A STAGE OF VARIABLE PITCHING BLADES FOR A TURBOMACHINE

Abstract

The present invention relates to a stage (1) of variable-pitch vanes (2) comprising: - a row of vanes (2), - a casing having an axis of revolution (X), and - a ring (5) for controlling the pitch of the vanes (2), comprising: - a body (50) which extends around the casing, - first members for connecting the body (50) to cylindrical pivots (22) of said vanes (2), - a second member (70) for connecting the body (50) to an actuator (7) carried by the casing, and - third members (8) for guiding the body (50) around and at a distance from the casing, in which said third members (8) comprise levers (80) which each comprise a first end (80a) articulated on the casing around a radial axis (R80a) relative to the axis (X), and a second end (80b) connected to the body (50). Figure for abstract: Figure 7

Description

Title of the invention: A STAGE OF VARIABLE PITCHING BLADES FOR A TURBOMACHINE Technical field

[0001] The present invention relates to a stage of variable-pitch blades for a turbomachine, in particular an aircraft. Technical background

[0002] A stage of pitch vanes for a turbomachine, in particular an aircraft one, comprises an annular row of variable-pitch stator vanes (also called VSV - English acronym for "Variable Stator Varies") which are carried by an external annular casing, generally of a compressor of the turbomachine. Each vane comprises a blade which is connected at its radially external end by a plate with a substantially circular contour to a radial cylindrical pivot. This cylindrical pivot defines the axis of rotation of the vane and is guided in rotation in a corresponding orifice of the external casing. The radially internal end of the blade of each vane generally comprises a second cylindrical pivot. This second cylindrical pivot extends along the axis of rotation of the vane and is guided in rotation in an orifice of an internal casing of the compressor.

[0003] The radially outer end of the cylindrical pivot of each blade is connected by a connecting rod (or a lever) to a control ring moved in rotation around the annular casing by actuators (such as cylinder actuating means or the like). The rotation of the control ring is transmitted by the connecting rods to the cylindrical pivots of the blades and causes them to rotate around their axes.

[0004] The angular setting of the VSV blades in a turbomachine is intended to adapt the geometry of the compressor to its operating point and in particular to optimize the efficiency and the pumping margin of this turbomachine and to reduce its fuel consumption in the different flight configurations.

[0005] Each of these blades is movable in rotation about its axis between a first “open” or “fully open” position in which each blade extends substantially parallel to the longitudinal axis of the turbomachine, and a second “closed” or “quasi-closed” position in which the blades are inclined relative to the axis of the turbomachine and thus reduce the air passage section through the blade stage.

[0006] The control ring must be centered and guided in rotation about its axis of rotation. In the current technique, the annular casing comprises tracks on which the internal periphery of the control ring can cooperate by friction. In particular, the control ring includes guide members, such as pads, on the tracks of the annular casing. These pads serve, on the one hand, to ensure the concentricity of the control ring around the annular casing by adjusting the clearance between the pads and the tracks, and on the other hand, to limit the deformation of the control panel caused by the aerodynamic forces on the variable-pitch blades exerted on the kinematics during operation. It is indeed important to have radial clearances, called cold clearances, between the pads and the tracks, in order to allow thermal expansion of the annular casing and allow good timing kinematics in all engine configurations. These radial clearances between the pads and the tracks may depend on the initial setting when mounting the control ring and the thermal conditions.As a result, the quality of the guidance of the control ring by these guide members may depend on the radial clearance between these pads and the tracks of the annular casing. Thus, this radial clearance (or otherwise known as the mounting clearance) between the pads and the tracks may be designed to optimize the clearance under the pads under thermal operating conditions (i.e. when hot), and to prevent the control ring from deforming, in particular in compression on the annular casing, for example during acceleration phases of the turbomachine in operation.

