Optimised fan section of an aeronautical propulsion system

Abstract

The invention relates to a fan section which has: a ratio of a product of a mean axial chord of the blade and a blade root axial chord to a product of the square of a blade height (47) and a blade tip axial chord (21), which ratio is greater than or equal to a difference between a product of a maximum rotational speed and the value 2.501×10-3 s / rad and the value 1.7148 × 10-1, and a ratio of the blade root axial chord to a product of the square of the mean axial chord, the blade height and the blade tip axial chord, which ratio is greater than or equal to a difference between a product of the maximum rotational speed and the value 0.2515 s / rad and the value 52.701.

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 (36),

[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: - a first 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 square of the blade height by the axial chord at the blade tip,

[0023] - a first speed parameter corresponding to a difference in the product of a maximum permissible speed of the blower rotor expressed in rad / s by the value 2.501x10' 3 s / rad and the value 1.7148x10' 1 ,

[0024] - a second axial chord parameter corresponding to a ratio of the axial chord at the blade root to a product of the square of the average axial chord by the blade height and by the axial chord at the blade tip, and

[0025] - a second speed parameter corresponding to a difference of a product of the maximum permissible speed of the blower rotor expressed in rad / s by the value 0.2515 s / rad and the value 52.701, the blower section being configured so that the blower rotor has a maximum permissible speed greater than or equal to 280 rad / s, the blade being configured so that the first axial chord parameter is greater than or equal to the first speed parameter and so that the second axial chord parameter is greater than or equal to the second speed parameter.

[0026] Such a blower section is advantageously and optionally complemented by the following various features, taken alone or in combination:

[0027] - the blade has a third speed parameter corresponding to a difference in the product of a maximum permissible speed of the blower rotor expressed in rad / s by the value 3.12x10' 3s / rad and the value 2, 14x10' 1 , a fourth speed parameter corresponding to a difference of a product of the maximum permissible speed of the blower rotor expressed in rad / s by the value 0.314 s / rad and the value 65.9, the blade being configured so that the first axial chord parameter is less than or equal to the third speed parameter, the second axial chord parameter being less than the fourth speed parameter;

[0028] - a blower rotor diameter 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 being the first blade of a plurality of blades included in the blower rotor, one number of the plurality of blades being greater than or equal to 17 and less than or equal to 22; and

[0029] - the blower includes a blower hub, each blade being mounted pivotally relative to the blower hub so as to present a variable pitch.

[0030] The presentation also covers an aeronautical propulsion system comprising:

[0031] - a movable drive shaft rotating around an axis;

[0032] - a blower shaft;

[0033] - a blower section such as has been described so far, with the blower rotor being driven in rotation by the blower shaft; and

[0034] - 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.

[0035] Such a system is advantageously and optionally complemented by the following features:

[0036] - the system is configured to provide a thrust between 80,068 N and

[0037] 331,374 N, preferably between 88,964 N and 155,688 N;

[0038] - a propulsion system dilution ratio greater than or equal to 10, for example between 10 and 35 inclusive, for example between 10 and 18 inclusive; and

[0039] - 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 located vertically above the axis, the nacelle comprising a lower upstream axial end located 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.

[0040] The presentation also covers a method for dimensioning a fan section of an aeronautical propulsion system, the fan section being enclosed and comprising:

[0041] - a blower housing,

[0042] - an inner platform and an inner face of the blower housing defining between them a gas flow channel,

[0043] - 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:

[0044] - an axial chord at the base of the blade extending along an axis of the fan from the leading edge to the trailing edge,

[0045] - an axial chord at the apex of the blade extending along the axis from the leading edge to the trailing edge, and

[0046] - an average 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 presents

[0047] - a first axial chord parameter corresponding to a ratio of a product of the average axial chord by the axial chord at the blade root to a product of the square of the blade height by the axial chord at the blade tip,

[0048] - a first speed parameter corresponding to a difference in the product of a maximum permissible speed of the blower rotor expressed in rad / s by the value 2.501x10' 3 s / rad and the value 1.7148x10' 1 ,

[0049] - a second axial chord parameter corresponding to a ratio of the axial chord at the blade root to a product of the square of the average axial chord by the blade height and by the axial chord at the blade tip, and

[0050] - a second speed parameter corresponding to a difference of a product of the maximum allowable speed of the blower rotor expressed in rad / s by the value 0.2515 s / rad and the value 52.701, the blower section being configured so that the blower rotor has a maximum allowable speed greater than or equal to 280 rad / s, the process comprising a blower section dimensioning step such that the first axial chord parameter is greater than or equal to the first speed parameter and the second axial chord parameter is greater than or equal to the second speed parameter.

