Aircraft turbomachine comprising a system for heating a heat transfer fluid and associated heating method

Abstract

A gas turbine engine for an aircraft comprising a heating system (6) for a heat transfer fluid (FC) configured to heat a fuel (Q) from the cryogenic tank, the heating system (6) comprising: a closed loop (61) for circulating the heat transfer fluid (FC) comprising a main portion (61A) mounted in an outer casing (5) of the turbine engine, and a heating module (63) mounted in the primary flow path (V1) and configured to heat the heat transfer fluid (FC) from calories transferred by the exhaust air flow (AE), the closed loop (61) comprising a forward circulation channel (64) and a return circulation channel (65) extending in the secondary flow path (V2), the forward circulation channel (64) and the return circulation channel (65) being concentric. Abstract figure: Figure 4

Description

Title of the invention: Aircraft turbomachine comprising a system for heating a heat transfer fluid and associated heating method Technical field

[0001] The present invention relates to the field of aircraft comprising a fluid stored in a cryogenic tank, for example fuel for powering an aircraft turbomachine.

[0002] Climate change is a major concern for many legislative and regulatory bodies around the world. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by various States. In particular, an ambitious standard applies both to new types of aircraft and those in circulation requiring the implementation of technological solutions in order to make them compliant with current regulations. Civil aviation has been mobilizing for several years now to make a contribution to the fight against climate change.

[0003] Technological research efforts have already made it possible to significantly improve the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products 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 and minimizing greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0005] This sustained research and development work covers new generations of aircraft engines, the lightening of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to ensure propulsion, and, as essential complements to technological progress, aeronautical biofuels.

[0006] To this end, the invention is the result of technological research aimed at very significantly improving the performance of aircraft and, in this sense, contributes to reducing the environmental impact of aircraft. For this, The invention relates to the field of aircraft comprising one or more turbomachine(s) powered by fuel stored in a cryogenic tank.

[0007] In a known manner, an aircraft comprises one or more turbomachines to enable its movement from the acceleration of an upstream to downstream air flow. For this, with reference to [Fig.l], the turbomachine 1 comprises a propulsion member 2, mounted on a propulsion shaft 3, which converts the rotational movement into an air flow which enables the aircraft to be propelled.

[0008] To drive the rotation of the propulsion member 2, the turbomachine 1 comprises successively along a longitudinal axis X, a compression stage 11, a combustion chamber 12, a turbine stage 13 and an exhaust nozzle 14. The compression stage 11 is configured to receive an incoming air flow and compress it so as to supply the combustion chamber 12. The combustion, in the combustion chamber 12, between a fuel flow and the compressed air flow generates an exhaust air flow AE which drives the turbine stage 13 in rotation and circulates in the exhaust nozzle 14 to generate the thrust of the aircraft. The rotation of the turbine stage 13 then drives the rotation of the propulsion shaft 3, which in turn drives the compression stage 11 and the propulsion member 2.

[0009] In the case of a dual-flow turbomachine 1, as shown in [Fig.l], the compression stage 11, the combustion chamber 12, the turbine stage 13 and the exhaust nozzle 14 form a primary vein VI for circulating a primary air flow FL. The primary vein VI is delimited externally by an intermediate casing 4. The turbomachine 1 also comprises a secondary vein V2 for circulating a secondary air flow F2 accelerated by the propulsion member 2 and which extends externally to the intermediate casing 4. The secondary vein V2 is delimited externally by an external casing 5, known under the designation "nacelle" and which forms a ducted turbomachine 1.

[0010] In the example of a turbomachine powered by fuel from a cryogenic tank, it is known to store the fuel, in particular hydrogen, in liquid form to limit the size and mass of the aircraft tanks. For example, the fuel is stored at a temperature of the order of -253 to -251°C (20 to 22 Kelvins) in a cryogenic tank of the aircraft.

[0011] In order to be able to be injected into the combustion chamber 12 of the turbomachine 1, the fuel must be conditioned, that is to say pressurized and heated, in order to allow optimal combustion. Conditioning is for example necessary to reduce the risk of icing / solidification of the water vapor contained in the air circulating in the turbomachine 1, in particular, at the fuel injectors.

