Aircraft propulsion system equipped with a propeller and a cyclic pitch control device for the propeller blades, and method for regulating the cyclic pitch control of the propeller blades

Abstract

Propulsion assembly (10) for aircraft, the propulsion assembly (10) extending along an axis (X) and comprising a propulsion module (20) having a propeller (22) provided with blades (22A), a straightener (24), and a propeller shaft (26) configured to drive the propeller (22) in rotation, the blades (22A) of the propeller (22) being in whole or in part made of composite material; a gas generator (30) having a drive shaft (33A); a speed reduction device (40) coupling in rotation the drive shaft (33A) and the propeller shaft (26), and configured to drive the propeller shaft (26) at a rotational speed lower than the rotational speed of the drive shaft (33A); and a cyclic pitch control device for the propeller blades (50). Figure for the summary: Fig. 2.

Description

Title of the invention: Aircraft propulsion system equipped with a propeller and a cyclic pitch control device for the propeller blades, and method for regulating the cyclic pitch control of the propeller blades. Technical field

[0001] The present exposition relates to an aircraft propulsion assembly equipped with a propeller and a cyclic pitch control device for propeller blades and a method for regulating the cyclic pitch control of the propeller blades.

[0002] For the purposes of this description, the term "aircraft propulsion system equipped with a propeller" refers to all gas turbine devices producing thrust necessary for the propulsion of an aircraft, particularly an airplane, by reaction to the high-speed ejection of gas, essentially by the propeller. The term "propeller" refers to an unducted fan rotor. Aircraft propulsion systems equipped with at least one propeller are also known by the English terms "open fan" or "unducted fan". Previous technique

[0003] Much research is focused on improving the efficiency of aeronautical engines, particularly aircraft engines, in order to reduce their environmental impact. One of the identified approaches for improving the propulsive efficiency of aircraft engines, reducing their fuel consumption, and the noise generated by the propulsion module is to increase the bypass ratio (BPR) as much as possible. The bypass ratio is the ratio of the mass flow rate of the secondary airflow to the mass flow rate of the primary airflow. One solution to achieve this is to indirectly couple the propulsion module and the gas generator, for example, via a speed reduction device. This allows for the independent optimization of the rotational speed of the moving parts of the gas generator and the rotational speed of the propulsion module's rotor.However, improving the bypass ratio generally also implies an increase in the diameter of the propulsion module, and consequently in the external dimensions of the propulsion system. This results in an increase in the mass, fuel consumption, and drag of the propulsion system, as well as more difficult integration within the aircraft.

[0004] Furthermore, among engines exhibiting a high bypass ratio, propeller propulsion assemblies are subject to a mechanical stress known as "IP" which is predominant and must be taken into account in the sizing and design of the engine. This IP stress (or IP moment) is the result of a pressure differential This stress is exerted by the airflow acting on the propeller blades, due to the propeller's rotation. Since the airflow is generally not parallel to the propeller's axis of rotation, a pressure differential is exerted by this airflow on diametrically opposed blades of the propeller. This generates a resultant stress on the entire propeller, as well as on the propeller shaft and the entire drive system transmitting motive power to the propeller. This stress is significant and must be considered when sizing the engine. Also, the propellers and the power transmission chain of aircraft engines generally have to be particularly strong from a mechanical point of view, and are sometimes equipped with a cyclic stalling mechanism, active or passive (see for example FR2997138, FR3067415 or FR3101664), but all this can significantly penalize the mass of the engine, and therefore the efficiency of the engine.Therefore, there is a need in this direction, for optimizations. Description of the invention

[0005] One embodiment relates to an aircraft propulsion system, the propulsion system extending along an axis and comprising a propulsion module having a propeller equipped with blades, a stator, and a propeller shaft configured to drive the propeller in rotation, the propeller blades being wholly or partly made of composite material; a gas generator having a drive shaft; a speed reduction device coupling the drive shaft and the propeller shaft in rotation, and configured to drive the propeller shaft at a rotational speed lower than the rotational speed of the drive shaft; and a cyclic pitch control device for the propeller blades. Generally, the cyclic pitch control device is associated on the one hand with a stator in the propulsion system and on the other hand with the propeller, which is connected to the propeller shaft associated with the speed reduction device.

[0006] Generally, in this description and unless otherwise indicated, upstream and downstream are defined with respect to the normal direction of fluid flow (upstream to downstream) through the propulsion assembly. Furthermore, the axial direction corresponds to the direction of the axis of the propulsion assembly, and a radial direction is a direction perpendicular to the axis. The azimuthal or circumferential direction corresponds to the direction describing a ring around the axial direction. The three directions axial, radial, and azimuthal correspond respectively to the directions defined by the ridge, radius, and angle in a cylindrical coordinate system. Finally, unless otherwise specified, the adjectives "inside" / "internal" and "outside" / "external" are used with reference to a radial direction such that the inner (i.e., radially inner) part of an element is closer to the axis than the outer (i.e., radially inner) part.radially outer) of the same element.

[0007] Hereafter, and unless otherwise indicated, "propulsive assembly" means "Propulsion system for aircraft". Hereafter, and unless otherwise indicated, "cyclic pitch control device" means "cyclic pitch control device for propeller blades".

[0008] A stator is a stationary blade wheel rotating about the axis of the propulsion assembly (i.e., the stator is a stator), while the propeller is a rotating blade wheel rotating about the axis of the propulsion assembly (i.e., the propeller is a rotor). The stator, generally located downstream of the propeller, has the function of straightening the airflow downstream of the propeller along the axis.

[0009] The gas generator can be a single, double or triple body gas generator, and comprises from upstream to downstream, along the axis, a compressor (or compressor section), a combustion chamber, and a turbine (or turbine section).

[0010] For the purposes of this disclosure, the term "composite material" means a material comprising reinforcing fibers, for example, long reinforcing fibers, for example, with a length greater than or equal to 1 cm (one centimeter), embedded in a polymeric matrix, for example, an epoxy resin. For example, the fibers may comprise strands of carbon fibers. As another example, the fibers may comprise strands of glass fibers within strands of carbon fibers. Examples of such materials are described in EP2588758, WO2022018353, or WO2022208002. For example, 3D woven or laminated blades are considered to be composite material blades.Each propeller blade can be made entirely of composite material, or it can comprise one part, for example the airfoil, made of composite material, and another part, for example the root and / or an internal spar extending longitudinally from the root inside the airfoil, made of another material, for example metal.