[0051] Such a process is advantageously and optionally complemented by the following features: the fan section is further dimensioned so that a 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 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.

[0052] The presentation finally focuses on a manufacturing process for a fan section of an aeronautical propulsion system, including the dimensioning of the fan section as presented so far and the manufacturing of the section.

[0053] PRESENTATION OF THE DRAWINGS

[0054] 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:

[0055] - Figure 1 schematically represents an aircraft including propulsion systems,

[0056] - Figure 2 schematically represents, in partial view and in cross-section, an example of a propulsion system in which the fan section is enclosed,

[0057] - Figure 3 schematically represents a first example of a planetary reduction mechanism,

[0058] - Figure 4 schematically represents a first example of an epicyclic reduction mechanism; - Figure 5 schematically represents, in partial view and in cross-section, an example of a blower section.

[0059] - Figure 6 schematically represents, in partial view and cavalier perspective, an example of a blower section, and

[0060] - Figure 7 schematically represents, in partial view and in section, an example of a blower section.

[0061] DETAILED DESCRIPTION OF A METHOD OF IMPLEMENTATION

[0062] 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.

[0063] Figure 2 schematically represents, in partial view and in cross-section, a first example of a propulsion system 1.

[0064] In this example, propulsion system 1 is a twin-shaft, shrouded-fan gas turbine engine.

[0065] 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.

[0066] 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.

[0067] Fan section 2 is enclosed. For example, the propulsion system 1 may include a nacelle 28 which surrounds fan section 2.

[0068] With reference to Figure 5, the fan section 2 comprises an inner platform 34 and an inner face 35 of the fan housing 12, defining between them a gas flow channel 36. The flow channel 36 thus extends from the inner platform 34 to the inner face 35. The term "platform" here refers to any element of the propulsion system 1 from which a fan blade is 4, or at least any element of the propulsion system from which a fan blade is capable of extending substantially radially. The platform may, in particular, be a hub or a housing that surrounds the axis X of the propulsion system 1. 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 1 and the inner face 35 of the housing. The inner platform 34 and the inner face 35 of the fan housing each define a flow-stream side surface.

[0069] 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.

[0070] 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.

[0071] The blower rotor 9 includes at least one blade 14 extending in the channel 36. The blade may include a carbon fiber composite material.

[0072] 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).

[0073] The blade 1 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.

[0074] The fan blade 14 also has a straight chord at the blade crest 39 extending from the leading edge 14a to the trailing edge 14b along the blade crest 21 and a straight curve at the blade root 38 extending from the leading edge 14a to the trailing edge 14b along the blade root 37. The term straight 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 straight chord at the blade crest 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.

[0075] 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).

[0076] An average axial chord can be defined as the average of the axial chord at the blade root and the axial chord at the blade tip. 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.

[0077] Blade 14 has a first axial chord parameter corresponding to the ratio of the product of the mean axial chord and the axial chord at the blade root to the product of the square of the blade height 47 and the axial chord at the blade tip. The first axial chord parameter is calculated by expressing its constituent terms in meters; the value of the first ratio is considered dimensionless. The first axial chord parameter can be denoted Pc1.

[0078] The first axial chord parameter obeys the relation Pc1 = (C xm x C x Oh 2 x C x ioo) in which C xm denotes the mean axial chord, C x o designates the rope at the base of the blade, h designates the blade height 47 and C x ioo refers to the rope at the top of dawn.

[0079] The blade 14 has a first speed parameter corresponding to a difference in the product of a maximum permissible speed of the blower rotor 9 expressed in rad / s by the value 2.501x10 3 s / rad and the value 1.7148x10' 1 The maximum permissible 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'. The first speed parameter can be denoted Pv1.

[0080] It respects the relationship Pvl = 2.501 x 10 -3 x X n - 1.7148 x 10 -1 which can also be written as Pvl = 0.002501 x X n - 0.17148 and in which X n refers to the maximum permissible regime.

[0081] Blade 14 exhibits a second axial chord parameter, corresponding to the ratio of the axial chord at the blade root to the product of the square of the mean axial chord, the blade height 47, and the axial chord at the blade tip. This second ratio is calculated by expressing its constituent terms in meters; the value of this second ratio is considered dimensionless. The second axial chord parameter can be denoted Pc2.

[0082] It respects the relation Pc2 = — .