[0012] Figures 1 and 2 show a conditioning system SC which comprises a fuel circuit CQ connected at the inlet to a cryogenic tank R and at the outlet to the combustion chamber 12 of the turbomachine 1. A fuel flow Q circulates in the fuel circuit CQ and successively passes through a mechanical pump P and a heat exchanger EC. The mechanical pump P is configured to circulate the fuel flow Q in the fuel circuit CQ. The heat exchanger EC is configured to provide calories to the fuel flow Q in order to heat it so that it can be injected into the turbomachine 1.

[0013] In practice, to be heated, the fuel Q draws calories from heat sources of the aircraft, such as for example the heat generated by the turbomachine 1 whose temperatures are high. The heat from the aircraft can also be used. In particular, in order to avoid the risk of contamination between the fuel and an oxidizing fluid, it is known to heat the fuel flow Q in the heat exchanger EC by means of a heat transfer fluid FC, which itself draws calories from the heat sources of the aircraft.

[0014] For this, the turbomachine 1 comprises, in a known manner, a heating system 100 for the heat transfer fluid FC shown in more detail in [Fig. 2] and which comprises a closed loop 101 for circulating the heat transfer fluid FC. In this closed loop 101, the heat transfer fluid FC is heated in a heating module 102 by the heat sources of the aircraft, before transferring its calories to the fuel flow Q in the heat exchanger EC. In particular, it is known to mount the heating module 102 in the primary vein VI of the turbomachine 1 at the exhaust nozzle 14 so as to heat the heat transfer fluid FC from the calories generated by the exhaust air flow AE at the outlet of the turbine stage 13.

[0015] However, to circulate the heat transfer fluid FC in the circulation loop 101, the heating system 100 comprises a heavy and bulky mechanical pump 103. Also, the latter is mounted, in a known manner, in the outer casing 5 of the turbomachine 1, in order to avoid any risk of disturbance of the primary air flow F1 which could affect the performance of the turbomachine 1.

[0016] In other words, such a heating system 100 extends both in the outer casing 5 and in the primary vein VI of the turbomachine 1 and requires circulating the heat transfer fluid FC in the secondary vein V2. In practice, the heating system 100 comprises two pipes for circulating the heat transfer fluid FC mounted in the secondary vein V2, to allow the circulation, on the one hand, of a cold heat transfer fluid FC from the outer casing 5 to the primary vein VI, in a forward pipe 105, and, on the other hand, of a hot heat transfer fluid FC from the primary vein VI to the outer casing 5, in a return pipe 106.

[0017] However, the cold heat transfer fluid FC0, which has been previously cooled in the heat exchanger EC by the frigories transferred by the fuel flow Q, generally has a temperature below -25° at the inlet of the forward pipe 105. As is known, in the secondary vein V2, the secondary air flow F2 is loaded with water particles, which, in contact with the cold forward pipe 105, can condense and form frost on the outer walls of the pipe. The accumulation of frost can increase the cross-section of the forward pipe 105, which can lead to a significant resistance surface which can disrupt the circulation of the secondary air flow F2 and affect the performance of the turbomachine 1.

[0018] To limit this drawback, an immediate solution would be to insulate the outgoing pipe 105. However, to sufficiently insulate the pipe from the cold temperatures of the heat transfer fluid FC, a significant quantity of insulating members would be necessary. This would lead to a significant increase in the cross-section of the pipe which could, similar to frost, disrupt the circulation of the secondary air flow F2.

[0019] The invention thus aims to eliminate at least some of these drawbacks by proposing a gas turbomachine comprising a reliable and efficient system for heating a heat transfer fluid. The invention aims in particular at a heating system mounted partly in the secondary vein without the circulation of the secondary air flow being affected. PRESENTATION OF THE INVENTION

[0020] The invention relates to a gas turbomachine for an aircraft configured to be powered by fuel from a cryogenic tank, the gas turbomachine comprising an inner primary stream and an outer secondary stream delimited radially by an intermediate casing, the secondary stream being delimited externally by an outer casing, the gas turbomachine comprising at least one compression stage, a combustion chamber, a turbine stage and an exhaust nozzle forming the primary stream, an exhaust air flow from the combustion chamber circulating in the exhaust nozzle, the gas turbomachine comprising a system for heating a heat transfer fluid configured to heat the fuel from the cryogenic tank so as to allow its injection into the combustion chamber, the heating system comprising: • a closed loop for circulating the heat transfer fluid, the closed loop comprising at least one main portion mounted in the outer casing, • a mechanical pump configured to circulate the heat transfer fluid in the circulation loop, • a heating module mounted on the closed loop in the primary vein, the heating module being configured to heat the heat transfer fluid from calories transferred by the exhaust air flow, • the closed loop comprising a forward circulation channel, configured to convey the heat transfer fluid from the main portion of the closed loop to the heating module, and a return circulation channel, configured to convey the heat transfer fluid from the heating module to the main portion, the forward circulation channel and the return circulation channel extending in the secondary vein, • the forward circulation channel and the return circulation channel being concentric.

