Assembly for an aircraft propulsion unit, comprising a heat exchanger with improved performance
Patent Information
- Application Number
- PCT/FR2026/050171
- Authority / Receiving Office
- WO · WO
- Patent Type
- Applications
- Current Assignee / Owner
- Priority Date
- 2025-03-10
- Filing Date
- 2026-03-06
- Publication Date
- 2026-09-17
Smart Images
Figure FR2026050171_17092026_PF_FP_ABST
Abstract
Description
[0001] Aircraft propulsion system assembly, including an enhanced performance heat exchanger
[0002] TECHNICAL FIELD
[0003] The present invention relates to the field of aircraft propulsion systems, comprising heat exchangers configured to be traversed by a gas flow, such as the secondary flow of a turbofan engine.
[0004] The invention applies to propulsion assemblies comprising turbomachines of any type, such as turbojets and turboprops, with single or double fan or propeller, shrouded or unshrouded.
[0005] STATE OF PRIOR ART
[0006] In the propulsion systems equipping aircraft, several gas flows are observed circulating both within and outside the system. Indeed, in addition to the external airflow circulating around the system, there are also one or more gas flows circulating within the system, such as a primary and a secondary flow in the case of a turbofan engine.
[0007] Radial boundary structures are arranged within the propulsion assembly to channel these gas flows, which generally flow in a direction from upstream to downstream of the assembly.
[0008] The evolution of propulsion systems and their turbomachinery is leading to an increased need for equipment to control their operation and enhance their performance. Furthermore, due to ever-increasing demands for compactness, the space available for installing this equipment is becoming increasingly limited. Some of this equipment is installed to interact with circulating gas flows, such as the secondary flow of a turbojet engine. In this case, the equipment includes, for example, ACOC (Air-Cooled Oil-cooled) heat exchangers. To ensure their proper operation, this equipment is generally mounted protruding from the secondary flow of the propulsion system. However, the presence of this protruding equipment is usually compounded by the presence of aerodynamic profiles through which the associated gas flow passes.In the case of a turbofan engine, in the secondary flow, these are outlet guide vanes, or OGVs (from the English "Outlet Guide Vanes"), through which the secondary flow passes.
[0009] To meet the aforementioned compactness requirements, equipment can be positioned axially close to the trailing edge of airfoils. However, the protrusions formed by this equipment can have detrimental effects on nearby upstream airfoils due to pressure surges. Indeed, adverse aerodynamic interactions can occur due to pressure surges caused by equipment protruding into the flow. For example, a heat exchanger located close to and downstream of an outlet guide vane can cause static pressure disturbances on the vane, impairing its aerodynamic performance.
[0010] Therefore, there remains a need to improve the design of heat exchangers, aiming to increase their performance while reducing their size. This would help to decrease pressure losses and the pressure surges described earlier.
[0011] DESCRIPTION OF THE INVENTION
[0012] To meet the need mentioned above, the invention first relates to an assembly for an aircraft propulsion system, comprising:
[0013] - a radial boundary structure for the circulation of a gas flow, the structure comprising a radial boundary surface oriented in a first direction of a radial direction of the assembly;
[0014] - a heat exchanger projecting from the radial boundary surface, in the first direction of the radial direction,
[0015] the exchanger being configured to be traversed by a first part of said gas flow, the exchanger also comprising, at a radial end thereof in the first direction, a hood having a first and a second opposing radial surfaces, the second radial surface being intended to be followed by a second part of said gas flow, the exchanger comprising, at its upstream end, an upstream inlet intended to be traversed by the first part of said gas flow, and, at its downstream end, a downstream outlet.
[0016] According to the invention, the casing comprises, between the upstream end and the downstream end of the exchanger:
[0017] - an intermediate inlet intended to be traversed by a sample taken from the second part of said gas flow;
[0018] - an intermediate outlet intended to be traversed by a primary portion of the first part of said gas flow, in order to join the second part of said gas flow,
[0019] said sampling and a secondary portion of the first part of said gas flow being intended to be extracted from the exchanger by said downstream outlet.