[0011] As a reminder, the pitch or pitch angle of a blade corresponds to the angle that a chord of the blade makes, the chord being an abstract geometric segment extending, at a given height of the airfoil, between the leading edge and the trailing edge of the blade, with the axis of rotation of the propeller in which the blade is mounted. A cyclic pitch adjustment device is a device configured to adjust the pitch of each propeller blade according to its angular position around the propeller axis, during the rotation of the propeller around its axis. The cyclic pitch of the blades differs from a possible collective pitch of the blades in that it is specific to each blade and different for each of the propeller blades, at least for the adjacent blades within the propeller. A collective pitch corresponds to a pitch common to all the propeller blades.In other words, a cyclic pitch angle can be considered as a compensation or correction pitch angle, specific to each blade taken individually and a function of the angular position of the blade. around the propeller's axis of rotation, a collective pitch angle (fixed or variable) common to all blades is used. A cyclic pitch control device can be configured to adjust only the cyclic pitch of the blades, or to adjust both the cyclic pitch and any collective pitch of the blades. A cyclic pitch control device can be separate from any collective pitch control device, or it can also form part of a collective pitch control device.

[0012] The inventors have identified that, in the context of an open-fan propulsion system, the cyclic pitch control device significantly reduces the mechanical stresses on each of the propeller blades related to IP stresses. This makes it possible to use blades made entirely or partially of composite material, which offers different mechanical strength than more traditional materials such as metal, and is also lighter. The combination of a cyclic pitch control device and blades made entirely or partially of composite material creates a synergy that reduces the mechanical stresses generated by IP stresses within the engine torque transmission chain to the propeller. This allows for more favorable sizing of the various components involved in terms of mass, and therefore improves efficiency.This also allows for better balance of the aircraft on which the propulsion system is mounted, for example by specifically managing the overhang of the propulsion system relative to the wing that supports it on the aircraft.

[0013] In some embodiments, the propeller shaft may be coaxial with the drive shaft.

[0014] In other words, the axis of the propeller shaft and the axis of the drive shaft can be coaxial and coincide with the axis of the propulsion assembly.

[0015] Such a configuration allows the use of a simpler and relatively lighter external housing for the gas generator. Furthermore, the stability of the gas generator's feed airflow is improved, while the portion of the propeller blades radially closer to the axis provides initial compression of the gas generator's feed airflow upstream of the gas generator, thereby improving efficiency.

[0016] In some embodiments, the cyclic timing device for the propeller blades may comprise exactly three or four cylinders.

[0017] The cylinders can be of the "single chamber" or "double chamber" type. All the cylinders can be of the same type, but not necessarily. Generally, they are associated with a stator in the propulsion assembly and allow control of a displacement associated with the propeller in the cyclic pitch control device of the propeller blades via components rotating in correlation with the propeller shaft.

[0018] Such a number of cylinders allows for optimization of the mass of the shimming device cyclic stability while ensuring a satisfactory level of safety. Indeed, three cylinders are sufficient to ensure the operation of the cyclic stabilization device. For example, if at least one of these three cylinders (therefore, by providing exactly three cylinders) is of the "double chamber" type, sufficient redundancy is obtained to ensure the required availability of the cyclic stabilization device with an acceptable impact on the overall system mass. According to an alternative, by using "single chamber" type cylinders, a fourth cylinder (therefore, by providing exactly four cylinders) ensures sufficient redundancy to guarantee the required availability of the cyclic stabilization device with an acceptable impact on the overall system mass.

[0019] In some embodiments, the cylinders can be regularly distributed circumferentially around the axis.

[0020] Such a configuration allows the mass of the cylinders to be distributed homogeneously within the propulsion assembly, a good distribution of the forces generated by the cylinders, and therefore an optimization which makes it possible to reduce the capacity, and therefore the mass of the cylinders, as much as possible.

[0021] In some embodiments, the cyclic pitching device for the propeller blades may include at least one cylinder, the at least one cylinder being configured to adopt a position within a total collective pitching stroke, and to allow a cyclic pitching stroke of between ± 40 mm (plus or minus forty millimeters), for example between ± 20 mm (plus or minus twenty millimeters), for example between ± 16 mm (plus or minus sixteen millimeters), for example between ± 9.6 mm (plus or minus nine millimeters and six tenths of a millimeter), around said position.For example, all the cylinders can be identical and configured to adopt a position within a total collective stalling stroke, and to allow a cyclic stalling stroke of ± 40 mm (plus or minus forty millimeters), for example between ± 20 mm (plus or minus twenty millimeters), for example between ± 16 mm (plus or minus sixteen millimeters), for example between ± 9.6 mm (plus or minus nine millimeters and six tenths of a millimeter), around said position.

[0022] It is understood that the cylinder may have a sliding rod, this sliding rod having a reference point, for example the distal end of the rod, and that this reference point is movable between two extreme positions defining the total stroke of the cylinder. This total stroke allows for collective adjustment according to a total collective adjustment stroke and cyclic adjustment for all cylinder positions within the total collective adjustment stroke. Such a configuration makes it possible to use cylinders with sufficient stroke and optimized mass.

[0023] In certain embodiments, the cyclic timing device for the propeller blades may include a timing ring that is ball-jointed around the axis and slides parallel to it. element to the axis.

[0024] It is understood that the shim ring is pivoted around the axis, for example to ensure cyclic shimage of the blades. The shim ring slides along the axis, for example to ensure collective shimage. Such a ring can reduce the overall mass of the system, improve guidance, the rigidity of the assembly, and the shimage accuracy.

[0025] In some embodiments, the shim ring can be mounted to slide on the propeller shaft or on a stator.

[0026] Mounting on the propeller shaft or on a stator optimizes the system's size and kinematic efficiency, and indirectly reduces the overall mass. Such a mounting can also limit the kinematic overhang, resulting in improved rigidity and precise alignment.

[0027] In some embodiments, the shoring ring may have an internal radius between 150 mm (one hundred and fifty millimeters) and 450 mm (four hundred and fifty millimeters), for example 225 mm (two hundred and twenty-five millimeters).

[0028] Such dimensions allow for optimization of the system's size and kinematic efficiency, and indirectly optimize the overall mass. For example, this allows the ring to be integrated into the space available under the propeller hub.

[0029] In some embodiments, the shim ring can be spherically articulated over an angular range between ± 30° (plus or minus thirty degrees of angle), for example between ±15° (plus or minus fifteen degrees of angle), for example between ± 10.5° (plus or minus ten degrees of angle and five tenths of degrees of angle), for example about ± 4.0° (plus or minus four degrees of angle).

[0030] Such a wide range of ball joint movement ensures the required efficiency for an "open fan" type propulsion system in its aerodynamic context, on an aircraft, for example an airplane, in flight, while being dimensioned as precisely as possible, which indirectly optimizes the overall mass. Furthermore, a large angular range can improve the system's accuracy.

[0031] In some embodiments, the cyclic pitching device for the propeller blades may include a pressure accumulator configured to provide safety control energy to bring the blades (i.e. all the blades) into feathering.

[0032] The "feather" position of the blades corresponds to the blade position that minimizes the propeller's frontal area. In other words, the pitch associated with the "feather" position is the pitch that minimizes the propeller's drag relative to the airflow passing through the propeller.