[0021] The heating system comprising a concentric forward circulation channel and a concentric return circulation channel makes it possible to reduce the number of pipes in the secondary vein, which limits the space requirement inside the secondary vein and thus limits the pressure losses by limiting the impact on the secondary air flow. The performance of the turbomachine is thus not affected, although it comprises a heat transfer fluid heating system which extends partly into the secondary vein.

[0022] The heating system according to the invention also makes it possible to avoid a circulation channel that is too cold in contact with the secondary flow which circulates in the secondary vein of the turbomachine, which limits the risk of frost formation and thus makes it possible to limit the risk of increasing the resistance surface of the circulation channels with respect to the secondary flow. This makes it possible to limit disturbances of the secondary flow and therefore to improve the performance of the turbomachine.

[0023] A concentric forward circulation channel and a return circulation channel advantageously make it possible to preheat the cold heat transfer fluid coming from the main portion of the closed loop and which has not yet been reheated in the heating module by means of the hot heat transfer fluid which has been reheated in the heating module. Such preheating makes it possible to limit the risk of frost forming on the walls of the circulation channels regardless of the temperature of the cold heat transfer fluid at the inlet of the forward circulation channel. This also makes it possible to limit the performance of the heating module, which can thus have a reduced mass and size, since the heat transfer fluid has a higher temperature at the inlet of the latter.

[0024] In parallel, the hot heat transfer fluid is advantageously cooled in the return circulation channel by the cold heat transfer fluid, which makes it possible to heat the heat transfer fluid more significantly in the heating module since it will be cooled before entering the external casing. Greater heating allows a lower flow rate of the heat transfer fluid in the closed loop, which makes it possible to mount a mechanical pump of smaller dimensions and which is therefore advantageously lighter and less bulky.

[0025] Furthermore, the heating system according to the invention makes it possible to limit the addition of insulating members around the circulation channels, which limits both the resistance surface with respect to the secondary flow and the mass of the turbomachine.

[0026] Two concentric circulation channels also make it possible to limit the risk of leakage of the heat transfer fluid into the secondary vein. Indeed, in the event of a leak from the internal circulation channel, the heat transfer fluid is discharged into the external circulation channel. The sealing of the closed loop is thus improved.

[0027] Preferably, the forward circulation channel extends radially internally to the peripheral return circulation channel, which makes it possible to ensure that the cold heat transfer fluid does not circulate in a pipe in contact with the secondary flow which circulates in the secondary vein. The channel in which the cold heat transfer fluid circulates is encircled by the channel in which the hot heat transfer fluid circulates, which makes it possible to avoid any risk of frost forming on the walls of the circulation channels in the secondary vein. Such an embodiment makes it possible to avoid any risk of increasing the resistance surface with respect to the secondary flow, which makes it possible to avoid any risk of disturbance of the secondary flow and makes it possible to guarantee the performance of the turbomachine by ensuring optimal efficiency.

[0028] Furthermore, a radially inner flow channel makes it possible to avoid adding an insulating member around the flow channels, which limits both the resistance surface with respect to the secondary flow and the mass of the turbomachine.

[0029] Such an embodiment thus makes it possible to ensure the durability of the pipelines while limiting maintenance operations, which makes it possible to reduce costs.

[0030] Preferably, the heat transfer fluid is an inert gas, making it possible to limit any risk of contact between the fuel flow and an oxidizing fluid.

[0031] In one embodiment, the heat transfer fluid is chosen from: nitrogen, helium, carbon dioxide, argon, neon or a synthetic oil.

[0032] In one embodiment, the heat transfer fluid has a temperature of less than or equal to -25°C at the inlet of the forward circulation channel. The heat transfer fluid can thus transfer a significant amount of calories into the main portion, for example to the fuel flow, which allows efficient heating. Thanks to the concentric circulation channels, even in the case of such a temperature, the heating system does not present any risk of frost formation.

