Assembly for an aircraft propulsion unit designed to facilitate the installation of elements close to one another

Abstract

The invention relates to an assembly (15) for an aircraft propulsion unit, comprising - a radial delimiting structure (34) for the circulation of a gas flow (12b), the structure (34) comprising a radial delimiting surface (34a) oriented in a first sense (S1) of a radial direction (R); - an equipment item (45) projecting from the radial delimiting surface (34a) in the first sense (S1); - an aerodynamic profile (30) also projecting from the radial delimiting surface (34a) in the first sense (S1), this profile (30) being arranged upstream of the equipment item (45). According to the invention, the radial delimiting surface (34a) comprises, in the first sense (S1), a radial boss (52) from which at least part of the aerodynamic profile (30) extends.

Description

[0001] ASSEMBLY FOR AN AIRCRAFT PROPULSION UNIT, DESIGNED TO FACILITATE THE INSTALLATION OF ELEMENTS IN PROXIMITY TO EACH OTHER

[0002] TECHNICAL FIELD

[0003] The present invention relates to the field of aircraft propulsion assemblies, comprising aerodynamic profiles and equipment intended to be located close to each other, in a zone of passage of a gas flow of this assembly. The invention applies to propulsion assemblies comprising turbomachines of any type, such as turbojets and turboprops, with single or double fan or propeller, ducted or unducted.

[0004] STATE OF PRIOR ART

[0005] In the propulsion systems equipping aircraft, several gas flows are observed circulating in and outside this system. Indeed, in addition to the external air flow circulating around the system, there are also one or more gas flows circulating inside the system, such as for example a primary flow and a secondary flow in the case of a double-flow turbojet.

[0006] Radial boundary structures are arranged within the propulsion assembly, in order to channel these gas flows, which generally flow in a direction from upstream to downstream of the assembly.

[0007] The evolution of propulsion systems and their turbomachinery leads to an increase in the need for equipment to control their operation and increase their performance. In addition, due to the ever-increasing need for compactness, the space available for the installation of this equipment is becoming increasingly reduced.

[0008] Some of these devices are installed to cooperate with circulating gas flows, for example the secondary flow of a turbojet engine. In this case, the devices are, for example, of the air intake scoop type, or heat exchangers of the ACOC (Air-Cooled Oil-cooler) type. To ensure their proper functioning, these devices form protrusions in the secondary flow of the propulsion system.

[0009] However, the presence of this protruding equipment is usually added to the presence of aerodynamic profiles crossed by the associated gas flow. In the case of a double-flow turbojet, in the secondary vein, these are outlet guide vanes, or OGVs (from the English "Outlet Guide Vanes"), crossed by the secondary flow.

[0010] To meet the aforementioned compactness requirements, the equipment may be brought axially closer to the trailing edge of the aerodynamic profiles. However, the protrusions formed by the equipment may produce detrimental effects on the aerodynamic profiles located nearby, upstream, due to pressure build-up. Indeed, harmful aerodynamic interactions may occur, due to pressure build-up caused by the equipment projecting into the flow. For example, equipment placed nearby downstream of an outlet guide vane may cause a static pressure disturbance on the latter, detrimental to its aerodynamic operation.

[0011] There is therefore still a need to improve the design of propulsion systems, aimed at limiting the aforementioned harmful aerodynamic interactions between the aerodynamic profiles and the equipment, while maintaining satisfactory overall compactness.

[0012] STATEMENT OF THE INVENTION

[0013] To meet this need, the invention relates to an assembly for an aircraft propulsion unit, comprising the characteristics of claim 1.

[0014] The proposed invention advantageously allows the equipment to be installed in axial proximity downstream of the aerodynamic profile, while greatly limiting the harmful aerodynamic interactions likely to occur between these two elements. Indeed, the boss makes it possible to locally modify the static pressure at the level of the aerodynamic profile, so as to compensate for the pressure rise created downstream by the equipment. In other words, the local performance of the aerodynamic profile likely to be impacted by the pressure rise can thus be restored, thanks to the local increase in the speed of the flow close to the delimitation surface. This in fact induces a reduction in the static pressure, which was increased by the presence of the obstacle formed by the equipment downstream.

