SYSTEM FOR COOLING A LIQUID FOR LUBRICATION OF AN AIRCRAFT TURBOME

Description

TECHNICAL FIELD OF THE INVENTION

[0001] The technical field of the invention is that of the cooling of a lubricating liquid such as lubricating oil, of gas turbine engines.

[0002] The present invention relates generally to turbomachines, i.e. gas turbine engines, and more particularly to a cooling device for transferring excess heat from a turbomachine. TECHNOLOGICAL BACKGROUND OF THE INVENTION

[0003] Gas turbine engines use pressurized oil to lubricate and cool various components (e.g., bearings, transmission means, etc.).

[0004] A lubricant distribution system is planned to ensure proper lubrication and cooling of these components.

[0005] The lubricant (oil) absorbs a significant amount of heat in the process which must be dissipated to maintain the oil temperature within acceptable limits.

[0006] Consequently, the lubricant must then also be cooled via a heat exchanger.

[0007] One known method for achieving this is to cool the lubricant by circulating it through an air / oil heat exchanger using air drawn from the secondary flow, the so-called cold flow, of the nacelle. Drawing and circulating air through this exchanger disrupts the airflow and results in additional, undesirable pressure drops. Another known cooling device consists of a heat exchanger mounted on a wall that delimits the secondary flow; the fluid is cooled by the air flowing from the secondary flow along the surface of the exchanger.

[0008] A gas turbine engine is also known, comprising a nacelle having an external surface exposed to an external airflow and a cooling apparatus arranged in the nacelle comprising a plurality of heat pipes in parallel with each other.

[0009] A turboshaft engine nacelle equipped with an engine oil cooling system is also known, for example, from document EP2819921. The cooling system comprises a first circuit of a heat transfer fluid, which may be a two-phase coolant, serving as the coolant to cool a second fluid, such as engine lubricant (oil), in a heat exchanger mounted in the nacelle. The first circuit includes numerous valves, a pump, and cooling channels along the nacelle walls for heat exchange with the outside air. The heat recovered by the heat exchange between the lubricant and the coolant causes the latter to evaporate at the beginning of its circulation.The evaporated coolant is then cooled to recondense, and gravity carries it down the nacelle where it is pumped back to the fluid / oil heat exchanger. A turboshaft engine nacelle incorporating an active cooling device is described, for example, in document EP3487764. The active cooling device comprises a shell that forms a trailing edge and is arranged at a downstream end of the nacelle. This shell is delimited by an external face exposed to an outside airflow and an internal face exposed to an internal airflow circulating in the nacelle's secondary duct. The cooling device includes a plurality of heat exchange channels within the shell, through which a heat transfer fluid flows to dissipate the heat carried by the fluid.The cooling system comprises an inlet manifold and an outlet manifold, and a source connected to both manifolds, each connected to one inlet and one outlet end of each channel. The active cooling system includes a pump to circulate the fluid from the source to the channels.

[0010] However, these cooling systems require the heat transfer fluid to be circulated through the channels, which typically necessitates a pump to circulate the fluid from the source to the channels. Furthermore, in these two-phase loops, there is a need to control superheat at the evaporator outlet as well as subcooling at the condenser outlet. A reservoir is required at the pump inlet to prevent the presence of vapor and thus ensure liquid pumping throughout all phases of flight, thereby avoiding any risk of cavitation. The pump and its associated reservoir increase the system's weight and energy consumption.

[0011] It is also known for avoiding direct oil circulation within the nacelle surfaces, employing a passive cooling system in which each heat pipe has an elongated outer wall with closed ends, trapping a heat transfer fluid. The engine further includes a heat exchanger mounted outside the crankcase, into which oil from the engine's lubrication system enters the heat exchanger via a recovery line. One end of this line is located within the heat exchanger. In this cooling system, each heat pipe is therefore independent, resulting in uneven work of the heat transfer fluid volumes within the different channels. This reduces the overall cooling performance for the total volume of heat transfer fluid compared to the previously described examples where the entire heat transfer fluid circulates through all the cooling channels (closed-loop system).

[0012] Document EP1884628 A2 discloses a cooling system for a lubricating fluid of an aircraft turbomachine known in the prior art.

