Annular wall of an aeronautical turbomachine and aeronautical turbomachine comprising such a wall

Abstract

Annular wall (34) of an aeronautical turbomachine (50) having an axis (X) and comprising an aerodynamic surface (34A) configured to delimit a gas passage (V1), the wall (34) having a concavity (35) and being provided with a heat exchanger (10) comprising a heat exchanger matrix (12) and an aerodynamic profile (14), a heat transfer fluid circuit (20) extending at least into the matrix (12), the matrix (12) having a first portion (12) housed within the concavity (25) and a second portion (12B) outside the concavity (35), the aerodynamic profile (14) having a leading edge (14A) and a trailing edge (14B), the aerodynamic profile (14) being fixed to the matrix (12) such that the leading edge (14A) and the trailing edge (14B) extend circumferentially and the leading edge (14A) is outside the concavity (35). Figure for the abbreviation: Fig. 4.

Description

Title of the invention: Annular wall of an aeronautical turbomachine and aeronautical turbomachine comprising such a wall. Technical field

[0001] The present exposition relates to an annular wall of an aeronautical turbomachine and an aeronautical turbomachine comprising such a wall.

[0002] The term "aeronautical turbomachine" refers to all gas turbine devices that produce motive power, among which a distinction is made in particular between turbojets, which provide the thrust necessary for propulsion by reaction to the high-speed ejection of gas, and turboshaft engines, in which motive power is supplied by the rotation of a drive shaft. For example, turboshaft engines are used as engines for helicopters. As another example, turboprop engines (turboshafts driving a propeller) are turboshaft engines used as aircraft engines. Previous technique

[0003] Aeronautical turbomachinery generally features heat exchangers on annular walls radially defining a gas passage to cool a fluid, usually oil, which is used to cool and / or lubricate various components of the turbomachine, for example, bearings or other rotating parts. One such example is known from WO2023 / 001379.

[0004] Among known heat exchangers, there are two main families: air-cooled oil coolers, also known by the acronym ACOC for "Air Cooled OU Cooler," and surface air-cooled oil coolers, also known by the acronym SACOC for "Surface Air Cooled OU Cooler." ACOC exchangers are generally thermally efficient but generate significant pressure drops in the gas flow and cause static pressure rises that degrade the performance of upstream components. SACOC exchangers generate fewer pressure drops but are less thermally efficient.

[0005] In light of climate change, carbon emission restrictions are a constant concern for the Holder of this disclosure. Therefore, in order to obtain the most energy-efficient and environmentally friendly aeronautical components and turbomachinery possible, with a view to reducing the environmental impact of the aeronautical sector, particularly civil aviation, the Holder takes into account all elements that may have an impact in this regard, and in particular heat exchangers.

[0006] There is therefore a need to improve the thermal efficiency / pressure drop and pressure rise ratio of heat exchangers in aeronautical turbomachinery. Description of the invention

[0007] One embodiment relates to an annular wall of an aeronautical turbomachine, the wall having an axis and comprising an aerodynamic surface extending axially along the axis and circumferentially around the axis, the aerodynamic surface being configured to radially delimit a gas passage, the wall having a concavity extending radially from the aerodynamic surface, the wall being provided with a heat exchanger comprising a heat exchanger matrix and an aerodynamic profile, a heat transfer fluid circuit extending at least within the heat exchanger matrix, the heat exchanger matrix having a first portion housed in the concavity and a second portion extending radially outside the concavity and projecting radially from the aerodynamic surface, the aerodynamic profile having a leading edge and a trailing edge,the aerodynamic profile being fixed to the second portion of the heat exchanger matrix such that the leading edge and trailing edge extend circumferentially and the leading edge is radially distant from the aerodynamic surface and outside (i.e., on the outside) of the concavity.

[0008] The heat exchanger can be viewed as an element attached to the annular wall. In other words, the embodiment defined above can be seen as an assembly comprising the annular wall and the heat exchanger.