[0020] The renewal of cooling gases, provided by the intermediate inlet and outlet on the heat exchanger, increases the heat exchange capacity. This improves the performance of the heat exchanger and can thus contribute to reducing its overall volume. This advantageously results in a reduction of pressure losses and static pressure rises caused by the heat exchanger. It also leads to a decrease in the overall mass of the assembly, contributing to a compact and lightweight propulsion system. The invention is therefore the result of technological research aimed at significantly improving aircraft performance and, in this respect, contributes to reducing the environmental impact of these aircraft (decarbonization).
[0021] The invention also preferably provides for at least one of the following optional features, taken individually or in combination.
[0022] Preferably, the intermediate inlet comprises several inlet ports, the intermediate outlet comprises several outlet ports, and the inlet and outlet ports are preferably arranged alternately in a circumferential direction of the assembly. However, other arrangements may be implemented for these ports without departing from the scope of the invention.
[0023] Preferably, the inlet openings are scoops facing upstream, and the outlet openings are facing downstream. More generally, the inlet and outlet openings can either protrude from the cowling or not, in the first direction of the radial axis.
[0024] Preferably, the intermediate inlet and outlet are arranged in an axially central area of the exchanger casing.
[0025] Preferably, the assembly is configured so that a gas renewal rate, defined by the ratio between the flow rate of the sample passing through the intermediate inlet, and the flow rate of the first part of said gas flow passing through the upstream inlet, is between 0.1 and 0.6, and preferably between 0.3 and 0.5.
[0026] Preferably, the assembly includes equipment located downstream of the exchanger, intended to be supplied by the gas flow, this equipment being arranged on the radial boundary surface, or projecting from this surface, and preferably, at least one fictitious longitudinal plane of the assembly passes through both the exchanger and the equipment.
[0027] Preferably, the equipment is a scoop or an exchanger, and for example a scoop intended to supply fresh air to a blade tip clearance control device, preferably for a turbine of the turbomachine.
[0028] Preferably, the radial boundary structure also delimits an equipment housing compartment.
[0029] Preferably, the exchanger is arranged downstream of a guide vane exiting the assembly.
[0030] Preferably, the casing includes, at a downstream end thereof, a mixer of the second part of said gas flow, with the sampling and secondary portion of the first part of said gas flow exiting the exchanger by said downstream outlet.
[0031] Such a mixture makes it possible to reduce the stratification effect of the flow near the radial boundary surface, to reduce the temperature difference in the radial direction, and to increase the flow velocity near this surface, in order to reduce the thickness of the boundary layer.
[0032] This results in a more homogenized flow downstream of the heat exchanger, which is particularly beneficial for any equipment located downstream and also intended to interact with the gas flow. The resulting mixture increases the dynamic pressure of the flow near and radially below the boundary surface, while simultaneously reducing the flow temperature in this same area. This improves the operating conditions of this downstream equipment, allowing it to operate with a sufficiently cool flow at a sufficiently high total pressure, without the need to radially offset the intake or reduce this radial offset.
[0033] With increased efficiency of sampling, equipment sizes can be reduced, facilitating their installation in constrained environments, in addition to reducing associated pressure losses.
[0034] Preferably, the mixer comprises, alternately in a circumferential direction of the assembly, hollows and protruding elements, the latter being preferably in the shape of chevrons, teeth, lobes or petals, the protruding elements being preferably provided at a density of between 3 and 30 elements per 90° sector.
[0035] The invention also relates to an aircraft propulsion system comprising at least one such assembly, the assembly comprising a turbomachine, preferably a turbofan engine, said radial boundary structure of the assembly being preferably an internal or external radial boundary structure of a secondary flow of the turbofan engine, through which a secondary flow is intended to circulate. Other applications are obviously possible, within the turbofan engine, or outside of it, or for other types of turbomachinery.