[0033] Such a pressure accumulator is a reliable safety system (particularly due to the physical proximity between the accumulator and the actuators or cylinders, which improves the chances of the system resisting extreme accidents such as the loss of a propeller blade) and relatively lightweight compared to other conceivable safety systems.

[0034] In some embodiments, the cyclic pitch angle can be between ± 30° (plus or minus thirty degrees of angle), e.g. ± 6° (plus or minus six degrees of angle).

[0035] Such a wide cyclic pitch angle range ensures the required efficiency for an "open fan" type propulsion assembly in its aerodynamic context, on an aircraft, for example an airplane, in flight, while being dimensioned as precisely as possible, which indirectly allows the mass of the assembly to be optimized.

[0036] In some embodiments, the propeller may have a diameter measured at the leading edge greater than or equal to 1.98 m (one meter and ninety-eight hundredths of a meter) and less than or equal to 6.12 m (six meters and twelve hundredths of a meter), for example greater than or equal to 1.98 m (one meter and ninety-eight hundredths of a meter) and less than or equal to 4.30 m (four meters and thirty hundredths of a meter).

[0037] Such blades make it possible to optimize the overall efficiency of the propulsion system mounted on an aircraft, for example an airplane, according to the needs associated with the aircraft category (e.g., number of seats), which indirectly optimizes the overall mass. Furthermore, a large propeller diameter can improve propulsive efficiency despite the resulting increase in mass.

[0038] In some embodiments, the propeller may comprise at least 10 (ten) blades and at most 18 (eighteen) blades.

[0039] Such a number of blades makes it possible to ensure the required aerodynamic and acoustic efficiency, and therefore a satisfactory overall efficiency of the propulsion system, while optimizing the mass of the propulsion system.

[0040] In some embodiments, the helix may have a hub-to-head ratio greater than or equal to 0.22 (twenty-two hundredths) and less than or equal to 0.35 (thirty-five hundredths), for example greater than or equal to 0.25 (twenty-five hundredths) and less than or equal to 0.35 (thirty-five hundredths), for example less than or equal to 0.27 (twenty-seven hundredths).

[0041] The hub-to-head ratio is the ratio of the propeller's inner radius divided by its outer radius. The inner radius corresponds to the radial distance between the propeller's axis of rotation and the point of intersection of the leading edge of the propeller blade airfoil with the aerodynamic surface of the inner inter-blade platform. The outer radius corresponds to the distance between the propeller's axis of rotation and the The hub-to-tip ratio is the point of intersection between the leading edge of the propeller blades' airfoil and the tips of the propeller blades (and corresponds to half the propeller diameter). The smaller the hub-to-tip ratio, the more efficient the propeller and the greater the mechanical load on the hub.

[0042] Such a hub-head ratio makes it possible to ensure the required aerodynamic and acoustic efficiency, and therefore a satisfactory overall efficiency of the propulsion system, while optimizing the mass of the propulsion system.

[0043] In some embodiments, the propeller may comprise at least 10 (ten) blades and at most 16 (sixteen) blades and the hub-to-head ratio may be greater than or equal to 0.25 (twenty-five hundredths) and less than or equal to 0.30 (thirty hundredths).

[0044] Such a combination of the number of blades and the hub-to-head ratio makes it possible to ensure aerodynamic and acoustic efficiency, and therefore an overall efficiency of the propulsion system optimized, which makes it possible to further optimize the mass of the propulsion system.

[0045] In some embodiments, the propeller may comprise at least 14 (fourteen) blades and at most 18 (eighteen) blades and the hub-to-head ratio may be greater than or equal to 0.30 (thirty hundredths) and less than or equal to 0.35 (thirty-five hundredths).

[0046] Such a combination of the number of blades and the hub-to-head ratio makes it possible to ensure aerodynamic and acoustic efficiency, and therefore an overall efficiency of the propulsion system, which makes it possible to further optimize the mass of the propulsion system.

[0047] In some embodiments, the reduction mechanism may have a reduction ratio greater than or equal to 2.5 (two and five tenths) and less than or equal to 11.0 (eleven), for example greater than or equal to 2.7 (two and seven tenths) and less than or equal to 6.0 (six), for example greater than or equal to 2.7 (two and seven tenths) and less than or equal to 3.6 (three and six tenths), for example around 3.0 (three).

[0048] Such a reduction ratio makes it possible to optimize on the one hand the efficiency of the gas generator turbine and on the other hand the efficiency of the propeller as a function of its outside diameter at the blade head, which indirectly makes it possible to optimize the mass and efficiency of the propulsion assembly.

[0049] In some embodiments, the gas generator may comprise a high-pressure body and a low-pressure body.

[0050] It is understood that the high-pressure unit comprises a high-pressure compressor rotationally coupled to a high-pressure turbine via a high-pressure shaft. The low-pressure unit comprises a low-pressure compressor located upstream of the high-pressure compressor, and a low-pressure turbine located downstream of the high-pressure turbine, and rotationally coupled to the low-pressure compressor via a low-pressure shaft. The low-pressure shaft can form The gas generator drive shaft. The gas generator compressor includes low-pressure and high-pressure compressors. The gas generator turbine includes low-pressure and high-pressure turbines.

[0051] This makes it possible to optimize the overall efficiency of the propulsion system, and in particular the thrust, consumption, compressor and turbine efficiency in an "open fan" context, which indirectly makes it possible to optimize the mass of the propulsion system.

[0052] In some embodiments, the low-pressure body may include a low-pressure turbine, the low-pressure turbine having at least 3 (three) stages and at most 8 (eight) stages.

[0053] This makes it possible to optimize the overall efficiency of the propulsion system, and in particular the low-pressure turbine in an "open fan" context, which indirectly makes it possible to optimize the mass of the propulsion system.

[0054] In some embodiments, the low-pressure body may include a low-pressure compressor, the low-pressure compressor having at least 2 (two) stages and at most 5 (five) stages.

[0055] This allows the overall efficiency of the propulsion system to be optimized, and in particular the low-pressure compressors in an "open fan" context, which indirectly allows the mass of the propulsion system to be optimized.

[0056] In some embodiments, the high-pressure body may include a high-pressure turbine, the high-pressure turbine having 2 (two) stages.

[0057] This allows the overall efficiency of the propulsion system to be optimized, and in particular the high-pressure turbine in an "open fan" context, which indirectly allows the mass of the propulsion system to be optimized.

[0058] In some embodiments, the high-pressure body may include a high-pressure compressor, the high-pressure compressor having at least 8 (eight) stages and at most 11 (eleven) stages.

[0059] This allows the overall efficiency of the propulsion system to be optimized, and in particular the high-pressure compressors in an "open fan" context, which indirectly allows the mass of the propulsion system to be optimized.

[0060] In certain embodiments, the cyclic pitch control device for the propeller blades can be configured to regulate the cyclic pitch of the propeller blades as a function of one or more aircraft parameters on which the propulsion assembly is configured to be mounted, for example at least one parameter among the aircraft angle of attack, roll and yaw.