[0009] Generally, the axial direction corresponds to the direction of the axis of the annular wall, which, when mounted on an aircraft turbomachine, corresponds to the axis of rotation of the gas generator. A radial direction is a direction perpendicular to the axis / axial direction. The azimuthal or circumferential direction corresponds to the direction describing a ring around the axis / axial direction. The three directions axial, radial, and azimuthal (or circumferential) correspond respectively to the directions defined by the ridge, radius, and angle in a cylindrical coordinate system. Furthermore, upstream and downstream are defined with respect to the normal flow direction of the fluid (from upstream to downstream) through the aircraft turbomachine. Finally, unless otherwise specified, the adjectives inside / internal and outside / external are used with reference to the radial direction, such that the internal part (i.e.The radially internal part of an element is closer to the axis defining the axial direction than the external (i.e. radially external) part of the same element.

[0010] Hereafter, and unless otherwise specified, "wall" means "annular wall of an aeronautical turbomachine," "exchanger" means "heat exchanger," "matrix" means "heat exchanger matrix," and "profile" means "aerodynamic profile." For those skilled in the art, the term "matrix" refers to an actual part of the exchanger, comprising channels through which the heat transfer fluid circulates to exchange heat with the surrounding fluid of the matrix.

[0011] The annular wall may be a sector of a ring or a complete ring. The complete ring may be formed by a ring made from a single piece, or by a plurality of ring sectors. The heat exchanger may extend continuously over the entire circumferential extent of the annular wall, or only over annular sectors spaced apart in the circumferential direction. In the latter case, the aerodynamic profile may have the same circumferential extent as the heat exchanger matrix, or it may have a greater circumferential extent than the heat exchanger matrix, for example, by extending symmetrically by the same amount in the circumferential direction on each side of the heat exchanger matrix.According to one variant, the matrix can extend only over annular sectors spaced apart in the circumferential direction, while the profile can extend continuously over the entire circumferential extent of the wall and be fixed to all the annular sectors of the matrix.

[0012] The annular wall may be a radially internal or radially external wall of a gas stream, or an external wall of a nacelle, such an external wall delimiting the passage of gas outside the nacelle (i.e., the passage of air circulating around the nacelle). The concavity may extend radially inward if the annular wall is a radially internal wall, or an external wall of a nacelle. The concavity may extend radially outward if the annular wall is a radially external wall. When the concavity extends radially outward, the second portion of the heat exchanger matrix protrudes radially inward from the annular wall.

[0013] The heat transfer fluid can be a fuel or a lubricant, for example oil, but not necessarily.

[0014] For example, the leading edge and trailing edge of the airfoil can extend parallel to the aerodynamic surface of the wall. The opposite sides of the airfoil in the radial direction can respectively form an intrados and an extrados extending from the leading edge to the trailing edge. For example, The lower surface of the airfoil can be fixed to the matrix. For example, the trailing edge can be radially distant from the airfoil surface and outside the concavity. Alternatively, the entire airfoil can be radially distant from the airfoil surface and positioned outside the concavity.

[0015] Viewed from the perspective of the gas flow circulating along the wall, the heat exchanger has a portion "recessed" within the wall, corresponding to the first part of the matrix located within the concavity, and a portion protruding from the wall, corresponding to the second part of the matrix equipped with the aerodynamic profile. Since the leading edge of the profile is radially distant from the wall, it can guide a portion of the gas flow towards the concavity and towards the matrix. Such guidance is less intrusive and less restrictive to the gas flow than a prior art flow deflection, thus minimizing pressure losses and static pressure rises induced by the heat exchanger.

[0016] Such a configuration can therefore make it possible to obtain a satisfactory thermal efficiency, by the combined use of the matrix and the profile which also participates in the heat exchange, while reducing pressure losses as much as possible.