[0007] It is known in the prior art to use mounting shims to adjust the radial clearance between the pads and the tracks. These shims have the thickness necessary to achieve the ideal theoretical radial clearance JT ([Fig. 1a]) to enable the two aforementioned functions, namely optimizing the clearance under the pads 81 and preventing deformation of the control ring 5. However, if the radial clearance, in particular adjusted by the shims, is too large or otherwise called “open” radial clearance Jo ([Fig. 1b]) for example compared to the theoretical radial clearance JT, the guidance of the control ring 5 by the pads 81 may be less well ensured. This may lead to dispersion of the blades between them. If the radial clearance, in particular adjusted by the shims, is too small or otherwise called “closed” radial clearance ([Fig. 1c]), there may be a risk of tightening the control ring 5 on the annular casing 4.This may lead to kinematic blocking of the variable-pitch blade stage 1 and / or deterioration and breakage of the control ring 5. Furthermore, since the pads 81 are also designed to rub the tracks 46 of the annular casing 4, the friction surface of these tracks 46 may wear and thus increase the radial clearance under the pads 81 and drift towards the open radial clearance profile Jo of [Fig.lb].

[0008] Another disadvantage of the pads as guide members is that they generally only allow the deformation of the control ring to be restricted in one direction (e.g. in a radial direction inwards, i.e. towards the annular casing). Thus, these pads may not be sufficient to restrict radial outward and / or axial deformations.

[0009] The number and position of the actuators can also impact the deformation of the control ring in operation. For example, the control ring 5 of [Fig. 2a] comprising a single actuator 7 can concentrate the forces at a single point and thus de-symmetrize the control ring 5. Figures 2b and 2c illustrate an example of two control rings 5 ​​comprising, respectively, first 50a and second 50b annular bodies with a connecting member 70 for a single actuator 7, in which the first annular body 50a has deformations (such as ovalization) relative to the second annular body 50b which is intact.

[0010] For example, [Fig. 3a] represents a modeling of the timing errors due to the deformation (such as an ovalization of the control ring) measured in degrees (°) on the ordinate with respect to first (namely the opening position represented by circles) and second (namely the closing position represented by triangles) positions of the blades on the abscissa, in the case of the control ring 5 with a single actuator 7. In this [Fig. 3a], the single actuator 7 represented by an arrow and being positioned at the level of the first and at approximately sixty-fifth blades of the stage, a dysmetry of the blades between their first and second blade positions is observed. This configuration increases the timing errors of the blades and can thus increase the deformation of the control ring.

[0011] Figures 3b and 3b also represent models of the timing errors due to deformation in the case of a control ring comprising two actuators 7 spaced apart from each other at, respectively, 180° and 135° (or 225°). In [Fig. 3b] with an equidistant distribution at 180° of the two actuators 7 which are represented by arrows at the level of the first and approximately the sixty-fifth blades and approximately the thirty-fourth blade, the first and second blade positions are substantially symmetrical and the timing errors are smaller than those of [Fig. 3a]. However, the integration of two actuators can cumbersome, weigh down and complicate the operation and assembly of the variable-pitch blade stage. In [Fig.3c] with a distribution of the two actuators at 135° (or 225°) which are represented by arrows at the level of the first and approximately the sixty-fifth blades and approximately the twenty-eighth blade, the first and second blade positions are asymmetrical and can further deform the control ring.

[0012] In this context, it is interesting to overcome the drawbacks of the prior art, by proposing a reliable solution and limiting the deformation of a control ring in a variable-pitch blade stage of a turbomachine, while reinforcing the centering and guidance of the control ring in operation. Presentation of the invention

[0013] The present invention provides a simple, effective and economical solution to at least some of the aforementioned problems.

[0014] The invention proposes a variable-pitch blade stage for an aircraft turbomachine, comprising: - an annular row of variable-pitch blades, each blade comprising a blade comprising a cylindrical pivot at its radially outer end, this pivot defining a pitch axis C of the blade, - an annular casing having an axis of revolution X and comprising housings for mounting said cylindrical pivots, and - a blade timing control ring, the ring comprising: - an annular body which extends around the annular casing, - first connecting members of the annular body to the cylindrical pivots of said blades, - a second member connecting the annular body to at least one actuator carried by the annular casing, and - third guide members of the annular body around and at a distance from the annular casing.

[0015] According to the invention, said third guide members comprise levers which each have a first end articulated on the annular casing around a radial axis R80a relative to the axis of revolution X, and a second end connected to the annular body.

