Annular wall of an aircraft turbine engine and aircraft turbine engines including such a wall

Abstract

The present invention relates to an annular wall (34) of an aircraft turbine engine (50), the wall (34) having an axis (X) and including an aerodynamic surface (34A) configured to define a gas passage (V1), the wall (34) having a recess (35) and a heat exchanger (10) provided thereon, the heat exchanger including a heat exchanger base (12) and a wing (14), a heat transfer fluid circuit (20) extending at least into the base (12), the base (12) having a first portion (12) received in the recess (35) and a second portion (12B) outside the recess (35), the wing (14) having a leading edge (14A) and a trailing edge (14B), the wing (14) being attached to the base (12) such that the leading edge (14A) and the trailing edge (14B) extend circumferentially and such that the leading edge (14A) is located outside the recess (35).

Description

Technical Field

[0001] This disclosure relates to an annular wall of an aircraft turbine engine and an aircraft turbine engine including such a wall.

[0002] The term "aircraft turbine engine" refers to all gas turbine devices that generate power, including turbojet engines, which provide the thrust needed for propulsion through the reaction of high-speed jets of gas, and turboshaft engines, which provide power through the rotation of a drive shaft. For example, turboshaft engines are used as engines for helicopters. As another example, a turboprop engine (a turboshaft engine that drives a propeller) is a turboshaft engine used as an aircraft engine. Background Technology

[0003] Aircraft turbine engines typically have heat exchangers on the annular walls that radially define gas passages to cool fluids, usually oil, used to cool and / or lubricate various components of the turbine engine, such as bearings or other rotating elements. Examples of such engines are known from WO2023 / 001379.

[0004] Among known heat exchangers, there are two main series: Air-Cooled Oil Coolers (ACOC) and Surface-Cooled Oil Coolers (SACOC). ACOC heat exchangers generally have high thermal efficiency, but they generate significant pressure losses in the gas flow path, leading to an increase in static pressure, which degrades the performance of upstream components. SACOC heat exchangers generate lower pressure losses, but have lower thermal efficiency.

[0005] In light of climate change, the limitation of carbon emissions is a matter of ongoing concern to the assignee of this disclosure. Therefore, in order to obtain the most energy-efficient and environmentally friendly aerospace components and turbine engines possible, with the aim of reducing the environmental impact of the aviation sector, particularly civil aviation, the assignee has considered all factors that may have an impact in this regard, especially heat exchangers.

[0006] Therefore, it is necessary to improve the ratio between thermal efficiency / pressure loss and pressure rise in heat exchangers used in aero-turbine engines. Summary of the Invention

[0007] One embodiment relates to an annular wall of an aero-turbine engine, the wall having an axis and including an aerodynamic surface extending axially along the axis and circumferentially around the axis, the aerodynamic surface being configured to radially define a gas flow channel, the wall having a recess extending radially from the aerodynamic surface, the wall being provided with a heat exchanger, the heat exchanger including a heat exchanger body and an aerodynamic profile, a heat transfer fluid loop extending at least within the heat exchanger body, the heat exchanger body including a first portion received in the recess and a second portion extending radially outward of the recess 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 body such that the leading edge and trailing edge extend circumferentially and such that the leading edge is radially spaced from the aerodynamic surface and is outside the recess (i.e., outside the recess).

[0008] A heat exchanger can be considered as an element mounted on an annular wall. In other words, the embodiment defined above can be viewed as a component including an annular wall and a heat exchanger.

[0009] Typically, the axial direction corresponds to the axis of the annular wall, which, when the annular wall is mounted on an aero-turbine engine, corresponds to the rotation axis of the gas generator. The radial direction is perpendicular to the axis / axial direction. The circumferential direction corresponds to the direction describing the ring around the axis / axial direction. Axial, radial, and circumferential directions correspond to directions defined by height, radius, and angle in cylindrical coordinates, respectively. Furthermore, upstream and downstream are defined relative to the normal flow direction of fluid through the aero-turbine engine (from upstream to downstream). Finally, unless otherwise stated, the adjectives "inner" and "outer" are used with reference to the radial direction, such that the inner (i.e., radially inner) portion of the element is closer to the axis defining the axial direction than the outer (i.e., radially outer) portion of the same element.