[0017] For example, the cross-sectional area considered at the trailing edge between the wall and the airfoil is larger than the cross-sectional area considered at the leading edge between the wall and the airfoil. This can help to reduce the induced disturbance in the gas flow and pressure losses as much as possible.

[0018] In some embodiments, the heat exchanger may include at least one fin extending radially from the airfoil and axially along the airfoil, opposite the heat exchanger matrix.

[0019] Hereafter and unless otherwise indicated, "fin" means "at least one fin".

[0020] The fin can extend perpendicularly to the airfoil, for example on its upper surface, between the leading and trailing edges. The fin can improve heat exchange, for example by convection, while its parallel extension to the direction of the gas flow induces a very small disturbance within the gas flow. In other words, combining the fins on the airfoil with the matrix increases heat exchange with minimal pressure drop and without the heat exchanger occupying significantly more space on the wall. With the fins, the length and azimuthal extent are essentially the same as with the matrix alone, but more heat exchange is achieved with the fins using less additional height than if the matrix were increased.

[0021] In some embodiments, the heat transfer fluid circuit can extend into the aerodynamic profile.

[0022] By "in the aerodynamic profile", it is meant that the heat transfer fluid circuit extends within the volume of the profile itself, between the lower surface, the upper surface, the leading edge and the trailing edge.

[0023] In some embodiments, the heat transfer fluid circuit can extend into the wall.

[0024] By "in the wall", we mean in the volume of the wall itself, and not in an element added to the wall such as the heat exchanger.

[0025] In some embodiments, the heat transfer fluid circuit can extend into at least one fin.

[0026] By "in at least one fin" is meant in the volume of at least one of the fins, and not in an adjacent element such as the profile.

[0027] Such a fluid circuit according to one or the other or a combination of the above configurations can improve heat exchange, without impacting the pressure drop induced by the exchanger.

[0028] For example, the heat exchanger can be manufactured in whole or in part by additive manufacturing. This can make it possible to optimize both the external geometric shape of the exchanger and the heat transfer fluid circuit that extends within the exchanger.

[0029] In some embodiments, the heat exchanger can extend only over annular sectors spaced circumferentially from each other.

[0030] In other words, considered in the circumferential direction, adjacent annular sectors of the heat exchanger are spaced apart. Each annular sector has a limited circumferential extent, each sector being distinct from the others. The annular sectors can be regularly distributed in the circumferential direction. According to one embodiment, all the annular sectors of the heat exchanger can be arranged in the same axial position, so that they are all aligned in the circumferential direction.

[0031] An exchanger comprising such annular sectors can exhibit satisfactory thermal efficiency while minimizing the induced pressure loss.

[0032] In some embodiments, a ratio between a maximum axial length of the aerodynamic profile and a maximum axial length of the heat exchanger matrix can be between 1.0 and 1.4, for example between 1.2 and 1.3.

[0033] In some embodiments, a ratio between a maximum radial thickness of the aerodynamic profile and a maximum radial thickness of the heat exchanger matrix can be between 0.05 and 0.35, for example between 0.10 and 0.25.

[0034] Such ratios can ensure a good balance between the thermal efficiency of the exchanger and the disturbance induced within the gas flow.

[0035] In some embodiments, the aerodynamic profile may have a wing profile shape.

[0036] An airfoil shape may have a leading edge with a rounded shape, i.e., having a non-zero radius, without geometric discontinuity or angle. For example, an airfoil shape may be a NACA-type airfoil shape.

[0037] Such a wing profile shape can allow part of the gas flow to be diverted towards the matrix while minimizing the disturbance induced in this flow, and in particular the pressure loss.

[0038] An embodiment relates to an aeronautical turbomachine comprising an annular wall of an aeronautical turbomachine according to any one of the embodiments described in this presentation.