[0016] The variable-pitch blade stage according to the invention makes it possible to replace the support and guide pads of the prior art with levers to limit the deformation of the control ring during operation. In particular, the levers make it possible to connect the control ring to the annular casing and to guide this control ring while minimizing (or even eliminating) the radial deformation(s) (inwards and / or outwards relative to the X axis) and / or axial deformation(s). Indeed, the levers form a strong connection (otherwise called rigid or stiff) between the control ring and the annular casing to best contain and control the different deformation modes regardless of the number and position of the actuator(s) in the blade stage.

[0017] In addition, the levers make it possible to optimize the blade setting precision, in particular by reducing the impact of thermal expansion on the blade setting precision thanks to the levers' ability to allow and transmit movements (such as pivoting, translation, rotation, etc.) between the control ring and the annular casing. Whereas it is the radial clearance between the pads and the annular casing which is the main contributor to the blade offset in the prior art.

[0018] Furthermore, the solution of the invention also allows a saving in assembly time (for example elimination of assembly shim and adjustment of radial clearance between the levers and the annular casing) and maintenance (for example removal of wear parts and re-adjustment of the radial clearance of the pads) of the control ring and consequently of the blade stage.

[0019] The invention therefore has the advantage of being based on a simple design, offering very high reliability, and with little penalty in terms of cost, size and environmental impact, particularly for the aircraft turbomachine.

[0020] The term "articulated" means that the first end can be pivotally mounted on the annular casing around the radial axis R80a. The first end of the levers can thus be assembled on the casing to allow them to rotate relative to each other around the radial axis R80a.

[0021] The variable-pitch blade stage according to the invention may comprise one or more of the following characteristics, taken in isolation from one another or in comparison with one another:

[0022] - the second end is articulated by a ball joint to a first fixing pin engaged in an orifice of the annular body;

[0023] - the first fixing pin has a radial orientation with respect to the axis of revolution X ;

[0024] - the first end comprises a mounting hole for a second fixing pin engaged in an orifice of the annular casing;

[0025] - the first end has a thickness or radial dimension greater than that from the second end;

[0026] - said third guide members are approximately two to twenty in number, preferably this number is between three and twelve;

[0027] - the levers are distributed regularly around the axis of revolution X;

[0028] - each of said third guide members has an external diameter greater than an external diameter of the first connecting members;

[0029] - the second end is located at a predetermined radial distance from the annular body;

[0030] - the radial axis of the first end is inclined at a predetermined angle a relative to an axis perpendicular to the axis of revolution X;

[0031] — the angle a is for example between 0° and 50°;

[0032] - the angle a is for example between 5° and 50°;

[0033] — at least one internal surface (or internal periphery) of the first end and / or the second end of the levers is or are covered with an anti-wear coating, for example a plastic lubricant;

[0034] — said anti-wear coating comprises at least one chemical element, for example molybdenum or graphite;

[0035] — the levers are each made of aluminum, steel or titanium;

[0036] — the variable-pitch vane stage is a rectifier stage of a low-pressure compressor pressure or high pressure compressor of the turbomachine.

[0037] The present invention also relates to a turbomachine, in particular for an aircraft, comprising at least one stage of variable-pitch blades according to one of the particular features of the invention.

[0038] The turbomachine can be a turbojet or a turboprop. Description of figures

[0039] The invention will be better understood and other details, characteristics and advantages of the invention will appear more clearly on reading the following description given by way of non-limiting example and with reference to the appended drawings in which:

[0040] [Fig. la] [Fig. la] very schematically represents a theoretical radial clearance called ideal between the pads of a control ring and the tracks of an annular casing of a stage of variable-pitch blades;

[0041] [Fig. 1b] [Fig. 1b] very schematically represents a so-called open radial clearance between the pads and the tracks of the variable-pitch blade stage;

[0042] [Fig. 1c] [Fig. 1c] very schematically represents a so-called closed radial clearance between the pads and the tracks of the variable-pitch blade stage;

[0043] [Fig.2a] [Fig.2a] is a schematic top view of a control ring with a single cylinder actuator according to the prior art;

[0044] [Fig.2b] [Fig.2b] is a schematic view of a first deformed control ring relative to an intact second control ring with a single actuator according to the prior art;