[0013] Therefore, there is a need for an improved cooling system that is simple, efficient, and reduces energy consumption and weight. SUMMARY OF THE INVENTION

[0014] The invention offers a solution to the problems mentioned above by enabling the use of a heat transfer loop incorporating a capillary pumping system. The capillary effect of a porous wick in an evaporator brings the heat transfer fluid to a gaseous phase and generates a capillary pressure that compensates for all pressure losses in the loop. The capillary pressure must be greater than the total pressure loss of the system; otherwise, mechanical pumping is necessary, as described in the prior art.

[0015] One aspect of the invention relates to a cooling system for a lubricating fluid in an aircraft turbomachine, comprising: an exchanger comprising a first cooling volume for the lubricating fluid, a first cooling circuit for a two-phase heat transfer fluid by means of a two-phase heat transfer fluid circulation loop with thermo-capillary pumping, comprising: a first condenser comprising: a vapor inlet for the heat transfer fluid in the vapor state, channels extending along a first external surface of a turbomachine nacelle to cool the channels, the channels being connected in parallel to each other, to allow the fluid circulating in the condenser to be cooled from a vapor state to a liquid state, a liquid outlet for the heat transfer fluid in the liquid state, a first capillary evaporator of the exchanger, to absorb the heat from the first volume in which the lubricating fluid circulates,The first evaporator comprises: a first volume including a liquid inlet of the heat transfer fluid connected to the liquid outlet of the first condenser; a second volume including a vapor outlet of the heat transfer fluid, connected to the vapor inlet of the first condenser; and a porous capillary wick located between the first and second volumes, allowing the two-phase fluid to be separated into a liquid phase and a gaseous phase by a capillary pressure jump.

[0016] Thanks to the invention, the use of a capillary pumping system allows for the imposition of a specific direction of fluid flow within each condenser. This increases the cooling capacity of the coolant and, consequently, the oil, by utilizing the space and surface area of ​​the nacelle during different phases of aircraft flight, while simultaneously reducing the weight and energy required for the operation of the primary heat transfer fluid circuit. Furthermore, using a two-phase fluid circulation loop instead of a single-phase one reduces the amount of heat transfer fluid required to circulate in the cooling circuit. This, in turn, decreases the total mass of the heat transfer fluid and, therefore, of the entire cooling system.

