Splitter for separating a first air flow and a second air flow for a multi-flow axial aircraft turbine engine, and multi-flow axial aircraft turbine engine

Abstract

The invention relates to a splitter (11) for separating a first air flow (106) and a second air flow (107) for a multi-flow axial aircraft turbine engine. The turbine engine comprises a compressor (120) arranged to be axially traversed by the first air flow. The compressor comprises a flow straightener (121) provided with a variable-pitch system (123) for blades that comprises a plurality of pivots (Pv) and / or levers (Lv), each coupled to a blade of the flow straightener (121) in order to modify an orientation thereof. The splitter comprises an inner surface (112) arranged to border the first air flow and an outer surface (113) opposite to the inner surface to border the second air flow. The outer surface has a corrugation formed by an alternation of troughs (Cr) and crests (Ct). The splitter is designed to be arranged at the flow straightener such that each pivot and / or lever is housed radially between the inner surface and a crest of the outer surface. The invention also relates to a multi-flow axial aircraft turbine engine (100).

Description

[0001] - -I -

[0002] SPOUT FOR SEPARATING A FIRST AIR FLOW AND A SECOND AIR FLOW FOR A MULTI-FLOW AXIAL AIRCRAFT TURBOMACHINE

[0003] AND MULTI-FLOW AXIAL AIRCRAFT TURBOMACHINE

[0004] Technical field

[0005] The invention relates to a flow separation nozzle for a multi-flow axial turbomachine.

[0006] Prior art

[0007] A dual-flow aircraft turbomachine generally includes a front-end fan to draw in ambient air and initiate the compression process.

[0008] Downstream of this fan, a separation nozzle divides the incoming air into two flows. The first flow is a hot air flow, called the "primary (air) flow." It passes axially through the aircraft turbomachine, being compressed by one or more compressors and then conveyed to a combustion chamber. The hot gases at its outlet drive one or more turbines that convert the thermal energy into mechanical energy. The second flow is a cold air flow, called the "secondary (air) flow," which is primarily intended to generate a thrust reaction necessary for the aircraft's flight. This flow is accelerated by the fan and circulates around the periphery of the turbomachine's compressors and turbines without passing through the combustion chamber.

[0009] In some compressors, particularly low-pressure compressors, fixed vane straighteners are provided to stabilize and direct the primary airflow. Some straighteners are equipped with a variable-pitch system (called "Variable Stator Vanes"), in order to adjust the angle of the vanes, in order to optimize the compression efficiency according to the flight conditions. The timing adjustment mechanism consists of a set of pivots and levers at the base of the vanes. Document FR3131600A1, for example, discloses such a configuration. The separation nozzle may have a geometry that allows these parts to be integrated.

[0010] This configuration, however, negatively impacts the flow of air flows in the turbomachine:

[0011] • upstream, at the level of the blower and / or a rectifier;

[0012] • downstream, at the level of a secondary air flow rectifier (known as “OGV”). This negative impact consists of the development of undesirable disturbances in the direction of the air flow, with a direct impact on the overall performance of the machine.

[0013] This type of problem is addressed in document FR3078101 B1, which proposes a splitter lip with a corrugated leading edge to reduce aerodynamic interaction noise upstream of the splitter lip.

[0014] Statement of the invention

[0015] In the context of the present invention, the turbomachine considered is typically axial, and extends along a corresponding engine axis. It comprises a compressor arranged to be axially traversed by a primary air flow. The compressor comprises a rectifier comprising a variable blade pitch system provided with a plurality of pivots and / or levers, which are typically arranged circumferentially as is known to a person skilled in the art. Each pivot and / or lever is then coupled to a vane of the rectifier to modify the orientation thereof.

[0016] An object of this invention is to provide a separation nozzle for a first air flow and a second air flow for an aircraft turbomachine improving the aerodynamics of the turbomachine with regard to the prior art described.

