Aerial thruster for improving aeroacoustics

Abstract

An aviation thruster (40) having a longitudinal axis (X) and comprising an upstream rotor row and a downstream stator row (16) of unducted blades (18). Two adjacent vanes of the downstream stator row (16) have an azimuthal spacing ([delta] [theta]) therebetween about the longitudinal axis (X), and at least some of these spacings differ from one another such that at least some of the vanes of the downstream stator have a heterogeneous distribution about the longitudinal axis (X).

Description

Field of the Invention

[0001] The present disclosure relates to the field of aviation propellers having a longitudinal axis, each propeller comprising a hub and (at least) two annular rows of ducted fans along the longitudinal axis, one row being upstream and the other row being downstream.

[0002] According to the foregoing and the following, throughout this text, the relative terms "upstream" and "downstream" are defined relative to each other with reference to the flow of gases in a turbomachine during the cruise flight phase in the longitudinal direction (i.e., the direction of the longitudinal axis).

[0003] An aviation propeller may include (at least) one heat engine, in particular a turbomachine, a turboprop engine, a turbojet engine, a turbofan engine and / or (at least) one electric engine, and / or (at least) one hydrogen engine, and / or (at least) one hybrid engine: a heat and / or electric and / or hydrogen engine. Background Art

[0004] In the following, the case of a turbomachine will be more specifically and thus non - restrictively referred to, since the type of engine included in the propeller is not a determining factor here. A turbomachine is understood here to mean a propeller in which there is an energy exchange between the fluid in flow and the rotor.

[0005] In this context, it will be recalled by way of example that a turbomachine with a "ducted - fan" (or a turboprop engine of the "paddle - fan" or "open - fan" or "open - rotor" or "counter - rotating open - rotor" type) is a type of turbomachine in which the fan extends outside the engine casing (or nacelle), which is different from a conventional turbomachine with a ducted fan (of the "turbofan" type).

[0006] The absence of a duct (such as in a ducted - fan turbomachine) results in an increase in the noise level emitted by the aviation propeller, which generally includes at least one upstream rotor row, the blades of which have an influence on the blades of a downstream stator row.

[0007] In fact, the noise generated by the annular row of ducted fans propagates in the free field. The main cause of the noise is related to the vortex structures generated in the air flow at the free radial outer tips of the blades in the rotor row. These blade - tip vortices can interact with the blades of the downstream stator row.

[0008] One of the challenges of these architectures is the certification of the noise level during take - off and landing operations. The noise level emitted by an aircraft is subject to increasingly strict regulations.

[0009] The main noise sources in a ducted - fan turbomachine are listed below:

[0010] - If the rotor is followed by a stator downstream and the leading edge of the upstream rotor interacts with the downstream stator blades, there is noise from the interaction of the resulting blade tip vortices with the rotor wake. This noise source contributes to an increase in the following noise:

[0011] -- Broadband noise because the turbulence level in the wake is often extremely high at the blade tip,

[0012] -- Tone noise, which is related to the periodic nature of the wake from the upstream rotor and the vortices during the rotation of the rotor blades,

[0013] - Blade noise due to the static loads on the blades (tone noise source in the rotor) and the extension of the boundary layer on the blades (rotor or stator); thus, a broadband noise source is generated when the disturbed boundary layer passes through the trailing edge of the blade; increasing the chord of the blades in a ducted fan engine increases the surface area of the boundary layer extension.

[0014] It should be noted that the absence of a nacelle on a ducted fan engine means a significant reduction in the surface area with acoustic treatment (honeycomb resonators, absorbent materials, specifically porous materials, etc.) and thus means a reduction in noise.

[0015] In addition, when the rotor is subjected to a non-uniform upstream flow that is not parallel to the engine axis - the aforementioned longitudinal axis - (flight angle of attack, with crosswind or installation-related effects), forces and moments appear in the propeller plane, called 1P forces.

[0016] In the case of a non-zero angle of attack, the upstream rotor followed by a downstream stator does not only provide traction in the horizontal axis of the forward movement of the aircraft (not necessarily the longitudinal axis (axis of the engine / aeroengine); it is the horizontal direction of the forward movement). For example, when an aviation thruster is mounted below the wing and / or the thruster can be oriented at a certain angle of attack relative to the upstream flow, the descending rotor blades will experience an increase in the angle of attack and thus experience an increased force, which is different from the ascending blades that are subjected to a reduced force above. Therefore, during one engine rotation, the same rotor blade will experience variable forces depending on its azimuthal position around the longitudinal axis.

[0017] The downstream stator blades will thus also have variable loads depending on their azimuthal position. Axially opposite to the upstream rotor blades, if the angle of attack experienced by the blade at the position of a given blade of the stator is more or less significant, the descending blades of the stator will experience less load and less twist to be corrected, while the ascending blades will experience more load and more twist to be corrected.

[0018] It should also be noted that the upstream angle of attack α (the angle of attack of the aircraft) is not completely filtered out by the upstream rotor. In addition to the variation of the 1P force, due to the angle of attack of the aircraft and the often-present paired pylons (or equivalents) / airfoils, the stator blades will experience various angles of attack depending on their azimuthal position.

[0019] The stator blades located in the upper part of the engine (the area near 12 o'clock) will have a positive angle of attack and thus more load; the stator blades located in the lower part of the engine (the area near 6 o'clock) will have a negative angle of attack and thus less load.

[0020] Conventionally, the 3 o'clock, 6 o'clock, 9 o'clock, and 12 o'clock positions are considered as positions on a clock face and are oriented clockwise when viewing the thruster or aircraft from the upstream / front.

[0021] However, various solutions of the prior art are often only relatively suitable in the isolated configuration of an aircraft propeller and at zero angle of attack. In fact, the presence of surrounding elements (struts, airfoils, fuselage, etc.), the non-zero angle of attack of the airflow sensed by the propeller, and the shape of the blades in the rotor row can, on the one hand, modify the contraction and axial symmetry of the longitudinal axis X of the flow tube of the downstream airflow around the rotor row, and / or on the other hand, modify the magnitude of the vortices present in the downstream airflow of the rotor row, such that the truncation of the blades of the downstream stator row defined based on the isolated configuration and at zero angle of attack can no longer prevent the interaction between the blades of the downstream stator row and the vortices formed by the blades of the upstream rotor row.

[0022] This specification aims to propose solutions to these drawbacks. Summary of the Invention

[0023] At this stage, it should be immediately stated that even though the above prior art relates to a turbomachine, the solution of the present invention is still applicable to any ducted and / or "open rotor" type of aircraft propeller, because part of the above problems is not necessarily specific to the type of aircraft propeller mentioned above.

[0024] In this context, therefore, an aircraft propeller is proposed, which here and generally has a longitudinal axis (X) and includes a housing, and an upstream rotor row of unducted rotor blades and a downstream stator row of unducted stator blades that extend around the housing and are spaced apart from each other along the longitudinal axis (X), and there is an azimuthal pitch (Δθ, Δθ i ) between two adjacent blades in the downstream stator row of the stator blades, which is defined by an angle between corresponding axes around the longitudinal axis (X, or stator axis), and the axes are:

[0025] -- the pitch angle adjustment axes of two adjacent blades, provided that when these axes are projected onto a plane perpendicular to the longitudinal axis (X) and when the two adjacent blades have variable pitch angles, or

[0026] -- axes that are radial to the longitudinal axis (X) and / or that respectively pass through the radially inner ends or the radially outer ends of the two adjacent blades or through their centers of gravity, provided that the two adjacent blades have fixed pitch angles, or

[0027] -- when one of the two adjacent blades has a variable pitch angle, for one of the corresponding axes, it is the pitch angle adjustment axis of the one blade, and when the adjacent blade has a fixed pitch angle, the other of the corresponding axes is radial to the longitudinal axis (X) and / or passes through the radially inner end or through the radially outer end or through the center of gravity of the adjacent blade, and

[0028] Around the longitudinal axis (X), the angular position at 12 o'clock is defined as being vertically upward relative to the longitudinal axis (X), and the angular position at 6 o'clock is defined as being vertically downward relative to the longitudinal axis (X).

[0029] This assembly is characterized in that at least some of the blades of the downstream stator row of the stator blades have a non-uniform distribution around the longitudinal axis (X) such that: at least two of the adjacent blades of the downstream stator row of the stator blades have an azimuthal pitch Δθ or Δθ i such that Δθ i ≠360° / V; Δθ i ≥(360° / V)+1° or Δθ i ≤(360° / V)-1°, where V defines the number of blades in the downstream stator row of the stator blades; and / or such that there are at least two azimuthal pitches between the stator blades of the downstream stator row such that when i≠j, Δθ i and Δθ j are different, where i, j ≤ V and i, j = 1, 2, ….

[0030] Preferably, Δθ i ≥(360° / V)+3° or Δθ i ≤(360° / V)-3°, or preferably Δθ i ≥(360° / V)+5° or Δθ i ≤(360° / V)-5° will be selected in order to optimize the expected effect; it has been observed that these values are particularly relevant with respect to the aforementioned desired effect, but different values can also be expected.

[0031] Δθ i and Δθj Define two different azimuthal pitch distances between any two adjacent blades respectively. i and j are indices (natural numbers: i, j = 1, 2, …), which are different when i ≠ j and less than or equal to the number of stator blades (i, j ≤ V), and V defines the number of blades in the downstream stator row of the stator blades. In other words: i and j are different and can take (any) integer values among 1, 2, 3, … and (at most) V.

[0032] In other words, at least some of the azimuthal pitch distances of the stator blades are different from each other.

[0033] Here, heterogeneous, non - homogeneous, irregular, and non - uniform are synonyms regarding this azimuthal distribution and thus are around the longitudinal axis X (non - homogeneous = varying, not unique everywhere). Conventionally, the pitch angle of a blade can be the angle formed by the chord of one of the profiles and a plane perpendicular to the longitudinal axis X, such as the rotational plane of the blade. Since the blade is twisted by design, conventionally we consider the pitch to be the pitch of the profile at 70% of the maximum radius.

[0034] This configuration makes it possible to consider aerodynamic, acoustic, and / or integration constraints, provided that they are considered only according to the compromise between them. Additionally, this configuration is specifically beneficial from an acoustic point of view in decorrelating the noise sources emitted by the stator or modifying the directivity of the sound radiated by the stator blades (i.e., the regions with maximum sound radiation).

[0035] Furthermore, increasing the number of stator blades in the azimuthal positions close to the pylon (strut or engine mount) and / or the airfoil or the fuselage can be beneficial in reducing the potential impact (pressure buildup) of this assembly on the upstream rotor; thus (relatively) locally increasing the number of stator blades can help “filter out” or reduce the pressure buildup at the upstream rotor, which is related to the presence of the strut, pylon, or engine mount, or the airfoil (all the lifting surfaces of the aircraft) or the fuselage.

[0036] From an acoustic point of view, the non - homogeneous distribution of the stator blades in the azimuthal direction (i.e., circumferentially around the axis around which the blades of the downstream stator row are arranged; in the example, it is the longitudinal axis X) makes it possible to modify the directivity of the interaction noise generated during the interaction between the upstream rotor wake and the downstream stator.

[0037] This makes it possible to define an angular range around the main axis of the aircraft propulsion unit within which the number of stator blades will be reduced in order to reduce the noise emitted towards the ground (“community noise”) and / or the noise emitted towards the passenger cabin (“cockpit noise”).