[0017] In addition to the characteristics mentioned in the preceding paragraph, the cooling system according to one aspect of the invention may have one or more additional characteristics from among the following, considered individually or in all technically possible combinations: In one embodiment, the heat transfer fluid comprises between 30 and 70% of the cooling circuit in a liquid state. In another embodiment, the cooling system includes a second condenser similar to the first condenser, in which the first and second condensers are located on either side of a vertical plane encompassing the axis of the nacelle. In an example of this embodiment, the first cooling circuit for the two-phase heat transfer fluid includes the second condenser, whose vapor inlet is connected to the vapor outlet of the second volume of the first evaporator. In this example, the first and second condensers are therefore in parallel (fluid circulation).In another embodiment of this system, the cooling system comprises: a second capillary evaporator of the heat exchanger, similar to the first capillary evaporator; a second two-phase heat transfer fluid cooling circuit similar to the first two-phase heat transfer fluid cooling circuit; and the second cooling circuit comprising a second condenser, which includes its vapor inlet connected to the vapor outlet of the second volume of the second evaporator. In one embodiment, the first condenser is located higher than the first evaporator so that the fluid, in a liquid state at the condenser outlet, flows by gravity to the evaporator. This allows gravity to be used to promote the circulation of the heat transfer fluid in the cooling circuit.In one embodiment, the cooling system comprises a steam line connecting the steam outlet of the second chamber of the first evaporator to the steam inlet of the first condenser, the steam line being located along the axis of the nacelle upstream of the inlet of the first evaporator. In another embodiment, the heat exchanger is located in a lower part of the nacelle situated below the turbomachine. In a variant of the preceding embodiment, the heat exchanger is located in a central compartment of the turbomachine surrounded by the nacelle, the cooling circuit comprising a steam line connecting the steam outlet of the second chamber of the first evaporator to the steam inlet of the first condenser, passing inside a support arm of the turbomachine. In one embodiment, the first condenser is located in an upper part of the nacelle situated above the axis of the turbomachine.In one embodiment that is a variant of the previous embodiment, the first condenser is located in a lower part of the nacelle situated below the axis of the turbomachine. In another embodiment, the first external surface of the nacelle is a surface of an outer wall of the nacelle forming the outer periphery of the nacelle. In yet another embodiment that is a variant of the previous embodiment, the first external surface of the nacelle is a surface of an inner wall of the nacelle that externally delimits a secondary flow path of the turbomachine.In one embodiment, the nacelle comprises an inner wall and an outer wall surrounding the inner wall. The outer wall includes the first outer surface, and the inner wall includes an outer surface. Channels extend between a first and a second wall comprising the first outer surface of the nacelle, and channels extend between the outer surface of the inner wall and the first outer surface of the nacelle. In one embodiment, the channels extend in contact with the outer wall of the nacelle. In another embodiment, the channels extend in contact with the inner wall of the nacelle. In yet another embodiment, the channels extend between the inner and outer walls of the nacelle. In one implementation, the channels are radially at substantially the same distance from the outer surface of the inner wall as from the outer surface of the outer wall.(By "approximately the same radial distance," we mean between +10% and -10% of the distance.) In one embodiment, the channels of the first capacitor are superimposed in the nacelle to form a grid that follows the curvature of the first external surface of the nacelle. In another embodiment, the capillary pressure drop in the porous capillary wick compensates for the sum of the pressure losses experienced by the fluid in the first capacitor circuit.Another aspect of the invention relates to a propulsion assembly comprising the aircraft turbomachine including a turbomachine and a nacelle surrounding the turbomachine, support arms connecting the nacelle to a central compartment, and a cooling system for a lubricating fluid of the turbomachine according to the first aspect of the invention (with or without the various possible combinations of embodiments and examples, described previously), the cooling system comprising a closed circuit of the lubricating fluid including the first cooling volume of the lubricating fluid in the exchanger, a volume of the engine connected to an inlet and an outlet of the first cooling volume of the lubricating fluid in the exchanger, and a pump mounted between the engine volume and the first cooling volume of the lubricating fluid of the exchanger, allowing the lubricating fluid to circulate in the closed circuit..

[0018] The invention and its various applications will be better understood by reading the following description and examining the accompanying figures. BRIEF DESCRIPTION OF THE FIGURES

[0019] The figures are presented for illustrative purposes only and are in no way limiting to the invention. [ Fig. 1 [ ] shows a schematic diagram of a cooling system according to the invention. Fig. 2 [ ] shows a schematic diagram of a condenser in a nacelle of a cooling system according to an example of the invention. Fig. 3 [ ] shows a schematic diagram of a cooling system according to a first example of a first embodiment of the invention comprising two condensers according to the example of the figure 2 . [ Fig. 4 [ ] shows a schematic diagram of a cooling system according to a second example of the first embodiment of the invention comprising two condensers according to the example of the figure 2 . [ Fig. 5 [ ] shows a schematic diagram of a cooling system according to a third example of the first embodiment of the invention comprising two condensers according to the example of the figure 2 . [ Fig. 6 [ ] shows a schematic diagram of a cooling system according to a first example of a second embodiment of the invention comprising two condensers according to the example of the figure 2 . [ Fig. 7 [ ] shows a schematic diagram of a cooling system according to a second example of a second embodiment of the invention comprising two condensers according to the example of the figure 2 . [ Fig. 8 [ ] shows a schematic diagram of different radial locations of the condenser of the figure 2 in a nacelle of a cooling system according to an example of the invention. DETAILED DESCRIPTION

[0020] The figures are presented for illustrative purposes only and are in no way limiting to the invention.

[0021] Unless otherwise specified, the same element appearing on different figures has a unique reference.

[0022] [ Fig. 1 ] shows a schematic representation, a principle diagram of a cooling system according to the invention of a lubricating fluid for an aircraft turbomachine.

