Assembly for an aircraft propulsion unit, comprising a device through which a gas flow passes that is fitted with a flow mixer

Abstract

The invention relates to an assembly (15) for an aircraft propulsion unit, comprising: - a radial bounding structure (34) for the flow of a gas flow (12b), comprising a radial bounding surface (34a) oriented in a first direction (S1); - a first device (45a) projecting from the radial bounding surface (34a) in the first direction (S1) and configured to have a first portion (12b1) of the gas flow pass therethrough, and wherein the first device comprises a cowling (50) that is intended to be followed closely by the first portion (12b1) of the gas flow exiting the first device (45a), and by a second portion (12b2) of the gas flow. According to the invention, the cowling (50) comprises a mixer (64) for mixing the second portion (12b2) with the first portion (12b1) of the gas flow.

Description

[0001] ASSEMBLY FOR AN AIRCRAFT PROPULSION UNIT, COMPRISING A

[0002] EQUIPMENT CROSSED BY A FLOW OF GAS AND EQUIPPED WITH A FLOW MIXER

[0003] TECHNICAL FIELD

[0004] The present invention relates to the field of aircraft propulsion assemblies, comprising equipment configured to be traversed by a gas flow, such as the secondary flow of a dual-flow turbojet.

[0005] The invention applies to propulsion units comprising turbomachines of all types, such as turbojets and turboprops, with single or double fans or propellers, ducted or unducted.

[0006] STATE OF PRIOR ART

[0007] 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.

[0008] 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.

[0009] 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.

[0010] Some of this equipment is installed to cooperate with circulating gas flows, for example the secondary flow of a turbojet. In this case, the equipment is, for example, ACOC (Air-Cooled Oil-cooler) type heat exchangers. To ensure their proper operation, this equipment is generally arranged protruding into the secondary flow of the propulsion unit.

[0011] The aforementioned equipment forms protrusions that may produce harmful effects on other equipment located nearby downstream, due to the wake effect. Indeed, this equipment can impact the flow of the gas stream downstream of the protrusions that they form, resulting in a stratification of the flow. This stratification of the flow causes a significant increase in the thickness of the boundary layer, as well as an increase in the temperature of the flow near the radial delimitation surface.

[0012] These two phenomena are obviously detrimental, particularly when the equipment arranged downstream is intended to cooperate with the gas flow. This is particularly the case when such equipment is of the exchanger or scoop type, for example a scoop intended to supply fresh air to a device for controlling the clearance at the blade tips.

[0013] The air sampling carried out by the downstream equipment is then negatively impacted, in particular due to the reduced total pressure caused by the thickening of the boundary layer, as well as due to the increase in the temperature of the sampled flow.

[0014] This reduces the efficiency of the samples taken, which may require oversizing the devices / equipment intended to be supplied by the gas flow, so that they can properly perform their functions. This oversizing, apart from an undesirable increase in the overall mass of the assembly, can also create additional pressure losses and cause installation problems in areas that are already heavily congested.

[0015] STATEMENT OF THE INVENTION

[0016] To at least partially resolve the drawbacks mentioned above, relating to the embodiments of the prior art, the invention firstly relates to an assembly for an aircraft propulsion unit, according to the characteristics of claim 1.

[0017] The mixing between the first and second parts of the gas flow, carried out by the mixer arranged in the manner specific to the present invention, makes it possible to reduce the stratification effect of the flow near the radial delimiting surface, to reduce the temperature difference in the radial direction, and to increase the velocity of the flow near this surface, in order to reduce the thickness of the boundary layer.

[0018] This advantageously results in a homogenization of the flow downstream of the first equipment, beneficial in particular for potential second equipment arranged downstream of the first equipment, and also intended to cooperate with the gas flow. Indeed, the mixture obtained is capable of increasing the dynamic pressure of the flow close to and radially lower than the delimitation surface, and of reducing the temperature of the flow in this same zone. This improves the operating conditions of these potential second downstream equipment, which can cooperate with a sufficiently cool flow at a sufficiently high total pressure, without the need to radially offset the sampling, or to reduce this radial offset.

[0019] With the efficiency of the samples being increased, the dimensions of the equipment can be reduced, which facilitates their installation in constrained environments, in addition to reducing the associated pressure losses. This also results in a reduction in the overall mass of the assembly, which contributes to obtaining a lighter and compact propulsion system. 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).

