System and method for heating fuel to power an aircraft turbomachine

Abstract

A fuel heating system (1) for supplying an aircraft turbomachine (M), the heating system (1) comprising a fuel circuit (2); at least one first heat exchanger (41) configured to be mounted in an internal flow (V) of the turbomachine (M), the first heat exchanger (41) being configured to heat a phase-transition heat transfer fluid (F) from calories from an exhaust flow (FE); at least one second heat exchanger (42) configured to be mounted externally to the internal flow (V) of the turbomachine (M), the second heat exchanger (42) being configured to heat the fuel flow (Q) from calories from the heat transfer fluid (F), and at least one heat pipe device (5) mounted between the first heat exchanger (41) and the second heat exchanger (42), the heat transfer fluid (F) circulating in the heat pipe device (5). Figure from the summary: Figure 2

Description

Title of the invention: System and method for heating fuel to power an aircraft turbomachine. Technical field

[0001] The present invention relates to the field of aircraft comprising turbomachines powered by fuel stored in a cryogenic tank.

[0002] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies to both new types of aircraft and those already in service, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively working for several years now to contribute to the fight against climate change.

[0003] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into account the factors impacting all phases of design and development in order to obtain aeronautical components and products that are less energy-intensive, more environmentally friendly, and whose integration and use in civil aviation have moderate environmental consequences, with the aim of improving the energy efficiency of aircraft.

[0004] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes that minimize greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0005] This sustained research and development work focuses on new generations of aircraft engines, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and aviation biofuels, which are essential complements to technological progress.

[0006] To this end, the invention is the result of technological research aimed at significantly improving aircraft performance and, in this sense, contributes to reducing the environmental impact of aircraft. For this purpose, the invention relates to the field of aircraft turbomachinery powered by fuel stored in a cryogenic tank.

[0007] It is known to store fuel, particularly hydrogen, in liquid form to limit the size and mass of aircraft tanks. For example, fuel is stored at a temperature of approximately -253 to -251°C (20 to 22 Kelvins) in a cryogenic tank on the aircraft.

[0008] In order to be injected into the combustion chamber of a turbomachine, the fuel must be conditioned, that is to say, pressurized and heated, to allow for optimal combustion. Conditioning is necessary, for example, to reduce the risk of icing / solidification of the water vapor contained in the air circulating in the turbomachine, in particular, at the turbomachine's fuel injectors.

[0009] With reference to [Fig.1], a prior art SAA heating system is shown comprising a fuel circuit 100 connected at the inlet to a cryogenic tank R and at the outlet to the combustion chamber of a turbomachine M. A fuel flow Q circulating from upstream to downstream in the fuel circuit 100 passes successively through a mechanical pump 101 and a heating module 102.

[0010] The mechanical pump 101 is configured to circulate the fuel flow Q in the fuel circuit 100. The heating module 102 is configured to supply calories to the fuel flow Q in order to warm it so that it can be injected into the turbomachine M.

[0011] In practice, the fuel heating stage requires extracting heat from heat sources within the aircraft. For example, heat generated by the turbomachine can be used (heat from the lubricating oil, heat from the turbine outlet, heat from the nozzle, etc.). Heat from the aircraft can also be used (air from the cabin, heat from electrical or electronic systems, etc.).

[0012] In the prior art, for example, patent application FR2005628A1 discloses an architecture, also shown in [Fig. 1], in which a heat transfer fluid F passes through a heat exchanger EX in which it extracts heat from the available hot sources C on board the aircraft (for example, heat from the turbomachine's lubricating oil M, heat from the turbine outlet, or heat from the nozzle). The heat transfer fluid F is then conveyed via a circulation loop BC to the heat module 102 to heat the fuel Q. This circulation loop BC for the heat transfer fluid F prevents the risk of contamination between the fuel and an oxidant in a heat exchanger, for example. A mechanical recirculation pump PR keeps the heat transfer fluid F moving within the circulation loop BC.

