FR3140122A1 - ASSEMBLY FOR AIRCRAFT TURBOMACHINE INCLUDING A SACOC TYPE HEAT EXCHANGER OF IMPROVED DESIGN - Google Patents

Abstract

The invention relates to an assembly (40) for an aircraft turbomachine comprising a structure (22) for radially delimiting an air circulation vein (14b) intended to be crossed by an air flow (12b), the structure (22) comprising a surface (46) radially delimiting the vein, the assembly also comprising an air-lubricant surface heat exchanger (42), comprising an exchanger body (44) through which a circuit is formed lubricant (52), as well as fins (48) projecting from a support surface (50) provided on the exchanger body (44), the heat exchanger being of passive design and fixed relative to to the structure (22). In addition, the surface (50) is set back from the surface (46), so as to form a recess (54) into which the air flow (12b) is intended to penetrate, and where at least several fins (48) are each housed at least partially. Figure 2.

Description

ASSEMBLY FOR AIRCRAFT TURBOMACHINE INCLUDING A SACOC TYPE HEAT EXCHANGER OF IMPROVED DESIGN

The present invention relates to the field of aircraft turbomachines, and more precisely to assemblies equipped with heat exchangers of the SACOC type (from the English, “Surface Air-Cooled Oil-cooler”).

The invention applies to all turbomachine designs, for example turbojet engines with a ducted or non-ducted fan, driven directly by a low-pressure body, or driven indirectly by a reduction gear.

STATE OF PRIOR ART

Aircraft turbomachines include numerous elements such as the rolling bearings supporting the shaft(s) of the turbomachine, which need to be both lubricated and cooled. Furthermore, in the case of a turbomachine comprising a fan driven by a reduction gear, the latter also needs to be cooled and lubricated.

To do this, it is known to supply these elements with “cold” lubricant, generally oil. One or more surface air-oil exchangers of the SACOC type are then used, placed in an air circulation vein of the turbomachine. In this SACOC type air-oil exchanger, a multitude of channels are provided in which the oil to be cooled circulates. The exchanger comprises a body from which cooling fins protrude, these usually having an isosceles trapezoidal profile extending axially, in the direction of circulation of the air flow. These fins increase the heat exchange surface between the oil to be cooled circulating in the channels of the body, and the air circulating in the vein. Such an embodiment is for example known from document FR 3 016 956 B1.

However, by protruding into the air circulation vein, usually the secondary vein of the turbomachine, these fins generate undesirable pressure losses which harm the overall efficiency of the turbomachine, and which thus negatively impact its specific consumption. They also create a pre-rise of static pressure which has an effect on the components located upstream of the fins, usually the fan.

However, climate change is a major concern for many legislative and regulatory bodies around the world. Indeed, various restrictions on carbon emissions have been, are or will be adopted by various states. In particular, an ambitious standard applies both to new types of aircraft but also to those in circulation requiring the implementation of technological solutions in order to make them comply with current regulations. The Applicant takes into consideration the impacting factors in all phases of design and development to obtain aeronautical components and products that consume less energy, are more respectful of the environment and whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.

The invention aims to respond to the disadvantages and environmental objectives mentioned above, with the aim of very significantly improving the performance of turbomachines and, in this sense, contributing to the reduction of the environmental impact of aircraft. For this, the invention firstly relates to an assembly for an aircraft turbomachine comprising a radial delimitation structure for the circulation of an air flow, the structure extending around a longitudinal central axis of the turbomachine and comprising a radial delimiting surface, the assembly also comprising an air-lubricant surface heat exchanger, the exchanger comprising an exchanger body through which a lubricant circuit is formed, as well as cooling fins by the air flow, the fins projecting from a support surface provided on the exchanger body, the heat exchanger being of passive design, and fixed relative to the radial boundary structure.

According to the invention, the support surface of the fins is set back from the radial delimiting surface, so as to form a recess into which the air flow is intended to penetrate, and in which at least several of said fins are each housed at least partially.

In this way, several fins of the exchanger, and preferably all of them, find themselves at least partly arranged in the recess, thus adopting a position called entirely or partially "buried" in the radial delimitation structure. This arrangement helps to limit the protruding distance of the fins in the air flow, while providing the same or better thermal performance. This results in a reduction in pressure losses on the air flow, and therefore an increase in efficiency as well as a reduction in specific consumption, which perfectly meets the environmental objectives specified previously.

Furthermore, in the case where the air flow is a secondary flow of the turbomachine, this arrangement of the cooling fins in this same air flow advantageously limits the effects of static pressure rise on the fan located upstream, producing thus a slight gain in operability.

The invention also preferably provides for at least one of the following optional characteristics, taken individually or in combination.