[0039] The aeronautical turbomachine can be a turbojet engine with a shrouded or unshrouded fan. For example, the fan can be driven directly by a gas generator or via a speed reducer. For example, the fan diameter, measured at the leading edge of the fan blades, can be between 50.8 cm and 508 cm (or between 20 inches and 200 inches). For example, the bypass ratio, also known as BPR, which is the ratio of the mass flow rate of the secondary airflow (i.e., the gas flow from the fan that does not enter the gas generator) divided by the mass flow rate of the primary airflow (i.e., the gas flow from the fan that enters the gas generator), can be between 8 and 40, for example, between 8 and 18 for shrouded fans, and between 18 and 40 for unshrouded fans.For example, the compression ratio, or "fan pressure ratio" in English, also known by the acronym FPR, which corresponds to the pressure ratio between the inlet of the fan module, upstream of the fan, and the outlet of the fan module, downstream of the rectifier, can be between 1.0 and 1.5.

[0040] In some embodiments, the aeronautical turbomachine may include a hot gas stream and a cold gas stream, in which the wall radially delimits at least a part of the cold gas stream.

[0041] The hot gas vein is sometimes called by those skilled in the art a "primary vein" while the cold gas vein is sometimes also called by those skilled in the art a "secondary vein". Brief description of the drawings

[0042] The purpose of this presentation and its advantages will be better understood upon reading the detailed description below of various embodiments given by way of non-limiting examples. This description refers to the attached figure pages, on which:

[0043] [Fig-1] Figure [Fig.1] represents an aircraft equipped with an aircraft turbomachine,

[0044] [Fig.2] Fig.2 schematically represents the main internal organs of the aircraft turbomachine of the aircraft of the [Fig.1],

[0045] [Fig.3] Fig.3 schematically represents different directory walls of the aircraft turbomachine of the aircraft of the [Fig.1],

[0046] [Fig.4] Figure [Fig.4] shows in more detail an annular portion of the heat exchanger thermal image of [Fig. 3], seen in perspective,

[0047] [Fig. 5] Fig. 5 represents a first variant of the heat exchanger, and

[0048] [Fig.6] The [Fig.6] represents a second variant of the heat exchanger. Description of the implementation methods

[0049] Figure 1 represents an aircraft 100, in this example an airplane, equipped with two aircraft turbomachines 50, in this example two propulsion units 50, in this example two turbojet engines 50, namely one aircraft turbomachine 50 per wing 101, with only one aircraft turbomachine 50 and one wing 101 being shown in Figure 1. According to one variant, the aircraft 100 can be equipped with more than one aircraft turbomachine 50 per wing 101, each wing 101 being fitted with the same number of aircraft turbomachines 50.

[0050] Figure 2 shows a schematic cross-sectional view of the main internal components of the aircraft turbomachine 50, according to plane II of Figure 1. The aircraft turbomachine 50 comprises a fan 52, which may be shrouded or unshrouded, and a gas generator 54 with axis of rotation X. In the example of Figure 1, the fan is shrouded, the shroud not being shown in Figure 2 but shown in Figure 3. In this example, the gas generator 54 comprises, from upstream to downstream, a compressor 54A (or compressor section 54A), a combustion chamber 54B, and a turbine 54C (or turbine section 54C). The fan 52 can be driven in rotation directly by a shaft of the gas generator 54, for example a shaft of a low pressure body, or via a speed reducer not shown.

[0051] The gas generator 54 may be of the twin-spool type and comprise a low-pressure spool 60A and a high-pressure spool 60B. The low-pressure spool 60A may comprise a low-pressure compressor 62A rotationally coupled to a low-pressure turbine 66A via a low-pressure shaft 63A. The high-pressure spool 60B may comprise a high-pressure compressor 62B disposed downstream of the low-pressure compressor 62A and upstream of the combustion chamber 54B, and a high-pressure turbine 66B disposed downstream of the combustion chamber 54B and upstream of the low-pressure turbine 66A, and rotationally coupled to the high-pressure compressor 62B via a high-pressure shaft 63B. The compressor 54A of the gas generator 54 may comprise the low-pressure and high-pressure compressors. 62A and 62B. The turbine 54C of the gas generator 54 may include the low and high pressure turbines 66A and 66B. [Fig.2] is schematic, each compressor and each turbine being able to have one or more stages, each stage comprising a rotating wheel and a stator or rectifier.