[0083] C xm xhxC xl00

[0084] The blade 14 has a second speed parameter corresponding to a difference in the product of the maximum permissible speed of the blower rotor 9, expressed in rad / s, and the value 2.515x10' 1 s / rad and the value 52.701. The second speed parameter can be noted as Pv2.

[0085] It respects the relationship Pv2 = 2.515 x 10 -1 x X n- 52.701 which can also be written as Pv2 = 0.2515 x X n - 52,701.

[0086] To optimize the performance of propulsion system 1:

[0087] - the first axial chord parameter is greater than or equal to the first velocity parameter, and

[0088] - the second axial chord parameter is greater than or equal to the second velocity parameter.

[0089] The first axial chord parameter and the second axial chord parameter are chosen here to be sufficiently large to significantly increase the margin with the natural mode of vibration in bending of the blade and the margin with the natural mode in torsion of the blade.

[0090] Bending here refers to 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 with the natural mode of vibration in bending depends in particular on the difference between the blade's rotational frequency at maximum permissible speed and the frequency of the natural mode of vibration in bending.

[0091] Torsion here refers to a deformation of the blade during which the blade tip rotates around a radial axis relative to the blade root, the blade curving around a direction radial to the drive axis, i.e., the X-axis of the propulsion system. The margin with the blade's natural torsional mode depends, in particular, on the difference between the blade's rotational frequency at maximum permissible speed and the frequency of its natural torsional vibration mode.

[0092] When the first axial chord parameter is greater than or equal to the first velocity parameter, and the second axial chord parameter is also greater than or equal to the second velocity parameter, the fan becomes more stable relative to the blade's natural mode of vibration in bending under greater external disturbances. Stable fan operation becomes possible by reducing the ratio of the length between the shroud inlet and the fan to the fan diameter. This reduces the engine mass and engine drag through the shroud drag.

[0093] We can also define a third speed parameter for Pv3 and a fourth speed parameter for Pv4.

[0094] Blade 14 presents the third speed parameter Pv3 which corresponds to a difference of a product of a maximum permissible speed of the blower rotor 9 expressed in rad / s by the value 3.12x10' 3 s / rad and the value 2, 14x10' 1 .

[0095] It respects the relationship Pv3 = 3.12 x 10 -3 x X n - 2.14 x 10 -1 which can also be written as Pv3 = 0.00312 x X n - 0.214.

[0096] Blade 14 presents the fourth speed parameter Pv4 which corresponds to a difference of a product of the maximum permissible speed of the blower rotor 9 expressed in rad / s by the value 3.14x10' 1 s / rad and the value 65.9.

[0097] It respects the relationship Pv4 = 3.14 x 10 -1 x X n - 65.9 which can also be written as Pv4 = 0.314 x X n - 65.9.

[0098] Advantageously, the first axial chord parameter is less than or equal to the third velocity parameter and the second axial chord parameter is less than or equal to the fourth velocity parameter.

[0099] 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.

[0100] 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 7, 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 7 shows a cross-section in this vertical plane.

[0101] 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

[0102] 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

[0103] 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.

[0104] 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.

[0105] Advantageously, it can also be required that this ratio be greater than or equal to 0.2.

[0106] 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.

[0107] The fan rotor 9 has an internal radius Ri and an external radius Re. The internal radius Ri corresponds to the distance between the axis of rotation X and the point of intersection between the leading edge 14a and the inner platform 34 at the inlet of the fan rotor 9 (and corresponds to the point where the leading edge 14a connects to the aerodynamic surface of a platform of the fan rotor 9). The external radius Re corresponds to the distance between the axis of rotation X and the point of intersection between the leading edge 14a and the blade tip 21 (and corresponds to half the fan diameter D). The ratio between the internal radius Ri and the external radius Re of the fan rotor 9 is called the hub ratio. The lower the hub-to-head ratio, the more efficient the fan rotor 9. However, decreasing the hub-to-head ratio of the fan rotor 9 results in an increase in the mechanical load on the hub 13 of the fan rotor 9.

[0108] The blower rotor 9 preferably comprises at least ten blower blades 14 and at most twenty-four blower blades 14, even more preferably at least seventeen blower blades 14 and at most twenty-two blower blades 14. Figure 6 illustrates the case where the blower rotor 9 comprises twenty-two blades.

[0109] 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.

[0110] 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.

[0111] 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.

[0112] The primary body 3 comprises a compressor section 29, a combustion chamber 6 and a turbine section 30.

[0113] Compressor section 29 includes a low pressure compressor 4 and a high pressure compressor 5.

[0114] The low-pressure compressor 4 comprises a rotor 41 designed to be driven in rotation relative to the housing 31 of the propulsion system 1 and a stator 42 fixedly mounted on the housing 31. 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.