[0061] The aircraft angle of attack, measured in degrees, corresponds to the angle formed between the axis of the aircraft fuselage and the aircraft velocity vector projected onto the aircraft's median plane extending between the wings. The roll, measured in degrees, Yaw corresponds to the angular position of the aircraft around the fuselage axis relative to the horizontal reference position. Yaw, measured in degrees, corresponds to the angle formed between the aircraft's fuselage axis and the aircraft's velocity vector projected onto the aircraft's plane, including the wings.

[0062] A regulation based on such parameters makes it possible to optimize the overall efficiency of the propulsion system, considered in its aerodynamic environment within an aircraft, for example an airplane, in flight.

[0063] In some embodiments, the cyclic pitch control device for the propeller blades can be configured to regulate the cyclic pitch of the propeller blades as a function of one or more parameters of the gas generator, for example as a function of at least one parameter among the speed, power, torque, for example the torque of a low-pressure body.

[0064] A regulation based on such parameters makes it possible to optimize the overall efficiency of the propulsion system.

[0065] In some embodiments, the cyclic timing device for the propeller blades may include at least one sensor, for example disposed on at least one of the propeller shaft, a propeller bearing support, a control element for the cyclic timing device for the propeller blades, for example a timing ring or a jack, the at least one sensor being configured to determine an IP moment.

[0066] Such a sensor can enable the cyclic trim device to be controlled without having to use aircraft-specific parameters.

[0067] In some embodiments, the cyclic pitch control device for the propeller blades may include an inertial measurement unit configured to determine the angle of attack, roll and yaw of the aircraft on which the propulsion assembly is mounted.

[0068] Such an inertial navigation system can enable the cyclic pitch control device to be controlled without resorting to aircraft-specific parameters.

[0069] An embodiment relates to a method for regulating the cyclic pitch of the propeller blades of the propulsion assembly according to any one of the embodiments described in this exposition, in which the cyclic pitch of the propeller blades is regulated as a function of at least one aircraft parameter on which the propulsion assembly is configured to be mounted and / or at least one parameter of the gas generator. Brief description of the drawings

[0070] The purpose of this presentation and its advantages will be better understood upon reading the detailed description below of various embodiments given by way of non-limiting examples. This description refers to the attached figure pages, on which:

[0071] [Fig. 1] Figure 1 represents an aircraft equipped with a propulsion system,

[0072] [Fig.2] Fig.2 represents a cross-sectional view of the propulsion assembly of Fig.1,

[0073] [Fig. 3] Fig. 3 represents a schematic cross-sectional view of a device planetary type speed reduction

[0074] [Fig.4] Fig.4 represents a schematic cross-sectional view of an epicycloidal type speed reduction device,

[0075] [Fig.5] [Fig.5] represents a schematic cross-sectional view of the cyclic pitch control device of the propulsion assembly of [Fig.1], according to a first variant,

[0076] [Fig.6] Fig.6 represents a schematic cross-sectional view of the cyclic pitch control device of the propulsion assembly of Fig.1, according to a second variant,

[0077] [Fig.7] Fig.7 is a graph representing the cyclic pitch of the propeller blades as a function of the angular position of the blades

[0078] [Fig.8] Fig.8 represents a schematic view of the propeller blades of the propulsion assembly of Fig.1, as seen from arrow VIII of Fig.1, and

[0079] [Fig.9] [Fig.9] represents steps of a cyclic timing regulation process of the propeller blades of the propulsion assembly of [Fig.1]. Description of the implementation methods

[0080] Figure 1 represents an aircraft 100, in this example an airplane, equipped with two propulsion units 10, namely one propulsion unit 10 per wing 101, with only one propulsion unit 10 and one wing 101 being shown in Figure 1. According to one embodiment, the aircraft 100 can be equipped with more than one propulsion unit 10 per wing 101, each wing 101 having the same number of propulsion units 10. The reference symbol "A" designates the axis of the fuselage 102 of the aircraft 100. The propulsion unit 10 can be configured to propel the aircraft 10 at a cruising speed between Mach 0.7 and Mach 0.9.

[0081] Figure 2 shows a schematic cross-sectional view of the propulsion assembly 10, according to plane II of Figure 1. The propulsion assembly 10 extends along an X-axis and comprises a propulsion module 20, a gas generator 30, a speed reduction device 40, and a cyclic pitch control device 50. When the propulsion assembly 10 is mounted on the aircraft 100, the X-axis is not necessarily parallel to the A-axis.

[0082] The propulsion module 20 has a propeller 22 provided with a plurality of blades 22A, a stator 24 provided with a plurality of blades 24A, and a propeller shaft 26 configured to drive the propeller 22 in rotation. The propeller shaft 26 can extend along the X-axis. The blades 22A of the propeller 22 can be made entirely or partially of composite material. The blades 24A of the stator 24 can be made entirely or partially of composite material. For example, the propeller 22 can comprise between 10 and 18 blades 22A and the stator 24 can be made entirely or partially of composite material. Straightener 24 may include a smaller number of 24A blades, for example between 8 and 16 24A blades. The setting of the 24A blades of straightener 24 may be fixed or variable.

[0083] The propeller 22 may have a diameter D measured at the leading edge greater than or equal to 1.98 m and less than or equal to 6.12 m, for example greater than or equal to 1.98 m and less than or equal to 4.30 m. The propeller 22 may have a hub-to-head ratio RI / RE greater than or equal to 0.22 and less than or equal to 0.35, for example greater than or equal to 0.25 and less than or equal to 0.35, for example less than or equal to 0.27. The internal radius RI corresponds to the radial distance between the axis of rotation X of the propeller 22 and the point of intersection of the leading edge 22A1 of the airfoil profile of the blades 22A of the propeller 22 with the aerodynamic surface SA of the internal inter-blade platform. The external radius RE corresponds to the distance between the axis of rotation X of the propeller 22 and the point of intersection between the leading edge of the aerodynamic profile of the blades 22A of the propeller 22 and the top of the blades of the propeller (and corresponds to half of the diameter D of the propeller 22).

[0084] The propeller 22 may comprise at least 10 blades 22A and at most 18 blades 22A. According to one variant, the propeller 22 may comprise at least 10 blades 22A and at most 16 blades 22A and the hub-head ratio RI / RE may be greater than or equal to 0.25 and less than or equal to 0.30. According to another variant, the propeller 22 may comprise at least 14 blades 22A and at most 18 blades 22A and the hub-head ratio RI / RE may be greater than or equal to 0.30 and less than or equal to 0.35.

[0085] The gas generator 30 has a drive shaft 33A. The drive shaft can extend along the X axis. The propeller shaft 26 can be coaxial with the drive shaft 33A, and their respective axes of rotation can coincide with the X axis of the propulsion assembly 10. This allows for an annular air inlet within the gas generator 30 coaxial with the X axis, thanks to which the outer casing of the gas generator has a relatively simple shape and exhibits a certain rotational symmetry, which tends to reduce possible airflow disturbances. In this example, the gas generator 30 comprises, from upstream to downstream, the gases flowing within the propulsion assembly 100, from upstream to downstream, a compressor 32 (or compressor section 32), a combustion chamber 34, and a turbine 36 (or turbine section 36).