[0013] However, in such an architecture, the heat transfer fluid must be able to circulate at different speeds and in particular for maximum engine operation, that is to say This means a high fuel flow rate in the fuel circuit and therefore a high flow rate of the heat transfer fluid in the circulation loop. A high flow rate necessitates the use of a large-capacity recirculation pump, which means a heavy and bulky pump. The heat transfer fluid must also be transported via specialized piping, which adds significant mass and size.

[0014] In summary, the circulation loop of a heat transfer fluid is complex to implement and has significant mass and size, which is undesirable in an aeronautical context that aims to limit aircraft mass to reduce their energy consumption and thus their environmental impact. Furthermore, the use of an oversized mechanical pump results in significant electrical consumption, which is undesirable for the same reasons.

[0015] Incidentally, a heat transfer device, known as a "heat pipe," is also known. This device operates in a closed loop and allows a fluid to be heated from a phase-transition heat transfer fluid. More specifically, in a heat pipe, the heat transfer fluid absorbs heat from a hot zone and changes from a liquid to a gaseous state, then flows to a cold zone where it releases heat to the fluid to be heated by returning to a liquid state. A heat pipe has the advantage of eliminating the need for a mechanical pump and has a limited mass. However, a heat pipe is commonly used to heat an air stream or cool engine oil and cannot be easily integrated into an aircraft fuel system to heat a cryogenic fuel stream.

[0016] The invention thus aims to eliminate at least some of these drawbacks by proposing a simple and efficient cryogenic fuel heating system that eliminates the need for a heat transfer fluid circulation loop. PRESENTATION OF THE INVENTION

[0017] The invention relates to a fuel heating system for supplying an aircraft turbomachine, the fuel being drawn from a cryogenic tank, the turbomachine comprising a combustion chamber and an internal circulation channel for an exhaust flow from the combustion chamber, the heating system comprising: • a fuel circuit configured to be connected at the inlet to the cryogenic tank and at the outlet to the turbomachine, with a fuel flow circulating from upstream to downstream in the fuel circuit, • at least one first heat exchanger configured to be mounted in the inner core of the turbomachine, the first heat exchanger being configured to heat a phase-transition heat transfer fluid using calories from the exhaust stream, • at least one second heat exchanger configured to be mounted externally to the inner core of the turbomachine, the second heat exchanger being configured to heat the fuel stream using calories from the heat transfer fluid, and • at least one heat pipe device mounted between the first heat exchanger and the second heat exchanger, the heat transfer fluid circulating in the heat pipe device and being configured to pass into a gaseous state in the first heat exchanger and into a liquid state in the second heat exchanger.

[0018] The heating system according to the invention efficiently heats the fuel stream from the cryogenic tank using the exhaust stream from the turbomachine's combustion chamber, which is at high temperatures. Advantageously, the oxygen-containing exhaust stream and the fuel stream do not come into contact, thus preventing any risk of fuel combustion outside the turbomachine's combustion chamber. The heat pipe device transfers heat from the exhaust stream to the fuel stream indirectly.

[0019] The heat pipe device can advantageously be mounted as close as possible to the heat exchangers, making it possible to do away with a heat transfer fluid circulation loop according to the prior art, which makes it possible to do away with the addition of a mechanical pump which is bulky and heavy and which consumes a lot of energy.

[0020] By heat pipe device, we mean a hermetically sealed heat conducting device in which a heat transfer fluid circulates in liquid-vapor equilibrium and is configured to transfer heat by phase transition, i.e. operating according to an evaporation-condensation principle.

[0021] In summary, the heating system according to the invention makes it possible to avoid the use of expensive, bulky and energy-consuming elements.

[0022] In one embodiment, the heating system comprises a plurality of heat pipe devices mounted between the first heat exchanger and the second heat exchanger, enabling the fuel flow to be heated efficiently by extracting more calories from the exhaust flow.

[0023] Preferably, the number of heat pipe devices is between two and one thousand heat pipe devices.