[0015] The boss specific to the invention thus makes it possible to locally correct the flow of the gas flow close to the radial delimitation surface, thus restoring the normal operation of the aerodynamic profile, while authorizing the installation of one or more pieces of equipment in its downstream proximity, in the secondary vein corresponding to the vein

[0016] This contributes to obtaining a compact propulsion unit, reducing its overall mass. The invention is therefore the result of technological research aimed at significantly improving aircraft performance and, in this sense, contributes to reducing the environmental impact of these aircraft (decarbonization).

[0017] The invention furthermore preferably provides at least any one of the following optional features, taken individually or in combination.

[0018] Preferably, the aerodynamic profile is arranged so that at least one imaginary longitudinal plane of the assembly, passing through a longitudinal central axis of this assembly, crosses both the aerodynamic profile and the equipment. This translates into a sort of overlap zone between the profile and the equipment, in the circumferential direction. Alternatively, the equipment could for example be arranged axially opposite and downstream of an inter-profile circumferential space, and limited in circumferential length so as not to extend beyond this space, in this circumferential direction. The advantage of such an arrangement lies in limiting or eradicating pressure rises on the flow circulating in the inter-profile space, and therefore participating in the restoration of the normal operation of the aerodynamic profiles, such as OGVs.Preferably, in section in said imaginary longitudinal plane of the assembly, the radial delimitation surface has, at its radial boss, a first inflection point as well as a second inflection point located downstream of the first inflection point, the first inflection point being preferentially located upstream or at the level of the aerodynamic profile, and the second inflection point being preferentially located between the aerodynamic profile and the equipment.

[0019] Preferably, the boss has a peak located, in a longitudinal direction of the assembly, at the level of the aerodynamic profile, for example by being centered relative to the latter.

[0020] Preferably, the equipment extends, in the first direction of the radial direction, beyond the top of the boss.

[0021] Preferably, the equipment is a scoop or a heat exchanger, even if other types of equipment can be envisaged, without departing from the scope of the invention.

[0022] Preferably, the aerodynamic profile is, for example, an outlet guide vane, a compressor rectifier vane, or a turbine distributor vane.

[0023] Within the assembly, the following dimensions are defined, in any fictitious longitudinal plane passing through the aerodynamic profile, the equipment, and also through the top of the radial boss, or their projections in this plane, in the circumferential direction:

[0024] - H: total radial height of the aerodynamic profile, at its trailing edge;

[0025] - h: height of the boss, between a fictitious base of the latter and its summit, according to a direction of the height of this radial boss;

[0026] - Lt: length of the boss, between an upstream end and a downstream end of this boss, corresponding to opposite ends of its fictitious base;

[0027] - Ls: length of a front part of the boss, defined between its upstream end, and the orthogonal projection of the top on the fictitious base of the boss;

[0028] - Dbf: distance between the upstream end of the boss, and the orthogonal projection, on the fictitious base, of the trailing edge of the aerodynamic profile,

[0029] - Dh: radial height differential between a front end of the equipment and the top of the boss;

[0030] - Ea: axial distance between the top of the boss and the front end of the equipment.

[0031] In addition, the assembly is preferably carried out in such a way as to meet at least one of the following parameters:

[0032] - h / H between 0.0001 and 0.2 - Lt / H between 0.00001 and 0.4

[0033] - Lt / h between 0.001 and 50

[0034] - Dbf / H between 0.0001 and 2

[0035] - Ls / Lt between 0.1 and 0.999

[0036] - Dh / Ea between -10 and 10

[0037] Preferably, the boss extends continuously in the circumferential direction, preferably along the entire length of the radial delimiting structure. It may then be an annular boss extending over 360°, or interrupted circumferentially at the same locations as the radial delimiting structure with which it is fitted, for example at the passage of the mast, beams, etc. Whatever the design chosen, this boss then extends over a high angular sector, preferably being centered on the longitudinal central axis of the assembly. The boss can thus cooperate with several aerodynamic profiles spaced circumferentially from one another. Alternatively, several bosses could be provided spaced circumferentially from one another, each cooperating with one or more aerodynamic profiles. This embodiment may also be considered as a non-continuous boss, interrupted circumferentially.