[0017] To this end, a nozzle for separating a first air flow and a second air flow for a multi-flow axial aircraft turbomachine is proposed; the nozzle comprising: an inner surface arranged to border the first air flow, an outer surface opposite the inner surface to border the second air flow; in which the outer surface has a corrugation formed of alternating troughs and crests, the nozzle being adapted to be arranged at the rectifier so that each pivot and / or lever is housed radially between the inner surface and a crest of the outer surface.

[0018] The separation slat (or simply "slat" in this document) improves the aerodynamics of the turbomachine. It provides an excellent compromise between, on the one hand, the impact of the separation slat on the machine's performance, and on the other hand, the space required within it to accommodate the variable valve timing system. The corrugation of the external surface of the slat allows the mechanics of the variable valve timing system to be accommodated, even if its radial height is not constant. It allows conforming to the positions in which the variable valve timing system is radially prominent, thus reducing disturbances in the airflow. Indeed, by positioning the troughs of the corrugation where the thickness of the slat can be minimized, it is possible to reduce the overall radial thickness of the slat, ensuring, thanks to the crests, sufficient space for housing the mechanical elements of the variable valve timing system.

[0019] For the purposes of this document, a "ripple" refers to any variation in shape comprising alternating peaks and troughs. These variations may be characterized in terms of amplitude, frequency, and / or waveform. The waveform may be, for example, sinusoidal, trapezoidal, triangular, etc. A person skilled in the art may easily implement other shapes without departing from the scope of the invention. The term "serrational" may refer to one embodiment of a ripple. Preferably, however, the outer surface is smooth and / or regular (in particular, in the mathematical sense).

[0020] As will be understood by those skilled in the art, the words "trough" and "ridge" are used in the sense that the undulation is such that the distance separating a trough from the driving axis (in other words, the radial thickness of the beak at a trough) is strictly less than the distance separating a crest from the driving axis (or in other words, the radial thickness of the beak at a crest).

[0021] The term "border" preferably corresponds to "border directly", that is to say without intermediary. Typically, the internal surface of the nozzle is therefore part of a vein for the circulation of the first air flow (or simply "first flow" in this document) and the external surface of the nozzle is part of a vein for the circulation of the second air flow (or simply "second flow" in this document).

[0022] In this document, unless otherwise indicated, the term "at the level of" may preferably be substituted by "axially at the level of". The use, in this document, of the verb "to understand", its variants, and its conjugations, cannot in any way exclude the presence of elements other than those mentioned. Similarly, the use, in this document, of the indefinite article "un", "une", or of the definite article "le", "la" or "I'", to introduce an element does not exclude the presence of a plurality of these elements.

[0023] The terms "first", "second", "third", etc. are used in this document exclusively to differentiate between different elements, without implying any order between these elements.

[0024] Each crest of the corrugation is preferably arranged to be aligned (radially) with a pivot axis of the variable pitch system. This allows for optimization of the space required for housing the mechanical parts, since the pivot is the element that generally presents the greatest risk of interference with the external surface of the nozzle. This design ensures that each mechanical element is ideally positioned, thus reducing potential interference.

[0025] In one embodiment, the undulation is present in a radial section of the outer surface of the spout. This configuration is characterized by a succession, preferably regular, for which each crest is followed by an adjacent trough in the radial section. By circumferentially alternating crests and troughs, an effective reduction in the average radius of the outer surface of the spout is obtained. This reduction can favorably influence the weight, the size and, consequently, the costs associated with the manufacture of the spout.

[0026] In one embodiment, the corrugation has a (substantially) sinusoidal shape in a radial section of the outer surface of the beak. This provides the advantage of a significantly smooth and / or regular surface, which improves the aerodynamic performance of the beak and facilitates its design.

[0027] Preferably, the undulation is of (essentially) constant amplitude. The amplitude is typically adapted to the space required for the arrangement of pivots and / or levers of the variable-pitch system (at the crests). This configuration allows for uniform interaction with the airflow, thus promoting stable and / or predictable aerodynamic performance by reducing potential local disturbances. In addition, it simplifies the manufacture of the slat since the radial section has periodically identical characteristics.

