Annular wall of an aeronautical turbine engine and aeronautical turbine engine comprising such a wall

Abstract

The invention relates to an annular wall (34) of an aeronautical turbine engine (50) having an axis (X) and comprising an aerodynamic surface (34A) configured to delimit a gas passage (V1), the wall (34) having a cavity (35) and being provided with a heat exchanger (10) comprising a heat exchanger matrix (12) and an aerofoil (14), a heat transfer fluid circuit (20) extending at least inside the matrix (12), the matrix (12) having a first portion (12) housed inside the cavity (25) and a second portion (12B) outside the cavity (35), the aerofoil (14) having a leading edge (14A) and a trailing edge (14B), the aerofoil (14) being attached to the matrix (12) so that the leading edge (14A) and the trailing edge (14B) extend circumferentially and so that the leading edge (14A) is outside the cavity (35).

Description

Description Title of the invention: Annular wall of an aeronautical turbomachine and aeronautical turbomachine comprising such a wall Technical Field

[0001] The present disclosure 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 producing motive power, including turbojets providing the thrust necessary for propulsion by reaction to the high-speed ejection of gas, and turboshafts in which motive power is provided by the rotation of an engine shaft. For example, turboshafts are used as engines for helicopters. In another example, turboprops (turboshafts driving a propeller) are turboshafts used as aircraft engines. Prior art

[0003] Aeronautical turbomachines generally have heat exchangers on annular walls radially delimiting a gas passage in order to cool a fluid, generally oil, which is used to cool and / or lubricate various elements of the turbomachine, for example bearings or other rotating elements. Such an example is known from W02023 / 001379.

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

[0005] Given climate change, carbon emissions restrictions are a constant concern for the Holder of this presentation. Also, in order to obtain the least energy-consuming and most efficient aeronautical components and turbomachines, as environmentally friendly as possible with a view to reducing the environmental consequences of the aeronautics sector, particularly civil, the Holder takes into account all elements that could have an impact to this end, and in particular heat exchangers.

[0006] There is therefore a need to improve the thermal efficiency / pressure drop and pressure rise ratio of aeronautical turbomachine heat exchangers. Statement 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 in 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 relative to the aerodynamic surface, the aerodynamic profile having a leading edge and a trailing edge,the aerodynamic profile being fixed on the second portion of the heat exchanger matrix so that the leading edge and the trailing edge extend circumferentially and so that the leading edge is radially distant from the aerodynamic surface and outside (i.e. outside) the concavity.,

[0008] The heat exchanger can be seen 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 speaking, the axial direction corresponds to the direction of the axis of the annular wall, which corresponds, when the latter is mounted on an aeronautical turbomachine, 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 coast, the radius and the angle in a cylindrical coordinate system. In addition, upstream and downstream are defined with respect to the normal flow direction of the fluid (from upstream to downstream) through the aeronautical turbomachine. Finally, unless otherwise specified, the adjectives interior / internal and exterior / external are used in reference to the radial direction so that the internal part (i.e. radially internal) of an element is closer to the axis defining the axial direction than the external part (i.e. radially external) of the same element.

[0010] Hereinafter and unless otherwise indicated, by "wall" is meant "annular wall of aeronautical turbomachine", by "exchanger" is meant "heat exchanger", by "matrix" is meant "heat exchanger matrix" and by "profile" is meant "aerodynamic profile". As a reminder, the term "matrix" designates for those skilled in the art an effective part of the exchanger, which comprises channels in which the heat transfer fluid circulates in order to exchange calories with the fluid surrounding the matrix.

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

[0012] The annular wall may be a radially inner or radially outer wall of a gas stream, or an outer wall of a nacelle, such an outer wall delimiting the gas passage external to the nacelle (i.e. the air passage which circulates around the nacelle). The concavity may extend radially inward if the annular wall is a radially inner wall, or an outer wall of a nacelle. The concavity may extend radially outward if the annular wall is a radially outer wall. When the concavity extends radially outward, the second portion of the heat exchanger matrix projects radially inward relative to the annular wall. When the concavity extends radially inward, the second portion of the heat exchanger matrix projects radially outward relative to 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 the trailing edge of the airfoil may extend parallel to the aerodynamic surface of the wall. The opposite sides of the profile in the radial direction may respectively form a lower surface and an upper surface extending from the leading edge to the trailing edge. For example, the lower surface of the airfoil may be fixed to the die. For example, the trailing edge may be radially distant from the aerodynamic surface and outside the concavity. For example, the entire airfoil may be radially distant from the aerodynamic surface and arranged outside the concavity.

[0015] Seen from the gas flow circulating along the wall, the exchanger has a "retracted" part within the wall, which corresponds to the first portion of the matrix which is arranged within the concavity, and a part projecting from the wall, which corresponds to the second portion of the matrix provided with the aerodynamic profile. The leading edge of the profile being 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 for the gas flow than a flow diversion of the state of the art, which makes it possible to minimize the pressure losses and the rises in static pressure induced by the exchanger.