[0038] The invention also relates to an aircraft propulsion assembly comprising at least one such assembly. This assembly also comprises a turbomachine, preferably a dual-flow turbojet engine, said radial delimitation structure of the assembly being an internal radial delimitation structure of the secondary vein of the turbojet engine, in which a secondary flow is intended to circulate.

[0039] Other advantages and characteristics of the invention will appear in the detailed non-limiting description below.

[0040] BRIEF DESCRIPTION OF THE DRAWINGS

[0041] This description will be made with regard to the attached drawings, among which;

[0042] [Fig. 1] represents a schematic longitudinal sectional view of an aircraft propulsion assembly; [Fig. 2] represents a more detailed longitudinal sectional half-view of an assembly for the propulsion assembly shown in the preceding figure, the assembly being in the form of a preferred embodiment of the invention, and the sectional plane corresponding to a fictitious longitudinal plane of the assembly; and

[0043] [Fig. 3] shows a view similar to the previous one, on which the references of several dimensions within the assembly have been entered.

[0044] DETAILED DISCLOSURE OF PREFERRED EMBODIMENTS

[0045] Referring firstly to FIG. 1, there is shown a propulsion unit 100 for an aircraft, comprising a turbomachine 1 and a nacelle 9 surrounding the turbomachine.

[0046] In the preferred embodiment of the invention which will be described, the turbomachine preferably corresponds to a double-flow, double-spool turbojet engine, comprising a single ducted fan. However, it could be a turbomachine of another type, for example a turboprop. More generally, it can be all types of turbojet engines and turboprops, with single or double fan or propeller, ducted or unducted.

[0047] Subsequently, the terms “upstream” and “downstream” are defined relative to a general direction 5 of flow of the gases through the propulsion unit 1 when the latter generates direct thrust, this direction being parallel or substantially parallel to the axis 2. These terms “upstream” and “downstream” could respectively be substituted by the terms “front” and “rear”, with the same meaning. In addition, the propulsion unit 1 is represented in a reference frame formed by three mutually orthogonal directions, namely the longitudinal direction L parallel to the axis 2 which will be defined below, the circumferential direction C, and the radial direction R.

[0048] The turbojet 1 has the longitudinal central axis 2 around which its various components extend. It comprises, from upstream to downstream along the main direction 5 of flow of the gases through this turbomachine, a fan 3, a low-pressure compressor 4, a high-pressure compressor 6, a combustion chamber 11, a high-pressure turbine 7 and a low-pressure turbine 8. The fan 3 can be driven directly by a low-pressure body comprising the compressor 4 and the turbine 8, or else be driven indirectly by a reducer (not shown).

[0049] Conventionally, after passing through the fan 3, the air divides into a central primary flow 12a and a secondary air flow 12b which surrounds the primary flow. The primary flow 12a flows in a main gas circulation vein 14a passing through the compressors 4, 6, the combustion chamber 11 and the turbines 7, 8. The secondary flow 12b flows in a secondary air vein 14b delimited radially outwards by a casing, surrounded by the nacelle 9, this secondary vein therefore being radially around the primary vein. More specifically, the casing comprises a fan casing 20 which surrounds the fan blades, this casing 20 being extended downstream by an outer shroud 22 of an intermediate casing 24. This intermediate casing 24 comprises a hub 26 centered on the axis 2, and which may include a flow separation nozzle 28.The intermediate casing is completed by radial arms 30, which form outlet guide vanes, making it possible to straighten the secondary air flow 12b and conventionally called OGV, as indicated previously. The arms 30 thus connect the outer shell 22 of the intermediate casing, to its hub 26, at the right of the low pressure compressor 4.

