EP4278075A1

EP4278075A1 - Propulsion assembly for an aircraft comprising a stator vane integrated into an upstream part of a mounting pylon of reduced height

Info

Publication number: EP4278075A1
Application number: EP22702295.1A
Authority: EP
Inventors: Eva Julie Lebeault; Anthony BINDER; Laurent SOULAT
Original assignee: Safran Aircraft Engines SAS
Current assignee: Safran Aircraft Engines SAS
Priority date: 2021-01-12
Filing date: 2022-01-05
Publication date: 2023-11-22
Also published as: CN116917610A; FR3118787A1; FR3118787B1; WO2022152994A1; US20240116644A1

Abstract

The invention relates to a propulsion assembly (200) for an aircraft comprising a double-flow turbomachine (1) equipped with a fan (15), an aerodynamic outer casing (5) forming a nacelle, and a mounting pylon, the propulsion assembly having a secondary duct (18) delimited by an outer radial delimiting surface (23a) formed by the casing (5), the turbomachine comprising stator vanes (30), and the mounting pylon (7) comprising a part housed in the secondary duct, termed upstream part (7b). According to the invention, the upstream part (7b) of the pylon extends radially from the inner radial delimiting surface (40a) over a radial pylon height (Hm) strictly less than a total radial height (Ht) of the secondary duct (18), and the upstream part (7b) of the pylon extends in the downstream direction from a root part (62a) of one of the stator vanes (30).

Description

DESCRIPTION

Title: PROPULSIVE ASSEMBLY FOR AIRCRAFT COMPRISING A RECTIFIER VANE INTEGRATED IN AN UPSTREAM PART OF A REDUCED HEIGHT ATTACHMENT MAST

TECHNICAL FIELD The present invention relates to the field of propulsion systems for aircraft. It relates more particularly to propulsion assemblies comprising a nacelle of reduced length, referred to as a “short nacelle”, such as that described in document EP 2 628919 Al.

STATE OF THE PRIOR ART In a propulsion assembly comprising a dual-flow turbomachine, a primary gas circulation stream is provided, as well as a secondary gas circulation stream delimited radially outwards by an aerodynamic outer casing forming a nacelle. . The turbomachine generally comprises a fan sucking in a mass of air which is then divided into a primary flow circulating in the primary stream, and a secondary flow circulating in the secondary stream.

The primary flow typically passes through one or more compressors, for example a low pressure compressor and a high pressure compressor, a combustion chamber, one or more turbines, for example a high pressure turbine and a low pressure turbine, then finally a nozzle of gas exhaust. In known manner, the high pressure turbine drives the high pressure compressor in rotation via a first shaft, called the high pressure shaft, while the low pressure turbine drives the low pressure compressor and the fan in rotation via a second shaft, called the low pressure shaft. In order to improve the propulsion efficiency of the turbomachine and to reduce its specific consumption as well as the noise emitted by the fan, turbojets with an increased bypass ratio have been proposed, which corresponds to the ratio between the flow rate of the secondary flow (cold), and the flow rate of the primary flow (hot) which crosses the primary body). To achieve such bypass ratios, the fan is decoupled from the low pressure turbine, thus making it possible to independently optimize their respective rotation speeds. Usually, the decoupling is achieved using a reducer such as an epicyclic reduction mechanism, placed between the upstream end of the low pressure shaft and the fan. The fan then becomes driven indirectly by the low-pressure shaft, via the reduction mechanism and an additional shaft, called the fan shaft, which is fixed between the reduction mechanism and the hub of the fan. This decoupling thus makes it possible to reduce the speed of rotation and the pressure ratio of the fan (“fan pressure ratio”), and to increase the power extracted by the low pressure turbine. Thanks to the reduction mechanism, the low-pressure shaft can thus rotate at higher rotational speeds than in conventional turbojet engines.

Still in the conventional embodiments, the turbomachine is equipped, downstream of the fan, with an annular row of stator vanes (also called outlet guide vanes, or OGV vanes, standing for “Outlet Guide Vanes”). These stator vanes are located in the cold part of the turbomachine, in the secondary stream. Their main purpose is to straighten the flow of cold air coming from the fan blades.