[0086] The gas generator 30 may be of the twin-body type and comprise a low-pressure body 30A and a high-pressure body 30B. The low-pressure body 30A may comprise a low-pressure compressor 32A rotationally coupled to a low-pressure turbine 36A via a low-pressure shaft 33A, which may form the drive shaft of the gas generator 30. The high-pressure body 30B may comprise a high-pressure compressor 32B disposed downstream of the low-pressure compressor 32A and upstream of the combustion chamber 34, and a high-pressure turbine The pressure compressor 36B, located downstream of the combustion chamber 34 and upstream of the low-pressure turbine 36A, is rotationally coupled to the high-pressure compressor 32B via a high-pressure shaft 33B. The compressor 32 of the gas generator 30 may include the low-pressure and high-pressure compressors 32A and 32B. The turbine 36 of the gas generator 30 may include the low-pressure and high-pressure turbines 36A and 36B. The low-pressure and high-pressure shafts 33A and 33B may be coaxial. The high-pressure shaft 33B may receive a portion of the low-pressure shaft 33A. According to one variant, the low-pressure shafts 33A and high-pressure shafts 33B can be co-rotating, i.e., configured to rotate relative to each other in the same direction around the X-axis. According to another variant, the low-pressure shafts 33A and high-pressure shafts 33B can be contra-rotating, i.e., configured to rotate relative to each other in opposite directions around the X-axis.The rotational speed of the low-pressure body 33A may be lower than the rotational speed of the high-pressure body 33B.

[0087] According to an unshown variant, the propulsion assembly may be of the three-shaft type. The turbine 36 may include an intermediate turbine arranged axially between the high-pressure turbine 36B and the low-pressure turbine 36A and configured to drive an intermediate compressor arranged axially between the low-pressure compressor 32A and the high-pressure compressor 32B via an intermediate shaft. The intermediate shaft may be located between the low-pressure shaft 33A and the high-pressure shaft 33B. The intermediate shaft and the low-pressure shaft 33B may be co-rotating or counter-rotating with respect to each other.

[0088] Each compressor 32A, 32B and turbine 36A, 36B may comprise a plurality of stages, each stage comprising a blade wheel, respectively 32AA, 32BA, 36AA, 36BA, movable in rotation about the X-axis (or rotor) and a blade wheel, respectively 32AB, 32BB, 36AB, 36BB, fixed about the X-axis (or stator). In this example, the low-pressure compressor 32A may have at least 2 stages and at most 5 stages, for example 2 stages, the high-pressure compressor 32B may have between 8 and 11 stages (only two stages being shown for clarity of the figure), the high-pressure turbine 36B may have 2 stages, and the low-pressure turbine 36A may have between 3 and 8 stages (only two stages being shown for clarity of the figure). A rectifier 37, or fixed paddle wheel rotating about the X axis, can be arranged downstream of the combustion chamber 34 and upstream of the high-pressure turbine 36B.

[0089] A speed reduction device 40 can indirectly couple the drive shaft 33A to the propeller shaft 26. The speed reduction device 40 can be configured to drive the propeller shaft 26 at a rotational speed lower than the rotational speed of the drive shaft 33A. The drive shaft 33A connects the low-pressure turbine 36A (or the low-pressure housing 30A) to a input of the speed reduction device 40 while the propeller shaft 26 connects an output of the speed reduction device 40 to the propeller 22. The propeller 22 is therefore driven by the low pressure turbine 36A (or the low pressure body 30A) via the drive shaft 33A (or low pressure shaft), the speed reduction device 40 and the propeller shaft 26. In this example, the speed reduction device 40 can be arranged, considered along the X-axis, between an upstream end of the drive shaft 33A and a downstream end of the propeller shaft 36.

[0090] For example, the speed reduction device 40 may be an epicyclic gear train reduction device, for example of the "epicyclic" or "planetary" type, according to the terminology sometimes used by those skilled in the art. Such a mechanism may comprise one stage, two stages, or more than two stages.

[0091] According to a first variant 40' shown schematically in [Fig.3], the reduction device 40 can be of the planetary or "star" type and include a sun pinion 40A, which forms the input of the reduction mechanism 40. The axis of rotation of the sun pinion 40A forms the axis of rotation of the reduction mechanism 40, and can be confused with the X axis of the propulsion assembly 10. The sun pinion 40A is configured to be driven in rotation by the drive shaft 33A. A ring gear 40B forms the output of the reduction mechanism 40. The ring gear 40B is coaxial with the sun gear 40A and configured to drive the propeller shaft 26 in rotation around the X-axis. Several satellite gears 40C, or planetary gears—only one satellite being shown in [Fig. 3]—are distributed circumferentially around the X-axis between the sun gear 40A and the ring gear 40B. Each satellite gear 40C is meshed with both the sun gear 40A and the ring gear 40B.The satellites 40C are mounted on a satellite carrier 40D which is fixed relative to a stator part 40E of the propulsion assembly 10, for example relative to a casing upstream of the compressor 32.

[0092] According to a second variant 40” shown schematically in [Fig. 4], the reduction device 40 can be of the epicyclic or “planetary” type. In this case, compared to the planetary type 40’ described with reference to [Fig. 3], the ring gear 40B is fixedly mounted on a stator portion 40E” of the propulsion assembly 10, and the propeller shaft 26 is driven in rotation by the planet carrier 40D” (which is therefore rotatable relative to the stator portion 40E” of the propulsion assembly 10, for example, relative to a housing upstream of the compressor 32). The stator portions 40E and 40E” can correspond to different parts of the same element, or to distinct elements.

[0093] Regardless of the 40' or 40'' configuration of the reduction device 40, the diameter of the ring 40B is greater than the diameter of the satellite carrier 40D, 40D” which is itself greater than the diameter of the solar pinion 40A, the satellites 40C are radially arranged between the solar pinion 40A and the crown 40B, and the rotational speed of the propeller shaft 26 is less than the rotational speed of the drive shaft 33A.

[0094] The reduction ratio of the reduction device 40 can be greater than or equal to 2.5 and less than or equal to 11.0, for example greater than or equal to 2.7 and less than or equal to 6.0, for example greater than or equal to 2.7 and less than or equal to 3.6, for example equal to 3.0.

[0095] In operation, an airflow F (see [Fig.2]) entering the propulsion assembly 10 passes through the propeller 22 and is then divided into a primary airflow Fl and a secondary airflow F2, which flow from upstream to downstream within the propulsion assembly 10.