[0050] The turbojet engine has a bypass ratio, or BPR, of the order of three to forty, and more precisely of the order of three to eighteen in the present case of a ducted fan corresponding to that of the preferred embodiment described. In the case of one or more unducted fans / propellers, this bypass ratio is rather of the order of eighteen to forty.

[0051] Whether the fan(s) / propeller(s) are ducted or unducted, their diameter is preferably of the order of 0.5 to 5 m, while the compression ratio is preferably of the order of 1 to 1.8.

[0052] In the case of a turbomachine equipped with a reducer, with the propeller(s) / fan(s) shrouded or unshrouded, the reduction ratio can be between 1.1 and 20.

[0053] The outer shroud 22 forms an outer radial delimitation structure of the secondary air stream 14b. It is extended downstream by another annular delimitation structure 31 of this type, integrated into a mobile cowl 32 of the thrust reverser. This structure 31 is also called OFS (from the English "Outer Fixed Structure"). It is located radially opposite an annular inner radial delimitation structure 34 of the stream 14b, also called IFS (from the English "Inner Fixed Structure"), and initiated from the upstream by the slat 28. The inner radial delimitation structure 34 comprises the hub 26 of the intermediate casing.

[0054] This structure 34 for radial delimitation of the vein 14b, comprises an internal radial delimitation surface 34a oriented in a first direction SI of the radial direction R, corresponding to the direction going from the inside to the outside of the assembly 1, in relation to the axis 2. It is therefore this surface 34a which delimits the secondary vein 14b, radially towards the inside.

[0055] Each of the fixed outlet guide vanes 30 projects from the inner radial delimiting surface 34a, in the first direction SI of the radial direction R, towards the outer shell 22 of the intermediate casing.

[0056] The blades 30, forming aerodynamic profiles crossed by the secondary flow 12b in the secondary vein 14b, form with the radial delimiting structure 34 an assembly 15 specific to the present invention. This assembly 15 is completed by equipment 45 of the propulsion unit 1, or preferably by several pieces of equipment.

[0057] In the preferred embodiment shown, only one piece of equipment 45 will be described, arranged axially in close proximity to one of the blades 30, downstream of the latter. This is, for example, an air sampling scoop, or even a heat exchanger of the ACOC (Air-Cooled Oil-cooler) type. In all cases, the equipment 45 is intended to cooperate with the secondary flow 12b, such that this equipment projects from the inner radial delimiting surface 34a, in the first direction SI of the radial direction R. The equipment 45 thus forms a radial protuberance in the secondary vein 14b, located in the downstream extension of the blade 30 in relation to a main direction of circulation of the secondary flow 12b, corresponding to the general direction 5 of flow of the gases through the propulsion unit 1.

[0058] As mentioned previously, the equipment 45 is here at least partly aligned with one of the blades 30 in the longitudinal direction L. Consequently, there is at least one imaginary longitudinal plane P of the assembly 15, passing through the axis 2 also corresponding to the longitudinal central axis of this assembly, and which passes through both the blade 30 and the equipment 45. One of these planes P corresponds to the section plane of FIGS. 2 and 3. Alternatively, the equipment 45 could for example be arranged axially opposite and downstream of an inter-OGV circumferential space, and limited in circumferential length so as not to extend beyond this space, in this circumferential direction C. This makes it possible to limit or eradicate pressure rises on the flow circulating in the inter-OGV space, and therefore to participate in restoring normal operation of the blades 30.Such a principle will now be described for the preferred case where the equipment 45 and one of the blades 30 are crossed by the same imaginary longitudinal plane P of the assembly. But in the case of a circumferential offset between these elements, a similar arrangement is preferentially observed in a imaginary longitudinal plane P of the assembly, with the circumferential projection of these two elements in this plane.

