EP4367022A1

EP4367022A1 - Aeronautical thruster

Info

Publication number: EP4367022A1
Application number: EP22751128.4A
Authority: EP
Inventors: Fernando GEA AGUILERA; Anthony BINDER; Adrien Clément Marcel DUBOIS; Mathieu Simon Paul GRUBER; Norman Bruno André JODET; Eva Julie Lebeault; Josselin David Florian REGNARD
Original assignee: Safran Aircraft Engines SAS
Current assignee: Safran Aircraft Engines SAS
Priority date: 2021-07-06
Filing date: 2022-07-04
Publication date: 2024-05-15
Also published as: CN117769516A; WO2023281192A1; FR3125090A1; FR3125090B1

Abstract

The invention relates to an aeronautical thruster (10) having a longitudinal axis (X) and comprising a hub (12) and at least two annular rows of non-streamlined blades (18) including an upstream annular row (14) and a downstream annular row (16) spaced apart from one another along said longitudinal axis (X), the upstream annular row (14) being rotatable about the longitudinal axis (X), and the downstream annular row (16) comprising a series of blades including a first blade (18a) and a second blade (18b) each extending in a radial direction from the hub (12) so as to define a radial dimension between the hub (12) and a radially outer end of the corresponding blade (18a, 18b), characterised in that the radial dimension of the first blade (18a) is greater than the radial dimension of the second blade (18b).

Description

Title: AERONAUTICAL THRUSTER

Technical area

The present disclosure relates to the field of longitudinal axis aeronautical thrusters each comprising a hub and (at least) two annular rows of unducted blades, one upstream, the other downstream, along the longitudinal axis .

[0002] In accordance with the foregoing and the following, throughout the text, the relative qualifiers "upstream" and "downstream" are defined relative to each other with reference to the flow of gases in the turbomachine in the longitudinal direction (i.e. the direction of the longitudinal axis).

[0003] The aeronautical propellant may comprise (at least) a heat engine, in particular a turbomachine, a turboshaft engine, a turbojet, a turbofan, and/or (at least) an electric motor, and/or (at least) a hydrogen engine, and / or (at least) a hybrid engine: thermal and/or electric and/or hydrogen. Prior technique

[0004] Reference will be made hereafter more particularly, and therefore on a non-limiting basis, to the case of turbomachines, since the type(s) of engine that the propellant comprises is not decisive here. By turbomachine is meant a propellant in which there is an exchange of energy between a flowing fluid and a rotor. [0005] In this context, it is recalled, by way of example, that a turbofan engine

"Unducted" (or "Propfan" or "Open rotor" or "Counter-Rotating Open Rotor" type turboprop) is a type of turbomachine in which the fan (or propeller) extends outside the engine casing (or nacelle ), unlike conventional turbomachines (“Turbofan” type) in which the fan is shrouded. An example of such a turbine engine is shown in Figures 1 and 2. The turbine engine 10 comprises a hub 12, defining the crankcase, and on which is mounted an upstream annular row 14 of unducted blades 18 and a downstream annular row 16 of unducted blades 18 which are spaced apart along a longitudinal axis X of the turbomachine 10. The orientation qualifiers, such as “longitudinal”, “radial” or “circumferential”, are defined by reference to the longitudinal axis X of the turbomachine 10.

The relative qualifiers “upstream” and “downstream” are defined relative to each other with reference to the flow of gases in the turbomachine 10 along the longitudinal axis X. Furthermore, the turbomachine 10 comprises, from upstream to downstream inside the crankcase, one (or more) compressor(s) 2, at least one combustion chamber 4, one (or more) turbine(s) 6 and at least one exhaust nozzle 8.

[0006] Among these turbomachines with an unducted fan, turbomachines of the “Unducted Single (or Stator) Fan” (USF) type are known, in each of which, as illustrated in FIGS. 1 and 2, the upstream annular row 14 of blades 18 not ducted is rotatably mounted around the longitudinal axis X and the downstream annular row 16 of non-ducted blades 18 is fixed. The direction of rotation of the blades 18 of the upstream annular row 14 is not decisive. The downstream annular row 16 can be centered on an axis that may or may not coincide with the longitudinal axis X. As illustrated in FIG. 1, the downstream annular row 16 is centered on the longitudinal axis X. Such a configuration of the upstream annular row 14 and the downstream annular row 16 makes it possible to enhance, through the downstream annular row 16, the gyration energy of the air flow coming from the upstream annular row 14. The efficiency of the turbomachine 10 is thus improved, in particular vis-à-vis a single rotary propeller in the case of a conventional turboprop. The upstream annular row 14 of unducted blades 18 is driven in rotation around the longitudinal axis X by the turbine(s) 6 which drive(s) themselves the ) compressor(s) 2. The turbomachine 10 generally comprises a speed reduction box (“gearbox” in English) in order to decouple the speed of rotation of the turbines 6 with respect to the speed of rotation of the upstream annular row 14. By moreover, one of the advantages of a USF type turbomachine compared to a "Counter-Rotating Open Rotor" type turbomachine is to reduce the tonal noise emitted by the turbomachine due to the fact that the downstream annular row 16 of unducted blades 18 is fixed.

As shown schematically in Figure 2, the turbine engine 10 may have a so-called "puller" configuration (upstream annular row 14 and downstream annular row 16 located at an upstream end portion of the turbine engine 10) or, as schematically in Figure 1, a so-called "pusher" configuration (upstream annular row 14 and downstream annular row 16 located at a downstream end portion of the turbine engine 10).

[0008] In the puller configuration, the upstream annular row 14 and the downstream annular row 16 can surround a section of the compressor(s) 2 of the turbomachine or of the speed reduction box. In the pusher configuration, the upstream annular row 14 and the downstream annular row 14 can surround a section of the turbine(s) 6 of the turbomachine 10.

[0009] The absence of a shroud leads to an increase in the level of noise emitted by the turbomachine 10. Indeed, the noise generated by the annular rows of unshrouded blades 18 propagates in free fields. A main cause of the noise emitted is related to vortex structures 19 generated in the air flow at the free radially outer ends of the blades 18 of the upstream annular row 14 which impact the blades 18 of the downstream annular row 16.

[0010] Too high a noise level is detrimental to the comfort of the passengers of the aircraft on which the turbomachine is installed. In addition, current standards impose a maximum noise threshold, particularly in the area close to the ground, i.e. during the take-off and landing phases. A known solution for reducing the level of noise emitted consists in uniformly reducing the radial dimension of each blade 18 of the downstream annular row 16. Thus, as shown in FIG. 3, the radially outer end of each blade 18 of the annular row upstream 14 is inscribed in a first circle 20 centered on the longitudinal axis X and the radially outer end of each blade 18 of the downstream annular row 16 is inscribed in a second circle 22 centered on the longitudinal axis X, the radius Re2 of the second circle 22 being less than the radius Re1 of the first circle 20. In this way, the impact of the vortices 19 formed at the radially outer ends of the blades 18 of the upstream annular row 14 on the blades 18 of the downstream annular row 16 is limited is limited in that these vortices pass radially outside the blades 18 of the downstream annular row 16. This solution is called "clipping", or "cropping", or "truncation", or even " clipping", of the blades 18 of the downstream annular row 16.

[0011] However, this solution is limited in that the reduction in the size of the blades 18 of the downstream annular row 16 causes a reduction in the efficiency of the turbomachine 10.