[0038] From an aerodynamic point of view, the non-uniform distribution of stator vanes makes it possible to better distribute the aerodynamic loads on the stator during flight phases with a large angle of attack (e.g., the angle of attack is high during takeoff and / or landing), as well as the heterogeneity of the loads on the stator vanes associated with the 1P force (the force on the propeller blades in a direction perpendicular to the engine axis, which varies with the azimuthal position and modifies the flow downstream of the upstream rotor).

[0039] From an integration point of view, the non-uniform distribution of stator vanes makes it possible to bypass auxiliary equipment below the casing or hub, reduce the pressure accumulation at the airfoil / lift surface downstream of the stator vanes, and / or avoid the interaction between the stator wake and the airfoil, and, where appropriate, accommodate the integration of the pylon. It should be noted that for USF-type architectures, the pylon can be located upstream of the rotor / stator vanes (referred to as the "pusher" configuration), which is different from the preferred "tractor" configuration where the pylon and / or the stator vane row is at the level height of the stator vanes or downstream thereof. From an acoustic point of view, one advantage of the "tractor" configuration is that it avoids the influence of the wakes from the pylon, struts or engine mounts on the rotor blade assembly. From an aerodynamic point of view, one advantage of the "tractor" configuration is that it does not introduce twist or heterogeneity into the airflow upstream of the rotor, the introduction of which can degrade performance and increase the vibration response phenomenon on the rotor blades. The tractor configuration also allows the engine to be mounted below the wing, which is an advantage for aircraft manufacturers in terms of the center of gravity of the aircraft.

[0040] The above solution thus also has the advantage of being particularly suitable for USF-type aircraft propulsion units. In this context, while the "tractor" USF may be preferred, the "pusher" USF can also be considered.

[0041] In this document and in US 9242721, the associated technical effects and problems are different. The problem pointed out in US 9242721 is the formation of shock-wave related forces at the roots of the rotor blades when the number of upstream rotor blades increases. These forces cause aerodynamic losses, and the reduction in the number of rotor blades increases the noise because the blades experience more loads. To solve this problem of the counter-rotating rotor architecture (CROR), US 9242721 proposed an additional ring of stator vanes upstream of the rotor. The purpose (technical effect) of this stator ring is to avoid the formation of shock-wave related forces at the roots of the blades (to improve efficiency / aerodynamics) and to be able to increase the number of rotor blades (to reduce noise).

[0042] In contrast, in the solution of the present invention, a USF type thruster with a single annular row including ducted rotor blades may be preferred. The single annular row is the upstream row of the aforementioned rotor blades, thus avoiding weight problems, technical complexities associated with integration, and aerodynamic interactions between counter-rotating rotors.

[0043] The term "ducted" used with reference to the upstream rotor row and the downstream stator row indicates that the blades of the upstream rotor row and the blades of the downstream stator row are not surrounded by a nacelle (in other words, the stator blades have a radially outer free end), which is different from a conventional aviation thruster where the fan is ducted inside the nacelle.

[0044] In view of the above, the ducted upstream rotor row will extend around the hub, and similarly, the ducted stator row will extend around the (fixed) housing located downstream.

[0045] Some or all of the blades of the upstream rotor row and / or the downstream stator row may be variable pitch.

[0046] All the azimuthal spacings between two adjacent blades in the series of blades of the downstream stator blade row may be different from each other.

[0047] From an acoustic point of view, this may be beneficial for decorrelating the noise sources emitted by the stator or for modifying the directivity of the sound, i.e., the region of maximum sound radiation.

[0048] If all the blades of the downstream stator row have a homogeneous distribution around the longitudinal axis (X), a correlated situation may also occur, except in a single angular sector. This "homogeneous distribution" may be such that: for any azimuthal spacing between two (circumferentially) adjacent / continuous blades involved, Δθ ≤ 360° / V. The aforementioned angular sector will beneficially be limited between 15° and 75°, or preferably between 25° and 60°. It will thus be possible to limit the noise emitted by the stator blades in combination with a design of lower complexity than other "non-homogeneous" situations, while maintaining an optimized number of rotor and / or stator blades.

[0049] It will thus be possible to consider the presence of a bulge or non-axisymmetric hub extending between two blades of the downstream annular row of stator blades or axially adjacent to these two blades around the longitudinal axis X. In the presence of a pylon or engine mount or strut for attaching the thruster to an aircraft, it will be possible to have an increased azimuthal spacing between two stator blades flanking the bulge (e.g., the pylon).

[0050] This will be effective specifically for the number of stator blades V varying between 8 and 14 that are of concern.

[0051] In the above beneficial cases, the number of blades in the downstream stator blade row will effectively be greater than or equal to 5 for studying noise reduction and / or efficiency emitted by the stator blades: dynamic effects that are prone to generation / maintenance / convection.

[0052] An increase in the spacing will, if necessary, allow ensuring integration below the hub, typically in the nacelle, and thus downstream of the space for the attachment system and / or blade pitch (change) system for an aircraft propeller and / or for auxiliary maintenance ducts (oil, air, etc.).

[0053] For at least one of the upstream rotor row of rotor blades and the downstream stator row of stator blades, it may also be advantageous to have a C / E ratio of the chord C and the azimuthal spacing E between two consecutive blades around the longitudinal axis (X) such that C / E is less than 3 over the entire span.

[0054] Even more advantageously, at the radially outer ends of two blades of the same row (upstream and / or downstream) that are circumferentially or azimuthally consecutive or adjacent, the C / E ratio can be less than 1.

[0055] The number of stator blades (between 8 and 14), which may be optionally selected according to preference, will effectively satisfy this criterion. One advantage at low robustness (C / E ratio) (preferably less than 1 at the tip) is the reduction of blade - to - blade interaction. From an aerodynamic point of view, if C / E is large (greater than 3 or 4), the flow passage or flow area between the blades is reduced. This increases the velocity of the flow between the blades, which may generate shock waves (between the blades) and thus cause efficiency losses at certain operating points. From an acoustic point of view, the lower the robustness (C / E), the more the noise source correlation between the blades is reduced.

[0056] It will be recalled that conventionally, E is the length of the circumferential arc (in meters) between the axes of two adjacent stator blades (e.g., 180a / 180b below). It can be related to Δθ by the equation: E = r*Δθ, where Δθ is measured in radians and r corresponds to the radial position around the longitudinal axis (X).

[0057] It can also be stipulated that the upstream rotor row of rotor blades and the downstream stator row of stator blades have different numbers of blades.

[0058] In the prior art, one drawback is related to adding a stator upstream of the rotor blades. In fact, this creates a risk of generating new noise sources, such as the interaction between the upstream stator wake and the blades of the downstream rotor. Thus, the noise reduction that can be expected by increasing the number of blades of the upstream propeller may (at least partially) be offset by the new noise sources.

[0059] In contrast, in the solution of the present invention, it may be preferred that the number of blades of the upstream rotor is actually greater than the number of blades of the downstream stator, thus limiting unwanted or additional noise sources.

[0060] Thus, for example, in the case where the rotor blades and the stator blades are evenly distributed in the azimuthal direction, using the solution of the present invention, having twelve such blades in the upstream rotor row of the rotor blades and eight such blades in the downstream stator row of the stator blades will generate four rotor wakes that can interact with four stator blades simultaneously, which may increase the noise emitted by the aerospace propulsor. However, in the case of a non-uniform azimuthal distribution of the stator blades, this problem can be avoided while maintaining the above-optimized numbers of rotor and stator blades.

[0061] This is an example illustrating the advantages of a non-uniform azimuthal distribution of the stator, but it is not necessarily the preferred case. The preferred number of blades can actually be (rotor - stator): 12 - 10 or 14 - 12 or even 14 - 11.

[0062] At least for wake equalization / limitation and noise limitation, it is also proposed that there are at least two series of stator blades in the stator blade row, preferably at least three series of stator blades, and each series of stator blades includes one or more stator blades having the same geometric characteristics (at least including the chord (C), thickness (e), and height (such as L2 or L21 below)), where at least one of the geometric characteristics (at least chord, thickness, height) is different from the same geometric characteristics (chord, thickness, height) of the stator blades in another series of stator blades.

[0063] The set of geometric characteristics of the stator blades can include chord, thickness, and height.

[0064] The set of geometric characteristics of the stator blades can also include at least one of camber, sweep, and lean.

[0065] It is also proposed that there are at least three series of stator blades, preferably at least five series of stator blades, and each series of stator blades includes stator blades, where at least one of the geometric characteristics in the set of geometric characteristics is different from the same geometric characteristics of the stator blades in another series of stator blades.

[0066] It is also proposed that there are at least three series of stator blades, preferably at least five series of stator blades, and each series of stator blades includes at least two or exactly two stator blades having the same set of geometric characteristics, where at least one of the geometric characteristics in the set of geometric characteristics is different from the same geometric characteristics of the stator blades in another series of stator blades.

[0067] At least three series of stator vanes are also proposed, preferably at least five series of stator vanes, and each series of stator vanes includes one or more stator vanes. The stator vanes of the same series including a plurality of stator vanes have the same set of geometric characteristics, and at least one geometric characteristic in the set of geometric characteristics is different from the same geometric characteristic of the stator vanes in another series of stator vanes.

[0068] It can also be stipulated that the stator vane row includes at least three adjacent stator vanes, each belonging to different series of stator vanes. In other words, for three adjacent stator vanes each belonging to different stator vane series, it can be stipulated that the stator vane row includes three stator vanes, which are separated from each other by an azimuthal pitch Δθ i and Δθ i+1 and are each stator vanes belonging to different series, where i = 1, 2, ….

[0069] In addition, it is stipulated that:

[0070] - For i, j = 1, 2, …, when i ≠ j, max{Δθ i}-min{Δθ j}≤120°, preferably ≤75°, or even more preferably ≤50°, where i, j ≤ V and V ≥ 5 (V: the number of blades in the downstream row of the stator vanes), and / or

[0071] - ||Δθ i - Δθ 邻近 ||≤120°, preferably ≤75°, or more preferably ≤50°, where V ≥ 5,

[0072] This will prevent too large an angular sector from being completely free of any stator vanes, which may reduce the flow straightening downstream of the upstream rotor, thus resulting in:

[0073] - Aerodynamic problems related to thrust loss and / or efficiency, and / or

[0074] - Problems with the weight distribution and / or balance around the longitudinal axis X, and the aforementioned integration problems.

[0075] One benefit of the aforementioned optimization of the angular values is to ensure that, as defined, there will not be more different azimuthal pitches than stator vanes (physically impossible).

[0076] Δθ i and Δθ 邻近 are two angular sectors or azimuthal pitches (Δθ, Δθ i , Δθ j ), which are adjacent to each other in the circumferential direction, i.e., have a common stator axis (the axis around which the blades of the stator 16 are circumferentially arranged).

[0077] Of course, the stator axis (the axis about which the stator vanes are radially arranged) can typically be the longitudinal axis X.

[0078] In addition, it is provided that the number of different azimuthal spacings (Δθ, Δθ i , Δθ j ) across all the vanes of the downstream row of the stator vanes being between 2 and 6 will allow the vane geometry to be better adapted to the local nature of the flow (affected by the angle of attack) and / or a better distribution of the weight of the runner ring (or vane ring, these two terms being used interchangeably) about the longitudinal axis (X) of the thruster.