[0020] In addition, the first and second radial components, associated with the first and second projecting elements of the mixer, are in opposite radial directions. This principle further accentuates the mixing, by increasing the effect of pressure and temperature redistribution downstream of the first equipment, so as to limit stratification of the flow as much as possible.

[0021] The invention furthermore preferably provides at least any one of the following optional features, taken individually or in combination.

[0022] Preferably, the assembly comprises a second piece of equipment located downstream of the first piece of equipment, and intended to be supplied by the gas flow, this second piece of equipment being arranged on the radial delimiting surface, or projecting from this surface. In addition, preferably, at least one longitudinal imaginary plane of the assembly passes through both the first and second pieces of equipment.

[0023] Preferably, the second equipment is a scoop or an exchanger, and for example a scoop intended to supply fresh air to a device for controlling the clearance at the end of the blades, preferably for a turbine of the turbomachine.

[0024] Preferably, the radial delimiting structure also delimits an equipment housing compartment.

[0025] Preferably, the first equipment is a heat exchanger or a filter, even if other equipment can be envisaged, without departing from the scope of the invention.

[0026] Preferably, the equipment is arranged downstream of an outlet guide vane of the assembly.

[0027] Preferably, within the assembly, the following dimensions are defined, in any fictitious longitudinal plane passing through the first equipment:

[0028] - H: radial height of the vein in which said gas flow circulates, this height being considered at the level of a trailing edge of the mixer;

[0029] - L1: axial length of the mixer; and the assembly is carried out in such a way as to meet the following parameter:

[0030] 0.001 < Ll / H < 0.3

[0031] The invention also relates to an aircraft propulsion assembly comprising at least one such assembly, the assembly comprising a turbomachine, preferably a dual-flow turbojet, said radial delimitation structure of the assembly preferably being an internal or external radial delimitation structure of a secondary vein of the turbojet, in which a secondary flow is intended to circulate. Other applications obviously remain possible, within the turbojet, or outside it, or for other types of turbomachine.

[0032] Other advantages and characteristics of the invention will appear in the detailed non-limiting description below.

[0033] BRIEF DESCRIPTION OF THE DRAWINGS

[0034] This description will be made with regard to the attached drawings, among which;

[0035] [Fig. 1] represents a schematic view in longitudinal section of an aircraft propulsion unit

[0036] [Fig. 2] represents a more detailed longitudinal sectional half-view of an assembly for the propulsion unit shown in the previous figure, the sectional plane corresponding to a fictitious longitudinal plane of the assembly;

[0037] [Fig. 3] represents a perspective view of a part of the assembly shown in Fig. 2;

[0038] [Fig. 4] represents another perspective view of part of the assembly, from another angle of view; and

[0039] [Fig. 5] represents a schematic sectional view of a part of the assembly, according to a preferred embodiment of the invention.

[0040] DETAILED DISCLOSURE OF PREFERRED EMBODIMENTS

[0041] 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.

[0042] In the embodiments that will be described, the turbomachine preferably corresponds to a twin-spool, bypass 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 fans or propellers, ducted or unducted. 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.Furthermore, the propulsion unit 1 is represented in a frame of reference 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.

[0043] 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).

[0044] 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. More precisely, 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 centred 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 (from the English "Outlet Guide vanes"). 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.

[0045] 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 embodiment described. In the case of one or more unducted fans / propellers, this bypass ratio is rather of the order of eighteen to forty. 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.

[0046] 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.

[0047] 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.

[0048] 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.

[0049] 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.

[0050] The blades 30, forming aerodynamic profiles crossed by the secondary flow 12b in the secondary vein 14b, form with the radial delimitation structure 34 an assembly 15 specific to the present invention. This assembly 15 is completed by a first piece of equipment 45a and a second piece of equipment 45b of the propulsion assembly 1, even if the number of these pieces of equipment could be greater, without departing from the scope of the invention.

