the low-pressure turbine 8 drives not only the rotor 41 of the low-pressure compressor 4, but also the blower rotor 9, via the low-pressure shaft 11, the reduction mechanism 19, and the blower shaft 20. Thanks to the reduction mechanism 19, the blower rotor 9 is driven to rotate at a speed lower than the rotational speed of the rotor 41 of the low-pressure turbine 4.

[0102] The reduction mechanism 19 thus allows independent control of the rotation speed of the blower 22 and the rotation speed of the low pressure turbine 8 and the low pressure compressor 4.

[0103] The low-pressure turbine 8, the low-pressure shaft 11, the low-pressure compressor 4, the blower shaft 20, the reduction mechanism 19 and the blower 22 together form the "low-pressure body" of the propulsion system 1.

[0104] The propulsion system 1 further includes a high-pressure shaft 10 connecting the rotor 51 of the high-pressure turbine 5 to the rotor 71 of the high-pressure compressor 7, the high-pressure shaft 10 being mounted rotatably relative to the housing 31 around the longitudinal axis X. The high-pressure shaft 10 is coaxial with the low-pressure shaft 11 and extends around the low-pressure shaft 11.

[0105] When the propulsion system 1 is in operation, the rotor 81 of the low-pressure turbine 8 drives the rotor 51 of the low-pressure compressor 5 through the low-pressure shaft 11.

[0106] The high-pressure turbine 7, the high-pressure shaft 10 and the high-pressure compressor 4 together form the "high-pressure body" of the propulsion system 1.

[0107] The low-pressure shaft 11 and the high-pressure shaft 10 can be co-rotating, i.e., driven in the same direction of rotation around the longitudinal axis X. Alternatively, the low-pressure shaft 11 and the high-pressure shaft 10 can be contra-rotating, i.e., driven in opposite directions of rotation around the longitudinal axis X.

[0108] The dual-body propulsion system 1 may include, in particular, a single-stage high-pressure turbine 7, i.e. comprising exactly one stage, or a two-stage high-pressure turbine 7, i.e. comprising exactly two stages (as illustrated in the example in Figure 2).

[0109] The high-pressure compressor 5 comprises at least eight stages (as illustrated in the example in Figure 2) and at most eleven stages.

[0110] The low-pressure turbine 8 comprises at least three stages (as illustrated in the example in Figure 2) and at most seven stages.

[0111] The low-pressure compressor 4 comprises at least two stages and at most four stages. When the propulsion system is in operation, an airflow F entering the propulsion system 1 passes through the blower 22 and is then divided between a primary airflow F1 and a secondary airflow F2, which flow from upstream to downstream in the propulsion system 1 in the direction of gas flow through the propulsion system.

[0112] The secondary airflow F2, also called the "bypass airflow", flows in the secondary vein, 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.

[0113] The primary airflow F1 flows in a primary channel within the primary body 3, passing successively through the compressor section 29 (low-pressure compressor 4 and high-pressure compressor 5), the combustion chamber 6 where it is mixed with fuel to serve as an oxidizer, and the turbine section 30 (high-pressure turbine 7 and low-pressure turbine 8). The passage of the primary airflow F1 through the turbine section 30, receiving energy from the combustion chamber 6, causes the impellers 7a, 8a of the turbine section 30 to rotate, which in turn drives the impellers 4a, 5a of the compressor section 29 as well as the blower rotor 9.

[0114] To improve the propulsive efficiency of propulsion system 1 and reduce its specific fuel consumption and the noise emitted by the fan section 2, propulsion system 1 may have a high bypass ratio. A "high" bypass ratio is defined as a ratio greater than or equal to 10, for example, between 10 and 80 inclusive, preferably between 10 and 35 inclusive, preferably between 10 and 18 inclusive. The bypass ratio is defined as the ratio of the mass flow rate of the secondary airflow F2 to the mass flow rate of the primary airflow F1, these mass flow rates being measured when propulsion system 1 is stationary, uninstalled, in takeoff mode in a standard atmosphere (as defined by the International Civil Aviation Organization (ICAO) Manual, Doc 7488 / 3, 3rd edition) and at sea level.By "not installed", it is meant that the measurements are carried out when the propulsion system 1 is on a test bench (and not installed on an aircraft), the measurements then being simpler to carry out.