[0079] Regarding the aspect of "angle of attack", it should be noted that the angle of attack of the aircraft (angle α hereinafter) can be defined as the angle between the longitudinal axis of the fuselage (axis X1 hereinafter) and the direction of the flow upstream of the fuselage (or the forward direction of the aircraft). When the longitudinal axis X of the thruster and the longitudinal axis of the fuselage (hereinafter denoted as 33) or the longitudinal axis X1 of the aircraft are projected onto a vertical plane passing through the 12 o'clock and / or 6 o'clock positions and containing the longitudinal axis X of the thruster, there may be a non-zero angle (angle β hereinafter) between these axes (angle β is sometimes referred to as the "tilt angle" or "tilt angle"). This is the plane to be considered in what is mentioned hereinafter Figure 2 wherein angle β is shown (here, in a non-exclusive / non-limiting manner, in the case where the thruster is mounted below the wing (reference numeral 31 hereinafter) of the aircraft involved).

[0080] The longitudinal axis of the fuselage (or the aircraft, axis X1 hereinafter) can be defined as the roll axis of the aircraft, which can correspond to:

[0081] - an axis extending from the nose of the fuselage (upstream; reference numeral 33a hereinafter) to the tail (downstream), or alternatively

[0082] - in cruise flight, an axis passing through the most upstream and most downstream positions of the fuselage.

[0083] These axes X and X1 may not be parallel (β≠0°). For example, this can be useful for reducing the angle of attack and thus reducing the 1P force perceived by the rotor blades during takeoff. To minimize these adverse effects on the aerodynamic and mechanical performance of the blades, the absolute value of angle β can vary between 0.5° and 30°, preferably between 2° and 20°, or more preferably between 3° and 10°.

[0084] The absolute value aspect of the angle (||β||) is important because the tilt:

[0085] - will typically be downward in the case where the thruster is mounted below the aircraft wing, but

[0086] - when mounted towards the rear of the fuselage, it can be upward,

[0087] This is to limit the influence of the angle of attack during takeoff and / or landing.

[0088] In this regard, the present invention can also be relevantly applied to an aircraft (which will have a longitudinal axis (X1) and will include an aviation propeller, a fuselage with all or part of the features mentioned in the current text, and the wing to which the propeller will be fixed), and this aircraft advantageously enables:

[0089] -- the angle β (the absolute value ||β||) between the longitudinal axis (X) of the aviation propeller and the longitudinal axis (X1) of the aircraft will vary between 0.5° and 30°, preferably between 2° and 20°, or more preferably between 3° and 10°, and / or

[0090] -- d1 ≠ d2, and d1 or d2 is less than 0.75*D, preferably less than 0.5*D, or more preferably less than 0.3*D.

[0091] As mentioned above, the concern for the angle β relates to increased efficiency during takeoff and / or landing and / or during installation towards the rear of the fuselage; the concern related to d1 or d2 will cover the case where the aviation propeller is attached to the wing (specifically, below the wing) or more generally to any airfoil of the aircraft involved.

[0092] It should be noted that for the propeller according to the present invention, when facing / viewing it from the front, it is fixed in front of the airfoil of the aircraft:

[0093] - d1 can be defined as the axial distance (along the longitudinal axis X) between the trailing edge (TE) of the stator blade at the free end (hereinafter denoted as 25) and the leading edge (LE) of the airfoil (or wing). This is the case for the stator blade that is (azimuthally) closest to the leading edge of the airfoil (or wing) and is included in the angular sector located between 12 o'clock and 6 o'clock and including 9 o'clock (for example, in the Figure 10 mentioned below, d1 is measured relative to the circled blade denoted as d1), and

[0094] - d2 can be defined as the axial distance (along the longitudinal axis X) between the trailing edge of the stator blade at the free end and the leading edge of the airfoil (or wing). This is the case for the stator blade that is (azimuthally) closest to the leading edge of the airfoil and is included in the angular sector located between 12 o'clock and 6 o'clock and including 3 o'clock (for example, in the same Figure 10 mentioned below, d2 is measured relative to the circled blade denoted as d2).

[0095] In other words:

[0096] - d1 may relate to the radially innermost stator vane (as mentioned above, for a thruster located below the right wing of an aircraft, it is typically near 9 o'clock, or for a thruster located below the left wing, it is typically near 3 o'clock), and

[0097] - d2 may relate to the radially outermost stator vane (as mentioned above, for a thruster located below the right wing, it is typically near 3 o'clock, or for a thruster located below the left wing, it is typically near 9 o'clock).

[0098] To address the situation of inappropriate blade loading (see above), the present invention may be relevantly applied to the aircraft having the following specific features: the absolute value of the angle β (||β||) will vary between 0.5° and 30°, preferably between 2° and 20°, or more preferably between 3° and 10°.

[0099] In addition, the ratio of the following two is specified:

[0100] - the distance (S) along the longitudinal axis (X) between two central planes perpendicular to the longitudinal axis of the upstream rotor row of the rotor blades and the downstream stator row of the stator blades respectively, and

[0101] - the maximum diameter (D) of the aircraft propulsion unit at the radially outer ends of the blades of the upstream rotor row of the rotor blades or the downstream stator row of the stator blades

[0102] (In other words, the spacing (S) between the pitch angle adjustment axes of the two upstream / downstream rows or the axes where the centers of gravity of the blades are located, and the engine diameter (D), so it is S / D)

[0103] is between 0.01 and 0.8, and preferably between 0.15 and 0.35, making it possible to limit certain critical wake interferences between the two blade rows, and thus reduce noise, while limiting the axial length of the aircraft propulsion unit.

[0104] If ground noise is to be reduced, the number of stator blades extending towards the ground should be reduced. Therefore, the maximum angular azimuthal spacing (Δθ, Δθ i , Δθ j ) will be between the blades located at angular positions between 8 o'clock and 4 o'clock (in the case of a pusher configuration with pylons at 3 o'clock or 9 o'clock, there may be 16 stator blades between 10 o'clock and 2 o'clock), specifically or preferably at 2 o'clock and 4 o'clock and / or at 8 o'clock and 10 o'clock.

[0105] If it is desired to promote the balance of the propeller weight and avoid residual moments associated with the inhomogeneous distribution of the stator on the longitudinal axis X, the following preferences will be given:

[0106] - The distribution of the blades of the downstream stator blade row located between 2 o'clock and 4 o'clock and those between 8 o'clock and 10 o'clock, when viewed from upstream around the longitudinal axis (X), is symmetric with respect to the symmetry axis passing through the longitudinal axis (X) and through 12 o'clock and 6 o'clock, and

[0107] - The stator blades of the downstream row located in symmetric positions (θ and -θ) with respect to the axis passing through the longitudinal axis (X) and through 12 o'clock and 6 o'clock have the same blade thickness and height.

[0108] It will be recalled that:

[0109] - (Blade) thickness corresponds to the maximum length or distance between the pressure side and the suction side of the section of this blade, in a direction perpendicular to the straight line connecting the leading edge and the trailing edge of said section,

[0110] - (Blade) height is measured between the radially inner end 23 (at the hub or nacelle) and the radially outer end 25 (free end) of the blade in question.

[0111] In order to ensure that there is only one rotor blade wake interacting with the stator blades at a time, and thus reduce the noise source, the azimuthal spacing between the pitch angle adjustment axes of the blades of the upstream rotor row and the downstream stator row will be defined as follows: ||θ r,n -θ s,m || > 1° or preferably ≥ 2°,

[0112] where θ r,n and θ s,m correspond respectively to the times when the pitch change axes of the blades of the upstream rotor blade row and the blades of the downstream stator blade row are aligned when projected onto a plane perpendicular to the longitudinal axis (X), the angular positions of the pitch angle adjustment axes of the nth blade of the upstream rotor blade row and the mth blade of the downstream stator annular row, n being a natural number varying between 1 and B, B being the number of blades in the upstream rotor blade row, and m being a natural number varying between 1 and V. More generally, for stator blades with fixed or variable pitch, this will be the axis defined by the "stator axis".

[0113] In combination with this sole wake effect, at least some of the blades in the upstream rotor blade row and / or the downstream stator blade row can advantageously differ from each other in terms of their chord (C) and thickness (e).

[0114] And at least some of the blades in the upstream rotor blade row can also have an inhomogeneous distribution around the longitudinal axis (X).

[0115] In addition to the aircraft propellers mentioned above, the present specification also relates to an aircraft having a longitudinal aircraft axis (X1), said aircraft comprising at least one such aircraft propeller and a structure to which the aircraft propeller is attached.

[0116] In this case, the structure of the aircraft will generally comprise a fuselage and an angular sector about the longitudinal axis (X), where the maximum number of blades in the downstream stator blade annulus can be located in the upper part and / or in a position towards the fuselage.

[0117] Thus, the noise emitted towards the ground will be limited and the population near the airport will be protected. The noise of the blades located towards the top and / or towards the interior can have a reduced radiation towards the ground due to its azimuthal position and the possible shielding effect produced by the airfoil (if located below / above the wing), the fuselage and the struts, pylons or engine mounts used to attach the propeller to the aircraft.

[0118] In different methods, if it is desired to minimize the noise radiated towards the passenger cabin and the acoustic interaction with the fuselage, it will be preferable for the angular sector about the longitudinal axis (X) having the maximum number of blades in the downstream stator blade row to be located in the upper part and / or in the region of the downstream stator blade row furthest from the fuselage.

[0119] Furthermore, increasing the number of stator blades in the azimuthal positions close to the struts, pylons or engine mounts and close to the airfoil (if the airfoil is adjacent) can also be beneficial for reducing the potential effects (pressure build-up) at the upstream rotor. Subsequently, it will be advantageous to select the angular sector about the longitudinal axis (X) where the number of blades in the downstream stator blade row is maximum as the position where the distance between the trailing edge of the blades in the downstream stator blade row parallel to the longitudinal axis (X) and the leading edge of the airfoil is minimum.

[0120] Another possible consideration is that the stator facing the upstream rising rotor blades experiences more load and more twist to be corrected. For this purpose, considering the direction of rotation defined for the upstream rotor blade row, it is recommended that the angular sector about the longitudinal axis (X) where the number of blades in the downstream stator blade row is maximum is located on the side of the propeller where the relevant blades in the upstream rotor blade row are expected to rise.

[0121] This increase in the number of blades on said side will allow a better distribution of the stator load, which can also be beneficial for reducing noise.

[0122] Each blade in the upstream rotor row can extend in a radial direction from the hub so as to define a radial dimension (or blade height) between the hub and the radially outer end of the blade in question. The individual dimension of (possibly each) blade in the upstream rotor row is greater than the radial dimension of each blade in the downstream stator row in question between the housing and the radially outer end of the blade in question. In other words, compared with the blades of the upstream annular row, the blades of the downstream stator row can be truncated at their free ends. This limits the impact of the eddy current formed at the radially outer end of the blade in the upstream rotor row on the blades of the downstream stator row. "Truncated blade" means that the blade has a reduced radial dimension and / or a reduced radially outer end (or end surface area). Alternatively, it can be stipulated that at least one blade of the upstream row has a larger radial dimension than at least one blade of the downstream row. In another alternative, it can be stipulated that the radial dimension of at least one blade of the upstream rotor row is greater than the individual radial dimension of (possibly each) blade of the downstream row.

[0123] The radial dimension of a blade is measured between the radially inner end of the blade and the radially outer end of the blade, and the radially inner end is located (i.e., closest to) the hub (correspondingly, the housing) of the aviation propeller. The radially inner end of the blade can be longitudinally located at the leading edge of the blade (e.g., for a fixed blade) or at the pitch change axis of the blade in question. The radially inner end of the blade is also referred to as the "blade root".