[0033] Preferably, the heating system comprises a heat exchanger configured to transfer calories taken from the heat transfer fluid to the fuel flow, so as to heat the fuel flow to at least a vaporization temperature. The heat transfer fluid which has been heated in the heating module by the exhaust air flow can advantageously transfer calories to the fuel flow in the heat exchanger and heat it efficiently.

[0034] The invention also relates to an aircraft comprising at least one gas turbomachine as described above and a cryogenic tank in which a fuel is stored to supply the gas turbomachine.

[0035] Finally, the invention relates to a method for heating a heat transfer fluid in a gas turbomachine as described above, the method comprising the steps of: • convey the heat transfer fluid in the forward circulation channel from the main portion of the closed loop to the heating module, • heat the heat transfer fluid in the heating module from the calories transferred by the exhaust air flow, and • route the heat transfer fluid in the return circulation channel from the heating module to the main portion of the closed loop.

[0036] In a preferred embodiment, the routing of the heated heat transfer fluid in the heating module is carried out in the return circulation channel which extends externally to the supply circulation channel.

[0037] Preferably, the method comprises, after the heat transfer fluid has been conveyed into the return circulation channel, a step of reheating the fuel flow in a heat exchanger using the heat transfer fluid. PRESENTATION OF FIGURES

[0038] The invention will be better understood on reading the description which follows, 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.

[0039] [Fig.l] is a schematic representation of a gas turbomachine comprising a system for heating a heat transfer fluid according to the prior art.

[0040] [Fig.2] is a close-up view of the heating system of [Fig.l].

[0041] [Fig. 3] is a schematic representation of a gas turbomachine comprising a system for heating a heat transfer fluid according to one embodiment of the invention.

[0042] [Fig.4] is a close-up view of the heating system of [Fig.3].

[0043] [Fig. 5] is a view along a cross-sectional plane of the circulation channels of the heating system of [Fig.4] in the secondary vein of the turbomachine.

[0044] It should be noted that the figures set out the invention in detail to implement the invention, said figures can of course be used to better define the invention where appropriate. DETAILED DESCRIPTION OF THE INVENTION

[0045] In a known manner, an aircraft comprises one or more gas turbomachines to enable its movement from the acceleration of an upstream to downstream air flow. Such a turbomachine 1 is shown in [Fig. 3] and extends, in this example, longitudinally along an axis X. In other words, the air flow circulates in the turbomachine 1 from upstream to downstream along the longitudinal axis X. For this, the turbomachine 1 comprises a propulsion member 2, mounted on a propulsion shaft 3, which converts the rotational movement into an air flow which enables the aircraft to be propelled.

[0046] In practice, to drive the rotation of the propulsion member 2, the turbomachine 1 successively comprises, along the longitudinal axis X, a compression stage 11, a combustion chamber 12, a turbine stage 13 and an exhaust nozzle 14. It goes without saying that the turbomachine 1 could comprise several propulsion shafts 3. Likewise, it goes without saying that the turbomachine 1 could comprise several compression stages 11 as well as several turbine stages 13.

[0047] The compression stage 11 is configured to receive an incoming air flow and compress it so as to supply the combustion chamber 12. The combustion, in the combustion chamber 12, between a fuel flow Q and the compressed air flow generates an exhaust air flow AE which drives the turbine stage 13 in rotation. The rotation of the turbine stage 13 then drives the rotation of the propulsion shaft 3, which in turn drives the compression stage 11 and the propulsion member 2. Downstream of the turbine stage 13, the exhaust air flow AE circulates in the exhaust nozzle 14 and generates part of the thrust of the aircraft.

[0048] Still with reference to [Fig. 3], the turbomachine 1 according to the invention is a dual-flow turbomachine, in which the compression stage 11, the combustion chamber 12, the turbine stage 13 and the exhaust nozzle 14 form a primary vein VI for circulating a primary air flow FL. The turbomachine 1 also comprises a secondary vein V2 for circulating a secondary air flow F2 accelerated by the propulsion member 2. The primary vein VI and the secondary vein V2 are annular and extend radially around the propulsion shaft 3. An intermediate casing 4 is mounted between the primary vein VI and the secondary vein V2. Also, the air flow generated by the propulsion member 2 is divided into a primary air flow Fl, which circulates from upstream to downstream in the primary vein VI to be burned in the combustion chamber 12, and a secondary air flow F2 which circulates in the secondary vein V2.Likewise, an outer casing 5, also known to those skilled in the art under the designation “nacelle”, externally delimits the secondary vein V2 to form a shrouded turbomachine 1.