[0028] The same applies to the embodiment in which the hollows and the crests are successively equidistant two by two, which is furthermore adapted to the circumferential arrangement of the pivots and / or levers.

[0029] In one embodiment, each crest and each trough of the corrugation extends (at least) axially, from an upstream position to a downstream position. Thus, the corrugation has an axial extension. Preferably, this is at least 5 cm long, optionally at least 10 cm long, and / or preferably at most 30 cm long, measured along the axis. Preferably, the upstream position is adjacent a leading edge of the nozzle.

[0030] Preferably, each crest and each trough of the corrugation also extends circumferentially from the upstream position to the downstream position, preferably along a helical curve. The crests and troughs more preferably extend in a direction aligned with the flow of the second stream at a main performance point, for example on a helical curve that forms an extension angle of about 30 to 40° relative to the motor axis.

[0031] Advantageously, this embodiment makes it possible to create channels for the flow of the air flow, delimited by the walls formed by the extension of the corrugation on the external surface. The axial, and preferably circumferential, extension then acts as a rectifier, preferably guiding the second air flow and readjusting its parietal direction to bring it closer to its nominal trajectory. Thus, disturbances are minimized, and the flow is harmonized throughout the passage on the external surface of the nozzle.

[0032] Preferably, the upstream position is located at a leading edge of the nozzle. Thus, the second flow is guided by the undulation at the external surface of the nozzle as soon as it enters the vein intended for the circulation of this flow.

[0033] Preferably, the downstream position is arranged to be located at a rectifier (called "OGV") of the second air flow of the turbomachine. Thus, the second flow is guided in the vein provided for the circulation of this flow to the rectifier of the second flow. Preferably, the downstream position is arranged to be located at least up to the level of the leading edge of the vanes of the rectifier, and preferably, up to the level of the trailing edge of the vanes of the rectifier.

[0034] These designs are particularly advantageous because they take advantage of the undulating shape of the outer surface of the nozzle both to arrange the pivots and / or levers of the variable-pitch system, but also to adequately direct the second flow at the second-flow straightener. This latter effect is particularly evident at the foot of the straightener. In this way, the undulation defines ribs that nominally redirect the second flow at the outer surface to the leading edge of the straightener vanes. This offers a considerable gain in stability and performance of this second-flow straightener.

[0035] According to one embodiment, the beak comprises a leading edge on which the corrugation extends (axially). According to another distinct embodiment, the beak comprises a leading edge, preferably convex and / or which typically has an annular profile, from which the corrugation extends axially. In other words, the corrugation may originate from the leading edge (not included) or be present on the leading edge (in which case, it is not convex).

[0036] In the first case, the leading edge allows it to remain convexly regular and to obtain better aerodynamic control of the separation of the air flows in the circulation veins of the first and second flows. In the second case, providing the leading edge of the beak with the corrugation has the advantages described on page 2 of document FR3078101 B1, which are incorporated by reference, and can facilitate the design thereof. It is noted in the latter case that using the same corrugation extending over the leading edge and over the external surface of the beak is advantageous and simple to design.

[0037] The invention further provides a turbomachine comprising a nozzle according to any of the aforementioned embodiments. All of the preferred embodiments as well as all of the advantages of the nozzle according to the invention are transposed mutatis mutandis to the present turbomachine.

[0038] More specifically, the turbomachine is a multi-flow axial aircraft turbomachine comprising: a compressor arranged to be axially traversed by a first air flow and comprising a rectifier provided with a variable blade setting system comprising a plurality of pivots and / or levers, each coupled to a blade of the rectifier to modify an orientation thereof; and a nozzle separating a first air flow and a second air flow, arranged at the rectifier, and comprising:

[0039] • an internal surface designed to border the first air flow,

[0040] • an external surface opposite the internal surface to border the second air flow; in which the external surface has a corrugation formed of alternating troughs and crests, each pivot and / or lever being housed radially between the internal surface and a crest of the external surface.

[0041] In other words, the nozzle is according to the invention, and it is in particular arranged in the turbomachine as provided by the invention.

[0042] Preferably, the compressor is a low pressure compressor.