[0124] The angular position of each blade about the longitudinal axis can be identified by the angular position about the longitudinal axis of the inner end of the corresponding blade. The radially outer end of the blade is the end opposite to the radially inner end. The radially outer end of the blade can be the free end of the blade. The radially inner end and the individual radially outer ends of (possibly each) blade can be radially aligned, i.e., located at the same longitudinal position, or can be longitudinally offset from each other.

[0125] The downstream stator row can include 3 to 25 blades. The number of blades in the upstream rotor row can be different from the number of blades in the downstream annular row, and preferably B≥V + 1 or more preferably B≥V + 2. This further minimizes the noise level emitted by the aviation propeller.

[0126] As already indicated by reference to the C / E ratio, the robustness of the downstream annular row is defined as the ratio between the chord and the spacing between two circumferentially consecutive blades in the circumferential direction, and this ratio can be less than 3 over the entire radial dimension of each blade. Specifically, in a preferred embodiment, the robustness is less than 1 at the radially outer end of the blade.

[0127] The ratio between the distance between the central planes of each annular row perpendicular to the longitudinal axis in the longitudinal direction and the diameter of the aerospace propulsor may vary between 0.01 and 0.8 and preferably between 0.15 and 0.35. The central plane perpendicular to the respective longitudinal axis of each annular row may be the plane containing the respective pitch change axis of each blade of the corresponding annular row.

[0128] This limits or even avoids the interference between the annular rows of blades.

[0129] The upstream rotor row and the downstream stator row may be located at the upstream end portion of the aerospace propulsor in the longitudinal direction, or at the downstream end portion of the aerospace propulsor in the longitudinal direction.

[0130] The aerospace propulsor may have a so-called "tractor" configuration (the upstream rotor row and the downstream stator row are located at the upstream end portion of the aerospace propulsor) or a so-called "pusher" configuration (the upstream rotor row and the downstream stator row are located at the downstream end portion of the aerospace propulsor).

[0131] In the tractor configuration, the upstream rotor row and the downstream stator row may surround a section of the compressor or the reduction gearbox of the aerospace propulsor. In the pusher configuration, the upstream rotor row and the downstream stator row may surround a section of the turbine of the aerospace propulsor.

[0132] According to one aspect, the aerospace propulsor may sequentially include, from upstream to downstream along the longitudinal axis (X):

[0133] - at least one compressor,

[0134] - at least one combustion chamber,

[0135] - at least one turbine driving the compressor, and

[0136] - an air inlet to the compressor, the air inlet being located downstream of the upstream rotor row of rotor blades and upstream of the downstream stator row of stator blades, in other words, longitudinally along the propulsion device between the rotor blades and the stator blades.

[0137] According to another aspect, a propulsion assembly for an aircraft is described, which includes the aerospace propulsor as described above and an attachment hanger for attaching the aerospace propulsor to the aircraft, the attachment hanger being connected to one of the blades of the downstream stator row to form a single aerodynamic assembly.

[0138] According to another aspect, an aircraft is described, which includes the aerospace propulsor as described above or the propulsion assembly as described above. Description of the Drawings

[0139] Other features, details and advantages will become apparent by reading the following detailed description and by analyzing the attached drawings, in which all the blades are ducted-free and:

[0140] Figure 1 is a partial schematic cross-sectional view of a turbomachine that can be used in this context and thus has an upstream rotor and a downstream stator in a "thruster" configuration,

[0141] Figure 2 is a schematic view of a propeller in a configuration that can be a "tractor" configuration and which, in a phase that can be the take-off phase, thus has an angle of attack of the aircraft (angle α);

[0142] Figure 3 is a partial schematic cross-sectional view of a turbomachine that can be used in this context in a "tractor" configuration,

[0143] Figure 4 can be represented in a cross-sectional plane IV-IV (stator) perpendicular to the longitudinal axis X Figure 3 of a turbomachine, an example of a possible arrangement of an annular row of blades of a downstream stator,

[0144] Figure 5 is a schematic view from the front (viewed from upstream) in the same cross-sectional plane as in Figure 4 showing another arrangement of the annular row of blades of the downstream stator;

[0145] Figure 6 is a schematic view in the same cross-sectional plane showing another arrangement of the annular row of blades of the downstream stator;

[0146] Figure 7 is a schematic view in the same cross-sectional plane showing another arrangement of the annular row of blades of the downstream stator;

[0147] Figure 8 is a schematic view in the same cross-sectional plane showing another arrangement of the annular row of blades of the downstream stator;

[0148] Figure 9 is a schematic view in the same cross-sectional plane showing another arrangement of the annular row of blades of the downstream stator;

[0149] Figure 10 is Figure 9 a schematic front half-view (viewed from upstream) of a solution in which the propeller is attached under the wing;

[0150] Figure 11 is Figure 10 a schematic upper half-view in which the stator blade configuration can be Figure 4 ​​​​​​​​​​​a stator vane configuration of 9;

[0151] Figure 12 is also a schematic view in the same cross - sectional plane, showing another arrangement of the annular row of blades of the downstream stator; Figure 4

[0152] Figure 13 is also a schematic view in the same cross - sectional plane, showing another arrangement of the annular row of blades of the downstream stator;

[0153] Figure 14 also shows in the same cross - sectional plane the required azimuthal angular spacing between the blades of the upstream rotor and the blades of the downstream stator;

[0154] Figure 15 is a diagram integrating at least one downstream stator vane into an attachment system for fixing the thruster to the aircraft;

[0155] Figure 16 shows another solution where it is mounted via an engine mount between the thruster and the wing of the aircraft;

[0156] Figure 17 is a diagram of an aircraft equipped with two thrusters attached to the fuselage via struts, each thruster having a non - homogeneous azimuthal angular spacing of the blades in the downstream stator,

[0157] Figure 18 and

[0158] Figure 19 Figure 19 is a diagram showing the stator vane (downstream vane) and the way of considering the pitch angle of this vane, Figure 18 corresponding to Figure 2 cross - section XVIII - XVIII, this figure and

[0159] indicate the air flow around the thruster (lines with multiple arrows); Figure 20 schematically shows the angle or "azimuthal angular spacing" Δθ i or Δθ j between two consecutive stator vanes, and

[0160] Figure 21 Figure 2 can supplement and schematically represents in a side view the situation with the angle of attack of the aircraft, where the thruster is in a configuration that can be a "tractor" configuration, in a stage that can be the take - off stage, and thus has a non - zero angle β in the example. DETAILED DESCRIPTION

[0161] ​​​​​​​​​​​As an example, an aircraft propeller compatible with the content proposed by the present invention may be a turbomotor, such as Figures 1 to 3 the turbomotor in

[0162] Any propeller mentioned herein, such as the turbomotor 10, includes a hub 12 located upstream (AM) of the engine housing 13. An annular upstream rotor row 14 of ductedless blades 18 is mounted on the hub 12 (around), and an annular downstream stator row 16 of ductedless blades 18 is mounted on the engine housing 13 (around). The two rows are spaced apart from each other along the longitudinal axis X of the turbomotor 10.

[0163] The hub 12 and the engine housing 13 may be combined under a so-called nacelle 40, which is a structure around which the blades 18 of the rotor 14 and the stator 16 are arranged and extended. The nacelle 40 is fixed to the aircraft that the aircraft propeller mentioned herein will drive.

[0164] As is understood, orientation qualifiers such as "longitudinal", "radial" or "circumferential" are defined with reference to the longitudinal axis X of the propeller concerned, as in the case of the turbomachine 10. The longitudinal direction corresponds here to the forward direction of the propeller or the axis of rotation of the blades of the upstream rotor 14. Specifically, the longitudinal direction can coincide with the horizontal direction, i.e., perpendicular to the gravitational field. The relative qualifiers "upstream" (AM) and "downstream" (AV) are defined relative to each other with reference to the gas flow in the propeller along the longitudinal direction. The angular position of each of the blades 18 about the longitudinal axis X is identified relative to a clock face (here, for example, viewed from upstream), where the angular positions at 12 o'clock, 3 o'clock, 6 o'clock and 9 o'clock are positioned on said clock face in a conventional manner. Thus, the angular position at 12 o'clock is positioned vertically upward relative to the longitudinal axis X, and the angular position at 6 o'clock is positioned vertically downward relative to the longitudinal axis X. The angular position at 3 o'clock is positioned horizontally to the right relative to the longitudinal axis X, and the angular position at 9 o'clock is positioned horizontally to the left relative to the longitudinal axis X. Thus, the axis extending radially through the angular positions at 12 o'clock and 6 o'clock is perpendicular to the axis extending radially through the angular positions at 3 o'clock and 9 o'clock. Absolute position qualifiers such as the terms "up", "down", "left", "right", etc., or relative position qualifiers such as the terms "above", "below", "upper", "lower", etc., and orientation qualifiers such as the terms "vertical" and "horizontal" refer here to the orientation of the figure and are considered when the propeller is in the operating state, which is usually mounted on an aircraft placed on the ground. In this state of the turbomachine 10, the axis passing through the angular positions at 12 o'clock and 6 o'clock extends in the direction of the gravitational field, i.e., vertically. However, it can be deduced that a rolling motion during the flight of the aircraft on which the propeller is mounted will cause it to induce a rotation of the vertical and horizontal directions about the longitudinal axis X as considered in the figure. In the same way, a rolling motion during the flight of the aircraft on which the propeller is mounted above will cause the axis passing through the angular positions at 12 o'clock and 6 o'clock and the axis passing through the angular positions at 3 o'clock and 9 o'clock to rotate about the longitudinal axis X. The "lateral region" of the turbomachine 10 refers to the region circumferentially near the angular position at 3 o'clock or 9 o'clock. Similarly, the "upper region" and "lower region" of the propeller refer respectively to the region circumferentially near the angular position at 12 o'clock and the region circumferentially near the angular position at 6 o'clock.

[0165] Thus, the downstream stator row 16 (or stator) is fixed about the longitudinal axis X. In other words, the downstream stator row 16 does not rotate about the longitudinal axis X. This does not exclude the possibility that each blade 18 of the downstream stator row 16 can have a variable pitch.

[0166] If the aircraft propeller involved is (or includes) a turbomotor, then it will thus be a turbomotor that, parallel to the longitudinal axis (X), successively includes the following elements from upstream to downstream inside the nacelle 40 (including below the engine casing 13):

[0167] - one (or more) compressors 2,

[0168] - at least one combustion chamber 4,

[0169] - one (or more) turbines 6 driving the compressor, and

[0170] - at least one exhaust nozzle 8.

[0171] Among these turbomotors with an unducted fan, there are known turbomotors of the "unducted single (or stator) fan" (USF) type, as Figures 1 to 3 illustrated, in each of said turbomotors, the upstream rotor row 14 of the unducted blades 18 is mounted to rotate about the longitudinal axis X, and the downstream stator row 16 of the unducted blades 18 is fixed. The direction of rotation of the blades 18 of the upstream rotor row 14 (or rotor) is not a determining factor.

[0172] The downstream stator row 16 can be centered on an axis that may or may not coincide with the longitudinal axis X. In the presented example, the downstream stator row 16 is centered on the longitudinal axis X. This configuration of the upstream rotor row 14 and the downstream stator row 16 makes it possible to utilize the energy of the eddy air flow from the upstream rotor row 14 through the downstream stator row 16. Thus, the efficiency of the turbomotor 10 is improved, particularly compared to a single rotating propeller (such as 14) in the case of a conventional turboprop. The upstream rotor row 14 rotates about the longitudinal axis X by the turbine 6 driving the compressor 2. The turbomotor 10 generally includes a reduction gearbox in order to decouple the rotational speed of the turbine 6 from the rotational speed of the upstream rotor row 14. In addition, one of the advantages of the USF type turbomotor compared to the "counter-rotating open rotor" type turbomotor is the reduction of the tonal noise emitted by the turbomotor, since the downstream stator row 16 of the unducted blades 18 is fixed.