[0023] The cooling system according to the invention comprises a first cooling volume 201 for the lubricating fluid of a heat exchanger 213 and a first cooling circuit 230 for a two-phase heat transfer fluid via a two-phase fluid loop with thermo-capillary pumping. The lubricating fluid circulates in a closed circuit 120 and loses heat in the heat exchanger 213, which is recovered by a two-phase heat transfer fluid to cool the lubricating fluid.

[0024] Using a two-phase fluid loop instead of a single-phase fluid loop (where the fluid remains in a liquid state) allows the use of less dense working fluids, which also reduces the system's mass. Indeed, in a single-phase loop, the choice is limited to certain types of fluids that do not undergo a phase change across the entire pressure / temperature range corresponding to the different flight phases. Fluids that meet this criterion are generally dense.

[0025] On the figure 1The closed circuit 120 of a propulsion system including the turbomachine can be schematically represented. The closed circuit 120 comprises a motor 1 through which the turbomachine's lubricating fluid circulates, a heat exchanger 213, and a pump 12. The closed circuit 120 of the lubricating fluid thus includes the first cooling volume 201 of the lubricating fluid in the heat exchanger 213, a volume of the motor 1 connected to an inlet and outlet of the first cooling volume of the lubricating fluid in the heat exchanger 213, and the pump 12, which is mounted between the volume of the motor 1 and the first cooling volume 201 of the lubricating fluid in the heat exchanger 213, allowing the lubricating fluid to circulate in the closed circuit 120.Pump 12 is shown here between the outlet of the engine 1 volume and the inlet of the first volume 201 of the heat exchanger 213 because, in this example, the first volume 201 of the heat exchanger 213 can be located higher in the propulsion assembly mounted on the aircraft than the engine 1 volume. However, it could also be mounted between the outlet of the first cooling volume 201 of the heat exchanger 213 and the inlet of the engine 1 volume. Conversely, the engine 1 volume could be located higher in the propulsion assembly mounted on the aircraft than the first cooling volume 201 of the heat exchanger 213. The engine 1 may include a reservoir for the engine 1 lubrication fluid connected to pump 12 and one or more nozzles or lubrication fluid for mechanical parts of the engine.

[0026] By high or low in the demand, we refer to the direction of gravity for an aircraft propulsion system such that the deployed landing gear is lower than the aircraft propulsion system.

[0027] The first cooling circuit 230 of the two-phase heat transfer fluid includes a first condenser 3 and a first capillary evaporator 2 of the exchanger 213 comprising a porous capillary wick 200.

[0028] The capillary evaporator 2 generates the driving pressure responsible for circulating the two-phase heat transfer fluid throughout the loop. The heat to be dissipated from the lubricating fluid (engine oil, for example) is transferred to the heat transfer fluid within this capillary evaporator 2. The capillary evaporator 2 comprises a first volume 203 containing a liquid inlet for the heat transfer fluid and a second volume 223 containing a vapor outlet for the heat transfer fluid. The porous wick 200 is the element that separates the second volume 223 (the two-phase heat transfer fluid in its vapor state) from the first volume 203 (the two-phase heat transfer fluid in its liquid state). The porous wick 200 contains pores which, by fracturing the vapor-liquid interface, produce a capillary pressure jump corresponding to the difference between the liquid and vapor phase pressures.The overpressure of the vapor above the vaporization surface forces the migration of this vapor towards the condenser, causing the heat transfer fluid to move in the loop as soon as the capillary pressure jump compensates for the sum of the pressure losses experienced by the fluid throughout the loop. According to a first embodiment of the cooling system, the heat exchanger 213 can be integrated directly into a central compartment 5 delimited by a housing of the turbomachine engine 1, as explained in the examples of the first embodiment described below with reference to [reference to relevant examples]. figures 3 to 5 The heat exchanger 213, according to a second embodiment of the cooling system, can be located within a nacelle 4 of the propulsion assembly in the "lower" part, as explained later in the examples of the second embodiment with reference to figures 6 to 8 .

[0029] In the following, identical characteristics are designated by the same references and different characteristics of the second embodiment from those of the first embodiment are designated by the same number followed at the end of the same number by a quotation mark «'”.

