Turbine engine for an aircraft and method for cooling same

Abstract

The invention relates to a turbine engine (2) for an aircraft, the turbine engine comprising: - an annular flow path (V) for a gas flow (F3); - a motor (M); - at least one piece of equipment (E); - a first ACOC-type heat exchanger (210), which is capable of being swept by the gas flow (F3) and which comprises a first oil circuit (C1) connected to a cooling system (S1) of the motor; - a second ACOC-type heat exchanger (220), which is capable of being swept by the gas flow (F3) and which comprises a second oil circuit (C2) connected to a system (S2) for cooling the at least one piece of equipment (E); and - a third, OCOC-type heat exchanger (230), which comprises third and fourth oil circuits (C3, C4) that are capable of exchanging heat energy with one another.

Description

[0001] DESCRIPTION

[0002] TITLE: TURBOMACHINE FOR AN AIRCRAFT AND ITS COOLING PROCESS

[0003] Technical field of the invention

[0004] The present invention relates to a turbomachine for an aircraft, and a method of cooling this turbomachine.

[0005] Technical background

[0006] The technical background includes in particular documents US-A1 - 2012 / 085528, US-A1 -2023 / 332543, US-A-4,696,156 and US-A1 -

[0007] 2010 / 212857.

[0008] A turbomachine for an aircraft includes various components and / or equipment that need to be lubricated and / or cooled, such as rolling bearings, gears, etc. In addition, new generation turbomachines have on-board electrical machines for hybridization, these machines needing to be lubricated and cooled by oil. The heat released by these components is transported by the oil and removed to cold sources available on board.

[0009] Among the cold sources that exist on board, there are air and fuel. The oil can be cooled first by the fuel in a fuel / oil exchanger (generally known by the acronym FCOC, an acronym for the English expression "Fuel Cooled Oil Cooler"), which also has the function of heating the fuel before it enters the combustion chamber. FCOC exchangers are not sufficient to absorb all the heat dissipation because the fuel temperature is limited in view of safety constraints. Additional cooling can be provided by an air / oil exchanger (generally known by the acronym ACOC for "Air Cooled Oil Cooler"). The air constitutes a cold source for cooling the oil and can come for example from the secondary vein in the case of a double-flow turbomachine, or from the tertiary vein in the case of a triple-flow turbomachine as illustrated in figure 1 and described in document W0-A1 - 2023 / 099533.

[0010] ACOC exchangers are increasingly in demand in the future generation of engines due to the significant increase in heat dissipation, mainly due to:

[0011] - the engines of the future are larger, which increases the need for lubrication and cooling with oil;

[0012] - the presence of a speed reducer in new motor architectures, this reducer transmits very high mechanical power and needs to be lubricated and cooled by oil,

[0013] - adding electrical machines on board an engine for hybridization, these machines need to be lubricated and cooled by oil.

[0014] An engine may have several oil circuits, each with a defined function, for example a first circuit for cooling the oil dedicated to lubricating and cooling the engine, a second for lubricating and cooling the electrical machines, etc.

[0015] The oil temperature and flow rate must be controlled according to the corresponding circuit. For example, the oil used to lubricate and cool electrical machines has a different temperature range than that used for engine cooling. Therefore, the ACOC heat exchanger is preferably divided into several exchangers, each exchanger being dedicated to a given oil circuit. These different exchangers are installed in parallel in one of the turbomachine's veins, and occupy a part of this vein.

[0016] In Figure 2, the ACOC heat exchanger 1 can be used to cool the engine cooling oil. The ACOC heat exchanger 2 can be used to cool the oil of the electric machines for the hybridization of the turbomachine. Each of the ACOCs 1 and 2 occupies a given angular sector of the vein. The vein V is delimited by coaxial annular walls 101, 102 which are connected to each other by structural arms 100 to ensure mechanical strength. The angular sectors denoted by "A" are sectors where the air circulates without any obstacle (no ACOC exchanger).

[0017] Each of the two exchangers must be sized for the most sizing case in order to dissipate the thermal power over all other flight phases.