[0052] Figure 3 shows a schematic cross-sectional view of the aircraft turbomachine 50, according to Figure 1 in Figure 1, representing the annular walls of the turbomachine 50. The compressor 54A, in particular the moving and stator blades, the combustion chamber 54B, and the turbine 54C, in particular the moving and stator blades, are symbolized by hatched areas. The gas generator 54 is not shown in further detail for the sake of clarity.

[0053] In the example of [Fig. 3], the turbomachine 50 is provided with five annular walls 30, 32, 34, 36, and 38 with axis X and comprising, respectively, an aerodynamic surface 30A, 32A, 34A, 36A, and 38A extending axially along the axis X and circumferentially around the axis X (i.e., along the circumferential direction C). The annular wall 30 is an external wall of the nacelle 40 and delimits the passage of gas external to the nacelle 40. The walls 32 and 34 delimit a cold gas stream (or passage) VI, while the walls 36 and 38 delimit a hot gas stream (or passage) V2. Wall 32 is a radially external wall of the V1 vein while wall 34 is a radially internal wall of the VL vein. Wall 36 is a radially external wall of the V2 vein while wall 38 is a radially internal wall of the V2 vein.When the turbomachine 50 is operating, the V2 channel is configured to guide a primary gas flow Fl and the V1 channel is configured to receive a secondary gas flow F2.

[0054] In the example of [Fig.3] (see also [Fig.4]), the annular wall 34 has a concavity 35 extending radially (i.e. along the radial direction R) from the aerodynamic surface 34A. The wall 34 is provided with a heat exchanger 10 comprising a heat exchanger matrix 12 and an aerodynamic profile 14. A heat transfer fluid circuit 20 extends at least within the heat exchanger matrix 12. In this example, the circuit 20 is symbolized by channels 20A, and may also extend into the wall 34 and / or the profile 14. The heat exchanger matrix 12 has a first portion 12A housed in the concavity 35 and a second portion 12B extending radially outside the concavity 35 and projecting radially from the aerodynamic surface 34A. In [Fig.[4], the opening of the cavity 35 is symbolized by the dashed line 35A extending in the geometric continuation of the parts of the surface 34A upstream and downstream with respect to the cavity 35. The line 35A represents the radial limit of the concavity 35. The aerodynamic profile 14 has a leading edge 14A and a trailing edge 14B, the aerodynamic profile 14 being fixed on the second portion 12B of the heat exchanger matrix 12 so that. that the leading edge 14A and the trailing edge 14B extend circumferentially and that the leading edge 14A is radially distant from the aerodynamic surface 34A and outside the concavity 35. In other words, in this example the leading edge 14A is located outside the concavity 35, and adjoins the line 35A or is distant from the line 35A in the radial direction R. In this example, the trailing edge 14B is also radially distant from the aerodynamic surface 34A and outside the concavity 35. In this example, the entire airfoil 14 is located outside the concavity 35. In this example, the wall 34 forms a wall radially delimiting at least part of the cold gas stream VI. In this example, the concavity 35 and the exchanger 10 are located downstream of the rectifier 70, which is itself located downstream of the fan 52, and upstream of the SI output of the MF fan module, the MF fan module comprising the fan 52, the rectifier 70 and the nacelle 40.

[0055] As can be seen in [Fig. 4], the leading edge 14A and the trailing edge 14B of the airfoil 14 can extend parallel to the aerodynamic surface 34A of the wall 34 in the circumferential direction C. In this example, the airfoil 14 has an intrados 14C and an extrados 14D, each extending between the leading edge 14A and the trailing edge 14B and forming two opposite faces in the radial direction R of the airfoil 14. In this example, the airfoil 14 is attached to the matrix 12 via the intrados 14C. In this example, the airfoil 14 has a wing profile shape. In this example, the leading edge 14A has a rounded shape, i.e., a non-zero radius, without any geometric discontinuity or angle.