[0024] Preferably, the turbomachine comprising an exhaust nozzle downstream of the combustion chamber, the heating system is mounted at the level of a vein inside the exhaust nozzle, so as to mount the first heat exchanger directly in the inner vein of the exhaust nozzle and efficiently recover heat from the exhaust flow.

[0025] In one embodiment, the turbomachine comprising a primary airflow circulation channel, the first heat exchanger is mounted in the primary channel of the turbomachine.

[0026] Preferably, the phase transition heat transfer fluid is chosen from pentane, methanol, toluene, acetone, water, glycol water.

[0027] In one embodiment, the turbomachine includes at least one fixed guide vane mounted radially in the inner runner, and the first heat exchanger is integrated into the fixed guide vane of the turbomachine. This embodiment eliminates the need for an additional heat exchanger, thereby reducing the mass of the heating system.

[0028] Preferably, the guide vane comprising a hollow envelope defining an inner wall, the heat pipe device is configured to be positioned in contact with the inner wall, so as to efficiently recover heat from the exhaust flow circulating in the inner vein.

[0029] In one embodiment, the fixed guide vane comprising a hollow envelope defining an inner wall, the first exchanger comprises a plurality of conductive elements, mounted on the inner wall of the fixed guide vane, so as to increase the heat transfer from the exhaust flow to the heat transfer fluid by convection.

[0030] In one embodiment, the inner wall of the fixed guide vane is a porous wall so as to improve heat transfer with the exhaust flow.

[0031] In one embodiment, the second heat exchanger is mounted in a protection module comprising an inert gas, the protection module being configured to detect a leak in the second heat exchanger. Thus, any fuel leakage is prevented.

[0032] Preferably, the heat exchangers each include secondary heat exchange surfaces, for example grooves or fins, to promote heat exchange between the exhaust flow, respectively the fuel flow, and the heat transfer fluid.

[0033] Preferably, the second heat exchanger is positioned vertically above the first heat exchanger, so as to allow the heat transfer fluid to return in a liquid state by gravity into the heat pipe device.

[0034] In one embodiment, the heating system comprises: • at least two initial heat exchangers, each configured to be mounted in the inner core of the turbomachine, and • at least two heat pipe devices connecting the first two heat exchangers to the second heat exchanger respectively.

[0035] Such an embodiment allows the fuel flow to be heated gradually and more efficiently. The second heat exchanger is advantageously shared, which reduces the overall size.

[0036] The invention also relates to an aircraft turbomachine assembly comprising a combustion chamber and an internal circulation channel for an exhaust flow from the combustion chamber and a heating system as described above, the first heat exchanger being mounted in the internal channel of the turbomachine and the second heat exchanger being mounted externally to the turbomachine.

[0037] The invention also relates to an aircraft comprising a cryogenic tank in which fuel is stored and an assembly of a turbomachine and a fuel heating system as described above.

[0038] Finally, the invention relates to a method for heating fuel for supplying an aircraft turbomachine, the fuel being drawn from a cryogenic tank, the turbomachine comprising a combustion chamber and an internal circulation channel for an exhaust flow from the combustion chamber, the heating method being carried out by means of the heating system as described above and comprising the steps of: • Heat the heat transfer fluid in its liquid state in the first heat exchanger using heat from the exhaust flow circulating in the inner core of the turbomachine so that the heat transfer fluid is in a gaseous state, • To convey, within the heat pipe system, the heat transfer fluid in a gaseous state to the second heat exchanger, • Heat the fuel stream, in the second heat exchanger, using calories from the heat transfer fluid in the gaseous state so that the heat transfer fluid is in the liquid state. PRESENTATION OF THE FIGURES

[0039] The invention will be better understood upon reading the following description, given by way of example, and referring to the following figures, given by way of non-limiting examples, in which identical references are given to similar objects.

[0040] Fig. 1 is a schematic representation of a heating system according to the prior art.

[0041] The [Fig.2] is a schematic representation of a heating system according to one embodiment of the invention.

[0042] Fig. 3 is a schematic representation of the heating system of Fig. 2 mounted in an aircraft turbomachine according to a first embodiment of the invention.