[0115] In a propulsion system including a reduction mechanism 19 such as that illustrated in Figure 2, the decoupling between the rotational speed of the blower 22 and the rotational speed of the low-pressure turbine 8 makes it possible to reduce the rotational speed and the pressure ratio of the blower rotor 9 while increasing the power extracted by the low-pressure turbine 8. Indeed, the overall efficiency of the propulsion system 1 is conditioned to the first order by the propulsion efficiency, which is favorably influenced by minimizing the variation in kinetic energy of the air as it passes through the propulsion system 1.In a propulsion system with a high dilution 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 which the secondary airflow F2 undergoes when passing through the blower section 2. The propulsive efficiency and the pressure ratio of the blower section 2 are therefore linked: the lower the pressure ratio of the blower section 2, the better the propulsive efficiency. In order to improve 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 16 (or, in the absence of a stator 16, of the blower rotor 9) and the average pressure at the inlet of the blower rotor 9, is less than or equal to 1.70, preferably less than or equal to 1.50, for example between 0.90 and 1.45.The average pressures are measured here over the height of at least one of the blower blades 14, that is to say from the surface which radially delimits inside the airflow duct at the inlet of the blower rotor 9 to the top 21 of blower blade 14.

[0116] The peripheral speed at the top 21 of the blower blades 14 can also be between 260 meters per second (ms 1 ) and 400 meters per second (ms 1 ) included. The blower pressure ratio can then be between 1.20 and 1.45.

[0117] In a direct-drive propulsion system, the fan rotor 9 can alternatively be directly coupled to the low-pressure shaft 11, i.e., without a reduction mechanism. The low-pressure shaft 11 is then combined with the fan shaft 20 so that the fan rotor 9 is driven by the low-pressure shaft 11 at the same rotational speed as the rotor 81 of the low-pressure turbine 8.

[0118] The propulsion system 1 is configured to provide a thrust between 18,000 Lbf (80,068 N) and 70,000 Lbf (331,374 N), preferably between 20,000 Lbf (88,964 N) and 51,000 Lbf (226,859 N), or even between 20,000 Lbf (88,964 N) and 35,000 Lbf (155,688 N). The reduction mechanism 19 may include an epicyclic or planetary reduction mechanism, single-stage or two-stage.

[0119] For example, Figure 3 illustrates a planetary (or "star") type reduction mechanism 19. The reduction mechanism 19 comprises a sun pinion 19a (input of the reduction mechanism 19), centered on an 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 gear 19b (output of the reduction mechanism 19) coaxial with the sun pinion 19a and configured to drive in rotation the blower shaft 20 around its axis X of rotation, and a series of satellites 19c distributed circumferentially around the axis X of rotation of the rotor 9 of the blower section 2, between the sun pinion 19a and the ring gear 19b, each satellite gear 19c being internally meshed with the sun pinion 19a and externally with the ring gear 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.

[0120] In another example, Figure 4 illustrates an epicycloidal (or "planetary" in English) type reduction mechanism 19, 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 planet carrier 19d.

[0121] 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 rotor 9 of the blower section 2 is less than the rotational speed of the low pressure shaft 11.

[0122] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11. The reduction ratio can be greater than or equal to 2.7 and less than or equal to 3.5, typically around 3.0.

[0123] To further optimize the performance of the propulsion system 1, a strength of the blower rotor 9 can be chosen greater than or equal to 0.9 and less than or equal to 1.3. The strength is equal to the ratio between the direct chord at the blade tip 39 and an inter-blade pitch 23.The blade pitch 23 corresponds to the angular distance between the upstream intersection points P1 and P2 of two adjacent blades 14, that is, the distance along the circumferential axis 0 from point P1 to point P2; the blade pitch 23 is therefore equal to the external radius Re of the fan rotor 9 (half-diameter) multiplied by the angle between a first straight line D1, lying in a plane normal to the X-axis, which originates from the upstream intersection point P1 (see Figures 6 and 7) of a first blade 14 and intersects the X-axis, and a second straight line D2, lying in the plane normal to the X-axis, which originates from the upstream intersection point P2 of a second blade 14 immediately adjacent to the first blade 14 and intersects the X-axis. Since strength is a ratio of distances, it is measured when the propulsion system 1 is cold.