[0049] The turbomachine 1 is configured to be supplied by a fuel flow Q from a cryogenic tank R.

[0050] In this example, the fuel Q in the cryogenic tank R is stored at a temperature of the order of -253 to -251°C (20 to 22 Kelvins). At this temperature, the fuel stream Q is liquid. In particular, 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. In order to be introduced into the combustion chamber 12, the fuel Q, initially stored at cryogenic temperatures, must be reheated.

[0051] For this, a conditioning system SC, also shown in [Fig. 3], comprises a fuel circuit CQ configured to connect the cryogenic tank R to the combustion chamber 12 of the turbomachine 1. The fuel flow Q circulates in the fuel circuit CQ by means of a mechanical pump P and passes through a heat exchanger EC, in which it exchanges calories with a heat transfer fluid FC. In this example, the heat exchanger EC is configured to heat the fuel flow Q to a vaporization temperature. It goes without saying that the fuel flow Q could pass through more than one heat exchanger EC to heat, for example, the fuel flow Q gradually.

[0052] In order for the heat transfer fluid FC to transfer calories to the fuel flow Q, the turbomachine 1 according to the invention comprises a heating system 6 for said heat transfer fluid FC.

[0053] A heating system 6 according to one embodiment of the invention is shown in Figures 3 and 4.

[0054] The heating system 6 comprises a closed loop 61 for circulating the heat transfer fluid FC. The closed loop 61 comprises a main portion 61A mounted in the outer casing 5 of the turbomachine 1 and an auxiliary portion 61B which extends, in the turbomachine 1, outside the outer casing 5. In other words, the auxiliary portion 61B extends, in this example, in the secondary stream V2, in the intermediate casing 4 and in the primary stream VL. For the sake of clarity, the closed loop 61 is described only in the turbomachine 1, however, it goes without saying that the closed loop 61 could also extend outside the turbomachine 1, in particular to join the fuel circuit CQ.

[0055] Preferably, the heat transfer fluid FC is an inert gas, so as to limit any risk of contamination between the fuel Q and an oxidizing fluid. More preferably, the heat transfer fluid FC is chosen from: nitrogen, helium, carbon dioxide, argon, neon or a synthetic oil.

[0056] According to one aspect of the invention, the heating system 6 comprises a heating module 63, shown in [Fig. 4], mounted on the auxiliary portion 61B of the closed loop 61 and configured to heat the heat transfer fluid FC. In particular, the heating module 63 is mounted in the primary vein VI of the turbomachine 1, at the exhaust nozzle 14, and is configured to heat the heat transfer fluid FC. FC carrier from calories transferred by the exhaust air flow AE. The exhaust air flow AE preferably has, at the outlet of the turbine stage 13, a temperature between 300 and 600 °C, which allows efficient heating.

[0057] In this example, the heating module 63 is a tubular or plate heat exchanger. The heat exchange surfaces in the heating module 63 comprise, in one embodiment, additional exchange members, such as fins or grooves for example, in order to improve the overall aerothermal performance of the heating module 63. It goes without saying that the heating module 63 could be in a different form, for example in the form of an exchanger having an annular shape mounted partly in the primary vein VI or mounted in the entire primary vein VI, i.e. over 360°.

[0058] In a preferred embodiment, the heat transfer fluid and the exhaust air flow AE circulate in opposite directions, i.e. counter-current. Alternatively, the heat exchange between the heat transfer fluid FC and the exhaust air flow AE is carried out co-currently or by crossing the fluids.

[0059] In this document, a single heating module 63 is described, however it goes without saying that the heating system 6 could comprise several heating modules 63, mounted in the primary vein VI to gradually heat the heat transfer fluid FC from calories transferred by the exhaust air flow AE. Similarly, the heating system 6 could also comprise one or more heating modules 63 mounted on the main portion 61A and / or on the auxiliary portion 61B of the closed loop 61 to heat the heat transfer fluid FC from calories transferred by other heat sources of the turbomachine 1 or of the aircraft, such as for example the engine lubricating oil, the cabin air, the heat from electrical or electronic systems, etc.