[0036] Other advantages and features of the invention will appear in the detailed, non-limiting description below.
[0037] BRIEF DESCRIPTION OF THE DRAWINGS
[0038] This description will be made with reference to the attached drawings, among which are;
[0039] - [Fig. 1] represents a schematic longitudinal cross-sectional view of an aircraft propulsion system;
[0040] - [Fig. 2] represents a more detailed longitudinal sectional half-view of an assembly for the propulsion unit shown in the previous figure, the assembly being in the form of a preferred embodiment of the invention, and the section plane corresponding to a fictitious longitudinal plane of the assembly;
[0041] - [Fig. 3] represents a half-view in longitudinal section, even more detailed, of the assembly shown in the previous figure;
[0042] - [Fig- 4] represents a perspective view of part of the heat exchanger included in the assembly shown in figures 2 and 3.
[0043] DETAILED EXPLANATION OF PREFERRED METHODS OF IMPLEMENTATION
[0044] With reference first to Figure 1, a propulsion unit 100 for an aircraft is shown, comprising a turbomachine 1 and a nacelle 9 surrounding the turbomachine. In the preferred embodiments of the invention that will be described, the turbomachine preferably corresponds to a twin-spool, turbofan engine with a single ducted fan. However, it could be a turbomachine of another type, for example, a turboprop. More generally, it could be any type of turbojet or turboprop engine, with a single or twin fan or propeller, ducted or unducted.
[0045] Subsequently, the terms "upstream" and "downstream" are defined relative to a general direction 5 of gas flow through the propulsion unit 1 when it generates direct thrust, this direction being parallel or substantially parallel to axis 2. These terms "upstream" and "downstream" could respectively be replaced by the terms "front" and "rear," with the same meaning. Furthermore, the propulsion unit 1 is represented in a coordinate system formed by three orthogonal directions, namely the longitudinal direction L parallel to axis 2, which will be defined below, the circumferential direction C, and the radial direction R.
[0046] The turbojet 1 has a central longitudinal axis 2 around which its various components extend. It comprises, from upstream to downstream along the main direction 5 of gas flow 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 unit comprising the compressor 4 and the turbine 8, or indirectly by a reduction gear (not shown).
[0047] Conventionally, an airflow F arriving at an air inlet of the blower splits into an internal airflow 10a which enters the nacelle, and an external airflow 10b which follows the radially external surface of the nacelle 36a. This surface 36a is defined by an outer skin of the nacelle 36.
[0048] After passing through the fan 3, the internal airflow 10a splits into a central primary flow 12a and a secondary airflow 12b that surrounds the primary flow. The primary flow 12a flows into a main gas circulation channel 14a, passing through the compressors 4, 6, the combustion chamber 11, and the turbines 7, 8. The secondary flow 12b flows into a secondary air channel 14b, radially delimited outwards by a casing, which is surrounded by the nacelle 9. More specifically, the casing includes a fan casing 20 that surrounds the fan blades. This casing 20 is extended downstream by an outer ring 22 of an intermediate casing 24. This intermediate casing 24 has a hub 26 centered on the axis 2 and may include a flow separation nozzle 28.The intermediate casing is completed by radial arms 30, which form outlet guide vanes, allowing the secondary airflow 12b to be straightened and are conventionally called OGVs (from the English "Outlet Guide vanes"). The arms 30 thus connect the outer ferrule 22 of the intermediate casing to its hub 26, at the level of the low-pressure compressor 4.
[0049] The turbojet engine exhibits a bypass ratio, or BPR (Bypass Ratio), of approximately eight to forty, and more specifically, approximately eight to eighteen in the case of a shrouded fan corresponding to that of the preferred embodiment described. In the case of one or more unshrouded fans / propellers, this bypass ratio is more in the range of eighteen to forty.