[0045] [Fig.2c] [Fig.2c] is a schematic perspective and side view of the first and second control rings of [Fig.2b];

[0046] [Fig.3a] [Fig.3a] is a schematic view of the control ring with a single actuator of Figures 2a to 2c and a modeling of the blade timing errors on this ring;

[0047] [Fig.3b] [Fig.3b] is a schematic view of a control ring with two actuators distributed at an equidistant distance of 180° according to the prior art and a modeling of the blade timing errors on this ring;

[0048] [Fig.3c] [Fig.3c] is a schematic view of a control ring with two actuators distributed at a distance of 135° according to the prior art and a modeling of the blade timing errors on this ring;

[0049] [Fig.4] [Fig.4] is a partial schematic half-view in axial section of a stage of variable-pitch blades of a compressor module of a turbomachine according to the invention;

[0050] [Fig.5] [Fig.5] is a schematic perspective side view of a casing annular and control rings of the blade stage of [Fig.4];

[0051] [Fig.6] [Fig.6] is a schematic perspective view of the control ring of [Fig.4] or [Fig.5], comprising levers according to one embodiment of the invention;

[0052] [Fig.7] [Fig.7] is a schematic perspective and partial view of the floor of blades of [Fig.6];

[0053] [Fig.8] [Fig.8] is another schematic perspective and partial view of the floor of blades of [Fig.6];

[0054] [Fig.9] [Fig.9] is a schematic perspective and partial view of the floor of blades of [Fig.4] or [Fig.5], comprising the lever of [Fig.6] assembled on the control ring and an annular casing;

[0055] [Fig. 10] [Fig. 10] is a schematic view in axial and partial section of the blade stage of [Fig.9];

[0056] [Fig. 11] [Fig. 11] is another schematic view in axial and partial section of the blade stage of [Fig.9] or [Fig. 10].

[0057] Elements having the same functions in different implementations have the same references in the figures. Detailed description

[0058] Generally speaking, in the description below, the terms “longitudinal” and “axial” describe the orientation of structural elements extending in the direction of a longitudinal axis. This axis may be confused with an axis of rotation of a rotor of a turbomachine. The term “radial” describes an orientation of structural elements extending in a direction perpendicular to the longitudinal axis. The terms “inner” and “outer”, and “internal” and “external” are used with reference to a positioning relative to the longitudinal axis. Thus, a structural element extending along the axis has an inner face facing the axis and an outer surface, opposite its inner surface. Similarly, the terms “upstream” and “downstream” are defined relative to the direction of air circulation in the turbomachine.

[0059] Figures 1a to 1c, 2a to 2c and 3a to 3c have been described above in the technical background of the present application and represent at least a part of a stage of variable-pitch vanes 1 according to the prior art comprising pads 81 as guide members 8 of a control ring 5 relative to an annular casing 4.

[0060] The present invention applies generally and non-limitingly to a turbomachine 10, in particular an aircraft turbomachine. The turbomachine 10 may be a turbojet or a turboprop.

[0061] Figures 4 and 5 schematically represent, in a non-limiting manner, at least one stage 1 of variable-pitch blades 2 of a high-pressure compressor of the turbomachine 10. According to another variant, a low-pressure compressor of the turbomachine 10 may comprise one or more stages 1.

[0062] Floor 1 includes: - an annular row of variable-pitch blades 2, - an annular casing 4 having an axis of revolution X, and - a control ring 5 for the pitch of the blades 2.

[0063] Each stage 1 may also comprise an annular row of moving blades 3 carried by a rotor (not shown) of the turbomachine.

[0064] Each blade 2 comprises a blade 20 having a cylindrical pivot 22 at its radially outer end. The cylindrical pivot 22 defines a setting axis C of the blade 2.

[0065] The angular orientation of the blades 2 may be adjustable to optimize the gas flow in the compressor. In particular, the blades 2 may be movable in rotation about their axes C between a closed or quasi-closed position and an open or fully open position.