[0030] In this figure, the evaporator 2 is represented according to a schematic diagram comprising a porous wick forming a straight wall separating the two volumes forming chambers, but the porous wick 200 can, for example, be hollow and separate the first volume 203 by surrounding the second volume 223. For example, the porous wick 200 can be a hollow cylinder.

[0031] On the figures 3 to 5The capillary evaporator 2 according to the first embodiment allows for a reduction in the quantity of lubricating fluid in the closed circuit 120 and an increase in that of the heat transfer fluid. This has the advantage of reducing the total mass (part of the heat transfer fluid is in vapor form).

[0032] The cooling system may include two cooling circuits 230 1, 230 2, 230 1', 230 2' as shown in the examples of figures 3 And 7 respectively of the first and second embodiments, each comprising a capillary evaporator 2' 1 , 2' 2 , 2' 1 , 2' 2' . The references of the characteristics of the first cooling circuit, 230 1 , 230 1' , are supplemented by an index "1" and the references of the characteristics of the second cooling circuit 230 2 , 230 2' are supplemented by an index "2".

[0033] The first cooling circuit 230 may include two condensers 3 as in the various examples shown in each of the figures 3 to 8 of the first and second embodiments.

[0034] There figure 2Figure 3 represents an example of a condenser 3 in a nacelle 4. The condenser 3 includes a vapor inlet for the heat transfer fluid in its vapor state, that is, an inlet through which the heat transfer fluid enters the condenser in vapor phase, connected to a first vapor conduit 31, that is, a conduit in which the heat transfer fluid circulates in vapor phase. The condenser 3 further includes channels 32 extending along a first external surface of the turbomachine nacelle 4 for their cooling. The channels are connected in parallel. The channels 32 are superimposed in the nacelle 4 to form a grid that follows the curvature of the first external surface of the nacelle, thus cooling the fluid in its vapor state circulating in the condenser 3 and causing it to change to a liquid state before exiting the condenser.

[0035] The condenser 3 therefore includes a liquid outlet of the heat transfer fluid, that is to say an outlet through which the heat transfer fluid exits in liquid phase from the condenser, connected to a first liquid conduit 33, that is to say a conduit in which the heat transfer fluid circulates in liquid phase.

[0036] The first cooling circuit 230 therefore includes the liquid conduit 33 connected to the liquid outlet of the condenser 3 and to the liquid inlet of the first volume 203 of the capillary evaporator 2 to transfer the heat transfer fluid in the liquid state cooled in the condenser 3. The vapor conduit 31 is connected to the vapor outlet of the second volume 223 of the capillary evaporator 2 and to the vapor inlet of the condenser 3 to transfer the heat transfer fluid in the vapor state heated in the capillary evaporator 2.

[0037] The first example of the first embodiment represented in will now be described. figure 3 .

[0038] In this first example of this first embodiment, as described previously, the capillary evaporator 2 is mounted in a central compartment 5 delimited by a housing of the engine 1, and the first cooling circuit 230 comprises two condensers 31, 32 located on either side of a vertical plane comprising the axis of rotation of the turbomachine, surrounded by the nacelle 4. The nacelle 4 surrounding the turbomachine engine includes support arms (not shown) connected to the engine. In this embodiment, the two condensers 31, 32 are each mounted in an upper part of the nacelle 4. By upper part of the nacelle 4, we mean the part above a horizontal plane passing through the axis of rotation of the turbomachine. By horizontal plane, we mean that the axis of rotation of the turbomachine is aligned horizontally.

[0039] In this first example, the first cooling circuit 230 includes a main vapor conduit 331 and a main liquid conduit 332, each extending on either side along the periphery of the central compartment 5 and into one of the support arms extending upwards into the nacelle 4.

[0040] The main vapor line 331 is connected at one end to the outlet of the second volume 223 of the capillary evaporator 2 and at the other end to a first vapor line 311 and a second vapor line 312, each respectively connected to the first and second condensers 31, 32, to connect them to the outlet of the second volume 223 of the capillary evaporator 2.

[0041] The main liquid conduit 332 is connected on one end to the inlet of the first volume 203 of the capillary evaporator 2 and on the other end to a first liquid conduit 331 and a second liquid conduit 332, each respectively connected to the first and second condensers 31, 32.