[0050] Whether the blower(s) / propeller(s) are shrouded or unshrouded, their diameter is preferably in the order of 0.5 to 5 m, while the compression ratio is preferably in the order of 1 to 1.5, or even in the order of 1 to 1.75.
[0051] In a turbomachine equipped with a reduction gear, with the propeller(s) / fan(s) shrouded or unshrouded, the reduction ratio can be between 1.1 and 20. The outer shell 22 forms an outer radial boundary structure for the secondary air stream 14b. It is extended downstream by another annular boundary structure 31 of this type, integrated into a movable thrust reverser cover 32. This structure 31 is also called the OFS (Outer Fixed Structure). It is located radially opposite an annular inner radial boundary structure 34 for the stream 14b, also called the IFS (Inner Fixed Structure), and initiated upstream by the nozzle 28. The inner radial boundary structure 34 includes the hub 26 of the intermediate casing.
[0052] This radial delimitation structure 34 of the vein 14b includes 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, with respect to the axis 2. It is therefore this surface 34a which delimits the secondary vein 14b, radially towards the inside.
[0053] Each of the fixed exit guide vanes 30 protrudes from the inner radial boundary surface 34a, in the first direction SI of the radial direction R, towards the outer ferrule 22 of the intermediate housing.
[0054] The blades 30, forming aerodynamic profiles through which the secondary flow 12b passes in the secondary channel 14b, together with the radial boundary structure 34, form an assembly 15 according to a preferred embodiment of the invention. This assembly 15 is completed by a heat exchanger 45a, and preferably by one or more components 45b of the propulsion unit 1.
[0055] In the preferred embodiment shown in Figures 2 to 4, only one heat exchanger 45a will be described, arranged axially, preferably in the axial vicinity of one of the blades 30, downstream of the latter. This is, for example, an ACOC (Air-Cooled Oil-cooler) type heat exchanger. The heat exchanger 45a is intended to be traversed by a portion of the secondary flow 12b, and in particular by a first portion 12bl of this flow 12b. The various air flows involved in the heat exchange will be detailed with reference to Figures 3 and 4, which are more detailed.
[0056] The exchanger 45a protrudes from the inner radial boundary surface 34a, in the first direction SI of the radial direction R. It thus forms a radial protrusion in the secondary lava 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 gas flow through the propulsion assembly 1.
[0057] It is noted that the heat exchanger 5a can be semi-buried, as shown in Figures 2 to 4. To achieve this, locally at the level of this heat exchanger 45a, the inner radial boundary surface 34a can form a radial recess, in which the heat exchanger is partially housed. A fully buried configuration is also possible, with an even larger radial recess on the inner radial boundary surface 34a, from which the heat exchanger protrudes, while remaining entirely housed within this recess. Furthermore, a fully "intrusive" configuration, without burial, is also possible, in which the inner radial boundary surface 34a no longer requires a radial recess to fully or partially house the heat exchanger 45a.
[0058] The heat exchanger is connected to a heat exchange system 47 arranged wholly or partly within a compartment 49 designed to house various equipment. This compartment 49, also called the motor compartment or "core compartment," is radially bounded outwards by the structure 34, and is therefore located between the two ribs 14a and 14b. The compartment 49, accessible to operators, is shown schematically in dashed lines in Figure 2.
[0059] The heat exchange system 47 also includes means 48 for supplying and extracting the fluid to be cooled by air, in the exchanger 45a. This other fluid is preferably oil intended for the lubrication and / or cooling of turbojet components, such as bearing housings or the fan drive gearbox, when such a gearbox is provided.
[0060] The exchanger 45a can be aligned axially with one of the blades 30 along the longitudinal direction L, or offset circumferentially from these blades 30.