According to a preferred embodiment of the invention, all the fins of the exchanger are entirely housed in the recess, each of them having a free radial end flush or substantially flush with a fictitious surface reconstituting the delimitation surface radial at the level of the recess. In this scenario where the cooling fins are entirely buried, the pressure losses are advantageously minimal.

According to another preferred embodiment of the invention, at least some of the fins of the exchanger each extend radially into the recess and beyond the latter, so as to protrude into the air flow. The fins are then partially buried.

In the latter case, it is for example provided that for each fin among said at least certain fins, the projecting height in the air flow is constant or substantially constant over at least 75% of the length of the fin. This projecting height can for example go up to 50% of the total height of the fin from the support surface provided on the exchanger body.

According to an alternative, for each fin among said at least certain fins, the projecting height in the air flow is strictly decreasing in the direction of circulation of this air flow, over at least 75% of the length of the fin. This progressive burial of the fins makes it possible to avoid a premature exit of the air current lines outside the recess, and thus improves the exchanges between the air and the fins, the total height of which can remain constant or substantially constant over at least 75% of fin length. To do this, the support surface of the fins is preferably inclined so as to approach the longitudinal central axis going in the direction of circulation of the air flow. This characteristic is also applicable for a solution with fully buried fins.

Preferably, the heat exchanger extends in an annular manner circumferentially all along the radial delimiting structure of the air circulation vein, in a continuous or sectorized manner.

Preferably, the fins are parallel to the central longitudinal axis.

The invention also relates to an aircraft turbomachine comprising an assembly as described above.

This turbomachine preferably has a double vein design, including a secondary vein in which said air flow circulates, said radial delimitation structure being an interior or exterior radial delimitation structure of the secondary vein.

The invention thus applies to any type of turbomachine, and in particular to twin-vein turbojets with a ducted or non-ducted fan.

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

This description will be made with regard to the appended drawings including:

represents a schematic view in longitudinal section of an aircraft turbomachine according to the invention;

represents a half-view in longitudinal section, more detailed, of an assembly according to a first preferred embodiment of the invention, integrated into the turbomachine shown in the previous figure;

represents a cross-sectional view of the assembly, taken along line III-III of the previous figure;

represents a perspective view of a part of the heat exchanger integrated into the assembly shown in the previous figures;

represents a perspective view similar to that of the previous figure, with the assembly presented in the form of a second preferred embodiment of the invention;

represents a longitudinal sectional view of a part of the assembly, taken along the plane P of the previous figure; And

represents a longitudinal sectional view similar to that of the previous figure, with the assembly in the form of a third preferred embodiment of the invention;

represents a longitudinal sectional view similar to that of the previous figure, with the assembly in the form of a fourth preferred embodiment of the invention.

PRESENTATION OF THE MODES OF ACHIEVEMENT

With reference first to the , there is shown an aircraft turbomachine 1, according to a preferred embodiment of the invention. This is a dual-flow, dual-body turbojet, comprising a single ducted fan. However, it could be a turbomachine of another type, for example a turboprop, or even a turbomachine with one or more unducted fans/propellers.

The turbomachine 1 has a central longitudinal axis 2 around which its various components extend. It comprises, from upstream to downstream in a main direction 5 of flow of gases through this turbomachine, a fan 3, a low pressure compressor 4, a high pressure compressor 6, a combustion chamber 11, a high pressure turbine 7 and a low pressure turbine 8. The fan 3 can be driven directly by a low pressure body comprising the compressor 4 and the turbine 8, or else be driven indirectly by a reduction gear (not shown).

Conventionally, after passing through the fan 3, the air is divided into a central primary flow 12a and a secondary air flow 12b which surrounds the primary flow. The primary flow 12a flows in a main stream 14a for circulating the gases passing through the compressors 4, 6, the combustion chamber 11 and the turbines 7, 8. The secondary flow 12b flows in a stream of secondary air 14b delimited radially outwards by a casing, surrounded by a nacelle 9. More precisely, the casing comprises a fan casing 20 which surrounds the fan blades, this casing 20 being extended downstream by a shroud exterior 22 of an intermediate casing 24. This intermediate casing 24 comprises a hub 26 centered on the axis 2, and which may include a flow separation nozzle 28. The intermediate casing is completed by radial arms 30, which form blades fixed outlet directors, making it possible to straighten the secondary air flow 12b and conventionally called OGV (from the English “Outlet Guided Vanes”). The arms 30 thus connect the outer shell 22 of the intermediate casing to its hub 26.

The turbomachine has a dilution ratio, or BPR (from English “ByPass Ratio), of the order of eight to forty, and more precisely of the order of eight to eighteen in the present case of a fan faired. In the case of one or more non-ducted fans/propellers, this dilution rate is more of the order of eighteen to forty.