[0012] In addition, the current solution is not entirely satisfactory in that it allows effective noise reduction only in a configuration isolated from the turbomachine and at zero incidence. Indeed, the presence of surrounding elements (mast, fuselage, etc.), a non-zero incidence of the air flow perceived by the turbomachine 10 and the shape of the blades 18 of the upstream annular row 14 modify, on the one hand , the contraction and the axisymmetry around the longitudinal axis X of the air flow downstream of the upstream annular row 14, and/or on the other hand, the size of the vortices 19 present in the flow of air downstream of the upstream annular row 14 so that the truncation of the blades 18 of the downstream annular row 16 no longer prevents the interaction between the blades 18 of the downstream annular row 16 and the vortices 19 formed by the blades 18 of the upstream annular row 14.

The present description aims to propose a solution to these drawbacks.

Summary [0014] At this stage, it is immediately specified that, even if the preceding prior art therefore relates to an “Open Rotor” type turbomachine, the solution of the invention applies to any aeronautical thruster, since the aforementioned problem is not necessarily specific to the aforementioned type of aeronautical propellant. An “Open Rotor” of the CROR “Counter-Rotating Open Rotor” type is not excluded.

In this context, it is therefore here, and in general, proposed an aeronautical thruster with a longitudinal axis comprising a hub and at least two annular rows of unducted blades comprising an upstream annular row and a downstream annular row spaced apart l from each other along said longitudinal axis, the upstream annular row being rotatable around the longitudinal axis, said downstream annular row comprising at least a first blade and a second blade each extending in a radial direction from the hub so as to define a radial dimension between said hub and a radially outer end of the respective blade, characterized in that the radial dimension of the first blade is greater than the radial dimension of the second blade.

[0016] Such a configuration makes it possible to reduce the impact of the vortices formed at the radially outer end of the blades of the upstream annular row on the second blade of the downstream annular row. In particular, the second blade of the downstream annular row can advantageously be located in a predetermined circumferential zone which is conducive to the emission of a high level of noise. In addition, the radial dimension of the first blade can be greater, in particular with respect to the "clipping" solution known from the state of the art, thus increasing the performance of the aeronautical propellant without increasing, or even reducing the sound level emitted by the aeronautical propellant. In particular, the first blade can be located in a circumferential zone of the downstream annular row which is less conducive to noise emission.

[0017] Unlike the known configuration which is suitable for an aeronautical thruster of the CROR ("Counter-Rotating Open-Rotor") type, the solution has the advantage of being particularly suitable for an aeronautical thruster of the USF type. [0018] The term "non-ducted" used in reference to the upstream annular row and the downstream annular row indicates that the blades of the upstream annular row and the blades of the downstream annular row are not surrounded by a nacelle, unlike the conventional aeronautical thrusters in which the fan is shrouded inside a nacelle. The downstream annular row can be fixed around the longitudinal axis. In other words, the blades of the downstream annular row may not be driven in rotation around the longitudinal axis.

[0020] It may be noted that vis-à-vis in particular document US 2017/274993, which has already made a proposal within the aforementioned general framework, the above solution is not obvious, in particular because US 2017/ 274993 encourages those skilled in the art to move away from the claimed solution by using an aeronautical thruster in which the upstream annular row and the downstream annular row are both driven in rotation around the axis. Indeed, according to the embodiment of figure 13 of US 2017/274993, to obtain a reduction in the noise emitted by the thruster, it appears necessary that the blades of the upstream and downstream annular rows be driven in rotation in the same direction of rotation. rotation and that they are in phase alignment (see. g. paragraph [0099] of US 2017/274993). Consequently, those skilled in the art would not be encouraged to modify the aeronautical thruster described in this embodiment of figure 13 of US 2017/274993 to make the downstream annular row fixed around the longitudinal axis and thus orient towards the claimed solution.

[0021] This does not preclude the blades of the downstream annular row from being variable-pitch. The blades of the upstream annular row and/or of the downstream annular row can be of variable pitch. Each blade can thus be adjusted in rotation around a respective pitch change axis which extends radially. It is thus possible to adapt the pitch of the blades of the aeronautical thruster according to the operation of the aeronautical thruster and the phase of flight to improve aeronautical performance. The hub may also comprise a blade pitch variation system adapted to vary the incidence of the blades around the respective pitch change axis depending on the phase of flight.

[0022] Each blade of the upstream annular row can extend in a radial direction from the hub so as to define a radial dimension between said hub and a radially outer end of the blade in question, the dimension of each of the blades of the row upstream annular being greater than the radial dimension of the first blade of the downstream annular row. In other words, the first blade of the downstream annular row can be truncated with respect to the blades of the upstream annular row. This limits the impact of the vortices formed at the radially outer end of the blades of the upstream annular row on the first blade of the downstream annular row and in fact also on the second blade of the downstream annular row. The term "truncated blade" means that the blade has a reduced radial dimension. Alternatively, provision may be made for at least one blade of the upstream row to have a radial dimension greater than the radial dimension of the first blade of the downstream annular row. Alternatively again, provision may be made for at least one blade of the upstream annular row to have a radial dimension greater than the radial dimension of each of the blades of the downstream annular row.

[0023] The radial dimension of a blade is measured between a radially inner end of the blade, the latter being located at the level of (that is to say closest to) the hub of the aeronautical thruster, and an end radially outer of the blade. The radially inner end of a blade can be, longitudinally, at the level of a leading edge of the blade or at the level of the axis of change of pitch of the blade considered. The radially inner end of a blade is also called the "blade root". An angular position of each blade around the longitudinal axis can be identified by the angular position around the longitudinal axis of the inner end of the respective blade. The radially outer end of the blade is the opposite end from the radially inner end. The radially outer end of the blade may be the free end of the blade. The radially inner end and the radially outer end of each of the blades can be radially aligned, i.e. at the same longitudinal position. It is not excluded that the radially inner end and the radially outer end of each of the blades may be longitudinally offset relative to each other.

The first blade and the second blade of the downstream annular row may each have a radially outer radius passing through said radially outer end, the radially outer radius of the first blade being greater than the radially outer radius of the second blade. The radially outer radius of a blade can be thought of as the radial distance to the longitudinal axis from the radially outer end point of said blade. In other words, it is the maximum radius of the blade.

The first blade and the second blade can each have a radially internal radius. The radially inner radius of a blade can be thought of as the radial distance from the longitudinal axis of the radially inner tip of the blade. Each blade can be fixed to the hub of the aeronautical thruster at the level of the radially internal end. Each blade can be fixed to the hub close to the leading edge at the blade root or close to the pitch change axis at the blade root. [0026] Each blade of the upstream annular row may have a radially outer radius. A relative difference of the radially outer radius of any of the first blade and the second blade of the downstream annular row compared to the radially outer radius of any of the blades of the upstream annular row can be between -15% and 30%. The first blade and the second blade of the downstream annular row can be circumferentially consecutive. Alternatively, one (or more) blade(s) can be interposed circumferentially between the first blade and the second blade.

[0028] The radially outer end of each of the blades of the upstream annular row is inscribed in an outer casing of the upstream annular row. Similarly, the radially outer end of each of the blades of the downstream annular row is inscribed in an outer casing of the downstream annular row. The external envelope of the upstream annular row can surround the external envelope of the downstream annular row when these are projected in a common projection plane which is normal to the longitudinal axis.