[0018] The ACOC 1 heat exchanger is divided in Figure 3 into two parts:

[0019] - 1 a: part of the ACOC 1 exchanger used to ensure the cooling of the engine oil during all flight phases except the most dimensioning flight phases

[0020] - 1 b: part of the ACOC 1 exchanger which is added to part 1 a in order to ensure additional cooling during the most critical flight phases (case of an aircraft on the ground on an extremely hot day as an example).

[0021] The ACOC 2 heat exchanger is divided in Figure 3 into two parts:

[0022] - 2a: part of the ACOC 2 exchanger used to cool the oil of the electrical machines during all flight phases except the most dimensioning flight phases, and

[0023] - 2b: part of the ACOC 2 exchanger which is added to part 1a in order to ensure additional cooling in the most dimensioning flight phases

[0024] The two ACOC exchangers (ACOC 1 and ACOC 2) provide different cooling functions. This results in different dimensioning phases for each of the two exchangers and leads to oversized exchangers on the other phases and flight conditions, which has a direct impact on the air-side pressure drop (aerothermal performance) and on the mass of the exchangers. This has the effect of increasing the specific consumption (SFC) of the turbomachine and consequently the fuel consumption (FB for Fuel Burn). There is therefore a need to optimize the aerothermal performance of these exchangers.

[0025] The solution presented in EP-A1-2 472 067 proposes an ACOC exchanger buried in a compartment of the turbomachine. The heat exchanger is integrated in a cavity opening into a radially internal wall of the secondary stream. A portion of the secondary air flow, taken from the secondary stream, passes through the buried heat exchanger where the latter is reheated and is reinjected into the secondary stream. A controlled scoop, formed for example by a pivoting and / or translationally movable movable flap, is arranged at the inlet of the cavity so as to extend into the secondary stream and to produce an adjustable air flow towards the buried heat exchanger. The movable flap of the scoop can generate pressure losses in the secondary stream when it is open. The movable flap is controlled so as to close when the need for heat exchange in the buried heat exchanger becomes zero.

[0026] Patent WO-A1 -2022 / 123168A1 proposes a solution to optimize the aerothermal performance of the exchanger in a cavity and to reduce pressure losses thanks to a coupling between a movable flap and a movable member which allows adaptation to the different flight phases of the turbomachine and the aircraft. It relates to a heat exchange system for a turbomachine, comprising a cavity comprising an air inlet, a heat exchanger arranged in the cavity and comprising a first circuit in which a first fluid can circulate, a movable flap mounted at the air inlet and moving between two positions allowing or prohibiting, respectively, the circulation of an air flow in the cavity, a control device comprising a movable member intended to drive the movement of the movable flap.The control device is arranged in the supply circuit of the exchanger and is configured to authorize or prohibit the circulation of the first fluid towards the exchanger and to act simultaneously on one of the two positions of the movable flap. The present invention proposes a solution to the problems and needs of the prior art, which is simple, effective and economical.

[0027] Summary of the invention

[0028] The invention relates to a turbomachine for an aircraft, this turbomachine comprising:

[0029] - an annular flow vein for a gas stream,

[0030] - an engine,

[0031] - at least one piece of equipment,

[0032] - a first heat exchanger of the ACOC type, which is capable of being swept by said gas flow and which comprises a first oil circuit connected to an engine cooling system,

[0033] - a second heat exchanger of the ACOC type, which is capable of being swept by said gas flow and which comprises a second oil circuit connected to a cooling system of said at least one piece of equipment, characterized in that it further comprises:

[0034] - a third heat exchanger of the OCOC type, which comprises third and fourth oil circuits capable of exchanging calories between them, and

[0035] - connecting valves from the first circuit to the third circuit, and from the second circuit to the fourth circuit, the connecting valves being capable of adopting two configurations:

[0036] - a first configuration in which the third circuit is isolated from the first circuit, and the fourth circuit is isolated from the second circuit, and

[0037] - a second configuration in which the third circuit is connected in series to the first circuit for supplying oil to the first circuit by the third circuit, and the fourth circuit is connected in series to the second circuit for supplying oil to the second circuit by the fourth circuit.