[0010] In the following text, and unless otherwise stated, the term “wall” should be understood as “annular wall of an aircraft turbine engine,” the term “exchanger” should be understood as “heat exchanger,” the term “base” should be understood as “heat exchanger base,” and the term “profile” should be understood as “aerodynamic profile.” As a reminder, for those skilled in the art, the term “base” refers to the effective portion of the exchanger, including the channels through which the heat transfer fluid circulates to exchange heat with the fluid surrounding the base.

[0011] The annular wall can be annular sectors or a complete ring. A complete ring can be formed from a single-piece manufactured ring or from multiple annular sectors. The heat exchanger can extend continuously over the entire circumference of the annular wall or only over annular sectors spaced apart from each other in the circumference. In the latter case, the aerodynamic profile can have the same circumferential extent as the heat exchanger base, or it can have a larger circumferential extent than the heat exchanger base, for example, by extending the same circumferential extent symmetrically on each side of the heat exchanger base. According to a variation, the base can extend only over annular sectors spaced apart from each other in the circumference, while the profile can extend continuously over the entire circumference of the wall and be fixed to all annular sectors of the base.

[0012] The annular wall can be a radially inner or radially outer wall of the gas flow path, or it can be an outer wall of the cabin that defines the external gas flow passage (i.e., the air passage flowing around the cabin). If the annular wall is a radially inner wall or an outer wall of the cabin, the recess can extend radially inward. If the annular wall is a radially outer wall, the recess can extend radially outward. When the recess extends radially outward, a second portion of the heat exchanger base protrudes radially inward relative to the annular wall.

[0013] The heat transfer fluid can be fuel or lubricant, such as oil, but not necessarily.

[0014] For example, the leading and trailing edges of the aerodynamic profile can extend parallel to the aerodynamic surface of the wall. Opposite sides of the profile in the radial direction can respectively form a pressure surface and a suction surface extending from the leading edge to the trailing edge. For example, the pressure surface of the aerodynamic profile can be fixed to the substrate. For example, the trailing edge can be radially spaced from the aerodynamic surface and located outside the recess. For example, the entire aerodynamic profile can be radially spaced from the aerodynamic surface and arranged outside the recess.

[0015] From the perspective of the airflow circulating along the wall, the heat exchanger includes a "hidden" portion within the wall, corresponding to a first portion of the substrate disposed within a recess, and a protruding portion extending from the wall, corresponding to a second portion of the substrate with the aforementioned aerodynamic profile. Because the leading edge of the profile is radially spaced from the wall, it can direct a portion of the airflow towards the recess and the substrate. Compared to flow deflection in the prior art, this guidance is less intrusive and less restrictive to the airflow, thus minimizing pressure loss and static pressure rise caused by the heat exchanger.

[0016] Therefore, this structure can achieve satisfactory thermal efficiency by combining a matrix and profile that also participate in heat exchange, while minimizing pressure loss.

[0017] For example, the flow cross-section considered at the trailing edge between the wall and the profile is larger than the flow cross-section considered at the leading edge between the wall and the profile. This can minimize the disturbance and pressure loss in the airflow.

[0018] In some embodiments, the heat exchanger may include at least one fin that extends radially from and axially along the aerodynamic profile, opposite the heat exchanger substrate.

[0019] In the following text, unless otherwise stated, “fin” should be understood as “the at least one fin”.

[0020] Fins can extend perpendicular to the profile, for example, on its suction surface, between the leading and trailing edges. Fins can improve heat exchange, for example, through convection, while simultaneously causing very little disturbance within the airflow due to their extension parallel to the airflow direction. In other words, combining fins on the profile with the substrate increases heat exchange with minimal pressure loss, and the heat exchanger does not occupy significantly more space on the wall. Using fins, essentially the same length and circumferential range as a standalone substrate can be achieved, but with greater heat exchange achieved through fins having a smaller additional height than would be required to increase the substrate's height.

[0021] In some embodiments, the heat transfer fluid loop may extend within the aerodynamic profile.

[0022] By “within the aerodynamic profile”, it should be understood that the heat transfer fluid loop extends within the volume of the profile itself, between the pressure surface, suction surface, leading edge, and trailing edge.

[0023] In some embodiments, the heat transfer fluid loop may extend within the wall.

[0024] The phrase "within the wall" should be understood to mean within the volume of the wall itself, rather than within components mounted on the wall (such as heat exchangers).