In the short nacelle propulsion assembly configurations, the engine attachment pylon of the turbomachine may have to penetrate partly into the secondary stream, downstream of the stator vanes, thereby generating aerodynamic losses by friction, with a negative impact on the overall performance of this propulsion system.

DISCLOSURE OF THE INVENTION

To respond to the drawback mentioned above, relating to the embodiments of the prior art, the invention firstly relates to a propulsion assembly for an aircraft comprising a turbofan engine equipped with a fan, an outer casing aerodynamics forming a nacelle arranged around the fan, as well as an attachment pylon intended to ensure the attachment of the turbomachine to a wing element of the aircraft, the propulsion assembly having a primary flow path gases, as well as a secondary gas circulation stream delimited by an inner radial delimitation surface as well as by an outer radial delimitation surface formed by the aerodynamic outer casing, the turbomachine further comprising an annular row of vanes stator arranged in the secondary stream downstream of the fan, each stator vane extending through the secondary stream with a head end connected to the outer aerodynamic casing, as well as a root end connected to the surface of internal radial delimitation of the secondary stream, the attachment pylon comprising a part housed in the secondary stream, called the upstream part, as well as a downstream part arranged downstream from a trailing edge of the aerodynamic outer envelope, itself even intended to be arranged entirely upstream of a leading edge of the wing element.

According to the invention, the upstream part of the attachment mast housed in the secondary vein extends radially from the internal radial delimitation surface, over a radial height of the mast strictly less than a total radial height of the secondary vein, and further, the upstream portion of the engine mount extends downstream from a root portion of one of the stator vanes.

The invention thus provides an attachment pylon which extends only over part of the radial height of the secondary stream into which it penetrates, in order to advantageously limit the aerodynamic losses by friction in this same secondary stream. As a complementary measure to the previous one, the invention provides for the integration of the root part of one of the stator vanes with the upstream part of this attachment strut, in order to form an aerodynamic continuity between these two parts, in the axial direction. Such an integration makes it possible to further minimize aerodynamic losses, and also to reduce the transverse heterogeneities of static pressure (distortion), rising from the downstream towards the fan.

The combination of these measures makes it possible overall to improve the propulsion efficiency of the turbomachine.

The invention preferably provides at least any one of the following optional features, taken alone or in combination. Preferably, the radial height of the mast represents locally, for example, 20 to 70% of the total radial height of the secondary stream, and more particularly 30 to 60% of the total radial height of the secondary stream. This percentage can obviously change along the upstream part of the mast, and therefore not remain constant. Preferably, the stator vane integrated into the attachment pylon comprises, radially outwards from the upstream part of this pylon, a free trailing edge extending as far as the outer aerodynamic casing. This free trailing edge thus extends radially over part of the height of the secondary vein complementary to the radial height of the mast. Preferably, the vane integrated into the engine mount comprises the following parts, succeeding one another radially from the inside outwards:

- the foot part integrated into the upstream part of the attachment mast;

- a transition part; and

- a head part. According to a first preferred embodiment of the invention, the trailing edge of the transition part has a transverse thickness which increases going radially towards the root part of the integrated blade. This makes it possible to obtain a smooth transition between the usually fine thickness of the trailing edge of the tip part of the integrated stator vane, and the much greater thickness of the fictitious trailing edge of the root part of this blade, this trailing edge blending into the front end of the attachment mast.

Alternatively, a sudden change in thickness could be provided in the radial direction, between the root part of the integrated blade and its transition part, the thickness of which could then be identical or similar to that of the tip part of dawn. According to a second preferred embodiment of the invention, the transition part has a chord of greater length than that of the head part. By locally increasing the length of the chord, it is possible to keep a thin trailing edge limiting the losses of base, while providing a more substantial blade thickness upstream of this trailing edge, at the place where it connects radially with the thick fictitious trailing edge of the root part of this blade. This advantageously makes it possible to limit the thickness differential between the root part and the transition part of the integrated blade, and therefore to soften the connection with the consequence of gains in terms of aerodynamic performance.