[0043] Fig. 4 is a schematic representation of a heat pipe device of the heating system of Fig. 2.

[0044] Fig. 5 is a schematic representation of the heating system of Fig. 2 mounted in an aircraft turbomachine according to a second embodiment of the invention.

[0045] Fig. 6 is a close-up view of the heating system and a turbine blade of the turbomachine of Fig. 5 along a section plane A:A.

[0046] Fig. 7 is a close-up view of the heating system and a turbine blade of the turbomachine of Fig. 5 according to a second section plane B: B.

[0047] Fig. 8 is a schematic representation of the heating system of Fig. 2 according to a third embodiment of the invention.

[0048] Fig. 9 is a schematic representation of the heating system of Fig. 2 mounted in an aircraft turbomachine according to a third embodiment of the invention.

[0049] The [Fig. 10] is a diagram of the steps of a fuel heating process according to an embodiment of the invention.

[0050] It should be noted that the figures set out the invention in detail to implement the invention, said figures being of course able to serve to better define the invention where appropriate. DETAILED DESCRIPTION OF THE INVENTION

[0051] With reference to [Fig. 2], a heating system 1 for a fuel Q is shown for supplying an aircraft turbomachine M, the fuel Q being drawn from a cryogenic tank R. The turbomachine M is configured to provide propulsion for the aircraft, in particular, by driving at least one propulsion unit (not shown in [Fig. 2]). In this example, the fuel Q is liquid hydrogen, but the invention applies to other types of fuel, for example, liquid methane or liquefied natural gas.

[0052] In this example, the fuel Q in the cryogenic tank R is stored at a temperature of approximately -253 to -251°C (20 to 22 Kelvins). At this temperature, the fuel Q flow is liquid.

[0053] With reference to [Fig. 2], the turbomachine M extends along a longitudinal axis X and comprises a combustion chamber CC and a main internal flow V for the circulation of an exhaust stream FE from the combustion chamber CC. In this example, the turbomachine M includes an exhaust nozzle TE ([Fig. 3]), downstream of the combustion chamber CC, which forms the internal flow V for the circulation of the exhaust stream FE.

[0054] More specifically, as is known, the turbomachine M comprises a primary runner delimited by a shroud CA and supplied by an upstream fan (not shown), also referred to as the "rotor," mounted to rotate about the X-axis. The rotor accelerates an upstream to downstream airflow circulating in the primary runner. After passing through a low-pressure compressor and a high-pressure compressor for acceleration, the airflow is then introduced into the combustion chamber CC where it is brought into contact with the fuel Q for combustion. The exhaust flow FE is generated at the outlet of the combustion chamber CC into the primary runner and passes through the exhaust nozzle TE. The turbomachine M also comprises a secondary runner (not shown) which extends circumferentially outward from the primary runner. An inter-runner space is defined between the primary and secondary runners.

[0055] As described previously, in order to be introduced into the combustion chamber CC, the fuel Q, initially stored at cryogenic temperatures, must be heated.

[0056] The invention thus relates to a heating system 1 for the fuel Q. In this example, the heating system 1 is mounted at the exhaust nozzle TE of the turbomachine M. Preferably, as shown in [Fig.2], the heating system 1 extends vertically along an axis Z, substantially radial to the flow axis X of the turbomachine M.

[0057] According to the invention, and again with reference to [Fig.2], the heating system 1 comprises a fuel circuit 2 configured to be connected at the inlet to the cryogenic tank R and at the outlet to the turbomachine M, a first heat exchanger 41, a second heat exchanger 42 and a heat pipe device 5, mounted between the first heat exchanger 41 and the second heat exchanger 42. A flow of fuel Q circulates, by means of a mechanical pump 3, from upstream to downstream in the fuel circuit 2 from the cryogenic tank R to the combustion chamber CC of the turbomachine M.

[0058] In this first embodiment, the heating system 1 comprises a single first heat exchanger 41 and a single second heat exchanger 42; however, it is understood that the number of first heat exchangers 41 and / or second heat exchangers 42 may be different. In particular, the system heating 1 may include a plurality of first heat exchangers 41 associated with the same second heat exchanger 42, as will be described in more detail later.