[0043] Preferably, the turbomachine comprises: a fan (ducted or not); successively axially:

[0044] • the low pressure compressor, i.e. “the compressor” according to the invention;

[0045] • a high pressure compressor,

[0046] • a combustion chamber,

[0047] • one or more turbines, arranged to be crossed by the first air flow; a circulation vein of the second air flow around the low and high pressure compressors, the combustion chamber and the turbine(s); a rectifier of the second air flow arranged in the vein at the low pressure compressor.

[0048] A leading edge of the nozzle is preferably located axially between the fan and the rectifier of the second air flow, and more optionally between the fan and the low pressure compressor (typically in the case of a turbomachine with a ducted fan and double flow).

[0049] The external surface of the nozzle partly forms the circulation vein of the second air flow. The undulation then preferably extends axially, and more preferably in a helical manner around the engine axis, from the leading edge of the nozzle to the level of the rectifier of the second air flow, and preferably to the level of the trailing edge of the rectifier of the second air flow. The undulation thus also plays a role in orienting the second flow at the level of the rectifier of the second air flow as mentioned above.

[0050] Generally, those skilled in the art will understand that all embodiments relating to the nozzle apply to the turbomachine by positioning the nozzle in the manner provided by the invention.

[0051] Brief description of the figures

[0052] Other characteristics and advantages of the present invention will appear on reading the detailed description which follows, for the understanding of which reference will be made to the appended figures among which:

[0053] - figure 1 illustrates an axial section of a state-of-the-art turbomachine on which the nozzle according to the invention is intended to be arranged;

[0054] - figure 2 illustrates, among other things, a radial section of the spout according to one embodiment of the invention;

[0055] - figures 3a and 3b illustrate an axial section of the nozzle and the corresponding flow separation according to one embodiment of the invention.

[0056] The drawings of the figures are not to scale. Like elements are, in general, denoted by like references in the figures. For the purposes of this document, identical or similar elements may bear the same references. Furthermore, the presence of reference numbers or letters in the drawings cannot be considered as limiting, including when these numbers or letters are indicated in the claims. Detailed description of embodiments of the invention

[0057] This section presents a detailed description of preferred embodiments of the present invention. The latter is described with particular embodiments and references to figures but the invention is not limited thereby. In particular, the drawings and figures described below are only schematic and are not limiting.

[0058] References are illustrated in these figures as abstract geometric references essentially in order to quantify and / or visualize properties of embodiments of the invention. The reference X designates, for example, generally the “engine axis” of the aircraft turbomachine. This is directed from “upstream” to “downstream” as is known to a person skilled in the art, these terms being used to locate elements along the engine axis. The turbomachine comprises stages (or rows of blades) forming compressors and turbines of the aircraft turbomachine which are stacked essentially along the engine axis.

[0059] In the context of this document, reference is made to the directions "axial", "circumferential" and "radial" preferably corresponding and respectively in directions parallel to the motor axis, essentially circular around the motor axis, and direction perpendicular to the motor axis. References in Figures 1 and 2 illustrate these directions (provided with a direction) respectively denoted X, R and C. The direction and the direction of the motor axis and the vector X correspond. The terms "axially", "circumferentially" and "radially" are derived from the respective words "axial", "circumferential" and "radial" with a similar preferred meaning. The terms "circumferential" and "radial" further preferably refer to a polar coordinate system known to a person skilled in the art in each plane perpendicular (or "radial plane") to the motor axis.A radial section of a turbomachine or an element thereof designates in particular its intersection with a radial plane.

[0060] The detailed description presents the embodiment of integration of the nozzle according to the invention within a dual-flow axial turbomachine, said first and second (air) flows being respectively called primary (air) flow and secondary (air) flow. This embodiment applies in combination with all the other embodiments presented in the description of the invention. However, it is not limiting on the number of flows that the turbomachine can include nor on the arrangement of the nozzle between any two of these flows.