[0025] In some embodiments, the heat transfer fluid loop may extend within the at least one fin.

[0026] By “within at least one fin”, it should be understood that it refers to the volume of at least one fin, rather than the volume of adjacent elements (such as the profile).

[0027] Such a fluid circuit, based on one or a combination of the above structures, enables improved heat exchange without affecting the pressure loss caused by the exchanger.

[0028] For example, heat exchangers can be manufactured wholly or partially using additive manufacturing. This can optimize both the external geometry of the exchanger and the heat transfer fluid loops extending within it.

[0029] In some embodiments, the heat exchanger may extend only on annular sectors spaced apart from each other in the circumferential direction.

[0030] In other words, when considered in the circumferential direction, adjacent ring sectors of a switch are spaced apart from each other. Each ring sector has a finite circumferential range, and each sector differs from the others. Ring sectors can be regularly distributed in the circumferential direction. According to one variation, all ring sectors of the switch can be arranged at the same axial position, such that they are all aligned in the circumferential direction.

[0031] Exchangers including this type of ring sector can exhibit satisfactory thermal efficiency while minimizing the resulting pressure loss.

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

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

[0034] This ratio ensures a good balance between the heat exchanger's thermal efficiency and the disturbances caused by the airflow.

[0035] In some embodiments, the aerodynamic profile may have an airfoil profile shape.

[0036] An airfoil profile can have a rounded leading edge, meaning it has a non-zero radius, without geometric discontinuities or angles. For example, an airfoil profile can be a "NACA" type profile.

[0037] This airfoil profile shape allows a portion of the airflow to be deflected toward the base while minimizing disturbances in the airflow, particularly pressure loss.

[0038] One embodiment relates to an aircraft turbine engine that includes an annular wall of an aircraft turbine engine according to any embodiment of the present disclosure.

[0039] Aircraft turbine engines can be ducted or non-ducted turbojet engines. For example, the fan can be directly driven by the gas generator or driven via a reduction gearbox. For example, the fan diameter considered 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 (BPR), which corresponds to the ratio of the mass flow rate of the secondary airflow (i.e., the airflow from the fan that does not enter the gas generator) to the mass flow rate of the primary airflow (i.e., the airflow from the fan that enters the gas generator), can be between 8 and 40, for example, between 8 and 18 for ducted fans and between 18 and 40 for non-ducted fans. For example, the fan pressure ratio (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 guide vanes, can be between 1.0 and 1.5.

[0040] In some embodiments, an aero-turbine engine may include a hot gas flow path and a cold gas flow path, wherein the wall radially defines at least a portion of the cold gas flow path.

[0041] The hot gas flow path is sometimes referred to by those skilled in the art as the "primary flow path," while the cold gas flow path is sometimes referred to by those skilled in the art as the "secondary flow path." Attached Figure Description

[0042] The subject matter and advantages of this disclosure will be better understood by reading the following detailed description of various embodiments given by way of non-limiting example. This description refers to the accompanying drawings, in which: Figure 1 An aircraft equipped with a turbine engine is shown. Figure 2 schematically shown Figure 1 The main internal components of the aircraft turbine engine, Figure 3 schematically shown Figure 1 The different annular walls of the aircraft turbine engine, Figure 4 Showing more details Figure 3 A perspective view of an annular section of a heat exchanger. Figure 5 The first variant of the heat exchanger is shown, and Figure 6 A second variant of the heat exchanger is shown.

[0043] Detailed description Figure 1An aircraft 100 is shown, in this example an airplane, equipped with two aircraft turbine engines 50, in this example two propulsion units 50, in this example two turbojet engines 50, that is, one aircraft turbine engine 50 for each wing 101. Figure 1 Only one aircraft turbine engine 50 and one wing 101 are shown. According to a variant, the aircraft 100 may be equipped with more than one aircraft turbine engine 50 on each wing 101, with each wing 101 having the same number of aircraft turbine engines 50.