[0059] According to the invention, as shown in [Fig. 2] and [Fig. 3], the first heat exchanger 41 is mounted in the inner flow V of the turbomachine M. Preferably, the first heat exchanger 41 is mounted inside the inner flow V formed by the exhaust nozzle TE and is configured to recover the exhaust flow FE from the combustion chamber CC. In a preferred embodiment, the first heat exchanger 41 is mounted inside the primary flow of the turbomachine M, as described above, so as to recover a primary exhaust flow FE circulating in the primary flow.

[0060] According to the invention, the first heat exchanger 41 is configured to heat a phase-transition heat transfer fluid F from calories from the exhaust stream FE, as will be described in more detail later.

[0061] In this example, the exhaust flow FE circulates in the inner vein V and has a temperature between 300 and 600 °C at the inlet of the first heat exchanger 41.

[0062] In this example, the first heat exchanger 41 is a tubular or plate heat exchanger. The heat exchange surfaces in the first heat exchanger 41 comprise, in one embodiment, additional exchange elements, such as fins or grooves for example, in order to improve the overall aerothermal performance of the heat exchanger.

[0063] According to the invention, and with reference to Figures 2 and 3, the second heat exchanger 42 is mounted externally to the main internal flow V of the turbomachine M. More specifically, the second heat exchanger 42 is mounted, preferably, externally to the fairing CA delimiting the exhaust nozzle TE from the main internal flow V of the turbomachine M. In this example, the second heat exchanger 42 is mounted externally to the primary flow, for example in the inter-flow space or in the secondary flow of the turbomachine M.

[0064] Preferably, the second heat exchanger 42 is mounted on the turbomachine M radially opposite the first heat exchanger 41. In one embodiment (in the case of a gravity heat pipe device 5 which will be described in more detail later), as shown in Figures 2 and 3, the second heat exchanger 42 is positioned vertically along the Z axis above the first heat exchanger 41.

[0065] According to the invention, the second heat exchanger 42 is configured to heat the fuel stream Q from calories from the phase-transition heat transfer fluid F, as will be described in more detail later.

[0066] In this example, the second heat exchanger 42 is a tubular or plate heat exchanger. The heat exchange surfaces in the second heat exchanger 42 comprise, in one embodiment, additional exchange elements, such as fins for example, in order to improve the overall aerothermal performance of the heat exchanger.

[0067] With reference to figures 2 and 3, the heat pipe device 5 is mounted between the first heat exchanger 41 and the second heat exchanger 42.

[0068] In a preferred embodiment, as shown in Figures 2 and 3, the heating system 1 comprises a plurality of heat pipe devices 5 mounted between the first heat exchanger 41 and the second heat exchanger 42. In this example, the heating system 1 comprises four heat pipe devices 5; however, it is understood that the number of heat pipe devices 5 could be different. In particular, the heating system 1 preferably comprises a number of heat pipe devices 5 ranging from one to one thousand.

[0069] In this example in which the first heat exchanger 41 is mounted internally to the exhaust nozzle TE and the second heat exchanger 42 is mounted externally to the exhaust nozzle TE, each heat pipe device 5 preferably extends into the fairing CA delimiting the exhaust nozzle TE.

[0070] By definition, a heat pipe device 5 is a hermetically sealed heat-conducting device in which a heat transfer fluid F circulates in liquid-vapor equilibrium. The heat pipe device 5 operates in a closed cycle according to an evaporation-condensation principle, that is, it is configured to transfer heat by phase transitions. According to the invention, as shown in [Fig. 3], the heat transfer fluid F is configured to circulate in a loop and to pass into a gaseous state in the first heat exchanger 41 and into a liquid state in the second heat exchanger 42.

[0071] In this example, the heat transfer fluid F is configured to be at a temperature between 25 and 300 °C in both the liquid and gaseous states. The density ratio between the heat transfer fluid F in the gaseous state and the heat transfer fluid F in the liquid state is, in this example, between 10 and 1000.