[0061] At one radial end of the first direction SI, it includes a cowling 50, for example, in the shape of an aerodynamic profile. As shown in the figures, this cowling 50 extends axially upstream and downstream beyond the functional part 64 of the heat exchanger 45a, namely the structured part comprising channels or similar elements, enabling heat exchange between the two flows. At one upstream end of the heat exchanger 45a, it includes an upstream air inlet 52 for the first part 12bl of the secondary flow to pass through. This inlet 52 may be partially delimited by the upstream end of the cowling 50, which may extend axially upstream beyond the functional part 64 of the heat exchanger. At one downstream end of the exchanger 45a, it includes a downstream air outlet 54.This outlet 54 can also be partially delimited by the downstream end of the casing 50, which can extend axially downstream beyond the functional part 64 of the heat exchanger. The casing 50 has a first and a second opposing radial surfaces 56, 58. The first radial surface 56, oriented radially inwards, is designed to be followed by the first portion 12b1 of the secondary flow. The second radial surface 58, oriented radially outwards, is designed to be followed by a second portion 12b2 of the secondary flow. It should be noted that the two portions 12b1, 12b2 separate from each other at a leading edge of the casing 50, corresponding to its aforementioned upstream end. These two flows 12bl, 12b2 then remix within the secondary flow 12b, at the level of a trailing edge of this casing, corresponding to its downstream end.One of the distinctive features of the invention lies in the implementation of air renewal within the exchanger, which is therefore not only supplied with fresh air by the first part 12bl of the secondary flow. This principle will be described later.
[0062] The assembly preferably includes a component 45b, located downstream of the heat exchanger 45a, and intended to be supplied by the secondary flow 12b. The second component 45b is arranged on the radial boundary surface 34a, or projects from this surface in the radial direction SI. It could, for example, be another heat exchanger, or, as schematically illustrated, an air scoop. As mentioned above, this scoop 45b can be arranged to project radially to form a protrusion in the flow, or it can be arranged flush with the surface 34a.
[0063] The scoop 45b, for example, is designed to supply fresh air to a turbine blade tip clearance control device 62. This device is partially housed in compartment 49, which includes one or more air circulation ducts. This device 62 is also known as an active blade tip clearance control system for high and / or low-pressure turbines.
[0064] The heat exchanger 45a can be aligned axially with the equipment 45b along the longitudinal direction L, or circumferentially offset from it. In the case of such alignment, there is therefore at least one fictitious longitudinal plane P of the assembly passing through axis 2, which also corresponds to the longitudinal central axis of this assembly, and traversing both equipment 45a and 45b. One of these planes P corresponds to the longitudinal cross-sectional plane in Figure 2.
[0065] For information purposes, it should be noted that the aforementioned alignment case is not limited to perfect alignment, but rather applies to all cases where there is at least one overlap zone in direction C between the two devices 45a and 45b. A mixer for the second part 12b2 of the secondary flow can be implemented on the downstream end of the casing, connecting the airflow exiting through the outlet 54 of the heat exchanger. This mixer increases the pressure and temperature redistribution downstream of the heat exchanger 45a, thus limiting flow stratification. Device 45b can then draw air at a cooler temperature and higher pressure, which helps to enhance the efficiency of the scoop 45b and that of the entire clearance control device 62 into which it is integrated.As an indication, it is noted that the mixer (not detailed) may include an alternation, according to direction C, of hollow and protruding elements, the latter preferably being chevron-shaped.
[0066] Returning to figure 2, assembly 15 has the following dimensions, in any fictitious longitudinal plane P passing through the exchanger 45a:
[0067] - H: radial height of secondary vein 14b, in which secondary flow 12b circulates, this height being considered at the level of a trailing edge of the casing 50, and being preferably between 0.5 and 5 m;
[0068] - L1: axial length of the exchanger, along the direction L, this length being preferably between 5 and 25 cm;
[0069] - RI: radial height of the heat exchanger, along the R direction, this height preferably being between 5 and 40 cm. Furthermore, these dimensions preferably meet the following parameter:
[0070] 0.001 < Ll / H < 0.3
[0071] With more specific reference to Figures 3 and 4, the assembly 15 according to the invention proposes a design for the heat exchanger 45a that allows for partial renewal of the cooling air, thereby increasing the heat exchange capacity. This improves the performance of the heat exchanger and can thus contribute to reducing its overall volume. Consequently, this results in a reduction of pressure losses and static pressure rises caused by the heat exchanger 45.
