PROPULSION ASSEMBLY FOR AN AIRCRAFT COMPRISING A STATOR VANE INTEGRATED INTO AN UPSTREAM PART OF A MOUNTING PYLON OF REDUCED HEIGHT

A propulsion assembly for an aircraft comprising a dual-flow turbine engine equipped with a fan, an aerodynamic outer shroud acting as a nacelle as well as a mounting pylon, the propulsion assembly having a secondary flow path defined by an outer radial defining surface formed by the shroud, the turbine engine including stator vanes, and the mounting pylon comprising a part housed in the secondary flow path, referred to as upstream part. According to the invention, the upstream part of the pylon extends radially from the inner radial defining surface, along a radial pylon height strictly less than a total radial height of the secondary flow path, and the upstream part of the pylon extends in the downstream direction from a root part of one of the stator vanes.

TECHNICAL FIELD

The present invention relates to the field of propulsion assemblies for aircraft. It relates more specifically to propulsion assemblies comprising a nacelle of reduced length, referred to as “short nacelle”, such as that described in the document EP 2 628 919 A1.

PRIOR ART

In a propulsion assembly comprising a dual-flow turbine engine, a primary gas flow path is provided, as well as a secondary gas flow path defined radially outwardly by an aerodynamic outer shroud acting as a nacelle. The turbine engine generally comprises a fan drawing in an air mass which is subsequently split into a primary flow circulating in the primary flow path, and into a secondary flow circulating in the secondary flow path. 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 gas exhaust nozzle. In a known manner, the high-pressure turbine rotates the high-pressure compressor via a first shaft, referred to as high-pressure shaft, whereas the low-pressure turbine rotates the low-pressure compressor and the fan via a second shaft, referred to as low-pressure shaft. In order to improve the propulsive efficiency of the turbine engine and reduce the specific consumption thereof as well as the noise emitted by the fan, turbojet engines having an increased bypass ratio (corresponding to the ratio between the secondary (cold) flow rate, and the primary (hot) flow rate passing through the primary body) have been proposed.

To achieve such bypass ratios, the fan is uncoupled from the low-pressure turbine, thus making it possible to optimise the respective rotation speeds thereof independently. Usually, uncoupling is performed using a reduction gear such as an epicyclic gear mechanism, placed between the upstream end of the low-pressure shaft and the fan. The fan is then driven indirectly by the low-pressure shaft, via the gear mechanism and an additional shaft, referred to as fan shaft, which is fastened between the gear mechanism and the fan hub. This uncoupling thus makes it possible to reduce the rotation speed and the fan pressure ratio, and increase the power extracted by the low-pressure turbine. Thanks to the gear mechanism, the low-pressure shaft can thus run at higher rotation speeds than conventional turbojet engines.

Also in conventional embodiments, the turbine engine is equipped, downstream from the fan, with an annular row of stator vanes (also referred to as outlet guide vanes, or OGV vanes). These stator vanes are located in the cold part of the turbine engine, in the secondary flow path. They are essentially intended to straighten the cold air flow from the fan vanes.

In short-nacelle propulsion assembly configurations, the mounting pylon of the turbine engine may be required to partially enter the secondary flow path, downstream from the stator vanes, hence generating aerodynamic losses by friction, with a negative impact on the overall efficiency of this propulsion assembly.

DESCRIPTION OF THE INVENTION

To address the drawback mentioned above, in relation to the embodiments of the prior art, the invention firstly relates to a propulsion assembly for an aircraft comprising a dual-flow turbine engine equipped with a fan, an aerodynamic outer shroud acting as a nacelle arranged around the fan, as well as a mounting pylon for fastening the turbine engine to a wing element of the aircraft, the propulsion assembly having a primary gas flow path, as well as a secondary gas flow path defined by an inner radial defining surface as well by an outer radial defining surface formed by the aerodynamic outer shroud, the turbine engine further including an annular row of stator vanes arranged in the secondary flow path downstream from the fan, each stator vane extending through the secondary flow path while having a tip end connected to the aerodynamic outer shroud, as well as a root end connected to the inner radial defining surface of the secondary flow path, the mounting pylon comprising a part housed in the secondary flow path, referred to as upstream part, as well as a downstream part arranged downstream from a trailing edge of the aerodynamic outer shroud, in turn intended to be arranged entirely upstream from a leading edge of the wing element.