[0042] Thus, having the two condensers in the upper part allows gravity to be used for the return of the heat transfer fluid in liquid state to the capillary evaporator 2.

[0043] Furthermore, having the two condensers on either side of the vertical plane allows for good efficiency from at least one of the two condensers when the aircraft is moving in a long turn producing a rotation around the aircraft's roll axis.

[0044] The second example of the first embodiment of the cooling system shown in figure 4is different from this first example in that it includes a second cooling circuit 230 2 and a second capillary evaporator 2 2 of the exchanger 213, similar to the first capillary evaporator 2 1 '.

[0045] The exchanger 213 may include a second lubricating fluid cooling volume comprising this second capillary evaporator 22, separate from the first lubricating fluid cooling volume 201 or may include this second capillary evaporator 22 in contact with the first cooling volume 201 to cool it.

[0046] In this second example, each cooling circuit 230 1 , 230 2 differs from the first example in that they lack a main liquid line and a main vapor line.

[0047] Thus the first and second cooling circuits 230 1 , 230 2 respectively comprise a first vapor conduit 31' 1 and a second vapor conduit 31' 2 directly connected to the outlet of the second chamber 223 of the corresponding capillary evaporator 2 1 , 2 2.

[0048] The first and second cooling circuits 230 1 , 230 2 therefore also respectively a first liquid conduit 33' 1 and a second liquid conduit 33' 2 directly connected to the inlet of the first chamber 203 of the corresponding capillary evaporator 2 1 , 2 2.

[0049] The first and second steam conduits 31' 1, 31' 2 and the first and second liquid conduits 33' 1, 33' 2 therefore each extend on either side along the periphery of the central compartment 5 and into one of the support arms.

[0050] Having two cooling circuits 230 1 , 230 2 allows for two independent circuits, thus providing redundancy, but also improves performance in the event of a long turn of the aircraft (rotation around the roll axis).

[0051] The third example of the first embodiment of the cooling system shown in figure 5 is different from the first example in that the main steam line 331' 0, and the main liquid line 332' 0 extend from the central compartment 5 in one of the support arms to a lower part of the nacelle 4 and in that the first steam line 31" 1, the second steam line 31" 2, the first liquid line 33" 1, the second liquid line 33" 2 each extend on either side along the periphery of a lower part of the nacelle 4 towards the inlet and outlet of the corresponding condenser 3 1, 3 2.

[0052] The first example of the second embodiment of the cooling system shown in will now be described. figure 6 .

[0053] As explained previously, the second embodiment differs from the first embodiment in that the exchanger 213 is located in the lower part of the nacelle 4. Thus the capillary evaporator 2' is located in the lower part of the nacelle 4. The first and second condensers 31', 32', are also located in the lower part of the nacelle 4 such that the exchanger 213 is located between the first and second condensers 31', 32'.

[0054] In this first example, the cooling system has similarities with the second example of the first embodiment in that they are devoid of a main liquid conduit and a main vapor conduit.

[0055] Thus in this first example, the first cooling circuit 230 1 includes a first vapor conduit 31 1' and a second vapor conduit 31 2' in the lower part of the nacelle 4 each connected to an outlet of the second volume 223 of the capillary evaporator 2' respectively to the inlet of the first and second condenser 3 1' , 3 2' .

[0056] The second example of the second embodiment of the cooling system shown in figure 7 is different from this first example in that it includes a second cooling circuit 230 2' and a second capillary evaporator 2 2' of the exchanger 213, similar to the first capillary evaporator 2 1'.

[0057] In this second example the first cooling circuit 230 1' therefore includes a first vapor conduit 31' 1, connected on one side to the outlet of the second volume 223 of the first evaporator 2' 1' and on the other side to the inlet of the condenser 3 1' and a first liquid conduit 33' 1' connected on one side to the inlet of the first volume 203 of the first evaporator 2' 1' and on the other side to the outlet of the condenser 3 1'. The second cooling circuit 230 2' therefore includes a second vapor conduit 31' 2' connected on one side to the outlet of the second volume 223 of the second evaporator 2' 2' and on the other side to the inlet of the second condenser 3 2' and a second liquid conduit 33' 2' connected on one side to the inlet of the first volume 203 of the second evaporator 2' 2' and on the other side to the outlet of the second condenser 3 2'.