[0016] Such a configuration can therefore make it possible to obtain satisfactory thermal efficiency, through 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 passage section considered at the trailing edge between the wall and the profile is larger than the passage section considered at the leading edge between the wall and the profile. This can help to reduce as much as possible the disturbance induced in the gas flow and the pressure losses.

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

[0019] Hereinafter and unless otherwise indicated, “fin” means “at least one fin”.

[0020] The fin may extend perpendicular to the profile, for example on its extrados, between the leading edge and the trailing edge. The fin may improve heat exchange, for example by convection, while inducing, by its extent parallel to the direction of the gas flow, a very reduced disturbance within the gas flow. In other words, combining the fins on the profile with the matrix allows to increase the heat exchange with a minimal pressure drop and without the exchanger occupying significantly more space on the wall. With the fins, we have roughly the same length and azimuthal extension as the matrix alone, but we make more heat exchange thanks to the fins with less additional height than if we increased the matrix.

[0021] In some embodiments, the heat transfer fluid circuit may extend into the airfoil.

[0022] By "in the aerodynamic profile" we mean that the heat transfer fluid circuit extends within the volume of the profile itself, between the intrados, the extrados, the leading edge and the trailing edge.

[0023] In some embodiments, the heat transfer fluid circuit may 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 may extend into the at least one fin.

[0026] By "in the 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 exchanges, 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 the external geometric shape of the exchanger, but also the heat transfer fluid circuit that extends within the exchanger.

[0029] In some embodiments, the heat exchanger may extend only over annular sectors circumferentially spaced from one another.

[0030] In other words, considered in the circumferential direction, the adjacent annular sectors of the exchanger are distant from each other. Each annular sector has a limited circumferential extent, each sector being distinct of the other sectors. The annular sectors can be regularly distributed in the circumferential direction. Alternatively, all the annular sectors of the 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 of a maximum axial length of the airfoil to a maximum axial length of the heat exchanger matrix may 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 airfoil and a maximum radial thickness of the heat exchanger matrix may 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 airfoil may have a wing profile shape.

[0036] A wing profile shape may have a leading edge that is rounded, i.e., having a non-zero radius, without geometric discontinuity or angle. For example, a wing profile shape may be a "NACA" type profile 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 drop.

[0038] One embodiment relates to an aeronautical turbomachine comprising an annular aeronautical turbomachine wall according to any one of the embodiments described in the present disclosure.

[0039] The aeronautical turbomachine can be a turbojet with a ducted or unducted fan. For example, the fan can be driven directly by a gas generator or via a speed reducer. For example, the diameter of the fan, considered at the tip of the fan blades at the leading edge, can be between 50.8 cm and 508 cm (or between 20 inches and 200 inches). For example, the bypass ratio, or "bypass ratio" in English, also known by the acronym BPR, which corresponds to the ratio of the mass flow rate of the secondary air flow (i.e. the gas flow from the fan that does not enter the gas generator) divided by the mass flow rate of the primary air flow (i.e. the gas flow from the fan that enters the gas generator), can be understood between 8 and 40, for example between 8 and 18 for ducted fans, for example between 18 and 40 for unducted 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 comprise a hot gas stream and a cold gas stream, wherein the wall radially delimits at least a portion of the cold gas stream.

[0041] The hot gas stream is sometimes referred to by those skilled in the art as the "primary stream" while the cold gas stream is sometimes also referred to by those skilled in the art as the "secondary stream". Brief description of the drawings

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

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

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

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

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

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

[0048] [Fig. 6] Figure 6 shows a second variant of the heat exchanger. Description of the embodiments

[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 turbojets 50, namely one aircraft turbomachine 50 per wing 101, a single aircraft turbomachine 50 and a single wing 101 being represented on the figure 1. According to a variant, the aircraft 100 can be equipped with more than one aircraft turbomachine 50 per wing 101, each wing 101 being provided with the same number of aircraft turbomachines 50.

[0050] Figure 2 represents a schematic 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 an 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, the gases flowing within the aircraft turbomachine 50 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 by means of a speed reducer not shown.

[0051] The gas generator 54 may be of the double-body type and comprise a low-pressure body 60A and a high-pressure body 60B. The low-pressure body 60A may comprise a low-pressure compressor 62A rotatably coupled with a low-pressure turbine 66A via a low-pressure shaft 63A. The high-pressure body 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 rotatably coupled with the high-pressure compressor 62B via a high-pressure shaft 63B. The compressor 54A of the gas generator 54 may comprise the low- 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.Figure 2 is schematic, each compressor and each turbine being able to have one or more stages, each stage comprising a moving wheel and a stator or rectifier.

[0052] Figure 3 represents a schematic sectional view of the aircraft turbomachine 50, according to plane II of 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 more detail for the readability of the figure.