The outer shell 22 forms an outer radial delimitation structure of the secondary air stream 14b. It is extended downstream by another annular delimitation structure 31 of this type, integrated into a mobile thrust reverser cover 32. This structure 31 is also called OFS (from the English “Outer Fixed Structure”). It is located radially opposite an annular inner radial delimitation structure 34 of vein 14b, also called IFS (from the English “Inner Fixed Structure”).

One of the particularities of the invention lies in the implementation of an assembly 40 comprising the association of one of the annular structures 22, 31, 34 mentioned above, all centered on axis 2, and of an air-lubricant surface heat exchanger 42, here of the SACOC type. In the preferred embodiments which will be described below, the annular structure concerned corresponds to the outer shell 22 of the intermediate casing. In addition, the heat exchanger 42 also extends in an annular manner around the axis 2, all along the outer shell 22, continuously over 360° or substantially 360°, or alternatively in a sectorized manner depending on the circumferential direction 15 of the turbomachine.

Referring now to Figures 2 to 4, assembly 40 is shown according to a first preferred embodiment of the invention. Here, the heat exchanger 42 comprises an exchanger body 44 forming an integral part of the outer shell 22, being made in one piece with the latter, or with a part of this body attached to this shell, for example by overmolding. According to the radial direction 17 of the assembly 40 and the turbomachine, the exchanger body 44 is located outwards relative to an external radial delimitation surface 46 of the vein 14, defined by the ferrule 22 and channeling the circulation of flow 12b.

The heat exchanger comprises fins 48 for cooling by the secondary air flow 12b, these fins projecting radially inwards from a support surface 50 provided on the body 44. The cooling fins 48 are spaced circumferentially and parallel to each other, and they extend radially inwards while being parallel to axis 2. They also preferably extend over 360° or substantially 360°, in a continuous or circumferentially sectored manner.

The exchanger 42, which is of passive design and which remains fixed relative to the outer shell 22 during operation of the turbomachine, also includes an oil circuit 52 made through the exchanger body 44. The exchanger circuit oil 52 takes the form of “hot” oil circulation channels. In operation, the “hot” oil which arrives in the body 44 and which requires to be cooled, passes through the plurality of channels of the circuit 52. The heat released by the “hot” oil is transferred to each fin 48 which is -even cooled by part of the secondary air flow 12b, corresponding to “cold” air. Thus, the thermal energy “stored” in each fin 48 is transferred via a heat exchange surface of each fin to the “cold” air of the secondary flow.

One of the particularities of the invention lies in the fact that the support surface 50 of the fins 48 is set back relative to the delimiting surface 46 of the vein, radially outwards, so as to form a recess 54 in which the fins 48 are located. More precisely, this recess 54 forms a sort of radially external annular cavity in the outer shell 22, in which all the fins 48 of the exchanger are preferably housed. Consequently, part of the secondary air flow 12b is intended to penetrate into the recess 54 to its bottom formed by the support surface 50 from which the fins 48 project radially, so as to match the surface d heat exchange of these fins.

In this first preferred embodiment of the invention, the fins 48 are said to be entirely “buried” in the recess 54, in the sense that their free internal radial ends 48a do not project outside the recess 54 in the vein 14b. In fact, these free internal radial ends 48a of the fins 48 are preferably flush or substantially flush with a fictitious surface 56 reconstituting the vein 14b at the level of the recess 54. This surface 56 in fact corresponds to that which fictitiously closes the recess 54, by recreating the aerodynamic profile of the vein, without steps or differences in level.

In this first preferred embodiment, the absence of projection of the fins 48 in the secondary air stream 14b greatly limits the pressure losses on the flow 12b.

Within the recess 54, each fin 48 has a general trapezoid shape, with the small base of this trapezoid being located at the level of the support surface 50 from which it projects, radially inwards. The total axial length “L” of each fin 48 can be of the order of 5 to 25 cm. Its total height “Ht” can be of the order of 0.5 to 3 cm, it being understood that it remains preferably constant over a length “L1” corresponding to at least 75% of its total length L. This length L1, centered on the fin, corresponds in fact to the part of the trapezoid of constant height, namely without its opposite triangular lateral parts. The thickness “E” of the fins 48 can be of the order of a few millimeters.

In a second preferred embodiment of the invention shown in Figures 5 and 6, each fin 48 extends radially in the recess 54 and beyond the latter, until it projects into the vein 14b. The fins are then said to be partially “buried”. Here, the part of each fin 48 which is located in the recess 54 has a trapezoidal shape identical or similar to that described with reference to the first mode. It continues with a part projecting into the vein, preferably with a trapezoidal profile of inverted shape, that is to say with its small base located radially inwards and forming the free internal radial end 48a of the 'fin. Its projecting height “Hs” in the air stream 14b is constant or substantially constant over at least 75% of the total length L of the fin in the direction of axis 2. This projecting height Hs can correspond up to 'at 50% of the total height Ht of the fin.