According to the invention, the upstream part of the mounting pylon housed in the secondary flow path extends radially from the inner radial defining surface, along a radial pylon height strictly less than a total radial height of the secondary flow path, and furthermore, the upstream part of the mounting pylon extends in the downstream direction from a root part of one of the stator vanes.

The invention thus provides a mounting pylon which only extends along a part of the radial height of the secondary flow path wherein it enters, in order to advantageously limit aerodynamic losses by friction in the same secondary flow path. As an additional measure to the above, the invention provides the integration of the root part of one of the stator vanes with the upstream part of this mounting pylon, in order to form an aerodynamic continuity between these two parts, in the axial direction. Such an integration makes it possible to minimise the aerodynamic losses even further, and also reduce transversal static pressure heterogeneities (distortion), rising from the downstream direction to the fan.

The combination of these measures makes it possible overall to improve the propulsive efficiency of the turbine engine.

Preferably, the invention provides at least any one of the following optional features, considered separately or in combination.

Preferably, the radial pylon height represents locally for example 20 to 70% of the total radial height of the secondary flow path, and more particularly 30 to 60% of the total radial height of the secondary flow path. This percentage can obviously vary along the upstream part of the pylon, and therefore not remain constant.

Preferably, the stator vane integrated in the mounting pylon includes, radially outwards from the upstream part of this pylon, a free trailing edge extending to the aerodynamic outer shroud. This free trailing edge thus extends radially along a part of the height of the secondary flow path complementary with the radial height of the pylon.

Preferably, the vane integrated in the mounting pylon includes the following parts, in radial succession from the inside outwards:the root part integrated in the upstream part of the mounting pylon;a transition part; anda tip part.

According to a preferred embodiment of the invention, the trailing edge of the transition part has a transverse thickness which increases moving radially towards the root part of the integrated vane. This makes it possible to obtain a smooth transition between the usually thin thickness of the tip part of the integrated stator vane, and the much more substantial thickness of the imaginary trailing edge of the root part of this vane, this trailing edge merging in the front end of the mounting pylon.

Alternatively, an abrupt break in thickness could be provided in the radial direction, between the root part of the integrated vane and the transition part thereof, the thickness of which could then be identical or similar to that of the tip part of the vane. According to a second preferred embodiment of the invention, the transition part has a chord of greater length than that of the tip part. By increasing the length of the chord locally, it is possible to retain a thin trailing edge limiting base losses, while providing a more substantial thickness upstream from this trailing edge, at the position where it is radially connected with the thick imaginary trailing edge of the root part of this vane. This makes it possible advantageously to limit the thickness differential between the root part and the transition part of the integrated vane, and therefore soften the connection resulting in gains in terms of aerodynamic performance.

According to a third preferred embodiment of the invention, said transition part comprises a truncated trailing edge such that the chord of the transition part has an increasing length from the root part towards the tip part. Here again, the proposed solution makes it possible to soften the connection between the root part and the transition part of the integrated vane, since this connection is made at the position where the respective thicknesses are most similar, i.e. fully or partially upstream from the thick imaginary trailing edge of the root part of this vane. The radial thickness transition proves to be advantageously softer, with here also gains in terms of aerodynamic performance. Preferably, the ratio between the total axial length of the aerodynamic outer shroud, and the diameter of the fan, is less than 1.25. This ratio illustrates a so-called “short” nacelle design, generally associated with the fact that it does not integrate a thrust reversal system. In such a short nacelle, intended to be located axially facing and upstream from the leading edge of the wing element, the thrust reversal system is integrated in the fan which has for this purpose variable pitch rotary fans. This architecture wherein the thrust reversal function is fulfilled by the fan is known as “VPF (Variable Pitch Fan) architecture”. The invention also relates to an aircraft part comprising such a propulsion assembly, as well as a wing element, the aerodynamic outer shroud acting as the nacelle extending entirely upstream from a leading edge of the wing element.