[0173] As schematically shown in Figure 2 and 3 the propeller can have a "tractor" configuration (the upstream rotor row 14 and the downstream stator row 16 are located at the upstream end portion of the propeller), or as schematically shown in Figure 1 a "pusher" configuration (the upstream rotor row 14 and the downstream stator row 16 are located at the downstream end portion of the propeller).

[0174] In a tractor configuration, the upstream rotor row 14 and the downstream stator row 16 may surround a section of the compressor 2 of a turbomotor or a reduction gearbox. In a pusher configuration, the upstream rotor row 14 and the downstream stator row 16 may surround a section of the turbine 6 of a turbomotor 10.

[0175] Regardless of the type of pusher (turbomotor, hybrid, etc.), the attachment system 27 will allow the pusher to be fixed to an aircraft 29 equipped with said pusher and, more precisely, to its airfoil (wing) 31, or to its fuselage 33, or to any other suitable part. Generally, this can be done using:

[0176] - For the fuselage: struts 35 (as in the example of Figure 3 , 7 ), or

[0177] - For the attachment to a wing or airfoil: pylons 37 (e.g., as in Figure 3 , 11 ), or engine mounts 39 (e.g., as in Figure 16 ).

[0178] The blades 18 of the upstream rotor row 14 and / or the downstream stator row 16 may be variable pitch blades. Thus, the pitch of the blades 18 of the turbomotor 10 can be adjusted according to the operating point of the pusher or the flight phase. A pitch change system 38 can be provided, which is partly located in the nacelle 40 (hub 12 and / or housing 13) in order to adjust the angle of attack of the blades for each flight phase. Thus, the rotation of each blade 18 can be adjusted about a corresponding pitch change axis 19. The individual pitch change axis 19 of (possibly each) blade 18 is an axis that:

[0179] - Extends radially and / or is longitudinally positioned at the middle part of the corresponding blade, and

[0180] - The pitch angle of the blade can be adjusted about said axis.

[0181] In this regard, the present disclosure covers the cases where:

[0182] - The pitch change axis is perpendicular to the longitudinal axis X,

[0183] - The pitch change axis is not perpendicular to the longitudinal axis X, i.e., it is inclined; for example, in the case where the pitch change axis has a longitudinal component and / or a circumferential component with respect to the longitudinal axis X.

[0184] In order to more precisely (re)define the pitch angle of the blades if necessary, it is provided that each downstream stator blade 18 defines an aerodynamic profile. For this purpose, each downstream stator blade includes a stack of sections 30 along the radial direction. Figure 18One of the segments 30 is shown. Each segment 30 extends in a respective cross - sectional plane perpendicular to the radial extension direction of the corresponding downstream stator vane. Each segment 30 includes an upstream leading edge and a downstream trailing edge, and a pressure - side line 330 and a suction - side line 340 extend between the upstream leading edge and the downstream trailing edge. Each segment 30 defines an aerodynamic profile. Each segment 30 also includes a chord C defined by a straight - line segment connecting the leading edge to the trailing edge.

[0185] The pitch angle γ of each downstream stator vane 18 (see, for example Figures 18 - 19 ) will correspond to the angle formed between, on the one hand, a first axis A1, which is defined by the intersection point between the cross - sectional plane of a reference segment 30 among the stack of segments 30 of the downstream stator vane and a plane perpendicular to the longitudinal axis X, and the plane perpendicular to the longitudinal axis X may include the pitch axis AC of the downstream stator vane (this is usually the case, but not necessarily, when the pitch - change axis is perpendicular to the axis X); on the other hand, the chord C of the reference segment 30 of the downstream stator 16 vane. The pitch angle γ is measured on the upstream side of the plane perpendicular to the longitudinal axis X, and the plane perpendicular to the longitudinal axis X as described above may include the pitch axis AC of the downstream stator vane 18. The pitch angle γ is measured positively in the direction oriented from the first axis A1 towards the chord C of the reference segment 30, and more particularly in a direction consistent with the direction oriented from the pressure - side line 330 to the suction - side line 340.

[0186] Here, the reference segment 30 of each downstream stator vane 18 is located on the corresponding downstream stator 16 vane at a radial distance from the longitudinal axis X, and the radial distance corresponds to 75% of the outer radial radius of the corresponding downstream stator vane.

[0187] Each blade 18 of the upstream rotor row 14 and the downstream stator row 16 extends in the radial direction from the hub 12 so as to define the radial dimension between the hub 12 and the outer radial end of the corresponding blade 18. In other words, the radial dimension of the blade 18 corresponds to its height between its inner radial end 23 and its outer radial end 25. The inner radial end of each blade 18 is located at the hub 12 of the turbomachine 10. Specifically, each blade 18 may be fixed to the hub 12 of the turbomachine 10 at its inner radial end. Here, the outer radial end of each blade 18 is a free end (i.e., unshrouded). It is stipulated that the span of the blade 18 is thus the radial distance between its inner end 23 and its outer end 25 (see Figure 9 ), where:

[0188] For the blades of the upstream rotor row, L1 = Re1 - Ri1, and

[0189] For the blades of the downstream stator row, L2 = Re2 - Ri2.

[0190] In addition, the radial inner radius of each blade 18 of the upstream rotor row 14 and the downstream stator row 16 is Ri1 and Ri2 respectively, which is considered to be the radial distance from the longitudinal axis X of the radial inner end of the blade 18, for example located at (i.e., closest to) the hub 12 (rotor row) or the housing 13 (stator row). In Figure 3 the radial inner end 23 is close to the pitch change axis of the corresponding blade. Alternatively, the radial inner end of each blade may be close to the leading edge at the blade root. The radial outer radius of each blade 18, such as Figure 3 Re1 or Re2 in, is considered to be the radial distance from the longitudinal axis X of the radial outer end of the blade 18, i.e., the maximum radius of the blade.

[0191] For example, by viewing Figure 4 it can be understood that the radial outer ends 25 of the blades 18 of the upstream rotor row 14 and the downstream stator row 16 are respectively inscribed in the circumferential envelope 20 around the upstream rotor row 14 and the circumferential envelope 22 around the downstream stator row 16.

[0192] The projection of the circumferential envelope 20 of the downstream stator row 16 onto the cross-sectional plane IV-IV (see Figure 1 or 3) can define a circle with a radius of Re2 or a diameter of Ds, and the circle can be centered on the longitudinal axis X (Ds = 2 * Re2).

[0193] The diameter D or the circle with a radius of Re1 in the radial cross-sectional plane at the circumferential envelope 20 of the upstream rotor row 14 can represent the outer diameter of the considered propeller, which is the turbine engine 10 in the example (see Figure 1 or 3).

[0194] The radial dimension of each blade 18 of the downstream stator row 16 can be smaller than the individual radial dimension of (possibly each) blade 18 of the upstream rotor row 14, so as to limit the influence of the eddy current formed at the radial outer end of the blade 18 of the upstream rotor row 14 on the blade 18 of the downstream stator row 16. When these envelopes are projected onto a common projection plane perpendicular to the longitudinal axis X (here, for example, the cross-sectional plane IV-IV), the circumferential envelope 20 of the upstream rotor row 14 will then enclose the circumferential envelope 22 of the downstream stator row 16.

[0195] This must be compatible with the trimming within 360° (i.e., the truncation of the downstream stator blades, such as the solution associated with Figure 3 or 15 as a non-limiting example), and it is stipulated that a homogeneous trimming is not necessarily required. In other words, there may be at least one blade 18 of the downstream stator row 16 having a radius of Re2, which allows the definition of the circle 22, but other blades of the downstream stator row 16 may have a radius smaller than Re2.

[0196] Furthermore, here again, in order to promote a more balanced control of the blade loading and the generated noise, it is proposed that:

[0197] - Thus, in a non-limiting example of Figure 9 each blade of the downstream stator row 16 of the stator blades has a height L2 or L21 between a radially inner end 23 and a radially outer end 25,

[0198] - The corresponding heights, such as L2 and L21, of at least two blades 18 of the downstream stator row 16 are advantageously and beneficially different.

[0199] It should also be noted that the projection of the circumferential envelope of the downstream stator row 16 onto a common projection plane perpendicular to the longitudinal axis X (for example, the cross-sectional plane IV-IV in the example) can also define a circle, or even an ellipse as described above, where the center can be offset with respect to the longitudinal axis X, for example, in the direction of the axis passing through the angular positions at 12 o'clock and 6 o'clock. The radial distance between the center of the circumferential envelope 22 of the downstream stator row 16 having a circular shape and the longitudinal axis X can be between 0.005Ds and 0.2Ds.

[0200] The circle / ellipse defined by the circumferential envelope 22 of the downstream stator row 16 can have a radius Re2 (for example, the maximum radius in the case of an elliptical shape), and the radius Re2 is less than the radius Re1 of the circumferential envelope 20 of the upstream rotor row 14 (for example, the maximum radius in the case of an elliptical shape).

[0201] Therefore, the non-uniform distribution (in the azimuthal direction) of the blades 18 of the downstream stator 16 can be compatible with other noise reduction techniques such as, for example, "360° clipping". Thus, it is possible in at least one angular sector:

[0202] - Arrange the blades 18 of the downstream stator 16 non-uniformly (in the circumferential direction), and

[0203] - For the blades 18 of the downstream stator 16, each has or individually has a maximum radius (Re2) or height less than the maximum radius (Re1) or height of the blades 18 of the upstream rotor 14.

[0204] In this case, there may be shorter blades 18 of the stator 16 between 8 o'clock and 4 o'clock (in the case of a pusher configuration with a hanger at 3 o'clock or 9 o'clock, there may be blades of the stator 16 between 10 o'clock and 2 o'clock). However, specifically or preferably, it is possible to favor the stator blades in the lower part (between 4 o'clock and 8 o'clock) and the sides (between 2 o'clock and 4 o'clock or between 8 o'clock and 10 o'clock, towards the outside and / or towards the fuselage); all of this is to minimize the interaction noise during the phases with a larger angle of attack (landing / takeoff).

[0205] The center of the circle inscribed within the radially outer end 25 of each blade 18 of the downstream stator row 16 of stator blades can be offset relative to the longitudinal axis X, such that it is possible to adapt the configuration of these blades to their environment (position on the aircraft / type of noise to be controlled / fluid flow to be promoted / mechanical constraints to be met, etc.).

[0206] Having fewer blades in the stator 16 than in the upstream rotor 14 can also be used to combine noise reduction, aerodynamic efficiency, less stress on the load on certain blades from the downstream stator, and weight and size reduction. Recommendation: B ≥ V + 1, or preferably, B ≥ V + 2.

[0207] According to the important aspects mentioned above, for the reasons mentioned: aerodynamics, acoustics, and / or integration constraints, attention is thus focused here on the blades of the downstream stator 16 having a non - homogeneous azimuthal pitch.

[0208] Depending on the required trade - offs for the objectives (aerodynamics, acoustics, integration, etc.) or areas of professional concern, several embodiments can be envisaged.

[0209] As pointed out above, two adjacent blades, such as 18a, 18b, of the downstream stator row 16 of stator blades have an azimuthal pitch (Δθ, Δθ i , Δθ j ) around the longitudinal axis (X) defined by the angle between their respective axes 180a, 180b.