[0038] In the present application, ACOC exchanger means an air / oil exchanger, ACOC being the acronym for “Air Cooled Oil Cooler”, and OCOC means an oil / oil exchanger, OCOC being the acronym for “Oil Cooled Oil Cooler”.

[0039] The invention proposes to use an oil / oil exchanger (OCOC) which ensures the heat exchange between the oil of the first circuit of the first exchanger and the oil of the second circuit of the second exchanger.

[0040] This solution is applicable even in the case where the type of cooling oil is not the same for the first and second exchangers. Indeed, as the oils of the first and second exchangers circulate in circuits independent of the third exchanger, there is no risk of mixing these oils in the third exchanger.

[0041] The invention makes it possible to size the first exchanger to ensure cooling of the oil during all flight phases except for the most dimensioning phases, and to add the third exchanger to ensure additional cooling of the oil during the dimensioning phases of the first exchanger. This third exchanger is common to the first and second exchangers. The second exchanger is dimensioned to pass all the corresponding flight cases, including the most dimensioning case for cooling the equipment(s).

[0042] Since the dimensioning phases for the first and second exchangers are not identical, the third exchanger can be used to provide additional cooling for the most dimensioning phases for the first exchanger by ensuring an exchange between the oil of the first exchanger and the oil of the second exchanger, which makes it possible to transfer part of the thermal power of the oil of the first exchanger to the oil of the second exchanger via the third exchanger. The part of the thermal power given up by the oil of the first exchanger will be dissipated in the second exchanger. This is entirely feasible since the second exchanger does not operate at its maximum dissipation because the most dimensioning phases of the first exchanger do not generally correspond to those of the second exchanger. The invention makes it possible to reduce the size of the first exchanger and thus to reduce the pressure losses compared to the reference case.The size of the second exchanger can remain unchanged compared to the reference case.

[0043] The transfer of thermal power generally takes place from high temperature to low temperature, that is to say from the first exchanger to the second exchanger.

[0044] The turbomachine according to the invention may comprise one or more of the following characteristics, taken in isolation from one another, or in combination with one another; these characteristics having in particular the advantages of optimizing the compactness of the geared motor:

[0045] - in the second configuration, an output of the third circuit is connected to an input of the first circuit, and an output of the fourth circuit is connected to an input of the second circuit;

[0046] - the first circuit comprises a first inlet valve connected to an inlet of the first circuit, and the second circuit comprises a second inlet valve connected to an inlet of the second circuit;

[0047] - the connecting valves include:

[0048] - a first connecting valve mounted between a connecting point of the inlet of the first circuit, located upstream of the first inlet valve, and an inlet of the third circuit, and

[0049] - a second connecting valve mounted between a connecting point of the inlet of the second circuit, located upstream of the second inlet valve, and an inlet of the fourth circuit;

[0050] - the third circuit comprises an outlet connected to another connection point of the inlet of the first circuit, located downstream of the first inlet valve, and the fourth circuit comprises an outlet connected to another connection point of the inlet of the second circuit, located downstream of the second inlet valve;

[0051] - the third heat exchanger is located outside said vein;

[0052] -- said first heat exchanger has an angular sector shape and occupies only part of the angular extent of said vein; this integration model makes it possible to ensure better integration of the exchanger(s) in the vein and to reduce the pressure loss on the air flow side, but also to optimize the aerothermal performances;

[0053] -- said first heat exchanger has an angular extent of between 30 and 90°;

[0054] -- said second heat exchanger has an angular sector shape and occupies only part of the angular extent of said vein;

[0055] -- said second heat exchanger has an angular extent of between 30 and 90°;

[0056] -- said first and second heat exchangers are crossed by the same plane perpendicular to a longitudinal axis of the turbomachine;

[0057] -- part of the angular extent of said vein is free and is therefore not occupied by one or more exchangers, and in particular by said first and second exchangers;