[0029] A projection in a plane normal to the longitudinal axis of the outer casing of the upstream annular row can define a circle centered on the longitudinal axis. The circle defined by the projection of the external envelope of the upstream annular row in a plane normal to the longitudinal axis may have a diameter which represents the external diameter of the aeronautical thruster.

[0030] A projection in a plane normal to the longitudinal axis of the outer casing of the downstream annular row can define a circle. The center of the circle defined by the projection of the outer envelope of the downstream annular row can be offset from the longitudinal axis in the direction of an axis passing through the angular positions at 12 o'clock and 6 o'clock. In other words, the geometric center of the projection of the outer envelope of the downstream annular row (i.e. the center of the circle if the projection of the outer envelope defines a circle) can be offset from the longitudinal axis according to the direction of the axis passing through the angular positions at 12 o'clock and at 6 o'clock. The radial distance between the center of the circle defined by the projection of the external envelope of the downstream annular row and the longitudinal axis can be between 1/200th and 1/5th of the diameter of the circle defined by the projection of the envelope outside of the upstream annular row.

The hub may have, at the downstream annular row, a section normal to the longitudinal axis having the shape of a circle centered on the longitudinal axis. The hub may have an opening disposed, in whole or in part, longitudinally between the upstream annular row and the downstream annular row. The opening may be annular around the longitudinal axis. The opening may be intended to form an air inlet for the internal flow of the aeronautical propellant. Also, the radially outer radius of the hub may at the level of the downstream annular row may be greater than the radially outer radius of the hub at the level of the upstream annular row. The radially outer radius of the means at the upstream annular row and the downstream annular row may respectively coincide with the radially internal radius respectively of the blades of the upstream annular row and of the downstream annular row.

The downstream annular row may comprise at least one group of blades having the same radial dimension, including at least a first group comprising a plurality of first blades and/or a second group comprising a plurality of second blades. Alternatively, the downstream annular row may comprise at least one group of blades having the same radially outer radius, including at least a first group comprising a plurality of first blades and/or a second group comprising a plurality of second blades. The downstream annular row can comprise k groups of blades with k an integer greater than or equal to 1. The integer k can be less than or equal to the number of blades in the downstream annular row. The number of different blades to be manufactured is thus limited, making it possible to reduce the costs associated with the manufacture of such an aeronautical thruster.

The blades of said at least one group of blades of the downstream annular row can be arranged circumferentially contiguously in an angular sector around the longitudinal axis. In other words, the blades of each group of blades can all be consecutive two by two in said angular sector around the longitudinal axis. The manufacturing costs of the aeronautical propellant are further reduced.

In other words again, each group of blades can be associated with at least one angular sector around the longitudinal axis so as to form an angular sector consisting of blades of said group considered. There can therefore be provided several angular sectors of blades circumferentially adjacent to each other, each angular sector comprising blades having a given radial dimension or a given radially external radius, different from the radial dimension, respectively from the radially external radius, of the blades of an adjacent sector.

The first blade and the second blade can each be arranged angularly around the longitudinal axis at a respective angle, the angle being measured around the longitudinal axis in the clockwise direction with respect to an angular position at 12 o'clock, the radial dimension or the radially outer radius of the first blade and/or the radial dimension or the radially outer radius of the second blade being determined as a function of the respective angle according to a linear, parabolic, logarithmic, or exponential law.

[0037] The downstream annular row may comprise at least one set of blades arranged contiguously in an angular sector around the longitudinal axis, said set of blades possibly comprising the first blade and/or the second blade, each blade of the the set of blades being arranged angularly around the longitudinal axis according to a respective angle, the angle being measured around the longitudinal axis in a clockwise direction with respect to an angular position at 12 o'clock, each blade of the set of blades having a radial dimension or a radially outer radius determined according to the respective angle according to a linear, parabolic, logarithmic, or exponential law. The angular sector associated with said set of blades can extend between the angular position at 12 o'clock and an angular position at 6 o'clock.

The second blade can be positioned, angularly around the longitudinal axis, closer to an angular position at 6 o'clock than is the first blade. Conversely, the first blade can be positioned, angularly around the longitudinal axis, closer to an angular position at 12 o'clock than is the second blade. According to a particular embodiment, the first blade of the downstream annular row can be positioned angularly around the longitudinal axis, between an angular position at 12 o'clock and an angular position at 6 o'clock, and the second blade can be positioned angularly around the longitudinal axis between the first blade and the angular position at 6 o'clock. This configuration is particularly advantageous for reducing the influence of the vortices formed at the radially outer end of the blades of the upstream annular row on the blades of the downstream annular row when the incidence of the aeronautical propellant is high, i.e. when the longitudinal axis of the aeronautical thruster has a high inclination with respect to the horizontal, in particular a positive incidence during the take-off phases.

[0040] The downstream annular row may comprise a first angular extent of blades centered on the angular position at 6 o'clock and a second angular extent of blades centered on the angular position at 12 o'clock, the average radial dimension of the blades of the first angular extent being less to the average radial dimension of the blades of the second angular extent.

[0041] The downstream annular row may comprise at least one pair of blades whose angular positioning around the longitudinal axis is symmetrical with respect to a plane of symmetry comprising the longitudinal axis and an axis passing through angular positions at 6 o'clock and at 12 o'clock, the blades of said pair of blades having identical geometric parameters, in particular the same radial dimension. The downstream annular row may be symmetrical with respect to the plane of symmetry. It is understood by "symmetrical" that for each blade of the downstream annular row positioned angularly around the longitudinal axis according to an angle measured around the longitudinal axis in the clockwise direction with respect to the angular position at 12 o'clock and between 0 ° and 180° excluded, the downstream annular row comprises another blade positioned angularly around the longitudinal axis according to an opposite angle (ie the same angle but measured around the longitudinal axis in the counter-clockwise direction) and having identical geometric parameters. In particular, the blades of the downstream annular row positioned angularly around the longitudinal axis, respectively, at opposite angles with respect to the angular position at 12 o'clock may have the same radial dimension.

[0042] The downstream annular row may have a symmetry of revolution of order n with n an integer greater than or equal to 2. A symmetry of revolution corresponds to a discrete symmetry by rotation. Thus, an object presenting a symmetry of revolution of order n is invariant for any rotation of 1/nth of a turn, i.e. for any rotation of an angle of 2p/h. In such a configuration, two circumferentially adjacent blades each have a different radial dimension from each other. Thus, each blade has a different acoustic radiation from the circumferentially adjacent blades, thus promoting the decorrelation of noise sources and further reducing the noise generated by the aeronautical propellant. The downstream annular row can comprise n subsets of blades, each blade of a subset being associated with a group of blades.

The downstream annular row may comprise k ^* n blades. For example, the downstream annular row can comprise between 2 and 25 blades. The number of blades of the upstream annular row may be different from the number of blades of the downstream annular row. This makes it possible to further minimize the level of noise emitted by the aeronautical thruster.

[0045] The strength of the downstream annular row, defined as the ratio between the chord, and the spacing between two circumferentially consecutive blades in the circumferential direction, can be less than 3 over the entire radial dimension of each blade. . In particular, in a preferred embodiment, the strength is less than 1 at the radially outer end of the blades.