Preferably, an upper part of the shroud is located arranged axially facing the leading edge of the wing, and an apex of the outside of the aerodynamic outer shroud extends higher than the wing element considered in line with the connection of the mounting pylon with said wing element.

Finally, the invention also relates to an aircraft including at least one such part, and preferably two parts respectively integrating the two wings of the aircraft.

Other advantages and features of the invention will appear in the non-limiting detailed description hereinbelow.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

With reference toFIGS.1and2, a part100of an aircraft is shown, comprising a propulsion assembly200as well as a wing element202, here an aircraft wing. Preferably, two parts100are arranged laterally on either side of the fuselage204of the aircraft300shown inFIG.3(only one of the two propulsion assemblies200being represented in thisFIG.3).

The propulsion assembly200includes a dual-flow and dual-body turbojet engine1, an aerodynamic outer shroud5acting as a nacelle, as well as a mounting pylon7for mounting the turbojet engine1on the wing202. In the figures, the mounting pylon7is only represented with the outer contour thereof formed by one or more aerodynamic fairings. Inside these fairings, a so-called primary structure (not shown) is conventionally provided, for transferring the loads between the turbojet engine1and the wing202. More specifically, the primary structure of the pylon is generally fastened to a front wing spar208. In addition to containing the primary structure, the fairings of the mounting pylon7integrate a number of conventional elements connecting the engine to the aircraft, such as ducts, heat exchangers, electric cables, mechanical drive shafts, structural parts of the engine suspension system, etc.

The dual-flow and dual-body turbojet engine1has a high bypass ratio, for example greater than 15. Consequently, the outer diameter thereof is high, and for this reason, it is arranged in a substantially raised position in relation to the wing202carrying it, so as to retain a sufficient ground clearance despite the large diameter. As seen inFIGS.1to3, the outer shroud acting as a nacelle5is located entirely upstream in relation to a leading edge210of the wing202. In this regard, it is noted that the terms “upstream” and “downstream” are considered along a main direction14of gas flow inside the turbojet engine, when the latter is in the normal propulsion configuration. The terms “front” and “rear” are used in relation to a direction opposite the main gas flow direction14.

An upper part of the shroud5is arranged axially facing the leading edge210of the wing202, which conveys the raised nature of the short-nacelle propulsion assembly200. Furthermore, an apex of the outside of the aerodynamic outer shroud5extends higher than the wing element202considered in line with the connection of the mounting pylon7with said wing element.

The turbojet engine1conventionally includes a gas generator2on either side of which are arranged a low-pressure compressor4and a low-pressure turbine12, this gas generator2comprising a high-pressure compressor6, a combustion chamber8and a high-pressure turbine10. The low-pressure compressor4and the low-pressure turbine12form a low-pressure body, and are connected to one another by a low-pressure shaft11centred on a longitudinal axis3of the turbojet engine. Similarly, the high-pressure compressor6and the high-pressure turbine10form a high-pressure body, and are connected to one another by a high-pressure shaft13also centred on the axis3and arranged about the low-pressure shaft11.

The turbojet engine1further includes, upstream from the gas generator2and the low-pressure compressor4, a single fan15which is here arranged directly behind an air inlet cone of the engine. The fan15includes a ring of rotary fan vanes17about the axis3, this ring being surrounded by a fan casing9. The fan vanes17are variable-pitch, i.e. the incidence thereof can be controlled by a control mechanism20arranged at least partially in the inlet cone, and designed to pivot these vanes17about the respective longitudinal axes22thereof. This control mechanism20, of known mechanical, electrical, hydraulic, and/or pneumatic type design, is itself controlled by an electronic control unit (not shown), which makes it possible to order the value of the pitch angles of the vanes17according to the needs encountered, particularly to perform the thrust reversal function. As mentioned above, reference is made here to a propulsion assembly200wherein the thrust reversal function are actually integrated in the fan15, and not in the outer shroud acting as a nacelle5, as is more commonly encountered. This shroud5can consequently have a short length compared to the diameter of the fan15. Preferably, the ratio between the total axial length “Lt” of the shroud5, and the diameter “D” of the fan15, is less than 1.25.