According to a third preferred embodiment of the invention, said transition part comprises a truncated trailing edge so that the chord of the transition part has an increasing length going from the foot part towards the head part. Here again, the proposed solution makes it possible to soften the connection between the root part and the transition part of the integrated blade, since this connection is made at the place where the respective thicknesses are the most similar, that is that is to say in whole or in part upstream of the thick fictitious trailing edge of the root part of this blade. The radial thickness transition is advantageously smoother, here also with gains in terms of aerodynamic performance.

Preferably, the ratio between the total axial length of the aerodynamic outer casing, and the diameter of the fan, is less than 1.25. This report illustrates a so-called “short” nacelle design, generally associated with the fact that it does not incorporate a thrust reverser system. In such a short nacelle, intended to be located facing axially and upstream of the leading edge of the wing element, the thrust reversal system is integrated into the fan which has for this purpose rotary fan blades with variable pitch. This architecture in which the thrust reversal function is performed by the fan is known as “VPF architecture” (from the English “Variable Pitch Fan”).

The invention also relates to a part of an aircraft comprising such a propulsion assembly, as well as a wing element, the aerodynamic outer envelope forming the nacelle extending entirely upstream of a leading edge of the element. of sail. Preferably, an upper part of the casing is arranged axially opposite the leading edge of the wing, and an apex of the exterior of the aerodynamic outer casing extends higher than the wing element considered in line with the connection of the attachment pylon with said wing element.

Finally, the invention also relates to an aircraft comprising at least one such part, and preferably two parts respectively integrating the two wings of the aircraft. Other advantages and characteristics of the invention will appear in the non-limiting detailed description below.

BRIEF DESCRIPTION OF DRAWINGS

This description will be given with regard to the appended drawings, among which; [Fig. 1] shows a schematic view in longitudinal half-section of a propulsion assembly according to a first preferred embodiment of the invention, the section plane passing through the hourly positions at 12 o'clock and 6 o'clock of the turbojet equipping this propulsion assembly;

[Fig. 2] shows a front view of the propulsion assembly shown in Figure 1, the rectifier blades having been removed from the secondary stream for greater clarity, with the exception of one of them which is integrated into the mast , the fan blades having also been removed;

[Fig. 3] is a top view of an aircraft fitted with the propulsion assembly shown in the preceding figures; [Fig. 4] shows a sectional view taken along line IV-IV of FIG. 1, line IV-IV being parallel or substantially parallel to the longitudinal central axis of the turbojet engine, and crossing a radially inner end portion of an integrated vane transition portion;

[Fig. 5] shows a sectional view taken along line V-V of FIG. 1, line V-V being parallel or substantially parallel to the longitudinal central axis of the turbojet engine, and crossing a radially outer end portion of a foot part integrated dawn;

[Fig. 6] shows a sectional view taken along line VI-VI of Figure 5;

[Fig. 7] is a sectional view similar to that of FIG. 4, with the integrated blade in the form of an alternative embodiment;

[Fig. 8] is a sectional view similar to that of FIG. 5, with the integrated blade taking the form of the alternative embodiment shown in the preceding figure; [Fig. 9] shows a sectional view taken along line IX-IX of Figure 8; [Fig. 10] represents a sectional view similar to that of FIG. 1, less detailed, and showing the propulsion assembly according to a second preferred embodiment of the invention

[Fig. 11] shows a sectional view taken along line XI-XI of FIG. 10, line XI-XI being parallel or substantially parallel to the longitudinal central axis of the turbojet, and crossing a radially inner end portion of an integrated vane transition portion;

[Fig. 12] shows a sectional view taken along line XII-XII of FIG. 10, line XII-XII being parallel or substantially parallel to the central longitudinal axis of the turbojet engine, and crossing a radially outer end portion of an integrated blade root part;

[Fig. 13] represents a sectional view similar to that of FIG. 1, less detailed, and showing the propulsion assembly according to a third preferred embodiment of the invention; [Fig. 14] shows a sectional view taken along line XIV-XIV of FIG. 13, line XIV-XIV being parallel or substantially parallel to the central longitudinal axis of the turbojet, and crossing a radially inner end portion of an integrated vane transition portion; and

[Fig. 15] shows a sectional view taken along line XV-XV of FIG. 13, line XV-XV being parallel or substantially parallel to the central longitudinal axis of the turbojet, and crossing a radially outer end portion of an integrated blade root part.