[0061] With reference to Figure 1, the turbomachine 100 comprises a fan 110 provided with a plurality of blades. The fan is ducted in the case illustrated in Figure 1, but it can also be unducted, of the “propeller (open rotor)” type within the framework of the invention. The turbomachine 100 also comprises successively along the engine axis (parallel to the X axis), a low-pressure compressor 120, a high-pressure compressor 130, a combustion chamber 160, a high-pressure turbine 140 and a low-pressure turbine 150. These elements are known to a person skilled in the art. In operation, the mechanical power of the low 150 and high 140 pressure turbines is transmitted via the low 101 and high 102 pressure shafts to the low 120 and high 130 pressure compressors respectively, as well as to the fan 110 via the low pressure shaft 101.Although not systematically referenced in Figures 1, each compressor 120, 130 and each turbine 140, 150 comprises at least one stage, each such stage comprising a fixed blade and a mobile blade capable of being rotated around the engine axis X.

[0062] For the low-pressure compressor 120, these fixed and moving blades are denoted 121 and 122 respectively. The fixed blades of the compressors 120, 130 are generally referred to as "straighteners". The moving and fixed blades of the stages of such a compressor 120, 130 are alternated along the engine axis. Parameters such as the dimensions and surface geometry of the blades are determined so that the operating conditions of each stage are adapted to those of the upstream and / or downstream stages along the engine axis. In particular, the moving blades provide energy by increasing the relative flow velocity of an air stream passing through the compressor 120, 130, while the straighteners bring the flow back parallel to the engine axis while increasing the pressure and decreasing the absolute flow velocity. Each fixed blade has a given aerodynamic profile and has a pitch angle relative to the engine axis to impose a flow direction.The low 120 and high 130 pressure compressors thus make it possible to synergistically suck in and compress air so as to bring it to speed, pressure and temperature adapted to the inlet of the combustion chamber 160. These concepts are known to a person skilled in the art.

[0063] One or more rectifiers of the compressors 120, 130 are provided with a variable-pitch system for their blades to optimize the flow of the air flow between stages of these compressors 120, 130. As is well known to a person skilled in the art, each of the aforementioned variable-pitch rectifier blade systems preferably comprises a ring mechanically coupled to the blades and circumferentially adjusted externally around a casing of the associated compressor 120, 130, and a displacement unit, typically at least one cylinder, for moving this ring, and consequently modifying in a synchronized manner the pitch of the rectifier blades.

[0064] According to figures 1 and 2, the low pressure compressor 120 comprises at least one rectifier 121 (more simply called “the rectifier 121” in the following) provided with such a variable timing system referenced by 123. This rectifier 121 is preferably upstream of the low pressure compressor 120. Figure 1 is not limiting of the case where one or more of the other rectifiers would also be provided with such a variable timing system.

[0065] Vanes of the rectifier 121 can thus pivot relative to an air flow passing through the rectifier. By being more or less inclined relative to the air flow passing through the compressor 120, these vanes deflect this air flow more or less to orient the latter. As is known, each orientable vane comprises a pivot control axis 8 Pv which is linked to an actuating lever Lv. The control axis 8 is substantially oriented radially and materializes the geometric axis around which the vanes oscillate. A change in the positioning of a lever Lv causes, via the pivot Pv to which it is coupled, a rotation of the corresponding vane of the rectifier 121 around its control axis 8.

[0066] A first hot air flow 106, hereinafter called the primary air flow, passes axially through the aircraft turbomachine 100, thereby supplying the combustion chamber 160, while a second cold air flow 107, hereinafter called the secondary air flow, is primarily intended to generate a thrust reaction necessary for the flight of the aircraft. The primary 106 and secondary 107 flows are annular flows, separated from each other downstream of the fan 110 and upstream of the low-pressure compressor 120 by a flow separation nozzle 11 as illustrated in FIGS. 3a and 3b. This comprises an inner surface 112 bordering the primary air flow 106 and an outer surface 113 opposite the inner surface bordering the secondary air flow 107. The outer surface 113 is radially further from the engine axis X than the inner surface 112, as their names indicate. The separation beak 11 comprises a leading edge 12.