[0072] In a preferred embodiment, shown in [Fig. 4], each heat pipe device 5 has a tubular shape and comprises an outer tube 51 and an inner tube 52, mounted internally to the outer tube 51. The outer tube 51 and the inner tube 52 are preferably concentric. In this example, the heat transfer fluid F is configured to flow in the outer tube 51 in a liquid state and in the inner tube 52 in a gaseous state.

[0073] In this embodiment, the circulation of the heat transfer fluid F is achieved by capillary action. In such a heat pipe device 5, in contact with the first heat exchanger In the heat exchanger 41, through which the exhaust flow FE circulates, the heat transfer fluid F is configured to vaporize, inducing a slight overpressure that causes the vapor to move towards the second heat exchanger 42. In the second heat exchanger 42, the heat transfer fluid is configured to transfer its heat to the fuel flow Q; the vapor condenses and returns to the liquid phase. The condensed heat transfer fluid F circulates in a capillary network and returns to the first heat exchanger 41 under the effect of capillary forces.

[0074] Alternatively, in the example in which the second heat exchanger 42 is mounted vertically above the first heat exchanger 41, the circulation of the heat transfer fluid F is configured to be carried out by gravity in the heat pipe device 5. More specifically, the heat transfer fluid F is configured to pass into the gaseous state in the first heat exchanger 41, to rise by evaporation, to liquefy upon contact with the fuel flow Q by transferring its heat to it, to return to the liquid state and descend by gravity.

[0075] Such a heat pipe device 5 is known to the person skilled in the art and its operation will not be described in further detail in this document.

[0076] In a preferred embodiment, the phase-transition heat transfer fluid F is selected from pentane, methanol, toluene, acetone, water, glycol water.

[0077] Thanks to such a heat pipe device 5, the heating system 1 is free of mechanical pump unlike a heat transfer fluid loop according to the prior art.

[0078] In one embodiment, with reference to [Fig. 5], the turbomachine M comprises a fixed guide vane AG mounted radially in the inner duct V. The term "fixed guide vane" refers to a non-moving vane that may have structural and / or auxiliary functions. More specifically, the fixed guide vane AG extends radially between a central hub MC and the shroud CA of the turbomachine M, preferably vertically. The fixed guide vane AG is configured to straighten the exhaust flow FE or to support the central hub MC of the turbomachine M.

[0079] In this example, the fixed guide vane AG comprises a hollow casing ENV defining an inner wall PL

[0080] In this embodiment, the first heat exchanger 41 is integrated into the fixed guide vane AG of the turbomachine M, as shown in [Fig. 5]. In other words, the hollow casing ENV fulfills the function of the first heat exchanger 41 and allows heat exchange between the heat transfer fluid F circulating in the heat pipe device 5 and the exhaust flow FE circulating in the inner channel V. Each heat pipe device 5 thus extends inside the hollow casing ENV of the fixed guide vane AG to allow the fluid to evaporate heat transfer fluid F inside the heat pipe devices 5 and the conveyance of the heat transfer fluid F charged with calories to the second heat exchanger 42.

[0081] In this embodiment, with reference to [Fig.6] representing a longitudinal cross-sectional view (along a plane A:A) of the guide vane AG, each heat pipe device 5 is configured, preferably, to be in contact with the inner wall PI of the fixed guide vane AG, so as to allow convective heat transfer from the exhaust stream FE to the heat transfer fluid F circulating in each heat pipe device 5.

[0082] Preferably, with reference to [Fig.7] (representing a cross-sectional view of the guide vane AG along a plane B:B), the first heat exchanger 41 comprises a plurality of conductive elements 7, mounted on the inner wall PI of the fixed guide vane AG to improve heat exchange in the first heat exchanger 41. By way of example, each conductive element 7 is in the form of a continuous component made of a heat-conducting material (for example, copper, aluminum, ...) in thermal contact both with the inner wall PI of the fixed guide vane AG and with the heat pipe devices 5.