DETAILED DISCUSSION OF PREFERRED EMBODIMENTS

Referring to Figures 1 and 2, there is shown a part 100 of an aircraft, comprising a propulsion assembly 200 and a wing element 202, here an aircraft wing. Preferably, these are two parts 100 which are arranged laterally on either side of the fuselage 204 of the aircraft 300 shown in FIG. 3 (only one of the two propulsion assemblies 200 being represented in this FIG. 3). The propulsion assembly 200 comprises a turbofan engine 1 with a double spool, an outer aerodynamic envelope 5 forming a nacelle, as well as an attachment strut 7 intended for mounting the turbojet engine 1 on the wing 202. In the figures, the attachment strut 7 is shown only with its outer contour formed by one or more aerodynamic fairings. Inside these fairings, a so-called primary structure (not shown) is conventionally provided, intended to ensure the transfer of forces between the turbojet engine 1 and the wing 202. More specifically, the primary structure of the pylon is generally fixed on a forward wing spar 208 .

In addition to enclosing the primary structure, the fairings of the engine mount 7 incorporate a certain number of conventional elements connecting the engine to the aircraft, such as pipes, heat exchangers, electrical cables, drive shafts mechanics, structural parts of the engine suspension system, etc.

The bypass and double-spool turbojet engine 1 has a very high bypass ratio, for example greater than 15. Consequently, its outer diameter is large, and this is the reason why it is arranged in a position substantially raised by relative to the wing 202 which carries it, so as to retain sufficient ground clearance despite the large diameter. As can be seen in FIGS. 1 to 3, the outer casing forming the nacelle 5 is located entirely upstream with respect to a leading edge 210 of the wing 202. In this respect, it is noted that the terms “upstream and “downstream” are considered along a main direction 14 of gas flow within the turbojet, when the latter is in normal propulsion configuration. The terms “front” and “rear” are for their part employed in relation to a direction opposite to the main direction of flow of the gases 14. An upper part of the casing 5 is arranged facing axially from the leading edge 210 of the wing 202, which reflects the raised character of the propulsion unit 200 with a short nacelle. In addition, a top of the exterior of the aerodynamic outer casing 5 extends higher than the wing element 202 considered in line with the connection of the attachment strut 7 with said wing element. The turbojet engine 1 conventionally comprises a gas generator 2 on either side of which are arranged a low pressure compressor 4 and a low pressure turbine 12, this gas generator 2 comprising a high pressure compressor 6, a combustion chamber 8 and a high pressure turbine 10. The low pressure compressor 4 and the low pressure turbine 12 form a low pressure body, and are connected to each other by a low pressure shaft 11 centered on a central longitudinal axis 3 of the turbojet engine. Similarly, the high pressure compressor 6 and the high pressure turbine 10 form a high pressure body, and are connected to each other by a high pressure shaft 13 also centered on the axis 3, and arranged around the low pressure shaft 11. The turbojet engine 1 also comprises, upstream of the gas generator 2 and of the low pressure compressor 4, a single fan 15 which is here arranged directly behind an air inlet cone of the engine. The fan 15 comprises a ring of fan blades 17 rotating around the axis 3, this ring being surrounded by a fan casing 9. The fan blades 17 are variable-pitch, that is to say that their incidence can be controlled by a control mechanism 20 arranged at least partly in the inlet cone, and designed to cause these blades 17 to pivot around their respective longitudinal axes 22. This control mechanism 20, of known design of the type mechanical, electrical, hydraulic, and/or pneumatic, is itself controlled by an electronic control unit (not shown), which makes it possible to order the value of the pitch angles of the blades 17 according to the needs encountered, in particular to exercise reverse thrust function.