[0066] In the closed position, the blades 20 of the vanes are inclined relative to a longitudinal axis of the turbomachine and define between them a minimum air passage section in the duct. The longitudinal axis of the turbomachine corresponds substantially to the X axis of the annular casing 4. The vanes 2 are brought into this closed position when the turbomachine is at low speed or at idle, the air flow rate flowing into the compressor then having a minimum value. In the open position, the blades 20 of the vanes extend substantially parallel to the axis of the turbomachine so that the air passage section between the blades is maximum. The vanes 2 are brought into this open position when the turbomachine is at full throttle, the air flow rate flowing into the compressor then having a maximum value.

[0067] The blade 20 and the cylindrical pivot 22 may be connected by a disc or “plate” 28 extending perpendicular to the axis C of the blade in a first corresponding housing 44 of the annular casing 4. A radially internal surface 21 of the disc 28 may be aligned with an internal wall 42 of the annular casing 4 so as not to oppose the gas flow.

[0068] The cylindrical pivot 22 of each blade 2 can extend inside a second housing 40 (such as a radial cylindrical chimney) of the annular casing 4 and its radially external end can be connected to the control ring 5 by a connecting rod 24. This control ring 5 extends around the annular casing 4 in Figures 4 and 5.

[0069] The annular casing 4 may comprise projecting tracks 46 at its external periphery. allowing the centering and guiding of the control ring 5, which are schematically represented by dotted lines in [Fig. 4]. These tracks 46 may have a cylindrical shape, as visible in Figures 9 to 11. Each track 46 may have a first thickness E46 (or in other words a radial dimension) measured for example along a plane perpendicular to the axis X. By way of example, the first thickness E46 may be at most approximately 100 mm. This first thickness E46 may be variable depending on the dimensions of the stage 1 of variable-pitch blades (such as the size and / or shape of the control ring 5, the annular casing 4 and / or the variable-pitch blades 2). Each track 46 may be located on a dimension (such as a radius or the first thickness E46) smaller than that of the cylindrical pivot 22 of the corresponding blade 2. Each control ring 5 can surround its track(s) 46.

[0070] With reference to Figures 4 to 11, the control ring 5 comprises: - an annular body 50, - first members 26 for connection to the cylindrical pivots 22, - a second member 70 for connection to at least one actuator 7, and - third members 8 for guiding the annular body 50 and at a distance from the casing 4.

[0071] The annular body 50 can extend around a longitudinal axis A corresponding substantially to the axis X of the annular casing 4.

[0072] The first connecting members 26 may be located on a circumference, in particular around the axis A, of the annular body 50. Each of the first connecting members 26 may have a first external diameter D26 measured in particular along a plane perpendicular to the axis A.

[0073] The second member 70 for connection to the actuator 7 may comprise a yoke carrying an axis on which one end of a piston rod of the actuator 7 is articulated, when the latter is a jack for example.

[0074] The actuator 7 may comprise a (single) cylinder or several cylinders, for example hydraulic. This actuator 7 may comprise an element fixed to the annular casing 4 and a piston rod connected to the second member 70.

[0075] The actuator 7 makes it possible to rotate the control ring 5 in one direction or the other around the axis X to drive the blades 2 of an annular row in rotation around their axes C.

[0076] The annular body 50 may comprise first orifices 52 in which the first connecting members 26 (for example cylindrical pins) carried by the connecting rods 24 may be housed. Each connecting member 26 may generally be centered and guided in rotation in the first orifice 52 by at least one sleeve bearing mounted in this first orifice 52.

[0077] The body 50 may also comprise second orifices 54 in which the third guide members 8 can be housed. These second orifices 54 can be distributed regularly around the axis A. These second orifices 54 can be located between two first orifices 52.

[0078] One of the particularities of the invention is that the third guide members 8 comprise levers 80 (or otherwise known as connecting rods). Figures 6 to 11 illustrate a non-limiting example of the lever 80 according to the invention. As mentioned above, the levers 80 make it possible to articulate the control ring 5 relative to the annular casing 4 and vice versa to prevent deformation of the control ring 5 during operation.

[0079] Each of the levers 80 comprises a first end 80a articulated on the casing 4 around a first radial axis R80a relative to the axis of revolution X, and a second end 80b connected to the annular body 50. With reference to the example of FIGS. 7 to 11, the first end 80a is opposite the second end 80b and the first 80a and second 80b ends can be single-piece (i.e. made in one piece).