[0058] - said vein is a secondary flow vein for a secondary gas flow, the turbomachine further comprising a primary flow vein for a primary gas flow inside said engine;

[0059] - said vein is a tertiary flow vein of a tertiary gas flow, the turbomachine further comprising two annular veins, respectively primary and secondary, for flow of primary and secondary gas flow; the use of this third vein allows additional cooling during certain phases of use of the turbomachine, this vein having a reduced section it is more sensitive to aerodynamic losses; the annular integration therefore provides these benefits;

[0060] - the tertiary vein is intercalated radially between the primary and secondary veins;

[0061] -- the tertiary vein is annular;

[0062] - the third heat exchanger has a cooling capacity lower than that of the first heat exchanger, and lower than that of the second heat exchanger;

[0063] - said at least one piece of equipment comprises electrical machines;

[0064] -- said engine is a gas generator. The invention further relates to a cooling method in a turbomachine as described above, in which it comprises two steps corresponding to the two aforementioned configurations:

[0065] - a first step corresponding to the first configuration, in which the engine cooling system is supplied by oil leaving the first exchanger, and the cooling system of said at least one piece of equipment is supplied by oil leaving the second exchanger, and

[0066] - a second step corresponding to the second configuration, in which the engine cooling system is supplied by oil passing through the third exchanger then through the first exchanger, and the cooling system of said at least one piece of equipment is supplied by oil passing through the third exchanger then through the second exchanger.

[0067] The method according to the invention may comprise one or more of the following features or steps, taken in isolation from one another, or in combination with one another:

[0068] - the second stage occurs when the aircraft is on the ground and the ambient outside temperature is above a predetermined threshold,

[0069] - the first step occurs by default in other cases;

[0070] - the first heat exchanger is supplied with oil at a maximum temperature of 180°C, and the second heat exchanger is supplied with oil at a maximum temperature of 100°C.

[0071] Brief description of the figures

[0072] Other characteristics and advantages will emerge from the following description of a non-limiting embodiment of the invention with reference to the appended drawings in which:

[0073] [Fig.1] Figure 1 is a schematic axial sectional view of a turbomachine for an aircraft; [Fig.2] Figure 2 is a very schematic cross-sectional view of a flow vein of a gas flow, in which heat exchangers are located,

[0074] [Fig.3] Figure 3 is a view similar to that of Figure 2 and shows a breakdown of the heat exchangers,

[0075] [Fig. 4] Figure 4 is a schematic cross-sectional view of a flow vein of a gas stream, and three heat exchangers, and illustrates one embodiment of the invention,

[0076] [Fig.5] Figure 5 is a very schematic view of the three heat exchangers of Figure 4 and shows a first configuration of exchanger connection valves, and

[0077] [Fig.6] Figure 6 is a view similar to that of Figure 4 and shows a second configuration of the exchanger connecting valves.

[0078] Detailed description of the invention

[0079] Figure 1 shows a turbomachine 2 for an aircraft. This turbomachine 2 illustrates the prior art as described in document WO-A1 - 2023 / 099533, but can be used to describe an installation environment of the invention.

[0080] The turbomachine 2 moves in an air flow F whose movement relative to the turbomachine 2 is generated by the rotation of the propeller 4 and the advancement of the aircraft on which the turbomachine 2 is mounted.

[0081] The air flow F is separated by a first separation nozzle 10 into a radially internal air flow F' and a radially external air flow F2, called secondary flow F2. The propeller 4 can be arranged upstream of the first separation nozzle 10 or downstream.

[0082] The radially internal air flow F' passes through a moving wheel 12 which directs the latter towards a second separation nozzle 14 capable of separating the radially internal air flow F' into a primary flow F1 and a tertiary flow F3, the latter being distinct from the secondary flow F2. The first separation nozzle 10 comprises an internal wall forming a first external guide wall 11 of the radially internal air flow F', said first external guide wall 11 forming a convex profile seen from said radially internal air flow F'.