[0124] In a first option, the strength can be chosen to be greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.2. This first option is particularly suitable for a rotor 9 comprising twenty-two blades 14. In a propulsion system 1 with a reduction mechanism 19, the rotational speed of the fan rotor 9 is reduced, for example, to between 260 m / s and 400 m / s. Its pressure ratio can also be reduced and is, for example, between 1.05 and 1.5, or even between 1.1 and 1.45, so that the velocity difference between the outlet of the fan rotor 9 and the inlet of the fan rotor 9 is reduced while optimizing the efficiency of the fan section 2. The flow at the blade tip 14 through the fan rotor 9 is then supersonic. In particular, a supersonic shock is generated at the level of the 14 fan blades.However, this supersonic shock cannot be eliminated, especially during cruise: it must therefore be controlled to prevent it from deteriorating the efficiency of the fan section 2. By dimensioning and manufacturing the fan rotor 9 with twenty-two blades so that its strength is greater than or equal to 1.0 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.2, the inter-blade pitch 23 and the direct chord at the blade tip 39 are such that the supersonic shock is stable in maximum permissible regime and moves only in the divergent part of the inter-blade channel. The inter-blade channel corresponds to the passage between two adjacent blades 14 which extends between an inlet plane, which is normal to the airflow F into the inlet of the blower rotor 9 and passes through the leading edge 14a of the first blade 14, and an outlet plane which is parallel to the inlet plane and passes through the trailing edge 14b of the second blade 14.This channel, from upstream to downstream through the fan rotor 9, has a convergent section extending from the inlet plane to an intermediate plane corresponding to the channel's minimum cross-section, and a divergent section extending from the intermediate plane to the outlet plane. The Applicant observed that when the supersonic shock reached the convergent section of the inter-blade channel 24, it became unstable and reduced the efficiency of the fan section 2. Conversely, when the supersonic shock remains in the divergent section, for example, near the intermediate plane, the supersonic shock is stable at maximum permissible operating conditions.The longer the inter-blade channel, i.e., the greater the distance between the inlet and outlet planes, the easier it is to design a convergent-divergent channel and thus to maintain the supersonic shock in the divergent part of the channel; however, increasing the length of the inter-blade channel results in an increase in the direct chord at the blade tip 39, and therefore an increase in the mass of the fan rotor 9 (and consequently the specific consumption of the propulsion system 1) and a reduction in the efficiency of the fan section 2. It should be noted in particular that the length of the inter-blade channel and the direct chord are related to the inter-blade pitch by the strength of the fan rotor 9.For a 9-bladed fan rotor with twenty-two blades, a strength between 1.0 and 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.2, is then a good compromise between the length of the inter-blade channel 24 in order to control the supersonic shock in the 9-bladed fan rotor and the efficiency of the 2-bladed fan section.

[0125] In a second option, the strength can be chosen to be strictly less than 1.0. This second option is particularly suitable for a fan rotor 9 comprising between seventeen and twenty blades 14. Indeed, the direct chord at the blade tip is then reduced relative to the available space (high inter-blade pitch 23), which makes it possible to reduce the mass of the blade 14 and modify its frequency positioning, since modifying the direct chord at the blade tip has a significant impact on the first deformation mode of the blade 14 (at the same operating range). In fact, when the deformation mode of the blades 14 occurs in a stabilized regime of the fan rotor 9 (in particular, the maximum permissible regime), one solution may be to increase the thickness of the blade root in order to improve the margins and shift this deformation mode.This increase in the thickness of the blades 14 at the blade root, however, implies an increase in the overall size of the root, which influences the circumferential dimensions of the recesses that receive the root and therefore the design for managing the mechanical resistance of the disc. Increasing the thickness of the blades 14 is thus likely to increase the hub-to-head ratio of the fan rotor 9. However, increasing the hub-to-head ratio is detrimental to the aerodynamic efficiency of the fan rotor 9. Conversely, dimensioning the fan rotor 9 so that its strength is less than 1.0 makes it possible to reduce the direct chord at the blade tip and to shift the first mode of deformation of the blades 14 into an unstabilized operating range or outside the operating range of the fan rotor, without thickening the blades 14. Therefore, it is not necessary to increase the hub-to-head ratio.

[0126] The distances (length, radius, diameter, etc.) are measured at ambient temperature (approximately 20°C) when the propulsion system 1 is cold, i.e. 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 with respect to the conditions in which the propulsion system 1 is in takeoff mode.