According to a third preferred embodiment shown on the , the projecting height Hs in the vein 14b is strictly decreasing in the direction of circulation of the air flow 12b, over a length L1 corresponding to at least 75% of the total axial length L of this fin. Indeed, only one beveled upstream end of the fin, projecting into vein 14b, has an increasing projecting height Hs, but then, this height Hs in the radial direction 17 remains strictly decreasing until the downstream end of the fin, which merges into the surface 46 of the ferrule 22. At this downstream end, the fin 48 thus has a zero or substantially zero projecting height Hs.

To maintain a constant or substantially constant total fin height over a long length thereof, it is preferentially ensured that the fin support surface 50 is inclined so as to move away from the axis 2 going in the direction of circulation of the air flow 12b, comparable to the main direction 5 of gas flow through the turbomachine. Thus, over the length L1 of each fin 48 and going downstream, the projecting height Hs decreases in favor of the height of the buried fin. The recess 54 then has a longitudinal section of generally triangular shape as visible on the , in height widening towards the downstream. Finally, each free internal radial end 48a considered over the length L1, as well as the support surface 50, each preferably have a straight/plane shape.

According to a fourth preferred embodiment of the invention, represented on the , the fins 48 are said to be entirely “buried” in the recess 54, all having an increasing height going in the direction of circulation of the air flow 12b. To ensure that the free internal ends 48a of the fins 48a do not project outside the recess 54, and remain flush or substantially flush with the fictitious surface 56 for reconstitution of the vein 14b at the level of the recess 54, while presenting a decreasing height for these same fins 48, the support surface 50 is also inclined so as to move away from axis 2 going in the direction of circulation of the air flow 12b.

Of course, various modifications can be made by those skilled in the art to the invention which has just been described, solely by way of non-limiting examples and within the limit of the scope of the appended claims. For example, if the turbomachine described in the preferred embodiments above takes the form of a dual-flow turbojet with a ducted fan, any other type of turbomachine remains possible. In particular, this concerns turbomachines with single or multiple, non-ducted fans/propellers, such as turbomachines with two non-ducted counter-rotating propellers. In this regard, it is noted that the exchanger can in particular be installed on the internal or external radial boundary wall of the secondary vein in the case of a dual-flow turbojet, or even be installed on the external cover of a non-ducted turbomachine, such as a turboprop or a so-called “open rotor” turbomachine. In the latter case, the external radial delimitation surface 46, within the meaning of the invention, is a radially external surface matched by the external flow of the turbomachine.

Claims (10)

Assembly (40) for an aircraft turbomachine comprising a radial delimiting structure (22) for the circulation of an air flow (12b), the structure (22) extending around a longitudinal central axis (2) of the turbomachine and comprising a radial delimiting surface (46), the assembly also comprising an air-lubricant surface heat exchanger (42), the exchanger comprising an exchanger body (44) through which a circuit is formed of lubricant (52), as well as fins (48) for cooling by the air flow, the fins projecting from a support surface (50) provided on the exchanger body (44), the heat exchanger being of passive design and fixed with respect to said radial boundary structure (22),
characterized in that the fin support surface (50) is set back from the radial delimiting surface (46), so as to form a recess (54) into which the air flow (12b) is intended to penetrate, and in which at least several of said fins (48) are each housed at least partially. Assembly according to claim 1, characterized in that all the fins (48) of the exchanger (42) are entirely housed in the recess (54), each of them having a free radial end (48a) flush or substantially flush to a fictitious surface (56) reconstituting the radial demarcation surface (46) at the level of the recess. Assembly according to claim 1, characterized in that at least some of the fins (48) of the exchanger each extend radially into the recess (54) and beyond the latter, so as to protrude into the flow air (12b). Assembly according to claim 3, characterized in that for each fin (48) among said at least certain fins, the projecting height (Hs) in the air flow (12b) is constant or substantially constant over at least 75% of the length of the fin. Assembly according to claim 3, characterized in that for each fin (48) among said at least certain fins, the projecting height (Hs) in the air flow (12b) is strictly decreasing in the direction of circulation of this flow air (12b), over at least 75% of the length of the fin. Assembly according to any one of claims 2 to 5, characterized in that the surface (50) supporting the fins is inclined so as to move away from the central longitudinal axis (2) going in the direction of circulation of the air flow (12b). Assembly according to any one of the preceding claims, characterized in that the heat exchanger (42) extends in an annular manner circumferentially all along the radial boundary structure (22), in a continuous or sectorized manner. Assembly according to any one of the preceding claims, characterized in that the fins (48) are parallel to the longitudinal central axis (2). Aircraft turbomachine (1) comprising an assembly (40) according to any one of the preceding claims. Turbomachine according to claim 9, characterized in that it has a double vein design (14a, 14b), including a secondary vein (14b) in which said air flow (12b), said radial delimitation structure (22 ) being an interior or exterior radial boundary structure of the secondary vein.