[0115] 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.

[0116] The rotor 51 of the high-pressure compressor 5 includes movable wheels 5a and the stator 52 of the high-pressure compressor 5 includes fixed wheels 5b. The movable wheels 5a are arranged alternately with the fixed wheels 5b, thus forming a succession of high-pressure compressor stages.

[0117] Turbine section 30 includes a high-pressure turbine 7 and a low-pressure turbine 8.

[0118] 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.

[0119] 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.

[0120] 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.

[0121] 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.

[0122] 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.

[0123] 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. The propulsion system 1 further includes 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 in rotation 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.

[0124] Thanks to the reduction mechanism 19, the blower rotor 9 is driven into rotation at a speed lower than the rotational speed of the rotor 41 of the low pressure turbine 4.

[0125] 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.

[0126] 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.

[0127] 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.

[0128] 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.

[0129] 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.

[0130] The low-pressure shaft 11 and the high-pressure shaft 10 can be co-rotating, that is, 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, that is, driven in opposite directions of rotation around the longitudinal axis X. The twin-body propulsion system 1 can in particular include a single-stage high-pressure turbine 7, that is, comprising exactly one stage, or a two-stage high-pressure turbine 7, that is, comprising exactly two stages (as illustrated in the example in Figure 2).

[0131] The high-pressure compressor 5 comprises at least eight stages (as illustrated in the example in Figure 2) and at most eleven stages.

[0132] The low-pressure turbine 8 comprises at least three stages (as illustrated in the example in Figure 2) and at most seven stages.

[0133] The low-pressure compressor 4 comprises at least two stages and at most four stages.

[0134] When propulsion system 1 is in operation, an airflow F entering propulsion system 1 passes through the blower 22 and is then divided between a primary airflow F1 and a secondary airflow F2, which flow upstream to downstream in propulsion system 1 in the direction of gas flow through propulsion system 1.

[0135] The secondary airflow F2, also called the "bypass airflow", flows in the secondary vein, around the primary body 3. The secondary airflow F2 helps to cool the periphery of the primary body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0136] 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.

[0137] To improve the propulsive efficiency of propulsion system 1 and reduce its specific fuel consumption and the noise emitted by fan section 2, propulsion system 1 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.

[0138] In a propulsion system 1 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 1 with a high dilution ratio, the bulk 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 air flow duct at the inlet of the blower rotor 9 to the top 21 of blower blade 14.

[0139] 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.

[0140] In a direct-drive propulsion system 1, 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.

[0141] The propulsion system 1 is configured to provide a thrust of between 18,000 Ibf (80,068 N) and 70,000 Ibf (331,374 N), preferably between 20,000 Ibf (88,964 N) and 51,000 Ibf (226,859 N), or even between 20,000 Ibf (88,964 N) and 35,000 Ibf (155,688 N).

[0142] The reduction mechanism 19 may include an epicycloidal or planetary reduction mechanism, single-stage or two-stage.

[0143] 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.

[0144] 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.

[0145] Regardless of the configuration of the reduction mechanism 19, the diameter of the ring gear 19b and the satellite 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.

[0146] 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.

[0147] 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.

[0148] Example of a completed project

[0149] We now present two examples of the realization of a propulsion system, each comprising a blower section as previously described.

[0150] Common to both examples, propulsion system 1 is a twin-shaft gas turbine engine with a shrouded fan. The fan rotor comprises 18 composite material blades that are fixed relative to the fan hub.

[0151] The blower rotor 9 has a hub-to-head ratio of 0.27

[0152] The blower section includes a blower stator fixed to the blower housing. The blower stator comprises 40 fixed blades that are fixed relative to the blower housing.

[0153] The low-pressure and high-pressure shafts are co-rotating. The propulsion system includes a two-stage high-pressure turbine. The high-pressure compressor has 10 stages. The low-pressure turbine has 6 stages. The low-pressure compressor has 3 stages. The propulsion system has a bypass ratio of 10. The propulsion system provides a thrust of 120,000 N. The diameter of the blower rotor is 1.98 m.

[0154] In the first example, the engine does not meet the requirements for a short air intake (L / D < 0.4). If a short air intake were used on this engine, vibration levels in crosswind or high angles of attack would be unacceptable. In the second example, the engine is improved because it meets the requirements for a short air intake (L / D < 0.4).

[0155] A short air intake allows for a reduction in the mass and drag of the nacelle and therefore a reduction in fuel consumption.