The ratio between the distance in the longitudinal direction between a median plane of each annular row which is normal to the longitudinal axis, and the diameter of the aeronautical thruster can vary between 0.01 and 0.8. The median plane normal to the respective longitudinal axis of each annular row may be the plane containing a respective pitch change axis of each of the blades of the corresponding annular row. The trailing edge of each of the blades of the upstream annular row is located longitudinally upstream of a leading edge of each of the blades of the downstream annular row. Thus, interference between the annular rows of blades is limited, or even avoided.

The upstream annular row and the downstream annular row can be located at an upstream end portion of the aeronautical thruster in the direction longitudinal or at a downstream end portion of the aeronautical thruster in the longitudinal direction.

The aeronautical propellant can have a so-called "puller" configuration (upstream annular row and downstream annular row located at an upstream end portion of the aeronautical propellant) or a so-called "pusher" configuration (upstream annular row and downstream downstream annular located at a downstream end portion of the aeronautical thruster).

In the puller configuration, the upstream annular row and the downstream annular row can surround a section of the compressor(s) or of the gearbox of the aeronautical propellant. In the pusher configuration, the upstream annular row and the downstream annular row can surround a section of the turbine(s) of the aeronautical thruster.

According to another aspect, there is described a propulsion assembly for an aircraft, comprising an aeronautical thruster as described above and a pylon for fixing the aeronautical propellant to the aircraft, the fixing pylon being connected to the one of the blades of the downstream annular row so as to form a single aerodynamic assembly.

According to another aspect, an aircraft is described comprising an aeronautical thruster as described above or a propulsion assembly as described above.

Brief description of the drawings

[0052] Other characteristics, details and advantages will appear on reading the detailed description below, and on analyzing the appended drawings, in which:

[0053] Figure 1 is a partial schematic sectional view of an unducted fan turbine engine according to the prior art, in a "pusher" configuration;

[0054] Figure 2 is a schematic view of a turbine engine with an unducted fan, in a “puller” configuration, in a take-off phase;

[0055] Figure 3 is a schematic view illustrating an annular row of non-ducted fixed blades of the turbomachine of Figure 1 in the section plane l-l, according to the prior art;

[0056] Figure 4 is a partial schematic sectional view of an unducted fan turbine engine according to the present description, in a "pusher" configuration;

[0057] Figure 5 is a schematic view illustrating an annular row of non-ducted fixed blades of the turbomachine of Figure 4 in section plane IV-IV, according to a first embodiment of the present description; [0058] Figure 6 is a schematic view illustrating an annular row of non-ducted fixed blades of the turbomachine of Figure 4 in section plane IV-IV, according to a second embodiment of the present description;

[0059] Figure 7 is a schematic view illustrating an annular row of non-ducted fixed blades of the turbomachine of Figure 4 in section plane IV-IV, according to a third embodiment of the present description;

[0060] Figure 8 is a schematic view illustrating an annular row of non-ducted fixed blades of the turbomachine of Figure 4 in section plane IV-IV, according to a fourth embodiment of the present description;

[0061] Figure 9 is a schematic view illustrating an annular row of non-ducted fixed blades of the turbomachine of Figure 4 in section plane IV-IV, according to a fifth embodiment of the present description;

[0062] Figure 10 is a schematic view illustrating an annular row of non-ducted fixed blades of the turbomachine of Figure 4 in section plane IV-IV, according to a sixth embodiment of the present description;

[0063] Figure 11 is a schematic view of an unducted fan turbomachine of the present description, in a "puller" configuration;

[0064] Figure 12 is a schematic view of an unducted fan turbine engine of the present description according to an alternative embodiment;

[0065] Figure 13 is a schematic view of an unducted fan turbomachine of the present description according to another variant embodiment;

[0066] Figure 14 represents a schematic view of any aeronautical propellant according to the present description.

Description of embodiments

Reference is now made to Figure 4. Figure 4 shows, in section, a turbomachine 10 of longitudinal axis X which comprises, from upstream to downstream inside the crankcase, one (or more) compressor(s) 2, one (or more) combustion chamber(s) 4, one (or more) turbine(s) 6 and one (or more) exhaust nozzle(s) 8.

The turbomachine 10 comprises a hub 12 and two annular rows of unducted blades 18 including an upstream annular row 14 and a downstream annular row 16. The upstream annular row 14 and the downstream annular row 16 are spaced apart from one another. the other along the longitudinal axis X. The upstream annular row 14 is rotatable around the longitudinal axis X. The downstream annular row 16 is here fixed around the longitudinal axis X. In other words, the downstream annular row 16 is not driven in rotation around the longitudinal axis X. This does not exclude that each blade 18 of the downstream annular row 16 can be variable pitched as will be seen later. In addition, according to an alternative not shown, the downstream annular row 16 can be rotatable around the longitudinal axis X. In other words, the upstream annular row 14 is of the rotor type and the downstream annular row 16 is here stator type. In the configuration of FIG. 4, the upstream annular row 14 of unducted blades 18 is driven in rotation around the longitudinal axis X by the turbine(s) 6 which drive(s) it(they)- same (s) the compressor (s) 2. In addition, there may be provided a speed reduction box (or "gearbox" in English) between the (or) turbine (s) 6 and the row upstream annular ring 14. Depending on one application, the turbomachine can have a thrust of between 1,000 and 90,000 pounds, preferably between 2,500 and 50,000 pounds in the cruising phase or in the take-off phase.

The blades 18 of the upstream annular row 14 and/or of the downstream annular row 16 can be of variable pitch. It is thus possible to adapt the pitch of the blades 18 of the turbine engine 10 according to the operating point of the turbine engine 10 or the phase of flight. A system for changing the pitch integrated into the hub can be provided in order to adapt the incidence of the blades for each phase of flight. Each blade 18 can thus be adjusted in rotation around a respective pitch change axis according to the flight phases and conditions. The pitch change axis of each of the blades 18 is a radially extending and longitudinally positioned axis at a middle portion of the respective blade. The pitch change axis of each of the blades 18 of the downstream annular row 16 visible in FIG. 4 coincides with an axis passing through angular positions at 12 o'clock and at 6 o'clock around the longitudinal axis of the downstream annular row. [0070] In the rest of the description, the orientation qualifiers, such as "longitudinal",