As mentioned above, this is a propulsion unit 200 whose thrust reversal function is effectively integrated into the fan 15, and not into the outer casing forming the nacelle 5, as is most commonly meet. This casing 5 can therefore have a short length compared to the diameter of the fan 15. Preferably, the ratio between the total axial length "Lt" of the casing 5, and the diameter "D" of the fan 15, is less than 1.25.

In the remainder of the description of the propulsion assembly 200, reference is made to the longitudinal direction X parallel to the axis 3 of the turbojet, and also referred to as the axial direction, to the transverse direction Y also referred to as the lateral direction, and finally to the vertical direction Z also called the direction of the height, these three directions X, Y and Z being mutually orthogonal. Reference is also made to the radial direction R, to be considered in relation to axis 3.

The fan 15, of the VPF type, is not directly driven by the low pressure shaft 11, but only indirectly driven by this shaft, via a reduction mechanism 24, which allows it to rotate with a slower speed. Nevertheless, a solution with direct drive of the fan 15, by the low pressure shaft 11, falls within the scope of the invention.

In addition, the turbojet engine 1 defines a primary stream 16 of gas circulation, intended to be crossed by a primary flow 16a, as well as a secondary stream 18 of gas circulation, intended to be crossed by a secondary flow 18a located radially towards outside with respect to the primary flow. The flow from the fan 15 is thus divided at the level of a nozzle 26 for separating the flows.

As is known to those skilled in the art, the secondary vein 18 is delimited radially outwards in part by an outer shroud 23, integrated into the casing 5. The shroud 23 is preferably metallic, and it extends towards the behind the fan casing 9. More specifically, the shroud 23 internally has a surface 23a for external radial delimitation of the secondary stream 18. The shroud can alternatively be based on a composite having carbon fibers, such as known casings of module of blower.

In the radial direction R between the two veins 16, 18, there is provided an inter-vein compartment 44 in which several pieces of equipment/services 58 are arranged. internal radial delimitation of the secondary stream 18. Downstream of the fan 15, in the secondary stream 18, there is provided an annular row of stator vanes 30 centered on the axis 3, these stator vanes 30 also being called vanes OGV or outlet guide vanes.

Only one of these vanes 30 is visible in FIG. 1, that specific to the invention. Nevertheless, it is noted that conventionally, each of the vanes 30 of the annular row passes through the entire secondary stream 18 in the direction R, even if a slight inclination of these blades 30 in the X direction is possible, as shown in Figure 1.

Each stator vane 30 thus has a head end 31 connected to the outer shroud 23 of the casing 5 forming the nacelle, just as it has a root end 33 connected to the internal radial delimitation surface 40a of the secondary stream. 18. More specifically, this connection of the end of the foot 33 preferably takes place at an upstream part of the surface 40a defined by the outer shroud 40 of the inter-vein compartment 44, part at which it is formed an intermediate casing hub 32. In known manner, the stator vanes 30 are circumferentially spaced from each other, and make it possible to straighten the secondary flow after it has passed through the fan 15. In addition, these vanes 30 can also fulfill a function structural, ensuring the transfer of forces from the reducer 24 and the rolling bearings of the motor shafts and the fan hub, to the outer shroud 23. In other In other words, the entity 32 forms the hub of an intermediate casing, and the latter is completed by radial arms formed by the stator vanes 30, and also completed by the outer shroud 23 on which are fixed the head ends 31 of these dawns 30.

The inter-vein compartment 44 is also delimited by an inner shroud 42, configured to externally delimit the primary gas flow vein 16. The two ferrules 40, 42 extend downstream from the separation spout 26, which connects them. Downstream of the stator blades 30, a plurality of air discharge ducts 46 are provided, distributed around the axis 3. Each discharge duct 46 extends generally radially, possibly with an axial component going downstream , going from the inner shroud 42 to the outer shroud 40, so as to be able to communicate the primary stream 16 with the secondary stream 18. Each air discharge conduit 46 opens into the primary stream 16 through an orifice of inlet 48 equipped with a discharge valve VBV 50, the inlet orifice 48 being arranged axially between the low pressure compressor 4 and the high pressure compressor 6. Likewise, each discharge duct of air 46 opens into the secondary stream 18, through an outlet orifice 52 equipped with discharge fins 54.