[0096] The primary airflow Fl flows in a channel called the "primary channel", inside the gas generator 30, sometimes also called the primary body, passing successively through the low pressure compressor 32A, the high pressure compressor 32B, the combustion chamber 34, the high pressure turbine 36B, the low pressure turbine 36A, and then through the outlet nozzle 38. The expansion of the combustion gases downstream of the combustion chamber 34 within the turbine 36 provides the energy to drive the high and low pressure turbines 36B, 36A, and therefore the shafts 33A and 33B, into rotation.

[0097] The secondary airflow F2, sometimes also called the "bypass airflow", flows through the rectifier 24, then along the gas generator 30, outside the gas generator 30. This secondary airflow F2 provides by reaction the vast majority of the thrust generated by the propulsion assembly 10. The secondary airflow F2 can also be used to cool the gas generator 30 from the outside.

[0098] The bypass ratio (or BPR) of the propulsion assembly 10 is equal to the ratio of the mass flow rate of the secondary airflow F2 divided by the mass flow rate of the primary airflow Fl entering the gas generator 30. A high bypass ratio reflects the fact that most of the thrust is provided by the secondary airflow F2, and that the energy supplied by the primary airflow Fl is used primarily to generate the secondary airflow F2. In other words, in a propulsion assembly with a high bypass ratio, the primary airflow Fl is used primarily to generate the energy to drive the propulsion module, while the secondary airflow is used primarily to generate thrust. For example, the bypass ratio of the propulsion assembly 10 can be greater than or equal to 40, for example, greater than or equal to 40 and less than or equal to 80.

[0099] In a propulsion system with a high bypass ratio where the majority of the thrust is provided by the secondary flow F2, the kinetic energy of the secondary flow F2 is highly dependent on the compression produced by the propeller. One way to improve the propulsive efficiency, and therefore the overall efficiency of the propulsion system, is to reduce the pressure ratio of the propeller, and thus of the propulsion module. Such a reduction of the The pressure ratio can be obtained through a speed reduction device, which reduces the propeller's rotational speed but also increases the power extracted by the high-pressure turbine, further improving the overall efficiency of the propulsion system.

[0100] The propulsion assembly 10 can be configured to provide a thrust of between 18,000 Ibf (80,068 N) and 51,000 Ibf (22,241 N), for example between 20,000 Ibf (88,964 N) and 35,000 Ibf (15,568 N), when the propulsion assembly 10 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.

[0101] The cyclic timing device for the propeller blades 50 is described in more detail with reference to [Fig. 5]. The cyclic timing device 50 may comprise a plurality of cylinders 52, for example, exactly four cylinders 52 (only one cylinder being shown in [Fig. 5]). The cylinders 52 may be of the single-chamber type. According to an alternative not shown, the cyclic timing device 50 may comprise exactly three cylinders 52, at least one of the cylinders 52, or even all of the cylinders 52, being of the double-chamber type. The cylinders 52 may be regularly distributed circumferentially around the X-axis, for example, every 90° in the case where the cyclic timing device 50 comprises exactly four cylinders 52.The 52 cylinders can be configured to adopt a position P within a total collective shimming stroke CT, and to allow a cyclic shimming stroke CC between ± 40 mm, for example between ± 20 mm, for example between ± 16 mm, for example between ± 9.6 mm, around said position P. The 52 cylinders are for example hydraulic cylinders. Each cylinder 52 can be supplied with pressurized oil via a supply line 51, specific to each cylinder 52 and independent for each cylinder 52. These independent supply lines 51 can extend via arms 55A of the intermediate housing 55. In other words, the supply lines 51 can bypass the speed reduction device 40 and not extend through the speed reduction device 40. The pressurized oil supply circuit for the cylinders 52 can be independent and separate from the oil supply circuit for the speed reduction device 40.Such a configuration allows for a degree of modularity and facilitates maintenance as well as the assembly / disassembly of the cyclic camming device 50. Furthermore, the independent power supplies of the cylinders 52 reduce the risk of failure of the entire cyclic camming device 50 and prevent a risk of collective failure of the operability of the cylinders 52. In the event of a failure of one of the cylinders 52 or its power supply circuit, three cylinders 52 remain operational, which is sufficient to control the camming ring 54 described below.

[0102] The cylinders 52 can extend parallel to the X axis, but not necessarily. The cylinders 52 may each have a cylinder 52A mounted on a stator 70 of the propulsion assembly 10, for example on a bearing support housing 70A of the propeller shaft 26. The cylinders 52 may each have a rod 52B sliding relative to the cylinder 52A, each rod having for example a distal end, in this example a fixing eyelet, sliding over a total stroke C equal to the total collective trim stroke CT plus the cyclic trim stroke CC.

[0103] The cyclic shimming device 50 may include a shimming ring 54 which is stator-shaped like the cylinder 52 and which is ball-jointed about the X-axis and slides parallel to the X-axis. According to one variant 50' shown in [Fig. 5], the shimming ring 54 may be slidably mounted on the propeller shaft 26. According to another variant 50” shown in [Fig. 6], the shimming ring 54 may be slidably mounted on a stator 70, for example a bearing support housing 70A of the propeller shaft 26.

[0104] The shim ring 54 may include a ball-joint ring 54A slidably mounted about the X-axis on the propeller shaft 26 (see [Fig. 5]) or on the stator 70 (see [Fig. 6]), for example via rollers 53. These rollers 53 may be configured to allow the ball-joint ring 54A to slide about the X-axis and to rotate about the X-axis relative to the propeller shaft 26 (see [Fig. 5]) or relative to the stator 70 (see [Fig. 6]), and may thus form a sliding bearing. The ring 54A has a convex, ball-joint-shaped outer surface 54A1.

[0105] The shim ring 54 may include an inner ring 54B, for example forming an inner stator plate, arranged radially outside the ball-joint ring 54A. The inner ring 54B has a concave inner face 54B1 configured to cooperate by complementary shape with the convex outer surface 54A1 of the ball-joint ring 54A. This configuration is an example of providing the ball-jointed character about the X-axis of the shim ring 54. The inner ring 54B may have a plurality of arms 54B2 (only one arm being shown in Figures 5 and 6), extending axially and / or radially. For example, the inner ring 54B comprises as many arms 54B2 as there are cylinders 52. Each arm 54B2 can be mechanically connected to a (single) cylinder 52, for example to the distal end of a cylinder 52, for example via a pivot or ball joint.The inner ring 54B may include an outer face 54B3 assembled with a radially internal portion of a rolling bearing 56.

[0106] The shim ring 54 may include an outer ring 54C, for example forming an outer rotor plate, arranged radially outside the inner ring 54B. The outer ring 54C may include an inner face 54C1 assembled with a radially external portion of the bearing 56. In other words, the bearing 56 may be arranged radially between the inner ring 54B and the ring The outer ring 54C can thus be rotationally decoupled about the X-axis. In other words, the inner ring 54B and the outer ring 54C are rotationally movable relative to each other about the X-axis. The outer ring 54C can include an outer face 54C2 provided with a plurality of attachments 54C3, for example, eye brackets, to ensure a mechanical connection between the outer ring 54C and each of the blades 22A. For example, the outer ring 54C can include as many attachments 54C3 as there are blades 22A.