“Radial” or “circumferential” are defined with reference to the longitudinal axis X of the turbomachine 10. The longitudinal direction here corresponds to the direction of advancement of the turbomachine. In particular, the longitudinal direction can coincide with a horizontal direction, ie perpendicular to the gravity field. The qualifiers “upstream” and “downstream” are defined relative to each other with reference to the flow of gases in the turbomachine 10 in the longitudinal direction. The angular position of each of the blades 18 around the longitudinal axis X is marked with respect to a time dial (here seen from upstream for example) whose angular positions at 12 o'clock, 3 o'clock, 6 o'clock and 9 o'clock are positioned in a conventional manner . The angular position at 12 o'clock is therefore positioned vertically upwards with respect to the longitudinal axis X and the angular position at 6 o'clock is positioned vertically downwards with respect to the longitudinal axis X. The angular position at 3 o'clock is positioned horizontally to the right of the longitudinal axis X and the angular position at 6 o'clock is positioned horizontally to the left of the longitudinal axis X. An axis extending radially through the angular positions at 12 o'clock and 6 o'clock is thus perpendicular to an axis extending radially passing through the angular positions at 3 o'clock and 9 o'clock. Qualifiers of absolute position, such as the terms "top", "bottom", "left", "right", etc., or of relative position, such as the terms "above", "below", "upper", “lower”, etc., and the orientation qualifiers, such as the terms “vertical” and “horizontal” refer here to the orientation of the figures and are considered in an operational state of the turbomachine 10, typically when the latter it is installed on an aircraft on the ground. In this state of the turbomachine 10, the axis passing through the angular positions at 12 o'clock and at 6 o'clock extends in the direction of the gravity field, ie vertically. On the other hand, it can be deduced that a roll movement of the aircraft in flight on which the turbomachine 10 is mounted will be such as to cause a rotation of the vertical and horizontal directions as considered in the figures around the longitudinal axis X In the same way, a rolling movement of the aircraft in flight on which the turbomachine 10 is mounted will be such as to cause a rotation of the axis passing through the angular positions at 12 o'clock and 6 o'clock and of the axis passing through the 3 o'clock and 9 o'clock angular positions around the longitudinal axis X. A "lateral zone" of the turbine engine 10 refers to a zone which is circumferentially in the vicinity of the 3 o'clock angular position or the 9 o'clock angular position. Similarly, an “upper zone” and a “lower zone” of the turbomachine 10 refer, respectively, to a zone which is circumferentially close to the angular position at 12 o'clock and to a zone which is circumferentially close to the angular position at 6 a.m.

Each blade 18 of the upstream annular row 14 and of the downstream annular row 16 extends in a radial direction from the hub 12 so as to define a radial dimension between said hub 12 and a radially outer end of the blade 18 respectively. In other words, the radial dimension of a blade 18 is measured between a radially inner end of the blade 18 and a radially outer end of the blade 18. The radially inner end of each blade 18 is located at the hub 12 of the turbine engine 10. Each blade 18 can in particular be fixed to the hub 12 of the turbine engine 10 at the radially inner end. The radially outer end of each blade 18 is here a free end (i.e. not faired).

[0072] In addition, each blade 18 of the upstream annular row 14 and of the downstream annular row 16 has a radially internal radius respectively Ri1, Ri2 considered as the radial distance to the longitudinal axis X of the radially internal end of the blade 18, for example located at the level of (that is to say closest to) the hub of the turbomachine. The radially inner end is, in FIG. 4, close to the pitch change axis of the respective blade. The radially inner end of each blade can alternatively be close to the leading edge at the root of the blade. A radially outer radius Re of each blade 18 is considered as the radial distance from the longitudinal axis X of the radially outer end of said blade 18, that is to say, as the maximum radius of the blade.

As seen in Figure 5, which shows the turbine engine of Figure 4 in section plane IV-IV normal to the longitudinal axis X, the radially outer end of the blades 18 of the upstream annular row 14 and the downstream annular row 16 are inscribed, respectively, in an outer casing 20 of the upstream annular row 14 and an outer casing 22 of the downstream annular row 16. Here, a projection, in section plane IV-IV, of the casing 20 of the upstream annular row 14 defines a circle of radius Re1, or else of diameter D, centered on the longitudinal axis X. The diameter D of the projection, in the section plane IV-IV, of the outer envelope 20 of the upstream annular row 14 may represent the outer diameter of the turbomachine 10.

On the other hand, the radial dimension of each blade 18 of the downstream annular row 16 is less than the radial dimension of each of the blades 18 of the upstream annular row 14. In particular, the radially outer radius of each blade 18 of the downstream annular row 16 is less than the radially outer radius of each of the blades 18 of the upstream annular row 14. Thus, the outer casing 20 of the upstream annular row 14 surrounds the outer casing 22 of the downstream annular row 16 when these are projected into a common projection plane which is here the section plane IV-IV perpendicular to the longitudinal axis X. Each blade 18 of the downstream annular row 16 therefore has a truncation with respect to the blades 18 of the row upstream annular 14 so as to limit the impact of the vortices 19 formed at the radially outer end of the blades 18 of the upstream annular row 14. In other words, each blade 18 of the downstream annular row 16 is truncated p relative to the blades 18 of the upstream annular row 14.

[0075] FIG. 5 represents a first embodiment of the downstream annular row 16 in which the projection of the outer casing 22 of the downstream annular row 16 in the section plane IV-IV defines a circle whose center is offset relative to the longitudinal axis X along the direction of the axis passing through the angular positions at 12 o'clock and at 6 o'clock. The radial distance between the center of the outer casing 22 of the downstream annular row 16 in the shape of a circle and the longitudinal axis X can be between 0.005 D and 0.2 D. The circle defined by the outer casing 22 of the downstream annular row 16 has a radius Re2′ less than the radius Re1 of the outer casing 20 the upstream annular row 14. Thus, in the first embodiment as shown in Figure 5, the downstream annular row 16 is such that it comprises at least a first blade 18a and a second blade 18b so that the radial dimension of the second blade 18b is smaller than the radial dimension of the first blade 18a. The second blade 18b here has a radially outer radius smaller than the radially outer radius of each of the blades 18 of the upstream annular row 14. The first blade 18a here also has a radially outer radius smaller than the radially outer radius of each of the blades 18 of the upstream annular row 14.

Figure 5 illustrates, among the blades 18 of the downstream annular row 16, a particular combination of the first blade 18a and the second blade 18b in which the first blade 18a and the second blade 18b are circumferentially consecutive. However, it is not excluded that other combinations of the first blade 18a and the second blade 18b as described above are possible. In particular, the first blade 18a and the second blade 18b considered may not be circumferentially consecutive.

Such a configuration of the downstream annular row 16 allows greater truncation of the second blade 18b of the downstream annular row 16 to further reduce the interaction of the vortices 19 formed at the radially outer end of the blades 18 of the upstream annular row 14 with the second blade 18b of the downstream annular row 16. a high noise level. In addition, the first blade 18a of the downstream annular row 16 has a larger radial dimension than the second blade 18b, i.e. a smaller truncation, thus increasing the performance of the turbomachine 10.

According to another remarkable aspect of the first embodiment, the second blade 18b is positioned, angularly around the longitudinal axis X, closer to the angular position at 6 o'clock than is the first blade 18a. In particular, the first blade 18a and the second blade 18b can be determined so that whatever the angular position of the first blade 18a between the angular position at 12 o'clock and the angular position at 6 o'clock, the second blade 18b is positioned angularly around of the longitudinal axis X between the first blade 18a and the angular position at 6 o'clock. Here again, FIG. 5 illustrates, among the blades 18 of the downstream annular row 16, a particular combination of the first blade 18a and the second blade 18b in accordance with the arrangements described above. However, it is not excluded that other combinations of the first blade 18a and the second blade 18b in accordance with the arrangements described above are possible. The blades 18 of the lower part of the downstream annular row 16 therefore have a greater truncation than the blades 18 of the upper part of the downstream annular row 16. The blade 18 of the downstream annular row 16 having the radial dimension minimum is located, angularly around the longitudinal axis X, at (or close to) the angular position at 6 o'clock. Conversely, the blade 18 of the downstream annular row 16 having the maximum radial dimension is located, angularly around the longitudinal axis X, at (or close to) the angular position at 12 o'clock. Also, it can be defined a first angular extent of blades 18 centered on the angular position at 6 o'clock and a second angular extent of blades 18 centered on the angular position at 12 o'clock, the average radial dimension of the blades 18 of the first angular extent being lower to the average radial dimension of the blades 18 of the second angular extent.