Claims

DEMANDS 1. Aeronautical propulsion system (1) comprising: - a drive shaft (11) that rotates about an axis (X); - a blower shaft (20); - a fan section (2) of an aeronautical propulsion system (1), a fan rotor (9) being driven in rotation by the fan shaft (20); and - a reduction mechanism (19) coupling the drive 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 drive shaft (11), the blower section (2) being shrouded and comprising: - a blower housing (12), - an inner platform (34) and an inner face (35) of the blower housing (12) defining between them a gas flow channel (36), - the fan rotor (9) comprising a blade (14) extending in the duct (36) over a blade height (47) from a blade root (37) located on the inner platform (34) to a blade tip (21) located on the inner face (35), the blade (14) comprising a leading edge (14a) and a trailing edge (14b), the blade (14) having: - an axial chord at the base of the blade extending along an axis (X) from the blower of the leading edge (14a) to the trailing edge (14b), - an axial chord at the apex of the blade extending along the axis (X) from the leading edge (14a) to the trailing edge (14b), and - a mean axial chord equal to the average between the axial chord at the blade root and the axial chord at the blade tip, in which the blade (14) has an axial chord parameter corresponding to a ratio of a product of the mean axial chord by the axial chord at the blade root to a product of the blade height (47) by the axial chord at the blade tip and a velocity parameter corresponding to a difference of a product of a peripheral velocity at the blade tip expressed in m / s by the value 3.263x10' 3 s / m and the value 7.741x10' 1 , the peripheral speed being measured at maximum permissible operating conditions, the blade (14) being configured such that the axial chord parameter is greater than or equal to the speed parameter, the propulsion system comprising a nacelle (28) surrounding the fan section (2), the nacelle (28) comprising an upper upstream axial end (330) with respect to a direction of gas flow in the propulsion system, the upper upstream axial end (330) being located vertically above the axis (X), the nacelle (28) comprising a lower upstream axial end (331) located vertically below the axis (X),an average distance between a first distance (480) along the axis (X) separating the upper upstream axial end (330) and an intersection (450) between the inner face (350) of the housing above the axis (X) and the leading edge of the blade when the blade is above the axis (X) and a second distance (481) along the axis (X) separating the lower upstream axial end (331) and an intersection (451) between the inner face (351) of the housing below the axis (X) and the leading edge of the blade when the blade is below the axis (X), the first distance (480) being greater than the second distance (481), a ratio of the average distance to a diameter (D) of the blower rotor (9) being less than or equal to 0.

4.

2. Propulsion system according to claim 1, wherein the velocity parameter is a first velocity parameter, the blade having a second velocity parameter corresponding to a sum of a product of the peripheral velocity at the blade tip expressed in m / s by the value 3.263x10' 3 s / m and the value 9.2714x10 1 the axial chord parameter being less than or equal to the second velocity parameter.

3. Propulsion system according to claim 1 or 2, wherein a diameter (D) of the blower rotor (9) is greater than or equal to 177.8 cm and less than or equal to 304.8 cm, in particular greater than or equal to 213.36 cm and less than or equal to 266.7 cm and preferably greater than or equal to 215.9 cm and less than or equal to 266.7 cm.

4. Propulsion system according to any one of claims 1 to 3, wherein the blade (14) comprises a carbon fiber composite material.

5. Propulsion system according to any one of claims 1 to 4, the blade (14) being a first blade of a plurality of blades included in the blower rotor (9), a number of the plurality of blades being greater than or equal to 16 and less than or equal to 24.

6. Propulsion system according to claim 5, wherein the blower rotor (9) has a strength greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.2, where the strength is equal to a ratio between a direct chord at the blade tip extending from the leading edge (14a) to the trailing edge (14b) along the blade tip (21) and an interblade pitch (23) at the blade tip.

7. Propulsion system according to claim 5, wherein the blower rotor (9) has a strength strictly less than 1.0, where the strength is equal to a ratio between a direct chord (39) at the blade tip extending from the leading edge (14a) to the trailing edge (14b) along the blade tip (21) and an inter-blade pitch (23) at the blade tip.

8. Propulsion system according to any one of claims 1 to 7, configured to provide a thrust of between 80,068 N and 331,374 N, preferably between 88,964 N and 155,688 N.

9. Propulsion system (1) according to any one of claims 1 to 8, wherein 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.

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