[0107] An oil circuit not shown supplies oil to the sliding bearing formed by the rollers 53, the ball joint formed by the rings 54A and 54B, and the bearing 56 of the shim ring 54.

[0108] The shim ring 54 may have an internal radius RR, which corresponds to the internal radius of the ball-joint ring 54A, between 150 mm and 450 mm, for example 225 mm. The shim ring 54 may be ball-jointed over an angular range C2 between ± 30°, for example between ± 15°, for example between ± 10.5°, for example approximately ± 4.0°.

[0109] Each blade 22A is pivotally mounted about its radially extending Z-axis on the hub 22B of the propeller 22, so as to allow adjustment of the pitch angle of each of the blades 22A (see double arrow AC in Figures 5 and 6), i.e., of the airfoil profile 22A2 of each of the blades 22A. More specifically, in this example, the foot 22A3 of each of the blades 22A is mounted on the hub 22B of the propeller 22 via a bearing 22C. Note that Figures 5 and 6 are highly schematic and that all the details of mounting the blade 22A on the hub 22B, which are otherwise well known to those skilled in the art, are neither shown nor described. The hub 22B is rotationally coupled to the shaft 26 via an assembly, for example, by bolts 23.

[0110] The foot 22A3 of each of the blades 22 can be provided with a pitch control lever 22D, configured to rotate the blade 22 about its Z-axis. The pitch control lever 22D can be connected to the pitch control ring 54 via a connecting rod 25. The connecting rod 25 can be connected to the pitch control ring 54 via a pivot joint or a ball joint. The connecting rod 25 can be connected to the pitch control lever 22D via a pivot joint or a ball joint.

[0111] When the propeller 22 rotates, the blades 22A, via their respective levers 22D and connecting rods 25, drive the outer ring 54C in rotation about the X-axis. The inner ring 54B, being rotationally decoupled from the outer ring 54C via the bearing 56, and rotationally coupled about the X-axis to the stator 70 via the arms 54B2 and the actuators 52, remains stationary in rotation about the X-axis. Depending on the inclination (provided by an angular stroke less than or equal to the angular range C2) of the inner ring 54B about the ball-jointed ring 54A, imposed by the actuators 52, the outer ring 54C follows during its rotation The inner ring 54B is tilted around the X-axis. Thus, the pitch of each of the blades 22B varies during the rotation of the propeller 22. At each predefined angular position around the X-axis, the blades 22A, during their successive passage through these predefined angular positions during the rotation of the propeller 22, acquire a pitch associated with each of these predefined angular positions. This constitutes an example of a cyclic pitch control device.

[0112] The cyclic timing device 50 may include a pressure accumulator 58 (see [Fig. 2]) configured to provide safety control energy to bring all the blades to the feathered position. For example, the pressure accumulator 58 may be a vacuum reservoir configured to purge oil from the cylinders 52 so as to bring them to a neutral configuration corresponding to the feathered position of the blades 22A. For example, the pressure accumulator 58 may be disposed of in an enclosure E housing the speed reduction device 40.

[0113] In operation, all the cylinders 52 can be controlled simultaneously and according to the same command, so that the shim ring 54 is moved along the X axis (see double arrow Cl in the example of Figures 5 and 6), thereby changing the shim position of all the blades 22A to achieve an identical shim angle for all the blades 22A. This allows the collective shim positioning of the blades 22A to be controlled and regulated. Each cylinder 52 can also be controlled independently of the other according to a setpoint which is specific to it and different from that of the other cylinders, so that the ball-jointed shim ring 54 pivots around a radial direction (see for example the double arrow associated with the angular range C2 in figures 5 and 6), by which means the shim of each of the blades 22A is modified in a specific way according to its angular position during the rotation of the propeller 22. This makes it possible to control and regulate the cyclic shim of the blades 22A.The cyclic pitch angle can be within ±30°, e.g., ±6° around the collective pitch position. For example, the angular range of the ball joint C2 of the pitching ring 54 can be configured so that the cyclic pitch angle is within ±30°, e.g., ±6° around the collective pitch position. For example, the cyclic pitch stroke CC of each of the cylinders 52 can be configured so that the cyclic pitch angle is within ±30°, e.g., ±6° around the collective pitch position. In this example, the pitching device 50 can allow both cyclic and collective pitching.

[0114] Figure 7 shows a graph, on the ordinate, of the pitch angle of the blades 22A as a function of their angular position within the propulsion assembly 10 during the rotation of the propeller 22. The position of the x-axis on the y-axis corresponds to the collective pitch angle, associated, for example, with the position P of the cylinders 22 within the collective pitch stroke CT. The curve V corresponds to a An example of cyclic pitch angles within the propeller 22, associated with a predetermined displacement of each of the cylinders within the cyclic pitch stroke CC, with an amplitude specific to each of the cylinders 52 and distinct from that of the other cylinders 52. Thus, in this example, during a complete revolution of the propeller 22, the cyclic pitch angle of each blade 22A follows this curve A, around the collective pitch corresponding to the abscissa. Figure 8 represents an example of a configuration of the blades 22A with cyclic pitch: as the propeller 22 rotates, the cyclic pitch of the blades 22A changes so that each of the blades 22A successively adopts the configuration shown.

[0115] The cyclic pitch control device 50 can be configured to regulate the cyclic pitch of the propeller blades 22A 22 as a function of one or more parameters of the aircraft 100 on which the propulsion unit 10 is mounted, for example, at least one parameter among the aircraft angle of attack, roll, and yaw. For example, the cyclic pitch control device for the propeller blades 50 may include an inertial measurement unit 57 configured to determine the angle of attack, roll, and yaw of the aircraft 100 on which the propulsion unit 10 is mounted. As an alternative or in addition, the cyclic timing device 50 can be configured to regulate the cyclic timing of the blades 22A of the propeller 22 according to one or more parameters of the gas generator 30, for example according to at least one parameter among the speed, power and torque, for example the torque of a low pressure body.The cyclic timing device 50 may include at least one sensor 60 configured to determine a moment IP, which may be disposed, for example, on the propeller shaft 26, a propeller bearing support (not referenced), a control element of the cyclic timing device 50, for example the timing ring 54 or a cylinder 52. In the example of Figures 5, the sensor 60 is disposed on an arm 54B2 of the timing ring 54. According to an example not shown, the sensor 50 may measure the pressure within the hydraulic chambers of the cylinders 52.