The attachment mast 7 extends over a limited height in the radial direction R, also corresponding to the vertical direction Z in the area where this mast is located. Indeed, the mast 7 is conventionally arranged in a clock position at 12 o'clock, extending in length upstream in the direction X, from a lower portion of the wing 202, close to the front spar 208 and of the leading edge 210 of the wing 202.

The attachment pylon 7 comprises a downstream part 7a which extends forward from the wing 202, connecting along the outer ring 40, over a significant length of the latter. This downstream part 7a extends in the direction X up to near a trailing edge 35 of the casing 5 forming the nacelle. From this trailing edge 35, the mast continues continuously via an upstream part 7b which is completely housed in the secondary vein 18, still remaining connected full length to the outer shroud 40 of the inter-vein compartment 44. Thus, the mast 7 is closed radially inwards by the outer shroud 40, which it espouses continuously over an extended axial length in the direction X, which can go from the low pressure compressor 4 or even beyond that -ci upstream, to the low pressure turbine 12, or even beyond it downstream.

One of the particularities of the invention therefore resides in the reduced height of the upstream part 7b of the mast 7, in the radial direction R also corresponding here to the vertical direction Z. By reduced or partial height, it is understood that the part upstream 7b extends radially from surface 40a over a radial mast height "Hm", strictly less than a total radial height "Ht" of the secondary stream 18. All along this upstream part 7b of the mast, that -it is therefore never in contact with the surface 23a of outer radial delimitation of the secondary stream 18. Preferably, the radial height of the mast Hm locally represents 30 to 60% of the total radial height Ht of the secondary stream 18. More generally, the radial mast height Hm locally represents 20 to 70% of the total radial height of the secondary vein. This percentage is not necessarily identical all along the upstream part of the mast 7b, but on the contrary it can change locally, always preferably remaining within the range of values mentioned above. This change in percentage can be explained by a fairly variable radial height of the mast Hm along the upstream part 7b, whereas the total radial height of the vein Ht remains substantially constant, or slightly variable. In this regard, it is noted that the crest line 60 of the mast 7 is straight or substantially straight, preferably parallel or substantially parallel to the direction X. The radial height of the mast can be approximately 50% of the total radial height Ht of the secondary stream near the trailing edge 66 of the integrated blade 30. In addition, it may be of increasing height, for example with continuously variable curvature downstream. A second feature of the invention resides in the integration of one of the stator vanes 30 into the upstream part of the mast 7b. This is in fact the blade 30 located in the same clockwise position as that of the mast 7, and arranged axially upstream of the latter. Instead of having a discontinuity of material between the trailing edge of this blade 30, and the front end of the mast 7, provision is therefore made to integrate them into one another, thus resulting in an axial continuity of material between these two entities within the secondary stream 18. To do this, the integrated stator vane 30 comprises the following parts, succeeding each other radially from the inside out. It is first of all a foot part 62a integrated axially into the upstream part of the mast 7b, this foot part 62a comprising the end of the foot 33 connected to the delimitation surface 40a. Next, the integrated blade 30 comprises a transition part 62b, then a head part 62c ending with the head end 31 connected to the delimitation surface 23a. Thus, the specificity of this integrated blade 30 resides firstly in the root part 62a from which extends, axially downstream, the upstream part of the mast 7b. In other words, the foot part 62a has a fictitious trailing edge 64 which merges into the front end of the upstream part of the mast 7b, since no material discontinuity is observed between these two entities, according to direction X. At the level of the intrados and the extrados of this integrated assembly 62a, 7b, the continuity of material is achieved either by an aerodynamic wall in one piece, that is to say made of a single piece, or by the association of several walls presenting an acceptable aerodynamic junction, for example by covering with offsetting, or any other technique known in this field.