This further reduces the impact of the vortices formed at the radially outer end of the blades 18 of the upstream annular row 14 on the blades 18 of the lower part of the downstream annular row 16. In addition, the configuration of the downstream annular row 16 of the first embodiment makes it possible to retain a radial dimension of the blades 18 of the upper part of the downstream annular row 16 adapted to improve the performance of the turbomachine 10.

This configuration is particularly advantageous for reducing the impact of the vortices 19 formed at the radially outer end of the blades 18 of the upstream annular row 14 on the blades 18 of the downstream annular row 16 when the incidence of the turbomachine 10 is high, i.e. when the longitudinal axis X has a high inclination with respect to the horizontal, in particular a positive incidence during the take-off phases.

[0083] A relative difference of the radially outer radius of any of the blades 18 of the downstream annular row 16 with respect to the radially outer radius of any of the blades 18 of the upstream annular row 14 is preferably less than 30 %. The relative difference of the radially external radius of any one of the blades 18 of the downstream annular row 16 with respect to the radially external radius of any one of the blades 18 of the upstream annular row 14 is defined by (Re1-Re2') /Re1 ^* 100, where Re2' is the radially outer radius of any of the blades 18 of the downstream annular row 16 and Re1 is the radially outer radius of any of the blades 18 of the upstream annular row 14. particular, a relative difference of the radially outer radius of the blade 18 of the downstream annular row 16 having the maximum radial dimension with respect to the radially outer radius of any one of the blades 18 of the upstream annular row 14 is preferably between 2 % and 18%. A relative difference in the radially outer radius of the blade 18 of the downstream annular row 16 having the minimum radial dimension relative to the radially outer radius of the blade 18 of the downstream annular row 16 having the maximum radial dimension is less than 12%, preferably less than 2%, more preferably less than 1%.

Each blade 18 of the downstream annular row 16 is angularly disposed around the longitudinal axis X at an angle a, the angle a being measured around the longitudinal axis X clockwise with respect to an angular position at 12. In other words, the angle a is equal to 0° for an angular position at 12 o'clock and equal to 180° for the angular position at 6 o'clock. The angular position of each blade 18 around the longitudinal axis X can be identified by a pitch change axis of the blade considered, merged here with a stacking axis of the blade considered. Also, equivalently in the example of Figure 5, the angular position of each blade 18 around the longitudinal axis X can be identified by the angular position around the longitudinal axis X of the inner end of the blade 18 considered.

For example, the first blade 18a and the second blade 18b are each arranged angularly around the longitudinal axis X at a respective angle aa, ab. In the embodiment of FIG. 5, the radial dimension of each blade 18 of the downstream annular row 16 is determined as a function of the angle a of the blade considered according to a uniform and monotonous decreasing law for the angle a varying from 0° to 180°. Alternatively, it could be a linear, parabolic, logarithmic, or exponential law. Furthermore, in the example shown, each blade 18 of the downstream annular row

16 has the same radially internal radius Ri2. This is due to the fact that the hub 12 is axisymmetric about the longitudinal axis X at the level of the downstream annular row 16. In other words, the hub 12 has, at the level of the downstream annular row 16, a section normal to the longitudinal axis X which has the shape of a circle centered on the longitudinal axis X. Thus, the first blade 18a considered of the downstream annular row 16 has a radial dimension greater than the radial dimension of the second blade 18b considered only in that the radially external radius of the first blade 18a considered is greater than the radially external radius of the second blade 18b considered.

The downstream annular row 16 is here symmetrical with respect to a plane of symmetry P comprising the longitudinal axis X and the axis passing through the angular positions at 12 o'clock and at 6 o'clock. It is understood by "symmetrical" that, for each blade 18 of the downstream annular row 16 positioned angularly around the longitudinal axis X according to an angle a measured around the longitudinal axis in the clockwise direction with respect to the angular position at 12H and between 0° and 180° excluded, the downstream annular row 16 comprises another blade 18 positioned angularly around the longitudinal axis X at an angle -a and having identical geometric parameters. In particular, the blades 18 of the downstream annular row 16 positioned angularly around the longitudinal axis X, respectively, at an angle a and -a can have the same radial dimension. There is therefore a plurality of groups Gi comprising two identical blades 18 which are arranged angularly around the longitudinal axis X, respectively, at an angle a and -a.

According to a variant of the first embodiment of Figure 5, the projection of the outer casing 22 of the downstream annular row 16 in the IV-IV section plane can have any shape having a symmetry with respect to to the plane of symmetry P, for example an oval shape or even an ovoid shape having the plane P as its plane of symmetry.

The number of blades 18 of the upstream annular row 14 may be different from the number of blades 18 of the downstream annular row 16. This makes it possible to further reduce the noise level emitted by the turbine engine 10. For example, the annular row downstream 16 can comprise between 2 and 25 blades 18. The solidity of the downstream annular row 16, defined as the ratio between the chord, and the spacing in the circumferential direction between two circumferentially consecutive blades 18, can be less than 3 over the entire radial dimension of each blade 18. In particular, in a preferred embodiment, the strength is less than 1 at the radially outer end of the blades 18. The ratio between the distance in the longitudinal direction between a median plane normal to the respective longitudinal axis X of each annular row, and the diameter D of the turbomachine 10 can vary between 0.01 and 0.8. The median plane normal to the respective longitudinal axis X of each annular row 14, 16 is the plane containing the pitch change axis of each of the blades of the corresponding annular row 14, 16. The trailing edge of each of the blades 18 of the upstream annular row 14 is here located longitudinally upstream of a leading edge of each of the blades 18 of the downstream annular row 16. Thus, we limit, or even avoid, interference between the upstream and downstream annular rows 14, 16.

FIG. 6 represents a second embodiment of the downstream annular row 16 which differs from the first embodiment in that the downstream annular row 16 comprises a first group G1 of blades 18 and a second group G2 of blades 18. Each of the blades 18 of the first group G1 of blades 18, respectively of the second group G2 of blades 18, has the same radial dimension. The blades 18 of each of the first and second groups G1, G2 are arranged circumferentially contiguously in an angular sector around the longitudinal axis X. The number of different blades 18 to be manufactured, making it possible to reduce the costs associated with the manufacture of such a turbomachine 10. Of course, the number of groups of blades 18 is not limited to 2. The downstream annular row 16 can comprise k groups (s) of blades 18 with k an integer greater than or equal to 1. In the embodiment of FIG. 6, each of the first group G1 of blades

18 and the second group G2 of blades 18 is associated with at least one angular sector around the longitudinal axis X so as to form an angular sector consisting of blades 18 of said group G1, G2 considered. The first group G1 of blades 18 is here associated with a first angular sector S1 centered on the angular position at 12 o'clock. The second group G2 of blades 18 is associated with a second angular sector S2 centered on the angular position at 6 o'clock. The first angular sector S1 and the second angular sector S2 here each extend over 90° in the example shown. The downstream annular row 16 further comprises a plurality of groups Gi of blades 18 comprising two identical blades 18 which are arranged angularly around the longitudinal axis X, respectively, at an angle a and an angle -a. One of the two blades 18 of each group Gi is positioned angularly around the longitudinal axis X in a third angular sector S3 centered on the angular position at 3H and the other of the two blades 18 is positioned angularly around the axis longitudinal X in a fourth angular sector S4 centered on the angular position at 9 o'clock. The third angular sector S3 and the fourth angular sector S4 here also extend over 90°. Alternatively, the angular sectors S1, S2, S3, S4 can extend independently of each other over angular ranges greater or less than 90°.