[0072] To achieve this, the exchanger 45a includes, in addition to the upstream inlet 52 and the downstream outlet 54, an intermediate air inlet 52a, as well as an intermediate air outlet 54a.
[0073] These intermediate inlet and outlet 52a, 54a pass through the casing 50, being arranged axially between the upstream end and the downstream end of the exchanger, corresponding here to the upstream and downstream ends of this casing 50. More precisely, they are arranged in an axially central zone 68 of the casing 50, for example by being crossed by a transverse median plane of the exchanger, orthogonal to the axis 2.
[0074] The intermediate inlet 52a is designed to be traversed by a PI sampling point on the second part 12b2 of the secondary flow. This intermediate inlet 52a may be designed such that it protrudes from the surface 58 of the housing 50, in the SI direction. Alternatively, the intermediate inlet 52a remains flush with this surface 58.
[0075] The intermediate outlet 54a is designed to be traversed by a primary portion 12bla of the first part 12bl of the secondary flow exiting the functional part 64 before the downstream outlet 54, in order to join the second part 12b2 of this secondary flow. These two flows 12b2 and 12bla then mix in an annular space delimited radially inwardly by the surface 58 of the casing 50. Here too, the intermediate outlet 52b can be designed such that it protrudes from the surface 58 of the casing 50, in the SI direction. Alternatively, the intermediate outlet 54a remains flush with this surface 58.
[0076] Furthermore, the PI sample and a secondary portion 12blb of the first part 12bl of the secondary flow, passing through the functional part 64 of the exchanger, are intended to be extracted from it by the downstream outlet 54. Before this extraction, the two flows 12blb, PI can remain independent within the functional part 64, or be mixed within it.
[0077] After extraction of these two flows 12blb, PI by the downstream outlet 54, downstream of the trailing edge of the cowling 50, they mix with the two flows 12b2, 12bla mentioned above or with a flow corresponding to the mixture of these two, in order to reconstitute the entire secondary flow 12b continuing to circulate downstream in the secondary vein.
[0078] The sizing and orientation of the inlets and outlets 52, 52a, 54, 54a allow control of the cooling air renewal rate within the heat exchanger 45. This rate, defined by the ratio between the flow rate of the sample PI passing through the intermediate inlet 52a and the flow rate of the first part 12bl of the secondary flow passing through the upstream inlet 52, is preferably between 0.1 and 0.6, and even more preferably between 0.3 and 0.5. In this respect, it is preferable to ensure that the flow rate of the sample PI passing through the intermediate inlet 52a is identical or very close to the flow rate of the primary portion 12bla of the first part 12bl of the secondary flow, passing through the intermediate outlet 54a.
[0079] In the preferred embodiment shown, the intermediate inlet 52a comprises several upstream-facing scoop-shaped inlet outlets. The intermediate outlet 54a comprises several downstream-facing ramp-shaped outlet outlets. Furthermore, the scoops 52a and ramps 54a are preferably arranged alternately along the circumferential direction C, as shown in Figure 4. The radial height RI' of the scoops and ramps 52a and 54a is preferably such that the ratio Rl' / Rl is between 0.05 and 0.1.
[0080] Of course, various modifications can be made by a person skilled in the art to the invention just 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 ducted turbofan engine, 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 relate to the outer radial boundary wall of the secondary flow, similarly provided with several components 45a, 45b.
[0081] According to yet other possible applications, the assembly according to the invention could relate to a radial delimitation surface, internal or external, of the primary vein.