On the other hand, the integrated blade 30 has, radially outwards from the upstream part of the mast 7b, that is to say radially outwards from the foot part 62a, a trailing edge free 66 extending to the boundary surface 23a. The free trailing edge 66 thus corresponds to the trailing edge of the transition part 62b and of the head part 62c combined. The fictitious trailing edge 64 of the foot part 62a extends over the radial height Hm, while the free trailing edge 66 extends over a height corresponding to the differential between the heights Ht and Hm. In the first preferred embodiment shown in Figures 1 and 4 to 6, there is shown the fictitious trailing edge 64 of the foot part 62a with a substantial thickness in the direction Y, this thickness "E" referenced in the figure 6 being increasing going downstream in the direction of the upstream part of mast 7b. This transverse/lateral thickness E of the fictitious trailing edge 64 is much greater than that of the trailing edge 66 of the transition part 62b and of the leading part 62c. Indeed, the free trailing edge 66 of the head part 62c has a particularly thin conventional thickness "e", while the trailing edge 66 of the transition part 62b has a variable transverse thickness "e'", which increases going radially inwards so as to pass gradually from the value "e" to the value "E". To do this, two connecting spokes 68 can be respectively provided on the intrados side and on the extrados side of the transition part 62b, as is best visible in FIG. 6. This makes it possible to obtain a smooth transition between the thicknesses E e substantially different from the parts 62a, 62c, to limit the aerodynamic losses observed at this break in thickness. According to an alternative shown in Figures 7 to 9, the transition zone 62b has a profile identical or similar to that of the head part 62c, implying a reduced thickness "e" for its free trailing edge 66, identical or similar to the thickness of the trailing edge of the leading part 62c. This results in a sudden break in thickness between the foot part 62a and the transition part 62b, as is best seen in Figure 9. According to a second preferred embodiment represented in FIGS. 10 to 12, the transition part 62b has a chord “C” of greater length than that of the head part 62c. To do this, the transition part 62b is equipped with a trailing edge extension 72, for example of generally triangular shape and arranged so that the chord C has an increasing axial length an going radially from the inside towards the outside, that is to say from the head part 62c to the foot part 62a. The trailing edge extension 72 preferably has a transverse thickness that continuously decreases going downstream, as far as the free trailing edge 66 of the transition part 62b. A reduction in thickness is also observed going radially outwards, approaching the free trailing edge 66.

By thus locally increasing the length of the chord C within the transition part 62b, it is possible to keep a free trailing edge 66 of small thickness limiting the losses of base, while providing a more substantial blade thickness in upstream of this trailing edge 66, at the place where it joins radially with the thick fictitious trailing edge 64 of the foot part 62a. This advantageously makes it possible to limit the differential in transverse thickness between the parts 62a, 62b, and therefore to soften the radial connection, with the consequence of gains in terms of aerodynamic performance.

According to a third preferred embodiment represented in FIGS. 13 to 15, the transition part 62b comprises a free trailing edge 66 truncated, for example so as to form a notch 74 opening axially downstream. This notch 74 in the trailing edge 66 is preferably generally triangular in shape. It can extend into the head part 62c, as can be seen in FIG. 13. The shape retained for the notch 74 can be such that the free trailing edge 66 of the two successive parts 62b, 62c is substantially straight. , being inclined in the direction X as is also visible in FIG. 13, since its radially outer end is located further downstream than its radially inner end.

The notch 74 is made so as to truncate a downstream portion of the transition part 62b, so that the chord "C" of this part 62b has an increasing axial length going radially from the inside to the outside, c i.e. by going from the foot part 62b to the head part 62c. In other words, the chord length increases as one approaches the head part 62c, until finding a conventional chord length and identical or substantially identical to that of the other blades 30 of the annular row. This third preferred embodiment also provides a solution making it possible to soften the radial connection between the parts 62a, 62b, since it takes place at the place where the respective transverse thicknesses are the most similar, that is to say say in whole or in part upstream of the thick fictitious trailing edge 64 of the foot part 62a, as clearly shown by the alignment of Figures 14 and 15. With such a configuration, the radial transition in thickness advantageously proves to be more gentle, with here too gains in terms of aerodynamic performance.