[0156] According to this table, the inequality Pv1 < Pc1 is respected in the case of the second example but not in the case of the first example and the inequality Pv2< Pc2 is respected in the case of the second example but not in the case of the first example.

Claims

DEMANDS 1. Fan section (2) of an aeronautical propulsion system (1), the fan section (2) being enclosed 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), - a fan rotor (9) comprising a blade (1) 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 base of the blade and the axial chord at the top of the blade, in which the blade (14) presents: - a first 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 square of the blade height (47) by the axial chord at the tip of the blade, - a first speed parameter corresponding to a difference of a product of a maximum permissible speed of the blower rotor (9) expressed in rad / s by the value 2.501x10' 3 s / rad and the value 1.7148x10' 1 , - a second axial chord parameter corresponding to a ratio of the axial chord at the blade root to a product of the square of the average axial chord by the blade height (47) and by the axial chord at the blade tip, and - a second speed parameter corresponding to a difference of a product of the maximum permissible speed of the blower rotor (9) expressed in rad / s by the value 0.2515 s / rad and the value 52.701, the blower section being configured so that the blower rotor has a maximum permissible speed greater than or equal to 280 rad / s, the blade (1) being configured such that the first axial chord parameter is greater than or equal to the first velocity parameter and such that the second axial chord parameter is greater than or equal to the second velocity parameter.

2. Blower section (2) according to claim 1, wherein the blade (14) has: - a third speed parameter corresponding to a difference in the product of a maximum permissible speed of the blower rotor (9) expressed in rad / s by the value 3.12x10' 3 s / rad and the value 2.14x10' 1 , - a fourth speed parameter corresponding to a difference of a product of the maximum permissible speed of the blower rotor (9) expressed in rad / s by the value 0.314 s / rad and the value 65.9, the blade (14) being configured so that the first axial chord parameter is less than or equal to the third speed parameter, the second axial chord parameter being less than the fourth speed parameter.

3. Blower section (2) according to any one of claims 1 and 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. Blower section (2) according to any one of claims 1 to 3, 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 17 and less than or equal to 22.

5. Blower section (2) according to claim 4 in which the blower (2) comprises a blower hub (13), each blade (14) being pivotally mounted relative to the blower hub (13) so as to have variable pitch.

6. Aeronautical propulsion system (1) comprising: - a drive shaft (11) that rotates about an axis (X); - a blower shaft (20); - a blower section (2) according to any one of claims 1 to 5, the blower rotor (9) being driven in rotation by the blower 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).

7. Propulsion system (1) according to claim 6, configured to provide a thrust of between 80,068 N and 331,374 N, preferably between 88,964 N and 155,688 N.

8. Propulsion system (1) according to any one of claims 6 and 7, 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.

9. Propulsion system (1) according to any one of claims 6 to 8, comprising a nacelle (28) surrounding the blower 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 (L) 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 casing 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 casing below the axis (X) and the leading edge of the blade when the blade is below the axis (X), a ratio of the average distance (L) to a diameter (D) of the blower rotor (9) being less than or equal to 0.

4.

10. Method for dimensioning a fan section (2) of an aeronautical propulsion system (1), the fan section (2) being enclosed 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), - a fan rotor (9) comprising a blade (1) 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) presents - a first 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 square of the blade height (47) by the axial chord at the tip of the blade, - a first speed parameter corresponding to a difference of a product of a maximum permissible speed of the blower rotor (9) expressed in rad / s by the value 2.501x10' 3 s / rad and the value 1.7148x10' 1 , - a second axial chord parameter corresponding to a ratio of the axial chord at the blade root to a product of the square of the average axial chord by the blade height (47) and by the axial chord at the blade tip, and - a second speed parameter corresponding to a difference of a product of the maximum permissible speed of the blower rotor (9) expressed in rad / s by the value 0.2515 s / rad and the value 52.701, the blower section being configured so that the blower rotor has a maximum permissible speed greater than or equal to 280 rad / s, the process comprising a step of dimensioning the blower section (2) so that the first axial chord parameter is greater than or equal to the first speed parameter and the second axial chord parameter is greater than or equal to the second speed parameter.

11. Dimensioning method according to claim 10, wherein the blower section (2) is further dimensioned so that a thrust of the aeronautical propulsion system (1) is between 80,068 N and 331,374 N, preferably between 88,964 N and 155,688 N.

12. Sizing method according to claim 10 or 11, wherein the blower section (2) is further sized so that 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.

13. Method of manufacturing a fan section (2) of an aeronautical propulsion system (1) comprising dimensioning the fan section according to any one of claims 10 to 12 and manufacturing the fan section.

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