[0083] In a preferred embodiment, with reference to [Fig.8], the inner wall PI of the fixed guide vane AG is a porous wall PP so as to improve heat transfer between the exhaust flow FE and the heat pipe devices 5.

[0084] With reference to [Fig.3], in one embodiment, the heating system 1 includes a protection module 6 in which the second heat exchanger 42 is mounted. The protection module 6 includes an inert gas and is configured to detect a leak in the second heat exchanger 42.

[0085] A heating system 1 has been presented comprising a single first heat exchanger 41 and a single second heat exchanger 42. Alternatively, as shown in [Fig. 9], the heating system 1 may comprise a plurality of first heat exchangers 41, each mounted in the inner channel V of the turbomachine M. In this example, the heating system 1 comprises several plurality of heat pipe devices 5, each plurality of heat pipe devices 5 being mounted between one of the first heat exchangers 41 and the same second heat exchanger 42. The first heat exchangers 41 thus extend successively from upstream to downstream in the inner channel V and allow the fuel flow Q to be heated progressively in the second heat exchanger 42, while reaching higher temperatures.

[0086] It goes without saying that the heating system 1 could just as easily comprise a plurality of second heat exchangers 42, each mounted externally to the inner stream V of the turbomachine M. Each second heat exchanger 42 could exchange calories with one or more first heat exchangers 41 via different heat pipe devices 5.

[0087] A method for heating a fuel Q for supplying an aircraft turbomachine M by means of the heating system 1 as described previously, with reference to Figures 2 and 10, will now be described. The fuel Q is drawn from a cryogenic tank R. The turbomachine M comprises a combustion chamber CC and an internal channel V through which an exhaust flow FE from the combustion chamber CC circulates. In this example, the heating system 1 extends such that the second heat exchanger 42 is positioned radially opposite the first heat exchanger 4L. The heat transfer fluid F circulating in the heat pipe device 5 is initially in a liquid state at a temperature of 25 to 300 °C.For the sake of clarity and conciseness, the heating process will be described in this document for a single first heat exchanger 41 and a single second heat exchanger 42, connected by a single heat pipe device 5, the operation of several heat pipe devices 5 being analogous.

[0088] In a first step El, the heat transfer fluid F (circulating in the heat pipe device 5) is heated in the first heat exchanger 41 from the calories from the exhaust flow FE circulating in the inner channel V of the turbomachine M. In this example, the heat transfer fluid F then passes into the gaseous state.

[0089] In a second step E2, the heat transfer fluid F in the gaseous state is conveyed in the heat pipe device 5 from the first heat exchanger 41 to the second heat exchanger 42. In this example, the heat transfer fluid F circulates in the heat pipe device 5 via the inner tube 52.

[0090] The fuel stream Q, initially at a cryogenic temperature of approximately -251°C, is then heated in a third stage E3, in the second heat exchanger 42, using heat from the heat transfer fluid F in its gaseous state. In this example, the fuel stream Q is heated to a temperature of approximately -73 to 127°C. The heat transfer fluid F is then cooled to a liquid state.

[0091] The heat transfer fluid F in the liquid state circulates in this example by capillary action in the outer tube 51 of the heat pipe device 5 to reach the first heat exchanger 41 where it will be reheated by the exhaust flow FE.

Claims

Demands

1. A heating system (1) for a fuel (Q) for supplying an aircraft turbomachine (M), the fuel (Q) being drawn from a cryogenic tank (R), the turbomachine (M) comprising a combustion chamber (CC) and an internal flow (V) for circulating an exhaust stream (FE) from the combustion chamber (CC), the heating system (1) comprising: • a fuel circuit (2) configured to be connected inlet to the cryogenic tank (R) and outlet to the turbomachine (M), a fuel stream (Q) circulating upstream to downstream in the fuel circuit (2), • at least one first heat exchanger (41) configured to be mounted in the internal flow (V) of the turbomachine (M), the first heat exchanger (41) being configured to heat a phase-transition heat transfer fluid (F) from calories from the exhaust stream (FE),• at least one second heat exchanger (42) configured to be mounted externally to the inner runner (V) of the turbomachine (M), the second heat exchanger (42) being configured to heat the fuel stream (Q) using calories from the heat transfer fluid (F), and • at least one heat pipe device (5) mounted in direct contact with the first heat exchanger (41) and the second heat exchanger (42), the heat transfer fluid (F) circulating in the heat pipe device (5) and being configured to pass into a gaseous state in the first heat exchanger (41) and into a liquid state in the second heat exchanger (42).