[0210] These respective axes on the stator 16 are:

[0211] - The pitch - angle adjustment axes of the two adjacent blades (axis 19 above), provided that when these axes are projected onto a plane perpendicular to the longitudinal axis X and in the case where the two adjacent blades have variable pitch angles,

[0212] - Or axes respectively radial to the longitudinal axis X and / or passing through the radially inner end 23 or the radially outer end 25 ( Figure 4 the maximum radius Re2 in

[0213] ), or passing through the center of gravity of the two adjacent blades, provided that in the case where the two adjacent blades have fixed pitch angles,

[0214] - Or:

[0215] --When the adjacent blade, such as 18b, has a fixed pitch angle, another axis that is radial to the longitudinal axis X and / or passes through the radially inner end 23 or through the radially outer end 25 or through the center of gravity of the adjacent blade.

[0216] Thus, the axes 180a, 180b, 19 can be interchangeable in the presented situation, and specifically can be interchangeable in the figures.

[0217] For example, Figure 20 Schematically shows the angle, or "azimuthal pitch" Δθ, between two consecutive stator blades, such as blades 18a, 18b having respective radial axes 180a, 180b i or Δθ j . This is the smaller angle of the two stator blades circumferentially around the axis X here between the axes 180a, 180b.

[0218] In the case where one of the stator blades is fixed (for example, for integrated constraints, such as if there is a lack of space for an integrated pitch change system 38 below the hub, or to reduce weight), the main axis of the blade can thus be defined by a line perpendicular to the longitudinal axis X that passes through the leading edge (LE) at the blade root 23 or through the center of gravity of the blade or at the blade tip 25 (maximum radius, Re2).

[0219] In this context, in order to present the required non-uniform distribution around the longitudinal axis X, at least some of the blades 18 of the downstream stator row 16 are arranged such that there is an azimuthal pitch Δθ or Δθ between at least two of the adjacent blades (such as 18a, 18b) of the downstream stator row 16 i , such that:

[0220] Δθ i ≠360° / V;

[0221] Δθ i ≥(360° / V)+1° or Δθ i ≤(360° / V)-1°,

[0222] where V defines the number of blades 18 in the downstream stator row 16;

[0223] and / or such that (between the blades 18 of the downstream stator row 16) there are at least two azimuthal pitches such that when i≠j and i, j = 1, 2,... and i, j ≤ V, Δθ i and Δθ j have different values.

[0224] It should be noted that i and j are indices. i ≠ j, i, j ≤ V, and i, j = 1, 2, …, so V defines the number of blades in the downstream stator row of the stator blades; that is, i and j are different and can take any integer values among 1, 2, 3, …, and (at most) V. i or j can only take the value i = V or j = V when all the azimuthal spacings are different (and thus non - homogeneous).

[0225] Therefore, a difference of at least 1° is necessary to bring about a significant effect related to non - homogeneous azimuthal spacing, preferably ≥ 3°, or even more preferably ≥ 5°.

[0226] Generally speaking, and by way of Figure 3 non - limiting examples in, each of the azimuthal spacings is defined by the circumferential distance E between two consecutive blades 18a, 18b, and this distance varies as a function of the radial and azimuthal positions of the blades 18 involved. Thus, when these axes are projected onto a plane perpendicular to the longitudinal axis X of the aero - propeller, these azimuthal spacings can be characterized by the aforementioned angles Δθ, Δθ i or Δθ j As a reminder, E = r * Δθ (or E i = r * Δθ i ), where Δθ is measured in radians and r is the radial distance measured in meters relative to the longitudinal axis X.

[0227] In a highly general embodiment, all the spacings between two adjacent blades of the upstream and downstream stator rows 16 in the azimuthal direction can be different, as shown in Figure 5 . This can be beneficial from an acoustic point of view to decorrelate the noise sources emitted by the stator 16 or to modify the directivity of the sound, meaning there are regions of maximum sound radiation.

[0228] Even if only one or only some of the azimuthal spacings between two adjacent blades in the downstream stator row 16 are different and most of the other azimuthal spacings between the two adjacent blades are the same under the conditions specified above, it should be noted that in a preferred embodiment, the proposed solution satisfies one or more of the following parameters or characteristics:

[0229] -a) At least two adjacent blades 18 have an azimuthal spacing (Δθ or Δθ i ), such that where V is the number of blades in the downstream stator 16. This allows ensuring that at least one blade of the stator 16 is not positioned homogeneously around the longitudinal axis X of the aero - propeller,

[0230] -b) max{Δθ i}-min{Δθ j} ≤ 120°, preferably ≤ 75°, or more preferably ≤ 50°, where i, j = 1, 2, …, i ≠ j and i, j ≤ V (if the number of blades of the stator 16 allows, i.e., if for example V ≥ 5); this ensures that the difference between the azimuthal spacings of any two adjacent blades 18 is limited. For example, in Figure 4 this criterion means that Δθ7 - Δθ8 ≤ 120°; the problem is to avoid too large an angular sector being completely free of stator blades, which could reduce the straightening of the flow downstream of the rotor 14 and thus cause a loss of thrust and efficiency. This could also pose problems regarding the weight distribution about the longitudinal axis X and the integration of the thruster on the aircraft.

[0231] -c) ||Δθ i -Δθ 邻近 || ≤ 120°, preferably ≤ 75° or more preferably ≤ 50° (if the number of stator blades allows, i.e., if for example V ≥ 5); this ensures that the difference between two adjacent (consecutive) azimuthal spacings is limited. This criterion can be applied in particular to avoid excessive differences in the spacings in an angular sector. The problem is to avoid too large an angular sector being completely free of the blades of the stator 16, which could likewise reduce the straightening of the flow downstream of the rotor 14 and thus cause a loss of thrust and efficiency. Problems regarding the weight distribution about the longitudinal axis X and / or integration may also arise.

[0232] The above value ranges must also be sufficient to allow the modification of the sound directivity (and thus the reduction of the noise towards the ground / fuselage) by increasing the azimuthal spacing between the blades of the stator 16 in the desired angular zone about the longitudinal axis X.

[0233] First, we have explained the influence of the forces and moments called 1P forces in the plane of the rotor 14, as well as the unsteady aerodynamics and the forces depending on the azimuthal position of the blades 18. For example, when an aircraft thruster is mounted below the wing, the descending blades are subject to increased forces compared to the ascending blades of the rotor. However, it should be noted that this may be the opposite in the case where the aircraft thruster is mounted at the rear. In fact, the airfoil can generate a downward flow (''downwash'', negative angle of attack) downstream of its trailing edge. In this case, the ascending blades may be subject to increased forces compared to the descending blades. Thus, we can find cases of negative or positive angles of attack upstream. The blades of the stator 16 located downstream of the upstream propeller 14 will also have variable loads depending on their azimuthal position. The following choices can then be made:

[0234] - When the stator vanes opposite the downward blades of the rotor 14 experience less load and less twist to be corrected, it is considered that fewer vanes in the stator 16 are required in this area: for example, fewer vanes in the stator 16 are required in the right lateral area (from 45° to 135°, where 0° is at the 12 o'clock position) when the rotor 14 rotates clockwise when viewed from the front (upstream), and / or

[0235] - When the vanes of the stator 16 axially opposite the upward blades of the rotor 14 experience more load and more twist to be corrected, it is considered that more stator vanes are required in this area: for example, more vanes in the stator 16 are required in the left lateral area (from -45° to -135°, where 0° is at the 12 o'clock position) when the rotor 14 rotates clockwise when viewed from the front (upstream).

[0236] Preferably, the number of different azimuthal spacings / angles Δθ, Δθ i , Δθ j varies between 2 and 6. In fact, increasing the number of different spacings can increase the number of stator vanes to be designed (several vane series / groups can be envisaged). The design of each stator vane adapted to its azimuthal position may be necessary. For example, it may be necessary to make local chord modifications to minimize the azimuthal deviation of the stator vane load and the vane ring robustness, which is defined by the C / E ratio at a given radial position. Other geometric parameters of the vanes can also vary: thickness, camber, sweep, forward lean, etc. This will allow the vane geometry to better adapt to the local flow properties (affected by the angle of attack) and / or better distribute the weight of the vanes of the stator 16 around the longitudinal axis of the aircraft propulsion to promote engine balance.

[0237] Several embodiments that are compatible with the above characteristics and may be preferred are presented below.

[0238] It has been recognized that the noise generated by the interaction between the wake of the upstream rotor 14 and the downstream stator 16 creates "dipole" acoustic radiation on the stator vanes. This means that the interaction noise from the stator 16 is not axisymmetric, but depends on the azimuthal position of the vanes of the stator 16 (or even depends on the pitch in the presence). Therefore, by using at least one of the above features a), b), or c) that can thus be fully or partially combined, it will be possible to optimize the azimuthal position of the vanes of the stator 16 in order to reduce the noise towards the ground and / or towards the cockpit (fuselage) and passengers and / or in any desired direction, in order to take action according to at least one of the following criteria: limiting the noise disturbance, promoting the propeller aerodynamics, improving the performance and integration of the aircraft propulsion installed on the aircraft.

[0239] If we assume that the pitch deviation between the blades of stator 16 measured at 0.75xRe2 (reference radius) is negligible (~0°, e.g. when in-flight cruising in isolation configuration), the blade pitch angle of stator 16 is ~90°, and most of the noise is generated at the outer radial end 25 of the blade, then the blades that emit the most noise towards the ground are positioned at an angle that varies between [2 o'clock - 4 o'clock] and [8 o'clock and 10 o'clock]. Therefore, one of the preferred embodiments contemplates increasing the spacing between the blades (or limiting the number of blades) in these angular sectors (near 3 o'clock and 9 o'clock) at the sides of stator 16, as for example Figure 6 shown in.

[0240] In another configuration which is possible and may be preferred according to at least one of the above characteristics a), b) or c), as Figures 7 to 10 shown by the example in,

[0241] - a larger number of blades in stator 16:

[0242] -- are placed in the upper part (between 10 o'clock and 2 o'clock) and / or

[0243] -- are placed towards the inside, closer to the fuselage 33 (when viewed from the front / upstream, between 2 o'clock and 4 o'clock if the fuselage is on the right of the thruster, or between 8 o'clock and 10 o'clock if the fuselage is on the left of the thruster when viewed from the front / upstream);

[0244] This will limit the noise emission towards the community. In fact, the radiation of the noise of the blades of stator 16 positioned towards the top and / or towards the inside towards the ground may be reduced due to their azimuthal position and the possible shielding effect produced by the attachment system 27, the airfoil 31 (if located below / above the wing) or the fuselage 33,

[0245] - or, if it is necessary to minimize the noise radiated towards the passenger cabin (fuselage 33), a larger number of blades of stator 16 are placed towards the outside, diametrically opposite to the fuselage 33 (when viewed from the front / upstream, between 8 o'clock and 10 o'clock if the fuselage is on the right of the thruster, or between 2 o'clock and 4 o'clock if the fuselage is on the left of the thruster when viewed from the front / upstream), so that the most stator 16 blades are furthest from the fuselage.

[0246] The final choice will depend on the noise reduction objectives of the thruster architecture, as Figures 5 to 9 schematically shown in.

[0247] In all these cases, increasing the number of blades of stator 16 in the upper part of the stator (i.e., thus reducing the azimuthal spacing: Δθ or Δθ i or Δθj ) This can be beneficial for acoustics and aerodynamics, but it should be noted that if there are struts, hangers or engine mounts 27, or any structure or system for attaching the thruster to the aircraft, this may pose difficulties when integrating below the housing 13.