[0053] In the example of Figure 3, the turbomachine 50 is provided with five annular walls 30, 32, 34, 36 and 38 of axis X and respectively comprising an aerodynamic surface 30A, 32A, 34A, 36A and 38A extending axially along the axis X and circumferentially around the axis X (i.e. in the circumferential direction C). The annular wall 30 is an external wall of nacelle 40, and delimits the gas passage external to the nacelle 40. The walls 32 and 34 delimit a vein (or passage) of cold gas V1 while the walls 36 and 38 delimit a vein (or passage) of hot gas V2. The wall 32 is a radially external wall of the vein V1 while the wall 34 is a radially internal wall of the vein V1. Wall 36 is a radially outer wall of vein V2 while wall 38 is a radially inner wall of vein V2.When the turbomachine 50 is operating, the vein V2 is configured to guide a flow of gas called primary F1 and the vein V1 is configured to receive a flow of gas called secondary F2.

[0054] In the example of Figure 3 (see also Figure 4), the annular wall 34 has a concavity 35 extending radially (i.e. in 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 in the heat exchanger matrix 12. In this example, the circuit 20 is symbolized by channels 20A, and can also extend in the wall 34 and / or in 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 relative to the aerodynamic surface 34A.In Figure 4, the orifice of the cavity 35 is symbolized by the broken line 35A extending in the geometric extension of the parts of the surface 34A upstream and downstream relative 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 the leading edge 14A and the trailing edge 14B extend circumferentially and so 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 arranged outside the concavity 35, and adjoins the line 35A or is distant from 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 profile 14 is arranged outside the concavity 35. In this example, the wall 34 forms a wall radially delimiting at least one. part of the cold gas stream V1. In this example, the concavity 35 and the exchanger 10 are arranged downstream of the rectifier 70, itself arranged downstream of the fan 52, and upstream of the outlet S1 of the fan module MF, the fan module MF comprising the fan 52, the rectifier 70 and the nacelle 40.

[0055] As can be seen in Figure 4, the leading edge 14A and the trailing edge 14B of the aerodynamic profile 14 can extend parallel to the aerodynamic surface 34A of the wall 34 in the circumferential direction C. The profile 14 has in this example a lower surface 14C and an upper surface 14D, which each extend between the leading edge 14A and the trailing edge 14B and form two opposite faces in the radial direction R of the profile 14. In this example, the profile 14 is fixed to the matrix 12 via the lower surface 14C. In this example, the profile 14 has a wing profile shape. In this example, the leading edge 14A has a rounded shape, i.e. having a non-zero radius, without geometric discontinuity or angle.

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

[0057] For example, a LMP / LMM ratio between a maximum axial length of the aerodynamic profile LMP and a maximum axial length of the heat exchanger matrix LMM 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 die may 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 may be between 5.0 cm and 25.5 cm (or between 2 inches and 10 inches). The maximum axial length of the LMM die may be less than or equal to the maximum axial length of the LMP profile. The maximum axial length LMA of the fins 22 may 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 die.

[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 may 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 of the radial thickness of the profile 14 and the axial position of the maximum of the radial thickness of the matrix 12 do not necessarily coincide, as illustrated in the example of FIG. 4.

[0060] In the example of Figure 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 a variant not shown, the radial thickness EA of the fins 22 can be variable 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 22 fins, 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, in which the 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 section 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 Figure 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 die 12 by the same extent on each side of the die 12.

[0063] The heat transfer fluid circulating in the circuit 20 may be a lubricant, for example oil, which may for example be used for the lubrication and / or cooling of one or more bearings (not shown) of one or both low and high pressure shafts 63A, 63B. Instead of or in addition, the oil may 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 a variant 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 similar manner to the concavity 35 and the exchanger 10 described with reference to FIGS. 3 to 6. For example, in the case where the wall 36 and / or the wall 38 delimiting the hot gas stream, is / are provided with a concavity housing a heat exchanger, the concavity and the exchanger may for example be arranged upstream of the 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, projects radially outwards relative to the wall 34. Similarly, in the case where the wall 30 or the wall 38 are provided with a concavity and an exchanger, the concavity extends radially inwards and the exchanger, in particular the second portion of the matrix, projects radially outwards relative to the wall 30 or 38, respectively. Conversely, in the case where the wall 32 or the 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, projects radially inwards relative to the wall 32 or 36, respectively.

[0066] Although the present invention has been described with reference to specific embodiments, it is obvious that modifications and changes may 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 illustrated / mentioned embodiments may be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

Claims

Claims

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 (V1), 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 in 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 outwardly from the concavity (35) and projecting radially relative to the aerodynamic surface (34A),the aerodynamic profile (14) having 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 the leading edge (14A) and the trailing edge (14B) extend circumferentially and so that the leading edge (14A) is radially distant from the aerodynamic surface (34A) and outside the concavity (35), wherein 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 in the aerodynamic profile (14).,

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

3. Annular wall (34) of an aeronautical turbomachine (50) according to claim 1 or 2, in which 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, in which a ratio between a maximum axial length of the aerodynamic profile (LMP) and a 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, in which a ratio between a thickness maximum radial thickness of the aerodynamic profile (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, in which the aerodynamic profile (14) has a wing profile shape.

7. An aeronautical turbomachine (50) comprising an annular aeronautical turbomachine wall (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 (V1), in which the wall (34) radially delimits at least part of the cold gas stream (V1).

Images