2. Heating system (1) according to claim 1, comprising a plurality of heat pipe devices (5) mounted between the first heat exchanger (41) and the second heat exchanger (42).

3. Heating system (1) according to any one of claims 1 to 2, wherein the turbomachine (M) includes an exhaust nozzle (TE) downstream of the combustion chamber (CC), the system of heater (1) is mounted at the level of an inner vein of the exhaust nozzle (TE).

4. Heating system (1) according to any one of claims 1 to 3, wherein the phase transition heat transfer fluid (F) is selected from pentane, methanol, toluene, acetone, water, glycol water.

5. Heating system (1) according to any one of claims 1 to 4, wherein the turbomachine (M) includes at least one fixed guide vane (AG) mounted radially in the inner runner (V), the first heat exchanger (41) is integrated into the fixed guide vane (AG) of the turbomachine (M).

6. Heating system (1) according to claim 5, wherein the fixed guide vane (AG) includes a hollow casing (ENV) defining an inner wall (PI), the first heat exchanger (41) includes a plurality of conductive elements (7), mounted on the inner wall (PI) of the fixed guide vane (AG).

7. Heating system (1) according to claim 6, wherein the inner wall (PI) of the fixed guide vane (AG) is a porous wall so as to improve heat transfer with the exhaust flow (FE).

8. Heating system (1) according to any one of claims 1 to 7, wherein the second heat exchanger (42) is mounted in a protection module (6) comprising an inert gas, the protection module (6) being configured to detect a leak in the second heat exchanger (42).

9. Heating system (1) according to any one of claims 1 to 8, wherein the second heat exchanger (42) is positioned vertically above the first heat exchanger (41).

10. Heating system (1) according to any one of claims 1 to 9, comprising: • at least two first heat exchangers (41), each configured to be mounted in the inner core (V) of the turbomachine (M), and • at least two heat pipe devices (5) respectively connecting the first two heat exchangers (41) to the second heat exchanger (42).

11. Assembly of an aircraft turbomachine (M) comprising a combustion chamber (CC) and an internal runner (V) of circulation of an exhaust flow (FE) from the combustion chamber (CC) and of a heating system (1) according to any one of claims 1 to 10, the first heat exchanger (41) being mounted in the inner channel (V) of the turbomachine (M) and the second heat exchanger (42) being mounted externally to the turbomachine (M).

12. Aircraft comprising a cryogenic tank (R) in which is stored a fuel (Q) and an assembly of a turbomachine (M) and a fuel (Q) heating system (1) according to any one of claims 1 to 10.

13. A method for heating a fuel (Q) for supplying an aircraft turbomachine (M), the fuel (Q) being drawn from a cryogenic tank (R), the turbomachine (M) comprising a combustion chamber (CC) and an internal flow channel (V) for circulating an exhaust stream (FE) from the combustion chamber (CC), the heating method being carried out by means of the heating system (1) according to any one of claims 1 to 10 and comprising the steps of: • To heat (El) the heat transfer fluid (F) in the liquid state in the first heat exchanger (41) using calories from the exhaust flow (FE) circulating in the inner channel (V) of the turbomachine (M) so that the heat transfer fluid (F) is in the gaseous state, • To convey (E2), in the heat pipe device (5), the heat transfer fluid (F) in a gaseous state to the second heat exchanger (42), • Heat (E3) the fuel stream (Q), in the second heat exchanger (42), from the calories from the heat transfer fluid (F) in the gaseous state so that the heat transfer fluid (F) is in the liquid state.

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