[0248] Regarding acoustics, (in principle) it is the stator 16 blades in the upper part that radiate the least noise towards the ground. For aerodynamics, it is the stator 16 blades under the most load, because the upstream angle of attack (angle α / angle of attack of the aircraft, as Figure 2 in the example) is not completely filtered out by the upstream rotor 14.

[0249] Therefore, as Figures 10 - 11 shown in the example of Figure 11 , increasing the number of blades in the upper part of the stator 16 will allow for a better distribution of this load. In addition, increasing the number of stator 16 blades at the azimuthal positions close to the attachment system 27, the airfoil (wing) 31 or the fuselage 33 can also reduce the potential impact (pressure accumulation) of this obstacle (and / or strut and / or airfoil) towards the upstream rotor 14. In other words, increasing the number of stator 16 blades can thus help to "filter out" or reduce the pressure accumulation associated with the presence of said obstacle at the upstream rotor 14. Subsequently, it will be preferred to increase the number of stator 16 blades in the angular sector of the stator 16 where the distance between the trailing edge TE of the stator 16 blade and the leading edge 310 of the airfoil (or even the attachment system 27) is small. In this regard, recall the useful case where d1 < d2, as

[0250] shown in the solution illustrated as an example in Figure 12 . This is particularly relevant when d1 or d2 is less than 0.75*D, preferably 0.5*D or more preferably 0.3*D.

[0250] In yet another configuration, which is at least one of the above characteristics a), b) or c) and may be preferred, as Figure 12 shown in the example of

[0251] and mainly for aerodynamic reasons, the number of blades (or thus the azimuthal pitch) can be increased on the same side as the descending blades 18 of the upstream rotor 14. In fact, the blades of the stator 16 axially opposite to the descending blades of the upstream rotor 14 will have more load, and thus more twist to be corrected. Increasing the number of blades of the stator 16 on this side will allow for a better distribution of the load of these stator 16 blades, which can also contribute to reducing noise.

[0251] In yet another configuration, which is at least one of the above characteristics a), b) or c) and may be preferred, as Figure 13As shown in the example in [ ], it can be stipulated that the distribution of the blades 18 of the stator 16 on the right side (near 9 o'clock when viewed from the front / upstream of the thruster) and on the left side (near 3 o'clock when viewed from the front / upstream of the thruster) is symmetrical (for example, with respect to the symmetry axis passing through 12 o'clock and 6 o'clock).

[0252] This allows balancing the weight of the thruster and avoiding the residual moment on the longitudinal axis X associated with the non-uniform distribution of the blades of the stator 16.

[0253] In this case, the blades of the stator 16 located in symmetrical positions (θ and -θ) with respect to the vertical axis passing through 12 o'clock and 6 o'clock will advantageously exhibit the same geometric characteristics, specifically the thickness - see Figure 19 the example in [ ], the blade heights L2, L21 (or trimming),..., the blade pitch angles are not involved, as these can be variable in order to adapt the angle of attack of the blades to the local flow properties, which will allow a better distribution or smoothing of the loads on the blades in the azimuthal direction, specifically during the stages with a larger angle of attack (angle α).

[0254] Another key criterion that can be used to consider the above purposes relates to the azimuthal gap or spacing between the blades of the upstream rotor 14 and the blades of the downstream stator 16. This azimuthal spacing between the axes of the blades or the pitch change of the blades of the rotor 14 and the stator 16 is obtained through the following relationship:

[0255] ||θ r,n -θ s,m ||>1° or preferably ≥2°, where θ r,n and θ s,m correspond to the angular positions of the axes of the nth blade of the rotor 14 and the mth blade of the stator 16 at the moment when the axes of the blades or the pitch change of the blades of the rotor 14 and the stator 16 are respectively aligned when projected onto a plane perpendicular to the longitudinal axis X. Recall that n is a natural number and varies between 1 and B (the number of blades of the rotor 14), and m varies between 1 and V (the number of blades of the stator 16).

[0256] The utility of this criterion is to ensure that there is only one rotor 14 blade wake during the time of interaction with the stator 16 blades, which makes it possible to reduce the noise sources. Thus, the proposed embodiment allows to vary the periodicity of the interaction between the rotor 14 wake and the stator 16, which interaction causes interaction tone noise. The acoustic effect results in a reduction of the amplitude of the BPF (blade passing frequency - (noise level)), and thus in the appearance of the BPF compared to the broadband noise. Due to the reduction of the tone noise compared to the broadband noise, the noise during flight (effective perceived noise level EPNL according to acoustic regulations) can thus be lower. The total acoustic energy remains roughly the same, but is redistributed over the higher harmonics. Thus, if the number of blades of the rotor 14 and the stator 16 does not allow to satisfy this constraint with a uniform pitch (as in the case of Figure 14 the rotor 14 having twelve blades and the stator 16 having eight blades as shown), then the azimuthal position / pitch of certain blades of the rotor 14 and / or the stator 16 can be varied in a non-uniform manner. In this case, there may be variations in the chord and thickness of the blades in order to maintain a more or less constant average solidity C / E at each radial position. Here Figure 14 the upstream rotor 14 and the downstream stator 16 are viewed from the front upstream of the rotor, with the blades of the downstream stator partly hidden, and:

[0257] - The dashed line indicates the stacking axis or pitch change axis 19 of the blades of the rotor 14, and

[0258] - The dotted and dashed lines indicate the stacking axis or pitch change axis 1) of the blades of the stator 16.

[0259] At least in order to minimize the azimuthal deviation of the load on the stator blades and / or to correlate the noise sources between the blades, it is thus possible to define a law that bounds the fluctuations of the solidity in the azimuth for each given radius, such that it is possible to maintain and ensure the average azimuthal solidity for each radius:

[0260]

[0261] where represents the average solidity at a given radius r, П k (r) = C k (r) / E k (r) is the solidity between two adjacent (consecutive) blades of the stator 16 such as 18a, 18b at a given radius r, where V is the number of blades of the stator 16, is the angular pitch between these two blades at the same given radius r. This corresponds to the solidity C / E obtained by the weighted average with the azimuthal pitch; see Figure 8 for a non-limiting example.

[0262] The average azimuthal robustness can be specified to be less than 3 over the entire span and / or less than 1 at the radially outer end.

[0263] Thus, considering the angular pitch between two blades of the downstream stator 16 or between two adjacent axes, for example for as defined. The difference from Δθ i is such that:

[0264] - For and k = 1, 2,.., V; thus there can be two different indices having the same azimuthal spacing (i.e., the same value in radians or degrees). It should be noted that, for example, in Figure 8 and 9 there are several equal azimuthal spacings, all referring to Δθ1 in the same way;

[0265] - For Δθ1 and i = 1, 2,…, and i ≤ V: each index then corresponds to a different or a different value in radians / degrees of azimuthal spacing. This is the definition that is default used within the framework of the present disclosure.

[0266] Thus, it has been unexpectedly found that if П(r) satisfies the same criterion as C / E, the resulting performance increases by more than 3%, and the results support this.

[0267] Another aspect can be considered, namely integrating at least one blade 18 of the downstream stator 16 into the attachment system 27 (fork) in order to reduce the installation effects. This stator will then have a complementary structural function.

[0268] Figure 15 An example representing this situation is shown. This figure shows a propulsion assembly 24 for an aircraft. The propulsion assembly 24 includes a thruster 10 and an attachment system 27 for fixing the thruster 10 to the aircraft, such as a pylon 37. The attachment system 27 (pylon 26) is connected to one of the blades 18 of the downstream stator row 16 so as to form a single aerodynamic assembly. For this purpose, the attachment system 27 (pylon 37) can be connected to one of the blades 18 of the downstream stator 16 by way of an extension of its material. In other words, the attachment system 27 (pylon 26) can be integral with one of the blades 18 of the downstream stator 16.

[0269] Alternatively:

[0270] - The attachment system 27 (pylon 37) can be connected to one of the blades 18 of the downstream annular row 16 by one (or more) attachment members, or

[0271] - The part of the downstream stator 16 integrated into the attachment system may have a variable pitch. The attachment system 27 (pylon 37) also has an aerodynamic profile similar to that of the blades 18 of the downstream stator 16. The attachment system 27 (pylon 37) thus has the same effect on the air flow from the upstream annular row 14 as the blades 18 of the downstream stator 16. This arrangement allows further reduction of the noise emitted by the thruster 10.

[0272] If the stator blades are identical (i.e., belong to the same series of stator blades), the passage between the pylon 37 and the blades 18 of the downstream annular row 16 is reduced, which may generate shock waves and accelerate the flow, subsequently causing an increase in aerodynamic losses and thus a decrease in efficiency.

[0273] In the case where the pylon 37 is connected to one of the blades 18 of the downstream annular row 16 by one (or more) attachment members, the downstream annular row 16 may include blades 18 belonging to at least three series of stator blades.

[0274] When the pylon is connected to one of the blades 18 of the downstream annular row 16, it can be considered as part of a series of stator blades.

[0275] It can be provided that the pylon 37 and the two blades of the downstream annular row adjacent to the pylon 37 each belong to different series of stator blades. In other words, the two blades 18 located on both sides of the pylon 37 of the downstream annular row and the pylon 37 belong to three different series of stator blades.

[0276] This allows for a better adaptation of the aerodynamic operation of the stator blades and, specifically, avoids possible separation at the leading edge due to the change in the inflow at the leading edge of the stator blades caused by the presence of the pylon.

[0277] For example Figure 3 as shown in 15 or, another factor for noise reduction can be identified when

[0278] - The rotor 14 and the stator 16 are positioned towards the upstream end of the thruster ("tractor" configuration),

[0279] - The nacelle 40 has an opening 41 defining an air inlet, which can specifically be an inlet for the main air flow towards the turbomachine 10 (specifically towards the compressor 2), and

[0280] - On the nacelle 40, the opening 41 is axially located between the rotor 14 and the stator 16 and, more precisely and preferably, between the axes of the respective blades 18 of the rotor 14 and the stator 16.

[0281] This makes it possible to reduce the span of the downstream stator 16 (and thus to reduce Re2 - Ri2 by increasing Ri2), especially with respect to "clipping". The size of the blades and in particular their span creates radiated noise. This configuration will thus reduce the noise of the turbomachine.

[0282] On the nacelle, the air inlet 41 can be positioned along 360° (in a ring) or only along an angular sector. The air inlet 41 can have a protruding nozzle on the nacelle.

[0283] In this configuration, the turbomachine 10 (gas turbine / core engine) will operate conventionally so that the air entering the opening 41 will be accelerated and compressed by the compressor 2, then used in the combustion chamber and then passed into the turbine.

[0284] Independent of the nacelle, meaning that when the thruster includes a turbomachine 10 (gas turbine), it includes:

[0285] - a hub 12 provided with an upstream rotor row 14, and

[0286] - an engine casing 13 provided with a downstream stator row 16 downstream (AV) of the upstream rotor row 14, the air inlet bringing air to the compressor, such as 41, will advantageously be located:

[0287] -- downstream of the upstream rotor row 14 of the rotor blades, and

[0288] -- upstream of the downstream stator row 16 of the stator blades,

[0289] In other words, longitudinally along the thruster, between the rotor blades and the stator blades.

[0290] It should be understood that such a turbomachine can then successively include, from upstream to downstream along the longitudinal axis (X):

[0291] - at least one compressor 2,

[0292] - at least one combustion chamber 6,

[0293] - at least one turbine 4 driving the compressor, and

[0294] - the air inlet 41.