[0067] The annular vein for the circulation of the secondary air flow 107 surrounding, among other things, the low-pressure compressor 120 is provided with a rectifier 16 called "OGV" composed of fixed blades as is known to a person skilled in the art. This, although not shown in figure 1, is visible in figures 2, 3a and 3b. The purpose of the rectifier 16 is to straighten the secondary flow 107 relative to the engine axis X. Its blades are provided with a leading edge 161 and a trailing edge 162, as is known to a person skilled in the art.

[0068] As can be seen in Figure 2, the external surface 113 of the beak 11 has a corrugation formed by a circumferential alternation of peaks Ct and troughs Cr. The separating beak 11 is arranged at the level of the rectifier 121 so that each pivot Pv is housed between the internal surface 112 and a peak Ct of the external surface 113. In the embodiment illustrated in Figure 2, each peak Ct is radially aligned with the control axis 8 of a pivot Pv of the variable timing system 123. Although this arrangement is easy to implement and very preferred, variants could be envisaged such as a deliberate offset between the peaks Ct and the axes of the pivots Pv, according to optimization parameters or specific installation constraints.

[0069] The undulation in the radial section shown in the foreground of Figure 2 is characterized by a regular succession where each peak Ct is followed by an adjacent trough Cr. The undulation in said radial section is sinusoidal in shape. Advantageously, it is smooth and regular, contributing to the aerodynamics of the turbomachine 100 without unduly complicating the manufacture of the nozzle 11. This also leads to a modulation of the average radius of the external surface 113 of the nozzle 11.

[0070] In Figure 2, the corrugation has an amplitude A that remains invariable around the entire circumference of the nozzle 11. However, some applications might require an amplitude A adapted to specific variations in flows or structural constraints.

[0071] If the compressor 120 comprises several rectifiers 121 provided with such a variable-pitch system 123, the undulation may be such that each pivot Pv / or lever Lv of each rectifier 121 is housed between the internal surface 112 and a crest Ct of the external surface 113. This may be done by an axial extension of each crest Ct and trough Cr, or by an axial undulation of the preceding sinusoidal pattern in a radial plane. In this case, not shown, an axial section of the nozzle 11 would also comprise a corrugation for which, when a crest is at the level of a rectifier, a trough is at the level of a row of moving blades 122.

[0072] The primary flow rectifier 121 106, which is located in an axial position 60 which corresponds to the axial position of the row of pivots Pv (only the axial projection of the Pv pivot control axes 8 is shown in Figures 3a and 3b). Each pivot Pv is typically housed radially between the inner surface 112 and a crest Ct of the outer surface 113 of the nozzle 11.

[0073] With reference to Figure 3a, each crest Ct and each trough Cr of the corrugation extends at least axially along a proper curve, preferably helical around the engine axis, from an upstream position adjacent to or at the level of the leading edge 12 of the slat 11 to a downstream position. The leading edge 12 does not include the corrugation, and then has a circular profile, so that it is convex.

[0074] With reference to Figure 3b, the leading edge 12 has a corrugation on its axial trace originating from the axial extension of the corrugation on the leading edge 12. The secondary air flow 107 can thus be oriented axially in the direction of the primary air flow 106 up to the downstream position.

[0075] According to Figures 3a and 3b, the downstream position is located at the level of the rectifier 16 of the secondary air flow 107. It corresponds for example to the axial level of the leading edge 161 or the trailing edge 162 of the vanes of the rectifier 16. This makes it possible to have a corrugation extending sufficiently between the upstream position and the downstream position to guide the secondary air flow 107 at the foot of the vanes of the rectifier 16 as described in the description of the invention. In summary, the invention relates to a flow separation nozzle 11 for an axial aircraft turbomachine 100, making it possible to ensure good aerodynamic performance while adapting to the internal mechanics of a variable blade pitch system 123 of a rectifier 121 of a compressor 120.

[0076] The present invention has been described above in relation to specific embodiments, which are of purely illustrative value and should not be considered as limiting. It will be readily apparent to those skilled in the art that the invention is not limited to the examples illustrated or described above, and that its scope is more broadly defined by the claims introduced below.