Of course, various modifications can be made by those skilled in the art to the invention which has just been described, solely by way of non-limiting examples and the scope of which is defined by the appended claims. In particular, the different preferred embodiments described above can be combined with one another.

Claims

1. Propulsion assembly (200) for an aircraft comprising a turbomachine (1) equipped with a fan (15), an aerodynamic outer casing (5) forming a nacelle arranged around the fan, as well as an attachment strut (7) intended to ensure the attachment of the turbomachine (1) to a wing element (202) of the aircraft, the propulsion assembly having a primary flow path (16) for the gases, as well as a secondary flow path ( 18) for circulation of the gases delimited by an inner radial delimitation surface (40a) as well as by an outer radial delimitation surface (23a) formed by the aerodynamic outer casing (5), the turbomachine (1) further comprising a row ring of stator vanes (30) arranged in the secondary stream (18) downstream of the fan (15), each stator vane (30) extending through the secondary stream (18) presenting a leading end (31) connected to the aerodynamic outer casing (5), as well as q a foot end (33) connected to the internal radial delimitation surface (40a) of the secondary stream, the attachment pylon (7) comprising a part housed in the secondary stream, called the upstream part (7b), as well a downstream part (7a) arranged downstream of a trailing edge (35) of the aerodynamic outer casing (5), itself intended to be arranged entirely upstream of a leading edge (210) of the wing element (202), characterized in that the upstream part (7b) of the attachment pylon housed in the secondary section (18) extends radially from the internal radial delimitation surface (40a), over a radial mast height (Hm) strictly less than a total radial height (Ht) of the secondary stream (18), and in that the upstream part (7b) of the attachment mast (7) extends towards the downstream from a root part (62a) of one of the stator vanes (30).

2. Propulsion assembly according to claim 1, characterized in that the radial mast height (Hm) locally represents 20 to 70% of the total radial height (Ht) of the secondary stream (18).

3. Propulsion assembly according to claim 1 or 2, characterized in that the stator vane (30) integrated into the attachment pylon (7) comprises, radially outwards at from the upstream part (7b) of this mast, a free trailing edge (66) extending as far as the aerodynamic outer casing (5).

4. Propulsion assembly according to any one of the preceding claims, characterized in that the blade (30) integrated into the attachment pylon (7) comprises the following parts, succeeding one another radially from the inside outwards:

- the foot part (62a) integrated into the upstream part (7b) of the attachment mast;

- a transition part (62b); and

- a head part (62c).

5. Propulsion assembly according to claim 4, characterized in that the trailing edge (66) of the transition part (62b) has a transverse thickness (e') which increases going radially towards the foot part (62a) of the integrated vane (30).

6. Propulsion unit according to claim 4 or 5, characterized in that the transition part (62b) has a chord (C) of greater length than that of the head part (62c), or in that said transition part (62b) comprises a trailing edge (66) truncated so that the chord (C) of the transition part (62b) has an increasing length going from the foot part (62a) towards the head part (62c) .

7. Propulsion assembly according to any one of the preceding claims, characterized in that the ratio between the total axial length (Lt) of the aerodynamic outer casing (5), and the diameter (D) of the fan (15), is less than 1.25.

8. Propulsion assembly according to any one of the preceding claims, characterized in that the fan (15) comprises rotating fan blades (17) with variable pitch, and in that the aerodynamic outer casing (5) forming the nacelle is devoid of a thrust reverser system.

9. Aircraft part (100) comprising a propulsion assembly (200) according to any one of the preceding claims, as well as a wing element (202), the aerodynamic outer envelope (5) forming the nacelle extending entirely upstream of a leading edge (210) of the wing element (202).

10. Aircraft part according to the preceding claim, characterized in that an upper part of the envelope (5) is arranged axially opposite the leading edge (210) of the wing (202), and in what a vertex of the outside of the outer shell aerodynamic (5) extends higher than the wing element (202) considered in line with the connection of the attachment strut (7) with said wing element.

11. Aircraft (300) comprising at least one part (100) according to claim 9 or 10.