Hereinafter in the description of the propulsion assembly200, reference is made to the longitudinal direction X parallel with the axis3of the turbojet engine, and also known as axial direction, to the transverse direction Y also known as lateral direction, and finally to the vertical direction Z also known as the height direction, these three directions X, Y and Z being mutually orthogonal. Reference is also made to the radial direction R, to be taken into consideration in relation to the axis3.

The VPF type fan15, is not driven directly by the low-pressure shaft11, but merely driven indirectly by this shaft, via a gear mechanism24, which makes it possible to rotate with a slower speed. Nevertheless, a direct drive solution of the fan15, by the low-pressure shaft11, falls within the scope of the invention.

Furthermore, the turbojet engine1defines a primary gas flow path16, intended to be flowed through by a primary flow16a, as well as a secondary gas flow path18, intended to be flowed through by a secondary flow18alocated radially outwardly from the primary flow. The flow of the fan15is thus split at a flow separation nozzle26.

As known to a person skilled in the art, the secondary flow path18is defined radially outwardly partially by an outer shell23, integrated in the shroud5. The shroud23is preferably metallic, and extends to the rear of the fan casing9. More specifically, the shroud23has internally a radial outer defining surface23aof the secondary flow path18. The shroud can alternatively be composite-based having carbon fibres, like known fan module casings.

In the radial direction R between the two flow paths16,18, an inter-flow path compartment44is provided, wherein several items of equipment/utilities58are arranged. This compartment44is formed partially by an outer shell40, having externally a radial inner defining surface40aof the secondary flow path18.

Downstream from the fan15, in the secondary flow path18, an annular row of stator vanes30centred on the axis3is provided, these stator vanes30also being referred to as OGV vanes or outlet guide vanes.

Only one of these vanes30is visible inFIG.1, that specific to the invention. Nevertheless, it is noted that conventionally, each of the vanes30of the annular row passes through the entire secondary flow path18along the direction R, even if a slight inclination of these vanes30in the direction X is possible, as shown in theFIG.1. Each stator vane30thus has a tip end31connected to the outer shell23of the shroud5acting as a nacelle, in the same way as it includes a root end33connected to the radial inner defining surface40aof the secondary flow path18. More specifically, this connection of the root end33is preferably performed at an upstream part of the surface40adefined by the outer shell40of the inter-flow path compartment44, at which an intermediate casing hub32is formed.

In a known manner, these stator vanes30are spaced circumferentially apart from each other, and make it possible to rectify the secondary flow after it flows through the fan15. In addition, these vanes30can also fulfil a structural function, by transferring the loads from the reduction gear24and the bearings of the motor shafts and the fan hub, to the outer shell23. In other words, the entity32forms the hub of an intermediate casing, and the latter is supplemented by radial arms formed by the stator vanes30, and also supplemented by the outer shell23whereon the tip ends31of these vanes30are fastened.

The inter-flow path compartment44is also defined by the inner shell42, configured to outwardly define the primary gas flow path16. The two shells40,42extend in the downstream direction from the separation nozzle26, which connects them. Downstream from the stator guide vanes30, a plurality of air discharge ducts46, distributed about the axis3, is provided. Each discharge duct46extends globally radially, optionally with an axial component extending in the downstream direction, from the inner shell42to the outer shell40, so as to be able to connect the primary flow path16with the secondary flow path18. Each air discharge duct46opens into the primary flow path16through an inlet orifice48equipped with a VBV discharge valve50, the inlet orifice48being arranged axially between the low-pressure compressor4and the high-pressure compressor6. Similarly, each air discharge duct46opens into the secondary flow path18, through an outlet orifice52equipped with discharge fins54.