[0083] The second separation nozzle 14 comprises an external wall forming a second external guide wall 13 for the radially internal air flow F' having passed through the moving wheel 12, said second external guide wall 13 forming a convex profile seen from the tertiary flow F3.

[0084] The tertiary flow F3 enters a tertiary flow vein 16 radially external to said primary flow F1. The tertiary flow F3 passes through a heat exchanger 18 arranged in the tertiary flow vein 16.

[0085] The turbomachine 2 is illustrated symmetrically with respect to the longitudinal axis 8. Indeed, the tertiary flow vein 16 is annular and circumferentially continuous over 360° around the longitudinal axis 8. For this purpose, the tertiary flow F3 is a flow which passes annularly through the tertiary flow vein 16.

[0086] In this configuration, the tertiary flow F3 extends essentially in the axial direction and in a position radially between the primary flow F1 and the secondary flow F2.

[0087] The tertiary flow F3 extends in the tertiary flow vein 16 from the radially internal air flow F' downstream of the moving wheel 12 and to the secondary flow F2 after passing through the heat exchanger 18.

[0088] The turbomachine 2 further comprises a stator (not shown) arranged upstream of the heat exchanger 18 at the level of the tertiary flow vein 16. Advantageously, the stator makes it possible to straighten the tertiary flow F3 before the latter passes through the heat exchanger 18 in order to minimize the aerodynamic disturbances of the tertiary flow F3 which can be caused by the moving wheel 12, this makes it possible to optimize the heat exchange between the air and the oil.

[0089] The stator corresponds to a row of stator vanes arranged in the tertiary flow stream downstream of the separation nozzle 14. Alternatively, the stator may be arranged upstream of the separation nozzle 14 and downstream of the moving wheel 12.

[0090] The heat exchanger 18 extends radially and axially in an upstream section 20 of the tertiary flow vein 16, having a longitudinal section diverging in the direction of flow of the tertiary flow F3.

[0091] The heat exchanger 18 is arranged axially between the low pressure compressor 17 and the high pressure compressor 15 in the example shown.

[0092] A discharge channel of the “VBV” type 19 (Variable Bleed Valve) having an outlet passing through an internal wall of the tertiary flow vein 16 and arranged axially downstream of the heat exchanger 18, the “VBV” channel makes it possible to ensure a discharge function by returning a portion of the primary flow F1 to the tertiary flow F3, this makes it possible to evacuate, for example, any ice particles from the primary flow F1 to avoid clogging of the high-pressure compressor 15, in particular when the flow rate of the primary flow F1 becomes too low.

[0093] Advantageously, the arrangement of the outlet of the “VBV” channel 19 downstream of the heat exchanger 18 makes it possible to protect the latter from any possible risk of blockage.

[0094] The heat exchanger 18 can extend continuously over 360° in the upstream section 20 of the vein 16 around the longitudinal axis 8 of the turbomachine 2. Preferably, the heat exchanger 18 extends discontinuously over 360° around the longitudinal axis 8 by being subdivided into several angular segments and each can provide a heat exchange function between the air and the oil which can be different from one segment to another.

[0095] The heat exchanger 18 is of the “ACOC” type, an acronym for the English expression “Air Cooled Oil Cooler”, comprising oil passages which extend in the tertiary flow stream, said oil passages extending particularly in a radial and axial direction between an upper wall and a lower wall of said heat exchanger 18. Advantageously, the “ACOC” heat exchanger 18 allows heat exchange between the air and the oil, preferably cooling of the oil by the air. Indeed, the temperature of the oil can reach an operating temperature of up to 180°C and a flow rate reaching 30,000 l / h. In this regard, the exchanger 18 can ensure the cooling of the oil used in several components of the aircraft, in particular, an engine, a gearbox, an engine generator and any electronic component requiring cooling.

[0096] A single heat exchanger 18 can combine the cooling of several functions or oil circuits of the turbomachine, and this depending on different parameters linked to the need for cooling the oil, i.e. inlet temperatures, flow rates, required outlet temperature or air conditions, the different circuits can be put in thermal contact or insulated. The exchanger 18 and in particular its oil passages can withstand a low oil temperature of up to -54°C.