[0044] Figure 2 It shows along Figure 1 A schematic cross-sectional view of the main internal components of an aircraft turbine engine 50 in plane II. The aircraft turbine engine 50 includes a fan 52, which may be ducted or non-ducted, and a gas generator 54 having a rotational axis X. Figure 1 In the example, the fan is ducted, and the duct is in Figure 2 Not shown in the image, but... Figure 3 As shown in the figure. In this example, the gas generator 54 includes, from upstream to downstream: a compressor 54A (or compressor section 54A) through which the gas flows from upstream to downstream within the aircraft turbine engine 50, a combustion chamber 54B, and a turbine 54C (or turbine section 54C). The fan 52 may be driven to rotate directly by the shaft of the gas generator 54 (e.g., the shaft of the low-pressure rotor), or driven to rotate via a reduction gearbox not shown.

[0045] Gas generator 54 may be a dual-spool type and includes a low-pressure rotor assembly 60A and a high-pressure rotor assembly 60B. The low-pressure rotor assembly 60A may include a low-pressure compressor 62A rotatably connected to a low-pressure turbine 66A via a low-pressure shaft 63A. The high-pressure rotor assembly 60B may include 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 connected to the high-pressure compressor 62B via a high-pressure shaft 63B. The compressor 54A of gas generator 54 may include both the low-pressure compressor 62A and the high-pressure compressor 62B. The turbine 54C of gas generator 54 may include both the low-pressure turbine 66A and the high-pressure turbine 66B. Figure 2 This is a schematic diagram. Each compressor and each turbine may include one or more stages, and each stage includes a moving impeller and a stator or guide vanes.

[0046] Figure 3 It shows along Figure 1A schematic cross-sectional view of the aircraft turbine engine 50 in plane II, showing the annular wall of the turbine engine 50. The compressor 54A, particularly the rotating and stator blades, the combustion chamber 54B, and the turbine 54C, particularly the rotating and stator blades, are indicated by shaded areas. For clarity, the gas generator 54 is not shown in more detail.

[0047] exist Figure 3 In the example, the turbine engine 50 is provided with five annular walls 30, 32, 34, 36, and 38, which have an axis X and respectively include aerodynamic surfaces 30A, 32A, 34A, 36A, and 38A extending axially along the axis X and circumferentially (i.e., in the circumferential direction C) around the axis X. Annular wall 30 is an outer nacelle wall 40 and defines a gas flow passage outside the nacelle 40. Walls 32 and 34 define a cold gas flow path (or passage) V1, while walls 36 and 38 define a hot gas flow path (or passage) V2. Wall 32 is the radially outer wall of the cold gas flow path V1, while wall 34 is the radially inner wall of the cold gas flow path V1. Wall 36 is the radially outer wall of the hot gas flow path V2, while wall 38 is the radially inner wall of the hot gas flow path V2. When the turbine engine 50 is operating, the hot gas flow path V2 is configured to guide the so-called main airflow F1, and the cold gas flow path V1 is configured to receive the so-called secondary airflow F2.

[0048] exist Figure 3 In the example (see also) Figure 4 The annular wall 34 has a recess 35 extending radially (i.e., in the radial direction R) from the aerodynamic surface 34A. A heat exchanger 10 is provided in the wall 34, comprising a heat exchanger body 12 and an aerodynamic profile 14. A heat transfer fluid loop 20 extends at least within the heat exchanger body 12. In this example, the loop 20 is symbolically represented by a channel 20A, and may also extend within the wall 34 and / or the profile 14. The heat exchanger body 12 includes a first portion 12A received within the recess 35 and a second portion 12B extending radially outward from the recess 35 and projecting radially from the aerodynamic surface 34A. Figure 4In the diagram, the orifice of the recess 35 is represented by a dashed line 35A extending in the geometric continuation of the recess 35 in the upstream and downstream portions of the surface 34A. Line 35A represents the radial boundary of the recess 35. The aerodynamic profile 14 includes a leading edge 14A and a trailing edge 14B, which is fixed to a second portion 12B of the heat exchanger base 12 such that the leading edge 14A and trailing edge 14B extend circumferentially, and that the leading edge 14A is radially spaced from the aerodynamic surface 34A and located outside the recess 35. In other words, in this example, the leading edge 14A is arranged outside the recess 35 and adjacent to line 35A, or radially spaced from line 35A in the radial direction R. In this example, the trailing edge 14B is also radially spaced from the aerodynamic surface 34A and located outside the recess 35. In this example, the entire profile 14 is arranged outside the recess 35. In this example, the wall 34 forms a wall that radially defines at least a portion of the cold gas flow path V1. In this example, the recess 35 and the exchanger 10 are arranged downstream of the guide vane 70, which itself is arranged downstream of the fan 52 and upstream of the outlet S1 of the fan module MF, which includes the fan 52, the guide vane 70 and the nacelle 40.