The mounting pylon7extends along a limited height along the radial direction R, also corresponding to the vertical direction Z in the area where this pylon is located. Indeed, the pylon7is conventionally arranged in a 12 o'clock position, extending lengthwise in the upstream direction along the direction X, from a lower portion of the wing202, close to the front wing spar208and the leading edge210of the wing202.

The mounting pylon7includes a downstream part7awhich extends to the front from the wing202, connecting along the outer shell40, along a substantial length thereof. This downstream part7aextends along the direction X up to close to a trailing edge35of the shroud5acting as a nacelle. From this trailing edge35, the pylon is extended continuously by an upstream part7bwhich is entirely located housed in the secondary flow path18, once again while remaining connected along the entirety thereof on the outer shell40of the inter-flow path compartment44. Thus, the pylon7is closed radially inward by the outer shell40, which it moulds continuously along an axial length extended along the direction X, capable of going from the low-pressure compressor4or beyond in the upstream direction, to the low-pressure turbine12, or beyond in the downstream direction.

One of the specificities of the invention therefore lies in the reduced height of the upstream part7bof the pylon7, along the radial direction R also corresponding here to the vertical direction Z. By reduced or partial height, it is understood that the upstream part7bextends radially from the surface40aalong a radial pylon height “Hm”, strictly less than a total radial height “Ht” of the secondary flow path18. All along this upstream part7bof the pylon, it is therefore never in contact with the outer radial defining surface23aof the secondary flow path18. Preferably, the radial pylon height Hm locally represents 30 to 60% of the total radial height Ht of the secondary flow path18. More generally, the radial pylon height Hm locally represents 20 to 70% of the total radial height of the secondary flow path. This percentage is not necessarily identical all along the upstream pylon part7b, but it can on the other hand vary locally, once again while remaining preferably within the range of values mentioned above. This percentage variation can be explained by a relatively variable radial pylon height Hm along the upstream part7b, whereas the total radial flow path height Ht remains substantially constant, or relatively invariable. In this regard, it is noted that the crest line60of the pylon7is straight or substantially straight, preferably parallel or substantially parallel with the direction X. The radial height of the pylon can be approximately 50% of the total radial height Ht of the secondary flow path in the vicinity of the trailing edge66of the integrated vane30. Furthermore, it can be of increasing height, for example with continually variable curvature in the downstream direction.

A second specificity of the invention lies in the integration of one of the stator vanes30in the upstream pylon part7b. Indeed, it consists of the vane30located in the same hour position as that of the pylon7, and arranged axially upstream therefrom. Instead of having a material discontinuity between the trailing edge of this vane30, and the front end of the pylon7, it is therefore provided to integrate them into each other, thus giving rise to axial material continuity between these two entities within the secondary flow path18. For this purpose, the integrated stator vane30includes the following parts, in radial succession from the inside outwards. It consists firstly of a root part62aintegrated axially in the upstream pylon part7b, this root part62aincluding the root end33connected to the defining surface40a. Then, the integrated vane30includes a transition part62b, then a tip part62cending with the tip end31connected to the defining surface23a.

Thus, the specificity of this integrated vane30lies firstly in the root part62afrom which the upstream pylon part7bextends axially in the downstream direction. In other words, the root part62ahas an imaginary trailing edge64which merges in the front end of the upstream pylon part7b, since no material discontinuity is observed between these two entities, along the direction X. At the lower surface and the upper surface of this integrated assembly62a,7b, the material continuity is produced either by a one-piece aerodynamic wall, i.e. made of a single piece, or by the association of several walls having an acceptable aerodynamic junction, for example by swaged overlay, or any other technique known in this field.