[0097] Figures 2 and 3 have been described in the above.

[0098] Figure 4 illustrates an embodiment of the invention and shows very schematically an annular flow vein V of a gas flow in a turbomachine.

[0099] This turbomachine can be of the triple flow type as described above and illustrated in figure 1. It is then understood that the vein V is a tertiary flow vein 16 of a tertiary flow F3.

[0100] Alternatively, the turbomachine could be of the double-flow type, and vein V could be a secondary flow vein of a secondary flow.

[0101] It is therefore understood that the type of turbomachine is not limiting within the scope of the present invention.

[0102] The turbomachine according to the invention comprises, in addition to the vein V, an engine M and at least one piece of equipment E. The engine M is preferably the gas generator of the turbomachine and therefore comprises at least one compressor, an annular combustion chamber, and at least one turbine.

[0103] Said at least one piece of equipment E is, for example, an electrical machine or several electrical machines.

[0104] The vein V has an annular shape around an axis 8 which is generally the longitudinal axis of the turbomachine. The vein comprises two coaxial annular walls, respectively internal 101 and external 102, which are connected to each other by radial arms 100, four in number in the example shown.

[0105] The arms 100 divide the vein V into angular sectors, some of which are empty to allow the gas flow to pass without obstacle, and others are occupied by exchangers 210, 220, 230.

[0106] According to the invention, the turbomachine is equipped with two exchangers 210, 220 of the ACOC type and one exchanger of the OCOC type.

[0107] The first exchanger 210 is capable of being swept by the flow of gas flowing in the vein V and comprises a first oil circuit C1 connected to a system S1 for cooling the engine M.

[0108] The second exchanger 220 is capable of being swept by the gas flow and comprises a second oil circuit C2 connected to a system S2 for cooling the equipment(s) E.

[0109] The third exchanger 230 is not necessarily located in the vein because it is not intended to be swept by the gas flow. This third exchanger 230 comprises two independent circuits, called third and fourth oil circuits C3, C4.

[0110] In the example shown, each exchanger 210, 220 is located in the vein V and occupies a part of the passage section of this vein V. For example, the first exchanger 210 can occupy the entire radial extent or dimension of the vein, or only a part of this radial extent, and extend in the circumferential direction over a part, and for example 2 / 3, of an angular sector of the vein V between two arms 100. The second exchanger 220 can occupy the entire radial extent or dimension of the vein, or only a part of this radial extent, and extend in the circumferential direction over a part or the entirety of another angular sector of the vein V between two arms 100.

[0111] Each exchanger 210, 220 has, for example, an angular extent of between approximately 30 and 90° around the axis 8.

[0112] Preferably, the third exchanger 230 has a cooling capacity lower than that of the first heat exchanger 210, and lower than that of the second heat exchanger 220.

[0113] Figures 5 and 6 show the fluid connections between the exchangers 210, 220, 230 and in particular between the circuits C1, C2, C3, C4 of these exchangers.

[0114] The turbomachine according to the invention comprises valves 310, 320 for connecting the first, second and third circuits C1, C2, C3, which are capable of adopting two distinct configurations illustrated in Figures 5 and 6 respectively. According to a first configuration illustrated in Figure 5, the third circuit C3 is isolated from the first circuit C1, and the fourth circuit C4 is isolated from the second circuit C2.

[0115] According to a second configuration illustrated in Figure 6, the third circuit C3 is connected in series to the first circuit C1, and the fourth circuit C4 is connected in series to the second circuit C2.

[0116] Advantageously, in the second configuration, an output C3s of the third circuit C3 is connected to an input C1e of the first circuit C1, and an output C4s of the fourth circuit C4 is connected to an input C2e of the second circuit C2. In other words, oil circulates in the third exchanger 230 before supplying the first exchanger 210, and oil circulates in the third exchanger 230 before supplying the second exchanger 220.