[0058] In this second embodiment, having the exchanger in the lower part of the nacelle reduces the total pressure losses in the heat transfer fluid circulation loop.

[0059] There figure 8 represents different examples representing different radial locations of condenser 3 of the figure 2 in a nacelle 4 of a cooling system according to an example of this second embodiment of the invention but could also be for the first embodiment.

[0060] The nacelle 4 comprises an inner wall 40 and an outer wall 41 surrounding the inner wall 40, each comprising an inner surface opposite the other. The outer wall 41 comprises an external surface 410 in contact with the air surrounding the nacelle 4, and the inner wall 40 comprises an external surface 400 in contact with the air surrounding the nacelle 4.

[0061] The external surface 400 of the internal wall 40 is therefore inside the nacelle 4 and externally delimits a secondary flow channel with the turbomachine. The external surface 410 of the external wall 41 of the nacelle 4 is an external surface forming the outer periphery of the nacelle 4.

[0062] The condenser 31 is arranged so that the channels 32 extend along the outer surface 400 of the inner wall 40 of the nacelle 4, closer radially to the outer surface 400 of the inner wall 40 than to the outer surface 410 of the outer wall 41. In this example, the channels 32 are located against the inner surface of the inner wall 40, but in another example not shown, they can be located on the outer surface 400 of the inner wall 40 of the nacelle 4. In both examples, this allows the heat transfer fluid to be cooled by the air from the turbomachine's secondary flow even when the aircraft is on the ground or during takeoff.

[0063] The condenser 32 represents a position in which the channels 32 extend along the external surface 410 of the nacelle 4, and radially closer to the external surface 410 of the external wall 41 than to the external surface of the internal wall 40. The channels 32 are in this example located against the inner surface of the external wall 41 but can, according to another example not shown, be located on the external surface 410 of the external wall 41 of the nacelle 4. This allows the heat transfer fluid to be cooled by the external airflow licking the nacelle, particularly when the aircraft is in flight.

[0064] The condenser 3 represents a position in which the channels 32 extend between the inner wall 40 and the outer wall 41 of the nacelle 4, and in this case substantially as close radially to the outer surface 400 as to the outer surface 410. This allows the heat transfer fluid to be cooled by the external airflow licking the nacelle in particular when the aircraft is in flight but also by the secondary flow in particular when the aircraft is on the ground or taking off.

[0065] The channels can also be located in the inner wall 40 and / or the outer wall 41.

[0066] In general, a condenser 3 will preferably be placed at a higher height than a capillary evaporator 2, to take advantage of gravity in the circulation of the heat transfer fluid.

[0067] Unless otherwise specified, the same element appearing on different figures has a unique reference.

Claims

1. Cooling system for a lubricating liquid of an aircraft turbomachine comprising: 1) a heat exchanger (213) comprising a first cooling volume (201) for the lubricating liquid, 2) a first cooling circuit (230) for a two-phase heat transfer fluid by a circulation loop of the two-phase fluid, with thermocapillary pumping, comprising: a) o a first condenser (3, 31, 31') comprising: i) a vapor inlet for the heat transfer fluid in the vapor state, ii) channels (32) adapted to extend along a first external surface of a nacelle (4) of the turbomachine to cool the channels (32), the channels (32) being connected in parallel to each other, to allow cooling of the fluid circulating in the condenser (3) from a vapor state to a liquid state, iii) a liquid outlet for the heat transfer fluid in the liquid state, b) a first capillary evaporator (2, 2') of the heat exchanger (213), to absorb the calories from the first volume (201) in which the lubricating liquid circulates, the first evaporator (2, 2') comprising: i) a first volume (203) comprising a liquid inlet for the heat transfer fluid connected to the liquid outlet of the first condenser, ii) a second volume (223) comprising a vapor outlet for the heat transfer fluid, connected to the vapor inlet of the first condenser, iii) a porous capillary wick (200) located between the first volume (203) and the second volume (223), allowing separation of the two-phase fluid between a liquid phase and a gaseous phase by a capillary pressure jump.