The blades 18 of the first group G1 of blades 18 have a radial dimension greater than the radial dimension of the blades 18 of the second group G2 of blades 18. In the present case, the blades 18 of the first group G1 have a radially outer radius greater than the radially outer radius of the blades 18 of the second group G2. The projection in the IV-IV section plane of the outer casing 22 of the downstream annular row 16 here has an ovoid shape. Indeed, the projection in the section plane IV-IV of the outer casing 22 of the downstream annular row 16 has a first arcuate portion of radius Re2a at the level of the first angular sector S1 and a second arcuate portion of circle of radius Re2b at the level of the second angular sector S2. The radius Re2a of the first portion is greater than the radius Re2b of the second portion.

Furthermore, the blades 18 of the downstream annular row 16 arranged, around the longitudinal axis X, in each of the third angular sector S3 and the fourth angular sector S4 respectively form a set of blades whose radial dimension and/or where the radially outer radius of each blade 18 is determined as a function of the angle a of angular positioning of the blade 18 considered around the longitudinal axis X. The radial dimension and/or the radially external radius can be determined according to a uniform and monotonous law for the angle a varying in a range of values corresponding respectively to the third sector angular S3 and the fourth angular sector S4. Alternatively, it could be a linear, parabolic, logarithmic, or exponential law. Remarkably, said law can be chosen so that the calculated radially external radius is equal to Re2a for an angle a corresponding to the limit between the first angular sector S1 and the third angular sector S3, or between the first angular sector S1 and respectively the fourth angular sector S4. Similarly, said law can be chosen so that the calculated radially external radius is equal to Re2b for an angle a corresponding to the limit between the second angular sector S2 and the third angular sector S3, or respectively limit between the second angular sector S2 and the fourth angular sector S4.

[0096] Figure 7 shows a third embodiment of the downstream annular row 16 which differs from the second embodiment of Figure 6 in that the downstream annular row 16 does not have symmetry with respect to the plane of symmetry P. Indeed, a pair of blades 18 of the downstream annular row 16 positioned angularly around the longitudinal axis X, respectively, in the third angular sector S3 and the fourth angular sector S4, according to an angle a and an angle -a, are not identical. In particular, a pair of blades 18 of the downstream annular row 16 positioned angularly around the longitudinal axis X, respectively, in the third angular sector S3 and the fourth angular sector S4, according to an angle a and an angle -a have a radial dimension different from each other. This embodiment makes it possible to take into account the effects of installation, and in particular the relative position of the fuselage, which could have an influence on the positioning of the flow tube or air current tube generated by the upstream annular row 14 or even the effects of ascending and descending blades of the upstream annular row 14, which can modify the load of the blades 18 of the upstream annular row 14 and of the downstream annular row 16, as well as the level of truncation necessary on the third and fourth angular sectors S3 and S4 of the downstream annular row 16. The radial dimension of each blade 18 arranged in the third and fourth angular sectors S3 and S4 can be determined according to different laws depending on the angular sector considered.

FIG. 8 represents a fourth embodiment of the downstream annular row 16 in which the downstream annular row 16 comprises only the first group G1 of blades 18 and the second group G2 of blades 18. The first angular sector S1 associated with the first group G1 of blade 18 is centered on the angular position at 12 o'clock and extends over 260°. The second angular sector S2 associated with the second group G2 of blade 18 is centered on the angular position at 6 o'clock and extends over 100°.

FIG. 9 represents a fifth embodiment of the downstream annular row 16 in which the downstream annular row 16 has a symmetry of revolution of order 4. The downstream annular row 16 thus comprises 4 subassemblies E of blades 18 , each blade 18 of a sub-assembly being associated with a group of blades 18. Each blade 18 thus has a different acoustic radiation from the circumferentially adjacent blades 18, thus favoring the decorrelation of noise sources. In the fifth embodiment of FIG. 9, the downstream annular row 16 comprises a first group G1 of blades 18, a second group G2 of blades 18 and third group G3 of blades 18. Thus, each subset E of blades 18 comprises a first blade 18 (G1), a second blade 18 (G2) and a third blade 18 (G3) respectively associated with the first group G1, the second group G2 and the third group G3.

Here too, the symmetry of revolution of the downstream annular row 16 is not limited to an order equal to 4. The downstream annular row 16 can have a symmetry of revolution of order n with n an integer greater than or equal to to 2. The downstream annular row 16 can therefore have n subsets E of blades 18. As seen previously, the downstream annular row 16 can comprise k groups of blades with k an integer greater than or equal to 1, k being equal to 3 in the embodiment of FIG. 9. The downstream annular row 16 can therefore comprise k ^* n blades. Here the downstream annular row 16 comprises 12 blades in the embodiment of Figure 9.

According to a sixth embodiment shown in Figure 10, the blades 18 of the upper part of the downstream annular row 16 has a greater truncation than the blades 18 of the lower part of the downstream annular row 16. This configuration makes it possible to reduce the impact of the vortices formed at the level of the radially outer ends of the blades 18 of the upstream annular row 14 on the blades 18 of the upper part of the downstream annular row 16, in particular during the landing phases or to avoid effects related to the installation of the turbine engine 10 in the aircraft. This configuration could make it possible to reduce the level of noise emitted by the turbomachine 10 during the landing phases and/or when the turbomachine 10 is installed at the rear of the aircraft.

[0101] For each embodiment, the upstream annular row 14 and the downstream annular row 16 can be located at an upstream end portion of the turbomachine 10 in the longitudinal direction, as for the example of turbomachine 10 represented in FIG. 11. The turbomachine 10 has, in this case, a so-called “ pull”. In particular, the upstream annular row 14 and the downstream annular row 16 can surround a section of the compressor(s) or of the gearbox of the turbomachine 10.

Alternatively, as shown in Figure 4, the upstream annular row 14 and the downstream annular row 16 may be located at a downstream end portion of the turbine engine 10 in the longitudinal direction. The turbomachine 10 is then in a so-called “pusher” configuration. The upstream annular row 14 and the downstream annular row 16 can surround a section of the turbine(s) of the turbomachine 10.

The invention is not limited to the examples described above and is capable of numerous variants.

According to a variant shown in Figure 12, although the radial dimension of each blade 18 of the downstream annular row 16 is less than the radial dimension of each of the blades 18 of the upstream annular row 14, it can be provided that one or more blade(s) 18 of the downstream annular row 16 has a radially outer radius greater than the radially outer radius of each of the blades 18 of the upstream annular row 14. A relative difference in the radially outer radius of this blade 18 of the downstream annular row 16, Re2', relative to the radially outer radius of any of the blades 18 of the upstream annular row 14, Re1, that is to say (Re1-Re2')/Re1 ^* 100, is preferably between -15% and 0%.

Consequently, this (these) blade(s) 18 of the downstream annular row 16 has a radially internal radius greater than the radially internal radius of each of the blades 18 of the upstream annular row 14. This may be due to the hub 12 has a larger radially outer radius at the downstream annular row 16 than at the upstream annular row 14. downstream. Due to the trajectory or the contraction of the streamlines of the air flow circulating around the hub 12 of the turbomachine 10, the vortices 19 formed at the level of the radially outer end of each of the blades 18 of the annular row upstream 14 does not impact the blades 18 of the downstream annular row 16, even though the latter have a radially outer radius greater than the radially outer radius of each of the blades 18 of the upstream annular row 14.