[0295] This has the following result: the radial dimension of the blades 18 of the downstream annular row 16 can even be further reduced so as to allow the eddies formed at the tips of the blades 18 of the upstream annular row 14 to escape, which reduces the efficiency of the turbomachine 10.

Claims

1. An aircraft propulsor (10) having a longitudinal axis (X) and comprising a casing (13), and an upstream rotor row (14) of ducted rotor blades and a downstream stator row (16) of stator blades that are ducted and extend around the casing (13), the stator blades being spaced apart from one another along the longitudinal axis (X), and having an azimuthal pitch (Δθ, Δθ i , Δθ j ) defined by an angle between respective axes (180a, 180b, 19) around the longitudinal axis (X), the respective axes being: -- the pitch angle adjustment axes of two adjacent blades, provided that when these axes are projected onto a plane perpendicular to the longitudinal axis (X) and in the case where the two adjacent blades have variable pitch angles, or -- axes that are radial to the longitudinal axis (X) and pass respectively through the radially inner ends (23) or the radially outer ends (25) of the two adjacent blades or through their centers of gravity, provided that the two adjacent blades have fixed pitch angles, or -- for one of the corresponding axes when one of the two adjacent blades has a variable pitch angle, it is the pitch angle adjustment axis of said one blade, and when the adjacent blade has a fixed pitch angle, the other of the corresponding axes is radial to the longitudinal axis (X) and / or passes through the radially inner end (23) or the radially outer end (25) or through the center of gravity of the adjacent blade, and around the longitudinal axis (X), the angular position at 12 o'clock is defined as being vertically upward relative to the longitudinal axis (X), and the angular position at 6 o'clock is defined as being vertically downward relative to the longitudinal axis (X), This assembly is characterized in that at least some of the blades of the downstream stator row (16) of the stator blades have a non-uniform distribution around the longitudinal axis (X), such that: - At least two of the adjacent blades of the downstream stator row (16) of the stator blades have an azimuthal pitch Δθ therebetween i , such that Δθ i ≠360° / V; Δθ i ≥(360° / V)+1° or Δθ i ≤(360° / V)-1°, and / or - There are at least two azimuthal spacings such that when i ≠ j, the values of Δθ i and Δθ j are different, where i, j = 1, 2, … and i, j ≤ V where V defines the number of blades in the downstream stator row (16) of the stator blades, where there are at least three series of stator blades, preferably at least five series of stator blades, and each series of stator blades includes one or more stator blades, and the stator blades in the same series including several stator blades have the same set of geometric characteristics, where at least one geometric characteristic in said set of geometric characteristics is different from the same geometric characteristic of the stator blades in another series of stator blades.

2. The aero-propulsor (10) according to claim 1, which is of the USF type and includes a single annular row (14) of ducted rotor blades, and this annular row is the upstream rotor row of the rotor blades.

3. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, Δθ i ≥ (360° / V) + 3° or Δθ i ≤ (360° / V) - 3°, or preferably Δθ i ≥ (360° / V) + 5° or Δθ i ≤ (360° / V) - 5°.

4. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, All the azimuthal angular spacings between two adjacent blades in a series of blades in the downstream stator row (16) of the stator blades are different from each other.

5. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, For at least one of the upstream rotor row (14) of the rotor blades and the downstream stator row (16) of the stator blades, there is a C / E ratio of the chord C and the azimuthal angular spacing E around the longitudinal axis (X) between two consecutive blades of the downstream stator (16), such that C / E is less than 3 over the entire span.

6. The aircraft propeller (10) according to claim 5, characterized in that, The C / E ratio is less than 1 at the radially outer ends (25) of two blades of the same upstream rotor row and / or downstream stator row, and the blades are circumferentially or azimuthally consecutive or adjacent.

7. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, For at least one of the upstream rotor row (14) of the rotor blades and the downstream stator row (16) of the stator blades, the mean azimuthal stiffness is less than 3 over the entire span and / or less than 1 at the radially outer end, where represents the mean stiffness at a given radius r, П k (r) = C k (r) / E k (r) represents the stiffness between two adjacent stator blades at a given radius r, V is the number of stator blades, and Δθ k is the angular pitch between these two blades at the same given radius r.

8. The aircraft propulsion unit (10) according to any one of the preceding claims, characterized in that, The upstream rotor row (14) of the rotor blades and the downstream stator row (16) of the stator blades have different numbers of blades (18).

9. The aero-propulsor (10) according to any one of the preceding claims, characterized in that: - At least some of the blades (18) of at least one of the upstream rotor row of the rotor blades and the downstream stator row (16) of the stator blades are linked to a variable pitch system (38) that allows for changing its pitch angle by rotation about its respective axis (19, 180a, 180b), each said respective axis being radial to the longitudinal axis (X).

10. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, All the blades (18) of the downstream stator row (16) of the stator blades have a homogeneous distribution about the longitudinal axis (X), except at a single angular sector.

11. The aircraft propulsor (10) according to claim 10, characterized in that, All the blades (18) of the downstream stator row have a homogeneous distribution about the longitudinal axis (X) such that for any azimuthal angular pitch between two (circumferentially) adjacent / consecutive blades involved, Δθ ≤ 360° / V, except at a single angular sector where the azimuthal angular pitch is different and limited between 15° and 75° or preferably between 25° and 60°.

12. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, The number B of blades of the upstream rotor row (14) of the rotor blades is greater than the number V of blades (18) of the downstream stator row (16) of the stator blades, and preferably B ≥ V + 1, or more preferably B ≥ V + 2.

13. The aircraft propulsion unit (10) according to any one of the preceding claims, characterized in that, The radially outer ends of the blades (18) of the downstream stator row (16) of the stator blades are inscribed within an enclosing envelope (22), for which the projection onto a plane (IV-IV) perpendicular to the longitudinal axis (X) defines a circle.

14. An aircraft propulsor (10) according to any one of the preceding claims, characterized in that, The radially outer end of each blade (18) of the upstream rotor row (14) of the rotor blades is inscribed within a first circle (20), and the radially outer end of each blade (18) of the downstream stator row (16) of the stator blades is inscribed within a second circle (22), the radius (Re2) of the second circle (22) being less than the radius (Re1) of the first circle (20).

15. The aviation thruster (10) according to claim 13 or 14, characterized in that, The center of the circle within which the radially outer end of each blade (18) of the downstream stator row (16) of the stator blades is inscribed is offset relative to the longitudinal axis (X).

16. An aeronautical propulsor (10) according to any one of the preceding claims, characterized in that: - each blade of the downstream stator row (16) of the stator blades has a height (L2, L21) between a radially inner end (23) and a radially outer end (25), and - the respective heights (L2, L21) of at least two blades (18) of the downstream stator row (16) are different.

17. An aeronautical propulsor (10) according to any one of the preceding claims, characterized in that, max{Δθ i}-min{Δθ j} ≤ 120°, preferably ≤ 75°, or more preferably ≤ 50°, i, j = 1, 2, …, where i ≠ j and i, j ≤ V.

18. An aeronautical propulsor (10) according to any one of the preceding claims, characterized in that, ||Δθ i -Δθ 邻近 || ≤ 120°, preferably ≤ 75°, or more preferably ≤ 50°, i = 1, 2, …, where i ≤ V.

19. The aircraft propulsion unit (10) according to any one of claims 16 and 17, characterized in that, the number of blades (18) of the downstream stator row (16) of the stator blades is greater than or equal to 5, V ≥ 5.

20. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, The number of different azimuthal spacings (Δθ, Δθ i , Δθ j ) across all the blades (18) of the downstream stator row (16) of the stator blades is between 2 and 6.

21. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, The ratio of the following two: - the distance (S) along the longitudinal axis (X) between two central planes respectively perpendicular to the longitudinal axis of the upstream rotor row (14) of the rotor blades and the downstream stator row (16) of the stator blades, and - the maximum diameter (D) of the aero - engine (10) at the outer radial ends of the blades (18) of the upstream rotor row (14) of the rotor blades or of the downstream stator row (16) of the stator blades, is between 0.01 and 0.8, or preferably between 0.15 and 0.

35.

22. The aero - propulsion device (10) according to any one of the preceding claims, characterized in that, The azimuthal pitch between the axes (180a, 180b, 19) of the blades of the rotor (14) and the stator (16) is obtained by the following relationship: at the moment when the blade axes (180a, 180b, 19) of the rotor (14) blades and the stator (16) blades are aligned when projected onto a plane perpendicular to the longitudinal axis (X), ||θ r,n -θ s,m ||> 1° or preferably ≥ 2°.

23. The aircraft propulsion unit (10) according to any one of the preceding claims, characterized in that The number of blades (18) of the upstream rotor row (14) and of the downstream stator row (16) is respectively 12 and 10, or 14 and 12, or 14 and 11.

24. The aero - engine (10) according to any one of the preceding claims, which, along the longitudinal axis (X), successively comprises, from upstream to downstream: - at least one compressor (2), - at least one combustion chamber (4), - at least one turbine (6) driving the compressor, and - an air inlet (41) to the compressor (2), the air inlet (41) being located downstream of the upstream rotor row (14) of the rotor blades and upstream of the downstream stator row (16) of the stator blades.

25. The aircraft propulsion unit (10) according to any one of the preceding claims, characterized in that, The set of geometric characteristics of the stator blades includes chord, thickness and height.

26. The aviation thruster (10) according to the previous claim, characterized in that, The set of geometric characteristics of the stator blades further includes at least one of camber, sweep and forward sweep.

27. The aircraft propulsion unit (10) according to any one of the preceding claims, characterized in that, The stator blade row comprises at least three adjacent stator blades each belonging to a different series of stator blades.

28. The aircraft propulsor (10) according to any one of the preceding claims, characterized in that, The attachment system (27) is connected to one of the blades (18) of the downstream annular row by one or more attachment members.

29. The aviation propeller (10) according to the preceding claim, characterized in that, The attachment system (27) is a pylon (37).

30. The aviation thruster (10) according to the previous claim, characterized in that, The pylon (37) and two blades (18) of the downstream annular row adjacent to the pylon (37) each belong to a different series of stator blades.

31. An aircraft having a longitudinal axis (X1) and comprising an aero - engine (10) according to any one of the preceding claims, a fuselage (33) and a wing (31) fixing the aero - engine (10), wherein the absolute value of the angle (||β||) between the longitudinal axis (X) of the aero - engine and the longitudinal axis (X1) of the aircraft varies between 0.5° and 30°, preferably between 2° and 20° or more preferably between 3° and 10°.

32. An aircraft having a longitudinal axis (X1) and comprising an aero - engine (10) according to any one of claims 1 to 24, a fuselage (33) and a wing or airfoil (31) fixing the aero - engine (10), wherein d1≠d2 and d1 or d2 is less than 0.75*D, preferably less than 0.5*D, or more preferably less than 0.3*D, where: - d1 is defined as the axial distance (X) between the trailing edge (TE) of the stator blade at the free end (25) and the leading edge (LE) of the airfoil (or wing), which is the case for the stator blade azimuthally closest to the leading edge of the airfoil or wing and included in the angular sector located between 12 o'clock and 6 o'clock and including 9 o'clock, and -d2 is defined as the axial distance (X) between the trailing edge (TE) of the stator vane at the free end and the leading edge of the airfoil or wing, this being the case for the stator vane azimuthally closest to the leading edge of the airfoil or wing and included in the angular sector located between 12 o'clock and 6 o'clock and including 3 o'clock.

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