[0080] The first radial axis R80a may be inclined at a predetermined angle α relative to an axis perpendicular to the X axis. The inclination of the first radial axis R80a makes it possible in particular to further restrict the radial deviation of the control ring 5. This angle α may be between 0° and 50°. Advantageously, the angle α may be between 5° and 50°. In the example of FIGS. 9 to 11, the angle α may be approximately 10°. In a variant not illustrated in the figures, the angle α may be zero, in which case the first radial axis R80a may be perpendicular to the X axis.

[0081] The first end 80a may have a second thickness E80a and the second end 80b may have a third thickness E80b. These second E80a and third E80b thicknesses (or otherwise called radial dimensions) are in particular measured along a plane perpendicular to the axis A or the axis X. In the example of FIGS. 7 to 11, the second thickness E80a may be greater than the third thickness E80b.

[0082] In the example of Figures 7 to 11, the second end 80b may be located at a predetermined radial distance d80b relative to the annular body 50. This distance d80b may be less than 1 mm. This distance d80b may be zero when the second end 80b is mounted directly on the annular body 50 and an internal periphery of this second end 80b is in abutment and in contact with the annular body 50 (not illustrated in the figures).

[0083] When the second end 80b is mounted directly on the annular body 50, this second end 80b may comprise at its internal periphery (relative to the axis X), at least one second internal surface 800a covered with an anti-wear coating (not illustrated in the figures). This anti-wear coating may be a lubricant made of plastic material. The anti-wear coating may comprise at least one chemical element, for example molybdenum or graphite, in particular to restrict the wear friction between the annular body 50 and the second end 80b of the levers 80 and thus extend the service life of these levers 80.

[0084] The second end 80b can be articulated, in particular along a second radial axis Rsob relative to the axis X (or the axis A), on the annular body 50. The second radial axis R 80b can be substantially perpendicular to the axes A and X. In particular, the lever 80 can comprise a ball joint 82 and a first fixing pin 84. This ball joint 82 can be housed in a third mounting orifice 81 of the second end 80b ([Fig.l 1]). The second end 80b can thus be articulated by the ball joint 82 to the first fixing pin 84 engaged in the second orifice 54 of the annular body 50. The ball joint 82 and the first fixing pin 84 can be monobloc (i.e. made in one piece). The second 52 and third 81 holes can be radially aligned with each other ([Fig. 11]).

[0085] The first pin 84 may have a radial orientation with respect to the axis of revolution X. In particular, the first pin 84 may be located at the level of the second radial axis R80b of each of the second ends 80b.

[0086] The first end 80a may comprise a fourth orifice 85 for mounting a second fixing pin 86 engaged in a fifth orifice 460 of the annular casing 4. More particularly, the fifth orifice 460 may be provided in the track 46 of the annular casing 4. The fourth 85 and fifth 460 orifices may be aligned radially with each other ([Fig. 11]).

[0087] The first end 80a may comprise at its internal periphery (relative to the axis X), at least one first internal surface 800b covered with an anti-wear coating. This anti-wear coating may be a lubricant made of plastic material. The anti-wear coating may comprise at least one chemical element, for example molybdenum or graphite, in particular to restrict wear friction between the tracks 46 and the first end 80a of the levers 80 and extend the service life of these levers 80.

[0088] The third guide members 8 can be distributed regularly around the axis X. In particular, the levers 80 can be distributed regularly around the axis X.

[0089] The third guide members 8 can thus be located on a circumference, in particular around the axis A, of the annular body 50. Thus, each of the third guide members 8 (in particular the levers 80) can have a second external diameter D8 measured in particular along a plane perpendicular to the axis A. This second external diameter D8 can be greater than the first external diameter D26 of the first connecting members 26 (Figures 7 to 11). According to another variant not illustrated in the figures, the second external diameter D8 can be equal to the first external diameter D26 of the first connecting members 26. For this, the first thickness E46 of the track 46 may be reduced and / or the second E80a and third E80b thicknesses of the ends 80a, 80b of the levers 80 may be reduced.