[0049] like Figure 4 As shown, the leading edge 14A and 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. In this example, the profile 14 includes a pressure surface 14C and a suction surface 14D, each extending between the leading edge 14A and the trailing edge 14B and forming two opposing surfaces in the radial direction R of the profile 14. In this example, the profile 14 is fixed to the base 12 by the pressure surface 14C. In this example, the profile 14 has an airfoil profile shape. In this example, the leading edge 14A has a rounded shape, i.e., it has a non-zero radius and no geometric discontinuities or angles.

[0050] like Figure 4 As shown, the heat exchanger 10 may include at least one fin 22 extending radially from and axially along the aerodynamic profile 14, opposite the heat exchanger base 12. A heat transfer fluid loop 20 may extend within all or part of the fin 22 (not shown).

[0051] For example, the ratio LMP / LMM between the maximum axial length of the aerodynamic profile LMP and the maximum axial length of the heat exchanger substrate can be between 1.0 and 1.4, for example between 1.2 and 1.3.

[0052] In this example, the maximum axial length LMM of the substrate can be between 5.0 cm and 25.5 cm (or between 2 inches and 10 inches), while the maximum axial length LMP of the profile can be between 5.0 cm and 25.5 cm (or between 2 inches and 10 inches). The maximum axial length LMM of the substrate can be less than or equal to the maximum axial length LMP of the profile. The maximum axial length LMA of the fin 22 can be less than or equal to the maximum axial length LMP of the profile, and greater than or equal to the maximum axial length LMM of the substrate.

[0053] For example, the ratio EMP / EMM between the maximum radial thickness EMP of the aerodynamic profile and the maximum radial thickness EMM of the heat exchanger substrate can be between 0.05 and 0.35, for example, between 0.10 and 0.25. It should be noted that, as... Figure 4 As shown in the example, the axial position of the maximum radial thickness of profile 14 and the axial position of the maximum radial thickness of base 12 do not have to be the same.

[0054] exist Figure 4 In the example, the radial thickness EA of each fin 22 is constant along the entire axial direction of the fin 22 and can be between 1.25 cm and 5.00 cm (or between 0.5 inches and 2.0 inches). According to a variant not shown, the radial thickness EA of the fin 22 can vary in the axial direction X, and the average value of the radial thickness EA can be between 1.25 cm and 5.00 cm (or between 0.5 inches and 2.0 inches).

[0055] The maximum radial thickness EM of the heat exchanger 10, including the optional fins 22, can be between 5.0 cm and 13.0 cm (or between 2 inches and 5 inches).

[0056] Figure 5 It shows according to Figure 5 The first variation of the section V, in which the exchanger 10, in particular the base 12 and the profile 14, extends continuously in the entire circumference of the wall 34. Figure 6 It shows according to Figure 5 A second variation of section V, in which the heat exchanger 10, particularly the substrate 12 and profile 14, extends only on annular sectors 10' spaced apart from each other in the circumferential direction. In this second variation, the annular sectors 10' are regularly spaced apart from each other. In other words, the annular sectors 10' are regularly distributed in the circumferential direction C. Figure 6 As shown, for each annular sector 10', the circumferential extent of profile 14 is greater than the circumferential extent of base 12. In this example, profile 14 extends circumferentially beyond base 12 by the same amount on each side of base 12.

[0057] The heat transfer fluid circulating in loop 20 can be a lubricant, such as oil, which can be used, for example, to lubricate and / or cool one or more bearings (not shown) of the low-pressure shaft 63A and / or the high-pressure shaft 63B. Alternatively or additionally, the oil can be used, for example, to lubricate and / or cool any possible reduction gearbox that rotates the low-pressure shaft 63A and the fan 52.

[0058] In the example above, only wall 34 has a recess 35 for accommodating the heat exchanger 10. According to variations not shown, in addition to or instead of wall 34, walls 30 and / or wall 32 and / or wall 36 and / or wall 38 may be provided with recesses for accommodating the heat exchanger, in a manner similar to the reference. Figures 3 to 6 The recess 35 and the heat exchanger 10 are described. For example, if the wall 36 and / or wall 38 defining the hot gas flow path are provided with a recess for accommodating the heat exchanger, the recess and the heat exchanger may be arranged, for example, upstream of the compressor 54A.