[0056] As can be seen in [Fig.4], the heat exchanger 10 may include at least one fin 22 extending radially from the airfoil 14 and axially along the airfoil 14, opposite the heat exchanger matrix 12. The heat transfer fluid circuit 20 may extend into all or part of the fins 22 (not shown).

[0057] For example, an LMP / LMM ratio between a maximum axial length of the LMP airfoil and a maximum axial length of the LMM heat exchanger matrix can be between 1.0 and 1.4, for example between 1.2 and 1.3.

[0058] In this example, the maximum axial length of the LMM matrix can be between 5.0 cm and 25.5 cm (or between 2 inches and 10 inches), while the maximum axial length of the LMP profile can be between 5.0 cm and 25.5 cm (or between 2 inches and 10 inches). The maximum axial length of the LMM matrix can be less than or equal to the maximum axial length of the LMP profile. The maximum axial length LMA of the fins 22 can be less than or equal to the maximum axial length of the LMP profile and greater than or equal to the maximum axial length of the LMM matrix.

[0059] For example, an EMP / EMM ratio between a maximum radial thickness of the aerodynamic profile EMP and a maximum radial thickness of the heat exchanger matrix EMM can be between 0.05 and 0.35, for example between 0.10 and 0.25. It is noted that the axial position of the maximum radial thickness of the profile 14 and the axial position of the maximum radial thickness of the matrix 12 do not necessarily coincide, as illustrated in the example of [Fig.4].

[0060] In the example of [Fig. 4], the radial thickness EA of each of the fins 22 is constant over the entire axial extent of the fins 22, and can be between 1.25 cm and 5.00 cm (or 0.5 inch and 2.0 inches). According to an alternative not shown, the radial thickness EA of the fins 22 can vary along the axial direction X, and the average value of the radial thickness EA can be between 1.25 cm and 5.00 cm (or 0.5 inch and 2.0 inches).

[0061] The maximum radial thickness of the EM heat exchanger, including optional fins 22, can be between 5.0 cm and 13.0 cm (or between 2 inches and 5 inches).

[0062] Figure 5 represents a first variant seen along the section plane V of Figure 5 shows a second variant along the cross-sectional plane V of Figure 5, in which the heat exchanger 10, in particular the matrix 12 and the profile 14, extends continuously along the entire circumferential extent of the wall 34. Figure 6 represents a second variant along the cross-sectional plane V of Figure 5, in which the heat exchanger 10, in particular the matrix 12 and the profile 14, extends only over annular sectors 10' spaced circumferentially from one another. In this second variant, the annular sectors 10' are regularly spaced from one another. In other words, the angular sectors 10' are regularly distributed along the circumferential direction C. As can be seen in [Fig.6], for each annular sector 10', the circumferential extent of the profile 14 is greater than the circumferential extent of the matrix 12. In this example, the profile 14 extends circumferentially beyond the matrix 12 by the same extent on each side of the matrix 12.

[0063] The heat transfer fluid circulating in the circuit 20 can be a lubricant, for example oil, which can, for example, be used for the lubrication and / or cooling of one or more bearings not shown of one or both of the low and high pressure shafts 63A, 63B. Instead of or in addition to this, the oil can, for example, be used for the lubrication and / or cooling of a possible speed reducer coupling the low pressure shaft 63A and the fan 52 in rotation.

[0064] In the examples described above, only the wall 34 is provided with a concavity 35 housing a heat exchanger 10. According to an alternative not shown, in addition to or instead of the wall 34, the wall 30 and / or the wall 32 and / or the wall 36 and / or the wall 38 may be provided with a concavity housing a heat exchanger, in a manner similar to the concavity 35 and the heat exchanger 10 described with reference to Figures 3 to 6. For example, in the case where wall 36 and / or wall 38 delimiting the hot gas stream is / are provided with a concavity housing a heat exchanger, the concavity and the exchanger can, for example, be arranged upstream of compressor 54A.