[0090] The third guide members 8 may number from about two to twenty, preferably this number is between three and twelve. In particular, the levers 80 may number from about two to twenty, preferably from about three to twelve. With reference to [Fig. 6], the levers 80 may number six, distributed regularly around the axis A (or the axis X).

[0091] The levers 80 of the third guide members 8 may be made of a metallic material, such as titanium, steel or aluminum.

[0092] By way of example, Table 1 below summarizes the differences in blade timing measured in a blade stage configuration comprising ten prior art shoes 81 with a clearance of 0.03 mm and that comprising ten levers 80 of the invention with a clearance of 0.1 mm. These measured clearances may correspond to an axial clearance for example at the first radial axis R80a. Thus, the lever 80 can move with the clearance at maximum (or otherwise called total clearance) of approximately 0.1 mm around the first radial axis R 80a*

[0093] It is thus observed that for a clearance greater than the configuration of the pads 81, the timing differences are reduced in the configuration of the levers 80 according to the invention with a single or two actuators 7. This demonstrates that the levers 80 of the invention restrict the deformations of the rings to preserve good kinematics of the timing of the blades 2.

[0094] [Table 1] Configurations Clearance (measured in millimeters) Maximum timing difference (measured in degrees) Single actuator 7 Two actuators 7 Ten pads 81 0.03 0.58 0.161 Ten levers 80 0.1 0.32 0.118

Claims

Claims

1. Stage (1) of variable-pitch blades (2) for an aircraft turbomachine (10), comprising: - an annular row of variable-pitch blades (2), each blade (2) comprising a blade (20) comprising a cylindrical pivot (22) at its radially outer end, this pivot (22) defining a pitch axis (C) of the blade (2), - an annular casing (4) having an axis of revolution (X) and comprising mounting housings (40) for said cylindrical pivots (22), and - a ring (5) for controlling the pitch of the blades (2), the ring (5) comprising: - an annular body (50) which extends around the annular casing (4), - first members (26) for connecting the annular body (50) to the cylindrical pivots (22) of said blades (2), - a second member (70) for connecting the annular body (50) to at least one actuator (7) carried by the annular casing (4), and - third members (8) for guiding the annular body (50) around and at a distance from the annular casing (4),characterized in that said third guide members (8) comprise levers (80) which each comprise a first end (80a) articulated on the annular casing (4) around a radial axis (R80a) relative to the axis of revolution (X), and a second end (80b) connected to the annular body (50).,

2. Variable-pitch blade stage according to claim 1, characterized in that the second end (80b) is articulated by a ball joint (82) to a first fixing pin (84) engaged in an orifice (52) of the annular body (50).

3. Variable-pitch blade stage according to claim 2, characterized in that the first fixing pin (84) has a radial orientation relative to the axis of revolution (X).

4. Variable-pitch vane stage according to any one of the preceding claims, characterized in that the first end (80a) comprises an orifice (85) for mounting a second fixing pin (86) engaged in an orifice (460) of the annular casing (4).

5. Variable-pitch blade stage according to any one of the preceding claims, characterized in that the first end (80a) has a thickness (E80a) or radial dimension greater than that of the second end (80b).

6. Stage of variable-pitch blades according to any one of the preceding claims, characterized in that said third guide members (8) are approximately two to twenty in number, preferably this number is between three and twelve.

7. Variable-pitch blade stage according to any one of the preceding claims, characterized in that the levers (80) are distributed regularly around the axis of revolution (X).

8. Stage of variable-pitch blades according to any one of the preceding claims, characterized in that each of said third guide members (8) has an external diameter (D8) greater than an external diameter (D26) of the first connecting members (26).

9. A variable-pitch vane stage according to any preceding claim, characterized in that the second end (80b) is located at a predetermined radial distance (d80b) from the annular body (50).

10. Stage of variable-pitch blades according to any one of the preceding claims, characterized in that the radial axis (R80a) of the first end (80a) is inclined at a predetermined angle (a) relative to an axis perpendicular to the axis of revolution (X), the angle (a) is for example between 5° and 50°.

11. Aircraft turbomachine (10) comprising at least one stage (1) of variable-pitch blades (2) according to any one of the preceding claims.

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