[0051] In the embodiment shown in Figures 2 to 4, not covered by the invention, only one first piece of equipment 45a will be described, arranged axially in axial proximity to one of the blades 30, downstream of the latter. This is for example a heat exchanger of the ACOC type (from the English, “Air-Cooled Oil-cooler”). The first piece of equipment 45a is intended to be crossed by the secondary flow 12b, and more precisely by a first part 12bl of this flow 12b. The first equipment 45a projects from the inner radial delimiting surface 34a, in the first direction SI of the radial direction R. The equipment 45a thus forms a radial protuberance in the secondary vein 14b, located in the downstream extension of the blading 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.

[0052] It is noted that the first equipment 45a may be semi-buried as seen in Figures 2 to 4. To do this, locally at the level of this equipment 45a, the inner radial delimitation surface 34a may form a radial recess, in which the first equipment is partly housed. A fully buried configuration is also possible, with an even more substantial radial recess on the inner radial delimitation surface 34a, from which the first equipment protrudes, while remaining entirely housed in this recess. Furthermore, a fully “intrusive” configuration, without burial, is also possible, in which the inner radial delimitation surface 34a no longer requires a radial recess to fully or partially house the first equipment 45a.

[0053] The first equipment 45a, here in the form of a heat exchanger, is connected to a heat exchange system 47 arranged in whole or in part in a compartment 49 intended for housing several pieces of equipment. This compartment 49, also called the engine compartment or “core compartment”, is delimited radially towards the outside by the structure 34, and it is therefore arranged between the two veins 14a, 14b. The compartment 49, accessible to the operators, has been shown diagrammatically in dotted lines in Figure 2.

[0054] The first equipment 45a can be axially aligned with one of the blades 30 in the longitudinal direction L, or else circumferentially offset from these blades 30.

[0055] This equipment 45a comprises, at a radial end thereof in the first direction SI, a cowling 50, for example in the form of an aerodynamic profile. As has been shown in the figures, this cowling 50 extends axially upstream and downstream beyond the functional part of the equipment 45a, namely beyond an air inlet 52 and an air outlet 54 of this equipment.

[0056] The cowling 50 has first and second opposite radial surfaces 56, 58. The first radial surface 56, oriented radially inwards, is intended to be matched by the first part 12bl of the secondary flow exiting through the air outlet 54 of the equipment 45a. The second radial surface 58, oriented radially outwards, is intended to be matched by a second part 12b2 of the secondary flow. It is noted that the two parts 12bl, 12b2 separate from each other at a leading edge of the cowling 50, to then remix within the secondary flow 12b, at a trailing edge of this cowling, in a manner specific to the invention which will be described below. In addition, the assembly comprises a second equipment 45b, located downstream of the first equipment 45a, and intended to be supplied by the secondary flow 12b.The second equipment 45b is arranged on the radial delimiting surface 34a, or projects from this surface in the radial direction SI. It may for example be another heat exchanger, or, as has been shown schematically, an air scoop. As mentioned above, this scoop 45b may be arranged to project radially to form a protuberance in the vein, or else be arranged flush with the surface 34a.

[0057] The scoop 45b is for example intended to supply fresh air to a device 62 for controlling the clearance at the end of the turbine blades, this device being partly housed in the compartment 49, with in particular one or more air circulation pipes. This device 62 is also known as an active blade tip clearance control system, for high and / or low pressure turbines.

[0058] The first equipment 45a may be axially aligned with the second equipment 45b in the longitudinal direction L, or circumferentially offset from it. In the case of such an alignment, there is consequently at least one fictitious longitudinal plane P of the assembly passing through the axis 2 also corresponding to the longitudinal central axis of this assembly, and passing through both the first and second equipment 45a, 45b. One of these planes P corresponds to the longitudinal section plane of FIG. 2.

[0059] For information purposes, it is noted that the aforementioned alignment case is not limited to perfect alignment, but on the contrary relates to all cases in which there is at least one overlap zone in direction C, between the two pieces of equipment 45a, 45b.

[0060] One of the particularities of the invention lies in the implementation, on the downstream end of the cowling corresponding to its trailing edge, of a mixer 64 of the second part 12b2 of the secondary flow, with the first part 12bl of this same flow, and leaving the first equipment 45a. Thanks to this mixer, the redistribution of pressure and temperature is increased downstream of the first equipment 45a, thus limiting the effects of stratification of the flow. The second equipment 45b can then take air at a colder temperature and at a higher pressure, which contributes to reinforcing the efficiency of the scoop 45b, and that of the entire clearance control device 62 in which it is integrated.