On the other hand, the integrated vane30has, radially outwards from the upstream pylon part7b, i.e. radially outwards from the root part62a, a free trailing edge66extending to the defining surface23a. The free trailing edge66thus corresponds to the trailing edge of the transition part62band the tip part62ctogether. The imaginary trailing edge64of the root part62aextends along the radial height Hm, whereas the free trailing edge66extends along a height corresponding to the differential between the heights Ht and Hm.

In the first preferred embodiment shown inFIGS.1and4to6, the imaginary trailing edge64of the root part62ais represented with a substantial thickness along the direction Y, this thickness “E” referenced inFIG.6increasing in the downstream direction towards the upstream pylon part7b. This transverse/lateral thickness E of the imaginary trailing edge64is substantially greater than that of the trailing edge66of the transition part62band the tip part62c. Indeed, the free trailing edge66of the tip part62chas a particularly thin conventional thickness “e”, whereas the trailing edge66of the transition part62bhas a variable transverse thickness “e′”, which increases radially inwards progressively from the value “e” to the value “E”. For this purpose, two connection radii68can be respectively provided on the lower surface side and the upper surface side of the transition part62b, as seen more clearly inFIG.6. This makes it possible to obtain a soft transition between the substantially different thicknesses E,e of the parts62a,62c, to limit the aerodynamic losses observed at this break in thickness. According to an alternative shown inFIGS.7to9, the transition zone62bhas a profile identical or similar to that of the tip part62c, involving a reduced thickness “e” for the free trailing edge66thereof, identical or similar to the thickness of the trailing edge of the tip part62c. This results in an abrupt break in thickness between the root part62aand the transition part62b, as seen more clearly inFIG.9.

According to a second preferred embodiment represented inFIGS.10to12, the transition part62bhas a chord “C” of greater length than that of the tip part62c. For this purpose, the transition part62bis equipped with a trailing edge extension72, for example of general triangular shape and arranged in such a way that the chord C has an increasing axial length moving radially from the inside outwards, i.e. from the tip part62cto the root part62a. The trailing edge extension72preferably has a transverse thickness decreasing continually in the downstream direction, up to the free trailing edge66of the transition part62b. A reduction in thickness is also observed on moving radially outwards, closer to the free trailing edge66.

By thus increasing the length of the chord C locally in the transition part62b, it is possible to retain a thin free trailing edge66limiting base losses, while providing a more substantial thickness upstream from this trailing edge66, at the position where it is radially connected with the thick imaginary trailing edge64of the root part62a. This makes it possible advantageously to limit the transverse thickness differential between the part62a,62b, and therefore soften the radial connection, resulting in gains in terms of aerodynamic performance.

According to a third preferred embodiment shown inFIGS.13to15, the transition part62bcomprises a truncated free trailing edge66, for example so as to form a notch74opening axially in the downstream direction. This notch74in the trailing edge66is preferably of general triangular shape. It can extend up to into the tip part62c, as seen inFIG.13. The shape selected for the notch74can be such that the free trailing edge66of the two successive parts62b,62cis substantially straight, while being inclined along the direction X as also seen inFIG.13, since the radially outer end thereof is located more downstream than the radially inner end thereof.

The notch74is embodied so as to truncate a downstream portion of the transition part62b, such that the chord “C” of this part62bhas an increasing axial length moving radially from the inside outwards, i.e. moving from the root part62bto the tip part62c. In other words, the chord length increases on moving closer to the tip part62c, until a conventional chord length identical or substantially identical to that of the other vanes30of the annular row is obtained.

This third preferred embodiment also provides a solution making it possible to soften the radial connection between the parts62a,62b, since it is performed at the position where the respective transverse thicknesses are the most similar, i.e. fully or partially upstream from the thick imaginary trailing edge64of the root part62a, as shown clearly by the alignment ofFIGS.14and15. With such a configuration, the radial thickness transition proves to be advantageously softer, with here also gains in terms of aerodynamic performance.

Of course, various modifications may be made by the person skilled in the art to the invention as described, by way of non-limiting examples only, the scope of which is defined by the appended claims. In particular, the different preferred embodiments described above can be combined with each other.