[0059] Indeed, one of the particularities of the invention lies in the presence, on the inner radial delimiting surface 34a, of a radial boss 52. This boss 52 projects radially in the direction SI, and is therefore located upstream of the equipment 45. At least a portion of the blading 30 projects radially in the direction SI, from the boss 52. In the preferred embodiment which has been shown, it is the entirety of the blading 30 which extends radially from the boss, that is to say that this boss 52 has an upstream end 54a which is located upstream of a leading edge of the blading 30, as well as a downstream end 54b which is located downstream of the trailing edge of this blading 30, while preferably remaining upstream of the equipment 45. However, this equipment could in whole or in part protrude from the rear part of the boss 52, without departing from the scope of the invention.

[0060] The preferred principle is shown in Figures 2 and 3, the section plane of which corresponds to the imaginary longitudinal plane P of the assembly 15, also passing through a top 56 of the boss 52. This boss top is located, in the longitudinal direction L, at the level of the blading 30, and it is even preferably centered on the latter, or substantially centered. In the radial direction R, in the first direction SI, the equipment 45 extends beyond the boss top 56. In this regard, it is noted that the boss 52 extends in a direction 57 of the height of this boss, which is preferably the radial direction R, or slightly inclined relative to this direction R, as shown in the figures. This direction 57 of the height of the boss 52 is here preferably considered as orthogonal to a fictitious base of the boss, which will be specified below.In Figures 2 and 3, the two ends of boss 54a, 54b are shown connected to each other by a fictitious surface 58 for reconstituting the vein 14b at the level of the boss. This fictitious surface 58 corresponds to the fictitious part of the surface 34a which would have been apparent, in the absence of such a boss 52. Subsequently, it will be considered that this fictitious surface, recreating the aerodynamic profile of the vein 14b, without step or difference in level, corresponds to a fictitious base 58 of the boss.

[0061] The radial boss 52 extends continuously around the axis 2, in the circumferential direction C, along the entire length of the structure 34 which it equips. It can thus extend in an annular manner over 360°, or over a reduced angular sector, corresponding to that of the structure 34, interrupted for example by the passage of the mast, beams, etc.

[0062] A solution with one or more bosses 52, spaced circumferentially from each other, remains possible, without departing from the scope of the invention. In this respect, the boss is therefore not necessarily continuous in the circumferential direction.

[0063] In the preferred solution of annular boss, this therefore has a section which is preferably identical or similar, whatever the longitudinal cutting plane, which will therefore always pass through the vertex 56.

[0064] With more specific reference to Figure 2, the inner radial delimiting surface 34a has, at its boss 52, a first inflection point 60a as well as a second inflection point 60b located downstream of the first inflection point. The first inflection point 60a is preferably located upstream of the leading edge of the blade 30, and the second inflection point 60b is preferably located between the trailing edge of the blade 30, and the equipment 45, but it could alternatively be arranged longitudinally between the leading edge and the trailing edge, without departing from the scope of the invention. Indeed, the boss 52 can have a symmetrical shape with respect to a straight line passing through its apex 56, but preferably, its front part is longer than its rear part. This implies that the slope of the boss 52, still considered in the aforementioned section plane, is lower on the front part of the boss than on its rear part.Furthermore, at its apex 56, the boss 52 has a radius of curvature which is preferentially the smallest observed on the surface 34a, all along the vein 14b which it delimits.

[0065] Thanks to the aforementioned design, the boss 52 makes it possible to locally modify the static pressure at the level of the blading 30, so as to compensate for the pressure rise created downstream of this blading, by the equipment 45. This occurs due to the local increase in the speed of the flow 12b close to the delimiting surface 34a, this increase being observed on the first front part of the boss, as well as on the top part of the boss from which the blading 30 extends, radially in the direction SI. This in fact induces a reduction at these locations of the static pressure of the secondary air flow, this same static pressure which was increased by the presence of the obstacle formed by the equipment 45 downstream.

[0066] This principle specific to the invention makes it possible to bring the equipment 45 closer to the blades 30, in order to achieve increased compactness of the propulsion unit. This advantageously makes it possible to reduce its mass and improve its performance.