[0059] In the above example, wall 34 is a radially inner wall, recess 35 extends radially inward, and exchanger 10, particularly the second portion 12B of base 12, protrudes radially outward relative to wall 34. Similarly, when recesses and exchangers are provided in wall 30 or wall 38, the recesses extend radially inward, and the exchangers, particularly the second portion of base, protrude radially outward relative to wall 30 or 38, respectively. Conversely, when recesses and exchangers are provided in wall 32 or wall 36, which are radially outer walls, the recesses extend radially outward, and the exchangers, particularly the second portion of base, protrude radially inward relative to wall 32 or 36, respectively.

[0060] Although the invention has been described with reference to specific embodiments, it will be apparent 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, various features of the different illustrated / described embodiments may be combined in other embodiments. Therefore, the specification and drawings should be considered illustrative rather than restrictive.

Claims

1. An annular wall (34) of an aero-turbine engine (50), the annular wall (34) having an axis (X) and including an aerodynamic surface (34A) extending axially along the axis (X) and circumferentially around the axis (X), the aerodynamic surface (34A) being configured to radially define a gas passage, the annular wall (34) having a recess (35) extending radially from the aerodynamic surface (34A), the annular wall (34) being provided with a heat exchanger (10), the heat exchanger including a heat exchanger body (12) and an aerodynamic profile (14), a heat transfer fluid circuit (20) extending at least into the heat exchanger body (12), wherein... The heat exchanger base (12) has a first portion (12A) received in the recess (35) and a second portion (12B) extending radially outward from the recess (35) and projecting radially from the aerodynamic surface (34A). The aerodynamic profile (14) has a leading edge (14A) and a trailing edge (14B). The aerodynamic profile (14) is fixed to the second portion (12B) of the heat exchanger base (12) such that the leading edge (14A) and the trailing edge (14B) extend circumferentially, and that the leading edge (14A) is radially spaced from the aerodynamic surface (34A) and is located outside the recess (35). The heat exchanger (10) includes at least one fin (22) that extends radially from the aerodynamic profile (14), opposite to the heat exchanger base (12), and axially along the aerodynamic profile (14), and the heat transfer fluid circuit (20) extends within the aerodynamic profile (14).

2. The annular wall (34) of the aero-turbine engine (50) according to claim 1, wherein, The heat transfer fluid loop (20) extends into the annular wall (34).

3. The annular wall (34) of the aero-turbine engine (50) according to claim 1, wherein, The heat exchanger (10) extends only on annular sectors (10') spaced apart from each other in the circumferential direction.

4. The annular wall (34) of the aero-turbine engine (50) according to claim 1, wherein, The ratio between the maximum axial length (LMP) of the aerodynamic profile and the maximum axial length (LMM) of the heat exchanger substrate is between 1.0 and 1.

4.

5. The annular wall (34) of the aero-turbine engine (50) according to claim 4, wherein, The ratio between the maximum axial length (LMP) of the aerodynamic profile and the maximum axial length (LMM) of the heat exchanger substrate is between 1.2 and 1.

3.

6. The annular wall (34) of the aero-turbine engine (50) according to claim 1, wherein, The ratio between the maximum radial thickness (EMP) of the aerodynamic profile and the maximum radial thickness (EMM) of the heat exchanger substrate is between 0.05 and 0.

35.

7. The annular wall (34) of the aero-turbine engine (50) according to claim 6, wherein, The ratio between the maximum radial thickness (EMP) of the aerodynamic profile and the maximum radial thickness (EMM) of the heat exchanger substrate is between 0.10 and 0.

25.

8. The annular wall (34) of the aero-turbine engine (50) according to claim 1, wherein, The aerodynamic profile (14) has an airfoil profile shape.

9. An aircraft turbine engine (50) comprising an annular wall (34) of the aircraft turbine engine according to claim 1.

10. The aero-turbine engine (50) according to claim 9, comprising a hot gas flow path (V2) and a cold gas flow path (V1), wherein, The annular wall (34) radially defines at least a portion of the cold gas flow path (V1).

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