[0082] The assembly according to the invention could also relate to a radial delimitation surface, internal or external, of a tertiary flow, when such a flow is provided within the propulsion system. In other words, the invention applies equally to single, dual, or triple flow propulsion systems.
[0083] Finally, it is noted that the radial boundary surface of the assembly according to the invention could be the radially external surface of nacelle 36a, corresponding to the external surface of the propulsion unit. This would be particularly possible for a turbomachine with an unfaired receiver, such as a turboprop or an open-rotor turbomachine. In the latter case, the radial boundary surface, as defined in the invention, is a radially external surface, shaped by the external airflow of the propulsion unit.
Claims
DEMANDS 1. Assembly (15) for an aircraft propulsion system (1), comprising: - a radial boundary structure (34) for the circulation of a gas flow (12b), the structure (34) comprising a radial boundary surface (34a) oriented in a first direction (SI) of a radial direction (R) of the assembly; - a heat exchanger (45a) projecting from the radial boundary surface (34a), in the first direction (SI) of the radial direction (R), the exchanger (45a) being configured to be traversed by a first part (12bl) of said gas flow, the exchanger also comprising, at a radial end thereof in the first direction (SI), a cowling (50) having a first and a second opposing radial surfaces (56, 58), the second radial surface (58) being intended to be followed by a second part (12b2) of said gas flow, the exchanger comprising, at its upstream end, an upstream inlet (52) intended to be traversed by the first part (12bl) of said gas flow, and, at its downstream end, a downstream outlet (54), characterized in that the casing (50) comprises, between the upstream end and the downstream end of the exchanger: - an intermediate inlet (52a) intended to be traversed by a sampling (PI) on the second part (12b2) of said gas flow; - an intermediate outlet (54a) intended to be traversed by a primary portion (12bla) of the first part (12bl) of said gas flow, in order to join the second part (12b2) of said gas flow, said sampling (PI) and a secondary portion (12blb) of the first portion (12bl) of said gas flow being intended to be extracted from the exchanger by said downstream outlet (54).
2. Assembly according to claim 1, characterized in that the intermediate inlet (52a) comprises several inlet ports, in that the intermediate outlet (54a) comprises several outlet ports, and in that the inlet and outlet ports are preferably arranged alternately along a circumferential direction (C) of the assembly.
3. Assembly according to claim 2, characterized in that the inlet ports (52a) are scoops oriented upstream, and in that the outlet ports (54a) are oriented downstream.
4. Assembly according to any one of the preceding claims, characterized in that the intermediate inlet (52a) and the intermediate outlet (54a) are arranged in an axially central area (68) of the casing (50) of the exchanger.
5. Assembly according to any one of the preceding claims, characterized in that it is configured so that a gas renewal rate, defined by the ratio between the flow rate of the sampling (PI) through the intermediate inlet (52a), and the flow rate of the first part (12bl) of said gas flow through the upstream inlet, is between 0.1 and 0.6, and preferably between 0.3 and 0.
5.
6. Assembly according to any one of the preceding claims, characterized in that it comprises equipment (45b) located downstream of the exchanger (45a), and intended to be supplied by the gas flow (12b), this equipment (45b) being arranged on the radial boundary 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 exchanger and the equipment (45a, 45b).
7. Assembly according to claim 6, characterized in that the equipment (45b) is a scoop or an exchanger.
8. Assembly according to any one of the preceding claims, characterized in that the radial boundary structure (34) also delimits a compartment (49) for housing equipment.
9. Assembly according to any one of the preceding claims, characterized in that the exchanger (45a) is arranged downstream of an outlet guide blade (30) of the assembly.
10. 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 turbofan engine, said radial boundary structure (34) of the assembly being preferably an internal or external radial boundary structure of a secondary flow (14b) of the turbofan engine, in which a secondary flow (12b) is intended to circulate.