2. Cooling system for a lubricating liquid of an aircraft turbomachine according to claim 1, comprising a second condenser (32,32') similar to the first condenser (31, 31'), in which the first and second condensers (31, 31', 32, 32') are adapted to be located on either side of a vertical plane comprising the axis of the nacelle (4).

3. Cooling system for a lubricating liquid of an aircraft turbomachine according to the previous claim, in which the first cooling circuit (230) for the two-phase heat transfer fluid comprises the second condenser (32, 32') having its vapor inlet connected to the vapor outlet of the second volume (223) of the first evaporator (2, 2').

4. Cooling system for a lubricating liquid of an aircraft turbomachine according to claim 2, comprising: • a second capillary evaporator (22, 22') of the heat exchanger (213), similar to the first capillary evaporator (21, 21'), • a second cooling circuit (2302, 2302') for a two-phase heat transfer fluid similar to the first cooling circuit (2301, 2301') for a two-phase heat transfer fluid, the second cooling circuit (2302, 2302') comprising the second condenser which has its vapor inlet connected to the vapor outlet of the second volume of the second evaporator.

5. Cooling system for a lubricating liquid of an aircraft turbomachine according to one of the preceding claims, in which the first condenser (31, 31') is located higher than the first evaporator (21, 21') so that the fluid in a liquid state at the outlet of the condenser flows by gravity to the evaporator.

6. Cooling system for a lubricating liquid of an aircraft turbomachine according to one of the preceding claims, comprising a vapor conduit (31, 31'1, 31"1, 31'1', 311') connecting the vapor outlet of the second volume (223) of the first evaporator (2, 2', 21', 2'1') to the vapor inlet of the first condenser (31, 31'), the vapor conduit (31, 31'1, 31"1, 31'1', 311'') being adapted to be located along the axis of the nacelle (4) upstream of the inlet of the first evaporator (2, 2', 21', 2'1').

7. Cooling system for a lubricating liquid of an aircraft turbomachine according to one of the preceding claims, in which the heat exchanger (312) is adapted to be located in a lower part of the nacelle (4) situated below the turbomachine.

8. Cooling system for a lubricating liquid of an aircraft turbomachine according to one of claims 1 to 6, in which the heat exchanger (312) is adapted to be located in a central compartment (5) of the turbomachine surrounded by the nacelle (4), the cooling circuit comprising a vapor conduit (31, 31'1, 31"1, 31'1', 311') connecting the vapor outlet of the second volume (223) of the first evaporator (2, 2', 21', 2'1') to the vapor inlet of the first condenser (31, 31') passing, in operation, inside a support arm of the turbomachine.

9. Cooling system for a lubricating liquid of an aircraft turbomachine according to one of the preceding claims, in which the first condenser (31) is adapted to be located in an upper part of the nacelle (4) situated above the axis of the turbomachine.

10. Cooling system for a lubricating liquid of an aircraft turbomachine according to one of claims 1 to 8, in which the first condenser (31') is adapted to be located in a lower part of the nacelle (4) situated below the axis of the turbomachine.

11. Propulsive assembly for an aircraft comprising a turbomachine and a nacelle (4) surrounding the turbomachine, support arms connecting the nacelle to a central compartment (5) of the turbomachine, and a cooling system for a lubricating liquid of the turbomachine according to one of the preceding claims, the cooling system comprising a closed circuit (120) for the lubricating liquid comprising the first cooling volume (201) for the lubricating liquid in the heat exchanger (213), an engine volume (1) connected to an inlet and an outlet of the first cooling volume (201) for the lubricating liquid in the heat exchanger (213), and a pump (12) mounted between the engine volume (1) and the first cooling volume (201) for the lubricating liquid of the heat exchanger, allowing circulation of the lubricating liquid in the closed circuit (120).

12. Propulsive assembly for an aircraft according to the preceding claim, in which the first external surface (410) of the nacelle (4) is a surface of an outer wall (41) of the nacelle (4) forming the external periphery of the nacelle (4).

13. Propulsive assembly for an aircraft according to claim 11 or 12, in which the first external surface (400) of the nacelle (4) is a surface of an inner wall (40) of the nacelle (4) which externally delimits a secondary flow duct of the turbomachine.