[0067] To further improve these beneficial effects, the assembly 15 is made so as to meet at least one of the parameters which will be described below, and preferably several of them, or even all of them, these parameters being defined using the following dimensions, referenced on the section in the fictitious longitudinal plane of Figure 3. As indicated previously, this section plane passes through the blading 30, the equipment 45 downstream, and therefore also necessarily through the top 56 of the boss. In this section, the leading edge and the trailing edge of the blading are represented.

[0068] The dimensions are as follows:

[0069] - H: total radial height of the blade 30, at its trailing edge. This height is measured along the radial direction R, and it also corresponds locally to the radial height of the secondary vein 14b. - h: height of the boss, between its fictitious base 58 and its top 56, along the direction of the height 57 of this boss, preferably inclined relative to the radial direction R, by an angle preferably less than 30%.

[0070] - Lt: length of the boss 52, between its upstream and downstream ends 54a, 54b, corresponding to the two opposite ends of its fictitious base 58 of the boss. This length Lt is thus measured in a direction orthogonal to that of the height 57 of the boss.

[0071] - Ls: length of a front part of the boss 52, defined between its upstream end 54a, and the orthogonal projection 56a of the vertex 56 on the fictitious base 58 of the boss.

[0072] - Dbf: distance between the upstream end 54a of the boss, and the orthogonal projection 62, on the fictitious base 58, of the trailing edge of the blade 30.

[0073] - Dh: radial height differential between a front end of the equipment 45, and the top 56 of the boss. This height is also measured in the radial direction R.

[0074] - Ea: axial spacing between the top 56 of the boss, and the front end of the equipment 45.

[0075] The above mentioned parameters are as follows:

[0076] - h / H between 0.0001 and 0.2. This parameter makes it possible to estimate the coverage of the vein produced by boss 52.

[0077] - Lt / H between 0.00001 and 0.4. This parameter makes it possible to estimate the axial / longitudinal extension of the boss, in relation to the height of the secondary vein 14b.

[0078] - Lt / h between 0.001 and 50. This parameter makes it possible to estimate the curvature of the boss, and therefore the aggressiveness of the acceleration produced on the secondary flow 12b.

[0079] - Dbf / H between 0.0001 and 2. This parameter allows you to specify the distance over which the flow around the blade will be accelerated.

[0080] - Ls / Lt between 0.1 and 0.999. This parameter allows to characterize the positioning of the top 56 of the boss, in relation to its total length.

[0081] - Dh / Ea between -10 and 10, and preferably between 0.2 and 10. This parameter can be an indicator of the flow that will be able to enter the equipment. It makes it possible to characterize the capacity of bringing the equipment 45 closer upstream, as a function of the difference in radial height between this equipment and the top 56 of the boss. 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 within the limits of the scope of the appended claims. For example, if the turbomachine described in the preferred embodiment above takes the form of a turbojet with a bypass and a ducted fan, any other type of turbomachine remains conceivable. In particular, turbomachines with single or multiple fans / propellers, ducted or unducted, such as turbomachines with two unducted counter-rotating propellers, are concerned.In this regard, it is noted that the assembly according to the invention could alternatively relate to the outer radial delimiting wall of the secondary vein, provided in a similar manner with one or more pieces of equipment.

[0082] The invention applies equally to dual or triple flow propulsion units, but with the implementation of the invention in the secondary vein, which always constitutes the outermost vein of the unit, in the radial direction. Indeed, conventionally, even in a triple flow design, the tertiary vein of the propulsion unit is located radially between the primary vein and the secondary vein.

[0083] Finally, it is noted that the radial delimitation surface of the assembly according to the invention could be an external surface of the propulsion assembly. This would be possible in particular for a turbomachine with an unducted receiver, such as a turboprop or a so-called "open rotor" turbomachine. In the latter case, the radial delimitation surface, within the meaning of the invention, is a radially external surface of the secondary vein, followed by the secondary flow of the propulsion assembly.