[0065] In the examples described above, the wall 34 being a radially internal wall, the concavity 35 extends radially inwards and the exchanger 10, in particular the second portion 12B of the matrix 12, protrudes radially outwards relative to the wall 34. Similarly, in the case where the wall 30 or the wall 38 is provided with a concavity and an exchanger, the concavity extends radially inwards and the exchanger, in particular the second portion of the matrix, protrudes radially outwards relative to the wall 30 or 38, respectively. Conversely, in the case where wall 32 or wall 36, which are radially external walls, are provided with a concavity and an exchanger, the concavity extends radially outwards and the exchanger, in particular the second portion of the matrix, protrudes radially inwards relative to wall 32 or 36, respectively.

[0066] Although the present invention has been described with reference to specific embodiments, it is evident that modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, individual features of the various embodiments illustrated / mentioned can be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than a restrictive sense.

Claims

Demands

1. Annular wall (34) of an aeronautical turbomachine (50), the wall (34) having an axis (X) and comprising an aerodynamic surface (34A) extending axially along the axis (X) and circumferentially around the axis (X), the aerodynamic surface (34A) being configured to radially delimit a gas passage (VI), the wall (34) having a concavity (35) extending radially from the aerodynamic surface (34A), the wall (34) being provided with a heat exchanger (10) comprising a heat exchanger matrix (12) and an aerodynamic profile (14), a heat transfer fluid circuit (20) extending at least into the heat exchanger matrix (12), the heat exchanger matrix (12) having a first portion (12) housed in the concavity (25) and a second portion (12B) extending radially outside the concavity (35) and projecting radially from the aerodynamic surface (34A),the aerodynamic profile (14) having a leading edge (14A) and a trailing edge (14B), the aerodynamic profile (14) being fixed to the second portion (12B) of the heat exchanger matrix (12) such that the leading edge (14A) and the trailing edge (14B) extend circumferentially and the leading edge (14A) is radially distant from the aerodynamic surface (34A) and outside the concavity (35), in which the heat exchanger (10) comprises at least one fin (22) extending radially from the aerodynamic profile (14), opposite the heat exchanger matrix (12), and axially along the aerodynamic profile (14), the heat transfer fluid circuit (20) extending within the aerodynamic profile (14).

2. Annular wall (34) of aeronautical turbomachine (50) according to claim 1, wherein the heat transfer fluid circuit (20) extends into the wall (34).

3. Annular wall (34) of aeronautical turbomachine (50) according to claim 1 or 2, wherein the heat exchanger (10) extends only over annular sectors (10') spaced circumferentially from each other.

4. Annular wall (34) of an aeronautical turbomachine (50) according to any one of claims 1 to 3, wherein a ratio between a maximum axial length of the airfoil (MPL) and a The maximum axial length of the heat exchanger matrix (LMM) is between 1.0 and 1.4, for example between 1.2 and 1.

3.

5. Annular wall (34) of an aeronautical turbomachine (50) according to any one of claims 1 to 4, wherein a ratio between a maximum radial thickness of the airfoil (EMP) and a maximum radial thickness of the heat exchanger matrix (EMM) is between 0.05 and 0.35, for example between 0.10 and 0.

25.

6. Annular wall (34) of an aeronautical turbomachine (50) according to any one of claims 1 to 5, wherein the aerodynamic profile (14) has a wing profile shape.

7. Aeronautical turbomachine (50) comprising an annular wall of an aeronautical turbomachine (34) according to any one of claims 1 to 6.

8. Aeronautical turbomachine (50) according to claim 7, comprising a hot gas stream (V2) and a cold gas stream (VI), in which the wall (34) radially delimits at least a portion of the cold gas stream (VI).

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