[0117] In the example shown, the first circuit C1 comprises a first inlet valve 410 connected to an inlet C1 e of the first circuit C1 . The first circuit C1 could also comprise a first outlet valve (not shown) connected to the outlet C1 s of the first circuit C1 . The second circuit C2 comprises a second inlet valve 420 connected to an inlet C2e of the second circuit C2. The second circuit C2 could also comprise a second outlet valve connected to the outlet C2s of the second circuit C2.

[0118] The connecting valves 310, 320 are respectively mounted upstream of the inlets C3e, C4e of the third and fourth circuits E3, C4.

[0119] The connecting valve 310 can be mounted between a connecting point P1 of the inlet C1 e of the first circuit C1 , located upstream of the first inlet valve 410, and the inlet C3 e of the circuit C3. The outlet C3 s of the circuit can be connected to another connecting point P2 of the inlet C1 e of the first circuit C1 , located downstream of the first inlet valve 410.

[0120] The connecting valve 320 can be mounted between a connecting point P3 of the inlet C2e of the second circuit C2, located upstream of the second inlet valve 420, and the inlet C4e of the circuit C4. The outlet C4s of the circuit C4 can be connected to another connecting point P4 of the inlet C2e of the second circuit C2, located downstream of the second inlet valve 420.

[0121] The present invention also relates to a cooling method in the turbomachine according to the invention, this method comprising two steps corresponding to the two aforementioned configurations.

[0122] According to a first step corresponding to the first configuration, the cooling system S1 of the engine M is supplied by oil leaving the first exchanger 210 after having passed through only this first exchanger 210, and the cooling system S2 of the equipment(s) E is supplied by oil leaving the second exchanger 220 after having passed through only this second exchanger 220. This step and this configuration are illustrated in FIG. 5 in which the arrows in solid lines are seen which show the circulation of the oil in each of the exchangers 210, 220. The exchanger 230 is isolated and is not used in this step / configuration. This first step / configuration can be adopted by default in most of the operating phases of the turbomachine and of the flight of the aircraft.

[0123] According to a second step corresponding to the second configuration, the cooling system S1 of the engine M is supplied by oil which has passed through the third exchanger 230, then the first exchanger 210, and which leaves the first exchanger 210, and the cooling system S2 of the equipment(s) E is supplied by oil which has passed through the third exchanger 230, then the second exchanger 220, and which leaves the second exchanger 220. This step and this configuration are illustrated in Figure 6 in which we see the arrows in continuous lines which show the circulation of the oil in each of the exchangers 210, 220, 230. The inlet C1e of the first exchanger 210 is connected to the outlet C3s of the third circuit C3 of the third exchanger 230 so the oil which leaves this circuit C3 feeds the first exchanger 210 in order to be further cooled.The input C2e of the second exchanger 220 is connected to the output C4s of the fourth circuit C4 of the third exchanger 230, so the oil leaving this circuit C4 feeds the second exchanger 220.

[0124] This second step / configuration can be adopted when the aircraft is on the ground and the ambient outside temperature is above a predetermined threshold.

[0125] The first exchanger 210 can be supplied with oil at a maximum temperature of 180°C.

[0126] The second exchanger 220 can be supplied with oil at a maximum temperature of 100°C.

Claims

CLAIMS 1. Turbomachine (2) for an aircraft, this turbomachine comprising: - an annular vein (V) for the flow of a gas flow (F3), - an engine (M), - at least one piece of equipment (E), - a first heat exchanger (210) of the ACOC type, which is capable of being swept by said gas flow (F3) and which comprises a first oil circuit (C1) connected to an engine cooling system (S1), - a second heat exchanger (220) of the ACOC type, which is capable of being swept by said gas flow (F3) and which comprises a second oil circuit (C2) connected to a system (S2) for cooling said at least one item of equipment (E), characterized in that it further comprises: - a third heat exchanger (230) of the OCOC type, which comprises third and fourth oil circuits (C3, C4) capable of exchanging calories between them, and - valves (310, 320) for connecting the first circuit (C1) to the third circuit (C3), and from the second circuit (C2) to the fourth circuit (C4), the connecting valves (310, 320) being capable of adopting two configurations: - a first configuration in which the third circuit (C3) is isolated from the first circuit (C1), and the fourth circuit (C4) is isolated from the second circuit (C2), and - a second configuration in which the third circuit (C3) is connected in series to the first circuit (C1) for the supply of oil to the first circuit by the third circuit, and the fourth circuit (C4) is connected in series to the second circuit (C2) for the supply of oil to the second circuit by the fourth circuit.