[0106] Figure 13 shows another variant. FIG. 13 represents a propulsion assembly 24 for an aircraft. The propulsion assembly 24 includes a turbine engine 10 as described above and a pylon 26 for attaching the turbine engine 10 to the aircraft. The attachment pylon is connected to one of the blades 18 of the downstream annular row 16 so as to form a single aerodynamic assembly. For this purpose, the fixing pylon 26 can be connected to one of the blades 18 of the downstream annular row 16 by continuity of material. In other words, the attachment pylon 26 may be integral with one of the blades 18 of the downstream annular row 16. Alternatively, the attachment pylon 26 may be connected to one of the blades 18 of the downstream annular row 16 via one (or more) fixing means. The attachment pylon 26 also has an aerodynamic profile similar to an aerodynamic profile of the blades 18 of the downstream annular row 16. The attachment pylon 26 therefore has the same effect on the air flow from the upstream annular row 14 as the blades 18 of the downstream annular row 16. Such an arrangement makes it possible to further reduce the noise emitted by the turbine engine 10. by way of non-limiting examples in the figures, for example an electric, hydrogen or hybrid thruster. In this context, Figure 14 shows an aeronautical propellant therefore comprising, around the blades 18 of the two annular rows upstream 14 and downstream 16 and coaxially with the longitudinal axis X.

Claims

[Claim 1] Aeronautical thruster (10) of longitudinal axis (X) comprising a hub (12) and at least two annular rows of unducted blades (18) comprising an upstream annular row (14) and a downstream annular row (16 ) spaced apart along said longitudinal axis (X), the upstream annular row (14) being rotatable around the longitudinal axis (X), said downstream annular row (16) comprising a series of blades including a first blade (18a) and a second blade (18b) each extending in a radial direction from the hub (12) so as to define a radial dimension between said hub (12) and a radially outer end of the blade ( 18a; 18b) respectively, characterized in that the radial dimension of the first blade (18a) is greater than the radial dimension of the second blade (18b).

[Claim 2] Aeronautical propellant (10) according to claim 1, in which the downstream annular row (16) is fixed about the longitudinal axis (X).

[Claim 3] Aeronautical propellant (10) according to claim 1 or 2, in which the first blade (18a) and the second blade (18b) of the downstream annular row (16) each have a radially outer radius defined by the said radially outer end. external of the respective blade (18a; 18b), the radially external radius of the first blade (18a) being greater than the radially external radius of the second blade (18b).

[Claim 4] Aeronautical thruster (10) according to the preceding claim, in which each blade (18) of the upstream annular row (14) has a radially outer radius, and in which a relative difference in the radially outer radius of any of the first blade (18a) and the second blade (18b) of the downstream annular row (16) with respect to the radially outer radius of any one of the blades (18) of the upstream annular row (14) is between - 15% and 30%.

[Claim 5] Aeronautical propellant (10) according to any one of the preceding claims, in which the first blade (18a) and the second blade (18b) of the downstream annular row (16) are circumferentially consecutive.

[Claim 6] Aeronautical thruster (10) according to any one of the preceding claims, in which each blade (18) of the upstream annular row (14) extends in a radial direction from the hub (12) so as to define a radial dimension between said hub (12) and a radially outer end of the blade (18) considered, the radial dimension of each of the blades (18) of the upstream annular row (14) being greater than the radial dimension of the first blade (18a) of the downstream annular row (16).

[Claim 7] Aeronautical propellant (10) according to any one of the preceding claims, in which the radially outer ends of the blades (18) of the downstream annular row (16) are inscribed in an outer casing (22) of which a projection in a plane normal (IV-IV) to the longitudinal axis (X) defines a circle.

[Claim 8] Aeronautical thruster (10) according to the preceding claim, in which the center of the said circle is offset with respect to the longitudinal axis (X).

[Claim 9] Aeronautical thruster (10) according to any one of the preceding claims, in which the downstream annular row (16) comprises at least one group (G1; G2; G3; Gi) of blades (18) having the same dimension radial, including at least a first group comprising a plurality of first blades (18a) and/or a second group comprising a plurality of second blades (18b).

[Claim 10] Aeronautical thruster (10) according to the preceding claim, in which the blades (18) of the said at least one group (G1; G2) of blades (18) are arranged circumferentially in a contiguous manner in an angular sector (S1; S2 ) around the longitudinal axis (X).

[Claim 11] Aeronautical thruster (10) according to any one of the preceding claims, in which the first blade (18a) and the second blade (18b) are each disposed angularly about the longitudinal axis (X) at an angle ( aa; ab) respectively, the angle (aa; ab) being measured around the longitudinal axis (X) clockwise with respect to an angular position at 12 o'clock, the radial dimension or the radially outer radius of the first blade (18a) and/or the radial dimension or the radially outer radius of the second blade (18b) being determined as a function of the respective angle (aa; ab) according to a linear, parabolic, logarithmic or exponential law.

[Claim 12] Aeronautical thruster (10) according to any one of the preceding claims, in which the downstream annular row (16) comprises at least one set of blades (18) arranged contiguously in an angular sector (S3; S4) around the longitudinal axis (X), each blade (18) of the set of blades (18) being arranged angularly around the longitudinal axis (X) according to a respective angle a, the angle a being measured around the longitudinal axis (X) in the clockwise direction relative to an angular position at 12 o'clock, each blade (18) of said set of blades (18) having a radial dimension or a radially outer radius determined as a function of the angle a respective according to a linear, parabolic, logarithmic, or exponential law.

[Claim 13] Aeronautical thruster (10) according to the preceding claim, in which the angular sector (S3; S4) associated with said set of blades (18) extends between the angular position at 12 o'clock and an angular position at 6 o'clock.

[Claim 14] Aeronautical thruster (10) according to any one of the preceding claims, in which the second blade (18b) is positioned, angularly around the longitudinal axis (X), closer to an angular position at 6 o'clock than is not the first blade (18a).

[Claim 15] Aeronautical thruster (10) according to any one of the preceding claims, in which the downstream annular row (16) comprises at least one pair of blades (18) whose angular positioning about the longitudinal axis (X) is symmetrical with respect to a plane of symmetry (P) comprising the longitudinal axis (X) and the axis passing through angular positions at 6 o'clock and at 12 o'clock and in which the blades (18) of the said pair of blades have the same radial dimension.

[Claim 16] Aeronautical propellant (10) according to any one of the preceding claims, in which the downstream annular row (16) has a symmetry of revolution of order n with n an integer greater than or equal to 2.

[Claim 17] Aeronautical propellant (10) according to any one of the preceding claims, in which the upstream annular row (14) and the downstream annular row (16) are located at an upstream end portion of the aeronautical propellant (10) in the longitudinal direction or at a downstream end portion of the aeronautical propellant (10) in the longitudinal direction.

[Claim 18] Propulsion assembly (24) for an aircraft, comprising an aeronautical thruster (10) according to any one of the preceding claims and a mounting pylon (26) of the aeronautical thruster (10) to the aircraft, the pylon of attachment (26) being connected to one of the blades (18) of the downstream annular row (16) so as to form a single aerodynamic assembly.