[0116] Figure 9 represents a method for regulating the cyclic pitch of the blades 22A of the propeller 22 of the propulsion unit 10, in which the cyclic pitch of the blades 22A of the propeller 22 is regulated as a function of at least one parameter of the aircraft 100 on which the propulsion unit 10 is mounted and / or at least one parameter of the gas generator 30. For example, the regulation method may include a control loop comprising a first step E1 during which at least one parameter of the aircraft 100 on which the propulsion unit 10 is mounted is collected, for example at least one parameter among the aircraft angle of attack, roll and yaw, and / or at least one parameter of the gas generator 30 and / or a moment measured by the sensor 60. The method may include a second step E2 during which a cyclic pitch angle is evaluated. The process may include a third step E3 during which the cyclic timing of the blades 22A of the propeller 22 is adjusted according to the result of the second step E2.

[0117] The present disclosure also relates to a computer program comprising instructions which, when the program is executed by a computer, cause the computer to implement the steps, for example E1, E2 and E3, for the execution of the method of regulating the cyclic pitch of the blades 22A of the propeller 22 of the propulsion unit 10. This program may use any programming language, and be in the form of source code, object code, or code intermediate between source code and object code, such as in a partially compiled form, or in any other desirable form.

[0118] The present disclosure also relates to a computer-readable recording medium on which the computer program is recorded. The recording medium can be any entity or device capable of storing a program. For example, the medium can include a storage means, such as a ROM, for example a CD-ROM or a microelectronic circuit ROM, or a magnetic recording means, for example a floppy disk or a hard disk drive.

[0119] Although the present invention has been described with reference to specific embodiments, it is evident that modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, individual features of the various embodiments illustrated / mentioned can be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than a restrictive sense.

[0120] It is also evident that all the characteristics described with reference to a process are transposable, alone or in combination, to a device, and conversely, all the characteristics described with reference to a device are transposable, alone or in combination, to a process.

Claims

Demands

1. Propulsion assembly (10) for an aircraft (100), the propulsion assembly (10) extending along an axis (X) and comprising: - a propulsion module (20) having a propeller (22) provided with blades (22A), a stator (24), and a propeller shaft (26) configured to drive the propeller (22) in rotation, the blades (22A) of the propeller (22) being wholly or partly made of composite material, - a gas generator (30) having a drive shaft (33A), - a speed reduction device (40) rotationally coupling the drive shaft (33A) and the propeller shaft (26), and configured to drive the propeller shaft (26) at a rotational speed lower than the rotational speed of the drive shaft (33A), and - a cyclic blade adjustment device the propeller (50).

2. Propulsion assembly (10) for aircraft (100) according to claim 1, wherein the propeller shaft (26) is coaxial with the drive shaft (33A).

3. Propulsion assembly (10) for aircraft (100) according to claim 1 or 2, wherein the cyclic pitch control device for the propeller blades (50) comprises exactly three or four cylinders (52).

4. Propulsion assembly (10) for aircraft (100) according to claim 3, in which the cylinders (52) are regularly distributed circumferentially around the axis (X).

5. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 4, wherein the cyclic pitching device for the propeller blades (50) comprises at least one actuator (52), the at least one actuator (52) being configured to adopt a position (P) within a total collective pitching stroke (CT), and to permit a cyclic pitching stroke (CC) of between ± 40 mm, for example between ± 20 mm, for example between ± 16 mm, for example between ± 9.6 mm, around said position (P).

6. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 5, wherein the cyclic pitching device of the propeller blades (50) include a shim ring (54) ball-jointed around the axis (X) and sliding parallel to the axis (X).

7. Propulsion assembly according to claim 6, in which the shim ring (54) is mounted sliding on the propeller shaft (26) or on a stator (70).

8. Propulsion assembly (10) for aircraft (100) according to claim 6 or 7, wherein the shoring ring (54) has an internal radius (RR) between 150 mm and 450 mm, for example 225 mm.

9. Propulsion assembly (10) for aircraft (100) according to any one of claims 6 to 8, wherein the shim ring (54) is spherically articulated over an angular range (C2) between ±30°, for example between ±15°, for example between ±10.5°, for example about ±4.0°.

10. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 9, wherein the cyclic pitching device for the propeller blades (50) includes a pressure accumulator (58) configured to provide safety control energy to bring the blades (22A) into feathering.

11. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 10, wherein the cyclic pitch angle is between ± 30°, e.g. ± 6°.

12. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 11, wherein the propeller (22) has a diameter (D) measured at the leading edge (22A1) greater than or equal to 1.98 m and less than or equal to 6.12 m, for example greater than or equal to 1.98 m and less than or equal to 4.30 m.

13. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 12, wherein the propeller (22) comprises at least 10 blades (22A) and at most 18 blades (22A).

14. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 13, wherein the propeller (22) has a hub-to-head ratio greater than or equal to 0.22 and less than or equal to 0.35, for example greater than or equal to 0.25 and less than or equal to 0.35, for example less than or equal to 0.

27.

15. Propulsion assembly (10) for aircraft (100) according to claims 13 and 14, wherein the propeller (22) comprises at least 10 blades (22A) and at most 16 blades (22A) and the hub-to-head ratio is greater than or equal to 0.25 and less than or equal to 0.

30.

16. Propulsion assembly (10) for aircraft (100) according to claims 13 and 14, in which the propeller (22) comprises at least 14 blades (22A) and at most 18 blades (22A) and the hub-to-head ratio is greater than or equal to 0.30 and less than or equal to 0.

35.

17. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 16, wherein the reduction mechanism (40) has a reduction ratio greater than or equal to 2.5 and less than or equal to 11.0, for example greater than or equal to 2.7 and less than or equal to 6.0, for example greater than or equal to 2.7 and less than or equal to 3.6, for example around 3.

0.

18. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 17, wherein the cyclic pitch control device for the propeller blades (50) is configured to regulate the cyclic pitch of the propeller blades (22A) (22) as a function of one or more aircraft parameters on which the propulsion assembly (10) is configured to be mounted, for example at least one parameter among the aircraft angle of attack, roll and yaw.

19. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 18, wherein the cyclic pitch control device for the propeller blades (50) is configured to regulate the cyclic pitch of the propeller blades (22A) (22) as a function of one or more parameters of the gas generator (30), for example as a function of at least one parameter among the speed, power, torque, for example the torque of a low-pressure body.

20. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 19, wherein the cyclic pitch control device for the propeller blades (50) comprises at least one sensor, for example disposed on at least one of the propeller shaft (26), a propeller bearing support, a control element for the cyclic pitch control device (50), for example a pitch ring (54) or a jack (52), the at least one sensor (60) being configured to determine an IP moment.

21. Propulsion assembly (10) for aircraft (100) according to any one of claims 1 to 20, wherein the cyclic pitch control device for the propeller blades (50) includes an inertial measurement unit (57) configured to determine the angle of attack, roll and yaw of the aircraft (100) on which the propulsion assembly (10) is mounted.

22. A method for regulating the cyclic pitch of the blades (22A) of the propeller (22) of the propulsion assembly (10) according to any one of claims 1 to 21, wherein the cyclic pitch of the blades is regulated (22A) of the propeller (22) as a function of at least one aircraft parameter on which the propulsion assembly (10) is configured to be mounted and / or at least one parameter of the gas generator (30).

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