2. Turbomachine (2) according to claim 1, in which, in the second configuration, an output (C3s) of the third circuit (C3) is connected to an input (C1 e) of the first circuit (C1), and an output (C4s) of the fourth circuit (C4) is connected to an input (C2e) of the second circuit (C2).

3. Turbomachine (2) according to claim 1 or 2, in which the first circuit (C1) comprises a first inlet valve (410) connected to an inlet (C1e) of the first circuit (C1), and the second circuit (C2) comprises a second inlet valve (420) connected to an inlet (C2e) of the second circuit (C2).

4. Turbomachine (2) according to the preceding claim, in which the connecting valves (310, 320) comprise: - a first connecting valve (310) mounted between a connecting point (P1) of the inlet (C1 e) of the first circuit (C1 ), located upstream of the first inlet valve (410), and an inlet (C3e) of the third circuit (C3), and - a second connecting valve (320) mounted between a connecting point (P3) of the inlet (C2e) of the second circuit (C2), located upstream of the second inlet valve (420), and an inlet (D4e) of the fourth circuit (C4).

5. Turbomachine (2) according to the preceding claim, in which the third circuit (C3) comprises an outlet (C3s) connected to another point (P2) of connection of the inlet (C1 e) of the first circuit (C1), located downstream of the first inlet valve (410), and the fourth circuit (C4) comprises an outlet (C4s) connected to another point (P4) of connection of the inlet (C2e) of the second circuit (C2), located downstream of the second inlet valve (420).

6. Turbomachine (2) according to one of the preceding claims, in which the third heat exchanger (330) is located outside said vein (V).

7. Turbomachine (2) according to one of the preceding claims, in which said vein (V) is a secondary flow vein of a secondary gas flow, the turbomachine further comprising a primary flow vein of a primary gas flow inside said engine (M).

8. Turbomachine (2) according to one of claims 1 to 5, in which said vein (V) is a tertiary flow vein of a tertiary gas flow (F3), the turbomachine further comprising two annular veins, respectively primary and secondary, primary and secondary gas flow (F1, F2).

9. Turbomachine (2) according to the preceding claim, in which the tertiary vein is interposed radially between the primary and secondary veins.

10. Turbomachine (2) according to one of the preceding claims, in which the third heat exchanger (230) has a cooling capacity lower than that of the first heat exchanger (210), and lower than that of the second heat exchanger (220).

11. Turbomachine (2) according to one of the preceding claims, in which said at least one piece of equipment (E) comprises electrical machines.

12. Cooling method in a turbomachine (2) according to one of the preceding claims, in which it comprises two steps corresponding to the two aforementioned configurations: - a first step corresponding to the first configuration, in which the system (C1) for cooling the engine (M) is supplied by oil leaving the first exchanger (210), and the system (S2) for cooling said at least one piece of equipment (E) is supplied by oil leaving the second exchanger (220), and - a second step corresponding to the second configuration, in which the system (S2) for cooling the engine (M) is supplied by oil passing through the third exchanger (230) then through the first exchanger (210), and the system (S2) for cooling said at least one piece of equipment (E) is supplied by oil passing through the third exchanger (230) then through the second exchanger (220).

13. The method of claim 12, wherein: - the second stage occurs when the aircraft is on the ground and the ambient outside temperature is above a predetermined threshold, - the first step occurs by default in other cases.

14. Method according to claim 12 or 13, wherein the first heat exchanger (210) is supplied with oil at a maximum temperature of 180°C, and the second heat exchanger (220) is supplied with oil at a maximum temperature of 100°C.

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