[0061] In this embodiment, the mixer 64 extends in the direction C, over all or part of the trailing edge of the cowling 50. Consequently, it can, for example, extend over an angular sector of the order of 5 to 25°. The mixer 64 here comprises an alternation, in the direction C, of ​​hollows 66 and projecting elements 68, the latter preferably being in the form of chevrons. They are, for example, provided at a density of between 3 and 30 elements, per 90° sector. The hollows 66 and the projecting elements 68 extend in the longitudinal direction L, while being able to have a radial component, as will be described below.

[0062] In this regard, it is noted that such chevron mixers are known on turbomachines, in particular on the gas ejection cone, but for purposes different from those of the invention. Indeed, traditionally, the chevrons have an acoustic objective, to "break" the vortex structures present in the sheared jets, coming from adjacent flows circulating at different speeds. In the case of the invention, the objective is different since it consists of a redistribution of pressure and temperature downstream of the exchanger 45a crossed by the flow, so as to avoid a stratification of the flow, with a hot layer (first part of the flow) at low total pressure close to the surface 34a at the outlet of this exchanger, and a colder layer (second part of the flow) and at higher total pressure.

[0063] In the absence of such a mixer according to the invention, to maintain the efficiency of the sampling at the level of the second equipment 45b, it could be necessary to radially extend the length of the scoop 45b in the vein 14b, so that this scoop can collect cooler air, at higher pressure, above the hot boundary layer.

[0064] Such an extension obviously proves to be harmful in terms of mass and aerodynamic disturbances on the secondary flow 12b.

[0065] The projecting elements 68 are flat or very slightly curved, to follow the curvature of the vein 14b. Each forms, in section in any imaginary longitudinal plane P passing through it, an angle A with the direction L parallel to the axis 2, this angle A preferably being between 0 and 70°, in the first direction S1, or in a second direction S2 opposite to the first direction of the radial direction R. Each projecting element 68 has a direction D, forming the angle A, and the radial component C of which is here oriented in the second direction S2, so that the angle is said to be re-entrant. In the example shown, it is for example of the order of 20° in the second direction S2.

[0066] According to a preferred embodiment of the invention, shown in Figure 5, the projecting elements 68 comprise first and second elements 68a, 68b arranged alternately in the direction C, having radial components C1, C2 of opposite radial directions. Indeed, each first element 68a forms, in section in any imaginary longitudinal plane P passing through it, an angle A1 with the direction L, this angle A1 preferably being between 0 and 70°, in the first direction S1. Thus, each first projecting element 68a has a direction D1, forming the angle A1 with the direction L, and whose radial component C1 is oriented in the first direction S1, so that the angle is said to be outgoing. In the example shown, it is for example of the order of 20° to 45° in the first direction S1.

[0067] Similarly, each second element 68b forms, in section in any imaginary longitudinal plane P passing through it, an angle A2 with the direction L, this angle A2 preferably being between 0 and 70°, in the second direction S2. Each second projecting element 68b thus has a direction D2, forming the angle A2 with the direction L, and whose radial component C2 is oriented in the second direction S2, so that the angle is said to be re-entrant. In the example shown, it is for example of the order of 20° in the second direction S2.

[0068] Thanks to this alternation of inward and outward chevrons, the redistribution of pressure and temperature is further increased downstream of the first equipment 45a, thus limiting the effects of flow stratification, and reinforcing the efficiency of the second equipment 45b.

[0069] Returning to Figure 2, the assembly 15 has the following dimensions, in any fictitious longitudinal plane P passing through the first equipment 45a:

[0070] - H: radial height of secondary vein 14b, in which the secondary flow 12b circulates, this height being considered at the level of a trailing edge of the mixer 64;

[0071] - L1: axial length of the mixer 64, in the direction L.

[0072] These two dimensions respond to the following parameter:

[0073] 0.001 < Ll / H < 0.3

[0074] Of course, various modifications may 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 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, this concerns turbomachines with single or multiple fans / propellers, ducted or unducted, such as turbomachines with two unducted counter-rotating propellers. In this regard, it is noted that the assembly according to the invention could alternatively concern the outer radial delimiting wall of the secondary vein, provided in a similar manner with several pieces of equipment 45a, 45b.