Claims

CLAIMS 1. Assembly (15) for aircraft propulsion unit (1), comprising: - a radial delimitation structure (34) of a secondary vein for the circulation of a secondary flow (12b) of the propulsion assembly, the structure (34) comprising a radial delimitation surface (34a) of the secondary vein, oriented in a first direction (SI) of a radial direction (R) of the assembly; - equipment (45) of the propulsion assembly, this equipment projecting from the radial delimiting surface (34a), in the first direction (SI) of the radial direction (R); - an aerodynamic profile (30) also projecting from the radial delimitation surface (34a), in the first direction (SI) of the radial direction, this aerodynamic profile (30) being arranged upstream of the equipment (45) in relation to a main direction of circulation of the secondary flow (12b), characterized in that the radial delimitation surface (34a) comprises, in the first direction (SI) of the radial direction, a radial boss (52) from which at least part of the aerodynamic profile (30) extends.

2. Assembly according to claim 1, characterized in that the aerodynamic profile (30) is arranged so that at least one fictitious longitudinal plane (P) of the assembly, passing through a longitudinal central axis (2) of this assembly, crosses both the aerodynamic profile (30) and the equipment (45).

3. Assembly according to claim 2, characterized in that in section in said imaginary longitudinal plane (P) of the assembly, the radial delimitation surface (34a) has, at its radial boss (52), a first inflection point (60a) as well as a second inflection point (60b) located downstream of the first inflection point, the first inflection point (60a) being preferentially located upstream or at the level of the aerodynamic profile (30), and the second inflection point (60b) being preferentially located between the aerodynamic profile (30) and the equipment (45).

4. Assembly according to any one of the preceding claims, characterized in that the boss (52) has a top (56) located, in a longitudinal direction (L) of the assembly, at the level of the aerodynamic profile (30).

5. Assembly according to the preceding claim, characterized in that the equipment (45) extends, in the first direction (SI) of the radial direction, beyond the top (56) of the boss.

6. Assembly according to any one of the preceding claims, characterized in that the equipment (45) is a scoop or a heat exchanger.

7. Assembly according to any one of the preceding claims, characterized in that the aerodynamic profile is an outlet guide vane (30), a compressor rectifier blade, or a turbine distributor blade.

8. Assembly according to any one of the preceding claims, characterized in that within it, the following dimensions are defined, in any fictitious longitudinal plane (P) passing through the aerodynamic profile (30), the equipment (45), and also through the top (56) of the radial boss (52): - H: total radial height of the aerodynamic profile (30), at its trailing edge; - h: height of the boss (52), between a fictitious base (58) of the latter and its top (56), according to a direction of the height (557) of this radial boss; - Lt: length of the boss (52), between an upstream end (54a) and a downstream end (54b) of this boss, corresponding to opposite ends of its fictitious base (58); - Ls: length of a front part of the boss (52), defined between its upstream end (54a), and the orthogonal projection (56a) of the top (56) on the fictitious base (58) of the boss; - Dbf: distance between the upstream end (54a) of the boss, and the orthogonal projection (62), on the fictitious base (58), of the trailing edge of the aerodynamic profile (30), - Dh: radial height differential between a front end of the equipment (45), and the top (56) of the boss; - Ea: axial spacing between the top (56) of the boss and the front end of the equipment (45), and in that the assembly is carried out in such a way as to meet at least one of the following parameters: - h / H between 0.0001 and 0.2 - Lt / H between 0.00001 and 0.4 - Lt / h between 0.001 and 50 - Dbf / H between 0.0001 and 2 - Ls / Lt between 0.1 and 0.999 - Dh / Ea between -10 and 10 9. Assembly according to any one of the preceding claims, characterized in that the boss (52) extends continuously in the circumferential direction (C), preferably along the entire length of the radial delimiting structure (34).

10. Aircraft propulsion assembly (100) comprising at least one assembly (15) according to any one of the preceding claims, this assembly comprising a turbomachine (1), preferably a dual-flow turbojet.

Images