Claims

Claims 1. A nozzle (11) for separating a first air flow (106) and a second air flow (107) for a multi-flow axial aircraft turbomachine (100); the turbomachine (100) comprising a compressor (120) arranged to be axially traversed by the first air flow (106), and the compressor (120) comprising a rectifier (121) provided with a variable blade pitch system (123) comprising a plurality of pivots (Pv) and / or levers (Lv), each coupled to a blade of the rectifier (121) to modify an orientation thereof; the nozzle (11) comprising: an inner surface (112) arranged to border the first air flow (106), an outer surface (113) opposite the inner surface to border the second air flow (107);characterized in that the external surface (113) has a corrugation formed by alternating hollows (Cr) and crests (Ct), the beak (11) being adapted to be arranged at the level of the rectifier (121) so that each pivot (Pv) and / or lever (Lv) is housed radially between the internal surface (112) and a crest (Ct) of the external surface (113).; 2. Spout (11) according to claim 1, wherein the corrugation is present in a radial section of the external surface (113).

3. A spout (11) according to claim 2, wherein the undulation has a substantially sinusoidal shape in the radial section.

4. Spout (11) according to any one of claims 1 to 3, in which the undulation is of an essentially constant amplitude (A).

5. Spout (11) according to any one of claims 1 to 4, in which the hollows (Cr) and the crests (Ct) are successively equidistant two by two.

6. Spout (11) according to any one of claims 1 to 5, in which each crest (Ct) and each trough (Cr) of the undulation extends at least axially from an upstream position to a downstream position.

7. A spout (11) according to claim 6, wherein each crest (Ct) and each trough (Cr) of the undulation also extends circumferentially from the upstream position to the downstream position, preferably along a helical curve.

8. Spout (11) according to claim 6 or 7, wherein the upstream position is located at a leading edge (12) of the spout (11).

9. Nozzle (11) according to any one of claims 6 to 8, in which the downstream position is arranged to be located at a rectifier (16) of the second air flow (107) of the turbomachine (100).

10. Beak (11) according to any one of claims 1 to 9, comprising a leading edge (12) on which the undulation extends.

11. Beak (11) according to any one of claims 1 to 9, comprising a convex leading edge (12) from which the corrugation extends axially.

12. Axial multi-flow aircraft turbomachine (100) comprising: a compressor (120) arranged to be axially traversed by a first air flow (106) and comprising a rectifier (121) provided with a variable blade pitch system (123) comprising a plurality of pivots (Pv) and / or levers (Lv), each coupled to a blade of the rectifier (121) to modify an orientation thereof; and a nozzle (11) for separating a first air flow (106) and a second air flow (107), arranged at the rectifier (121), and comprising: • an internal surface (112) arranged to border the first air flow (106), • an external surface (113) opposite the internal surface to border the second air flow (107); characterized in that the external surface (113) has a corrugation formed by an alternation of hollows (Cr) and crests (Ct), each pivot (Pv) and / or lever (Lv) being housed radially between the internal surface (112) and a crest (Ct) of the external surface (113).

13. Turbomachine (100) according to claim 12, in which the nozzle (11) is according to any one of claims 2 to 11.

14. Turbomachine (100) according to claim 12 or 13, in which the compressor (120) is a low pressure compressor and in which the turbomachine (100) comprises: a fan (110); successively axially: • the compressor (120), • a high pressure compressor (130), • a combustion chamber (160), • one or more turbines (140, 150), arranged to be traversed by the first air flow (106); a circulation vein of the second air flow (107) around the low and high pressure compressors, the combustion chamber (160) and the turbine(s) (140, 150); a rectifier (16) of the second air flow (107) arranged in the vein at the low pressure compressor (120); a leading edge (12) of the nozzle (11) being located axially between the fan (110) and the rectifier (16) of the second air flow (107).

15. Turbomachine (100) according to claim 14, in which the corrugation extends axially from the leading edge (12) of the nozzle (11) to the level of the rectifier (16) of the second air flow (107).

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