[0075] According to still other possible applications, the assembly according to the invention could concern a radial, internal or external delimitation surface of the primary vein.

[0076] The assembly according to the invention could also relate to a radial delimitation surface, internal or external, of a tertiary vein, when such a vein is provided within the propulsion assembly. In other words, the invention applies equally to single, double or triple flow propulsion assemblies. 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, matched by the external air flow of the propulsion assembly.

Claims

CLAIMS 1. Assembly (15) for aircraft propulsion unit (1), comprising: - a radial delimiting structure (34) for the circulation of a gas flow (12b) of the propulsion assembly, the structure (34) comprising a radial delimiting surface (34a) oriented in a first direction (SI) of a radial direction (R) of the assembly; - a first piece of equipment (45a) of the propulsion unit, this first piece of equipment projecting from the radial delimiting surface (34a), in the first direction (SI) of the radial direction (R), the first piece of equipment (45a) being configured to be crossed by a first part (12bl) of said gas flow, and the first piece of equipment comprising, at a radial end thereof in the first direction (SI), a cowling (50) having first and second opposite radial surfaces (56, 58), the first radial surface (56) being intended to be matched by the first part (12bl) of said gas flow leaving the first piece of equipment (45a), and the second radial surface (58) being intended to be matched by a second part (12b2) of said gas flow, characterized in that the cowling (50) comprises, at a downstream end thereof, a mixer (64) of the second part (12b2) of said gas flow, with the first part (12bl) of the gas flow leaving the first equipment (45a),the mixer (64) comprising, alternately along a circumferential direction (C) of the assembly, hollows (66) and projecting elements (68, 68a, 68b), the latter preferably being in the form of chevrons, teeth, lobes or petals, the projecting elements (68, 68a, 68b) being preferably provided at a density of between 3 and 30 elements per 90° sector, and projecting axially downstream relative to the hollows (66), the projecting elements comprising first and second projecting elements (68a, 68b), arranged alternately along the circumferential direction (C) of the assembly, the first elements (68a) extending along a first direction (D1) comprising a first radial component (C1), and the second elements (68b) extending along a second direction (D2) comprising a second radial component (C1), radial component (C2) such that the first and second radial components (Cl, C2) are of opposite radial directions,each first and second projecting element (68a, 68b) forming, in section in any imaginary longitudinal plane (P) of the assembly passing through the projecting element concerned, an angle (A1, A2) with a longitudinal direction (L), the two angles (A1, A2) preferably being between 0 and 70°, and oriented respectively in the first direction (S1) and in a second direction (S2) opposite to the first direction of the radial direction (R)., 2. Assembly according to the preceding claim, characterized in that it comprises a second piece of equipment (45b) located downstream of the first piece of equipment (45a), and intended to be supplied by the gas flow (12b), this second piece of equipment (45b) being arranged on the radial delimitation surface (34a), or projecting from this surface, and in that preferably, at least one fictitious longitudinal plane (P) of the assembly passes through both the first and second pieces of equipment (45a, 45b).

3. Assembly according to claim 2, characterized in that the second equipment (45b) is a scoop or an exchanger.

4. Assembly according to any one of the preceding claims, characterized in that the radial delimiting structure (34) also delimits a compartment (49) for housing equipment.

5. Assembly according to any one of the preceding claims, characterized in that the first equipment (45a) is a heat exchanger or a filter, and in that it is preferably arranged downstream of an outlet guide vane (30) of the assembly.

6. 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 first equipment (45a): - H: radial height of the vein in which said gas flow (12b) circulates, this height being considered at the level of a trailing edge of the mixer (64); - L1: axial length of the mixer (64); and in that the assembly is carried out in such a way as to meet the following parameter: 0.001 < Ll / H < 0.3 7. Aircraft propulsion assembly (100) comprising at least one assembly (15) according to any one of the preceding claims, the assembly comprising a turbomachine (1), preferably a dual-flow turbojet, said radial delimitation structure (34) of the assembly preferably being an internal or external radial delimitation structure of a secondary vein (14b) of the turbojet, in which a secondary flow (12b) is intended to circulate.

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