[0060] In order to allow the circulation of the heat transfer fluid FC in the closed loop 61, the heating system 6 comprises a mechanical pump 62. Preferably, the mechanical pump 62 is mounted on the main portion 61A of the closed loop 61, that is to say in the outer casing 5, so as to limit the size of the primary vein VI and the secondary vein V2 in which the flow of the primary air flow F1 and the secondary air flow F2 could be disturbed.

[0061] According to one aspect of the invention, still with reference to [Fig. 4], the closed loop 61 comprises a forward circulation channel 64 and a return circulation channel 65 which connect the main portion 61A mounted in the outer casing 5 and the heating module 63 mounted in the primary vein VL. In other words, the forward circulation channel 64 and the return circulation channel 65 extend at least partly in the secondary vein V2 of the turbomachine 1. More precisely, the forward circulation channel 64 is configured to convey the heat transfer fluid FC from the main portion 61A of the closed loop 61 to the heating module 63, and the return circulation channel 65 is configured to convey the heat transfer fluid FC from the heating module 63 to the main portion 61A.

[0062] In this example, the forward circulation channel 64 and the return circulation channel 65 are in the form of cylindrical tubular pipes.

[0063] According to one aspect of the invention, the forward circulation channel 64 and the return circulation channel 65 are concentric, which limits the size of the closed loop 61 in the secondary vein V2 and thus limits the resistance and therefore the disturbances of the secondary air flow F2. Preferably, the return circulation channel 65 extends radially outwardly to the forward circulation channel 64, as shown in Figures 4 and 5, [Fig. 5] representing a sectional view of the circulation channels 64, 65 along a cross-sectional plane. In other words, the return circulation channel 65 is annular and extends around the forward circulation channel 64.

[0064] In practice, the heat transfer fluid FC circulating from the main portion 61A to the heating module 63 is configured to have previously transferred its calories to the fuel flow Q, it is therefore cold, and the heat transfer fluid FC circulating from the heating module 63 to the main portion 61A is configured to have been heated by the exhaust air flow AE in the heating module 63, it is therefore hot. Subsequently, a distinction is made between the cold heat transfer fluid FC-0 which circulates in the forward circulation channel 64 and the hot heat transfer fluid FC-1 which circulates in the return circulation channel 65.

[0065] In this example, the cold heat transfer fluid FC-0 has a temperature of less than or equal to -25°C at the inlet of the forward circulation channel 64. Preferably, the cold heat transfer fluid FC-0 has a temperature of between -100 and -25°C at the inlet of the forward circulation channel 64. Similarly, the hot heat transfer fluid FC-1 has, in this example, a temperature of between 50 and 250°C at the inlet of the return circulation channel 65.

[0066] Thanks to the forward circulation channel 64 mounted internally to the return circulation channel 65, the cold heat transfer fluid FC-0 which circulates in the forward circulation channel 64 does not cool the walls of the pipes of the closed loop 61 in contact with the secondary air flow F2 in the secondary vein V2, which makes it possible to avoid a risk of frost forming on the closed loop 61 in the secondary vein V2.

[0067] In one embodiment, with reference to [Fig. 5], the forward circulation channel 64 has an external radial diameter D64 of between 2 and 6 cm. The return circulation channel 65 preferably has an external diameter D65 of between 4 and 10 cm. The diameter of the pipes passing through the secondary vein V2 is thus limited, which limits the disturbances of the secondary air flow F2.

[0068] In order to heat the fuel flow Q, the heat exchanger EC described previously and mounted on the fuel circuit CQ is also mounted on the closed loop 61, so as to exchange calories between the heat transfer fluid FC and the fuel flow Q. In other words, the heat transfer fluid FC is configured to circulate in the closed loop 61 and pass through the heat exchanger EC to heat the fuel flow Q coming from the cryogenic tank R.

[0069] For the sake of clarity and conciseness, only the closed loop 61, the mechanical pump 62 and the heating module 63 are described in this document, however, it goes without saying that the heating system 6 may comprise other components known to those skilled in the art, such as valves, an expansion tank, etc.

[0070] A method for heating the heat transfer fluid FC in the heating system 6 described above will now be described. In this example, the forward circulation channel 64 is mounted internally to the return circulation channel 65. Similarly, in this example the heat transfer fluid FC is nitrogen.

[0071] The method comprises a first step E1 of conveying the heat transfer fluid FC into the forward circulation channel 64 from the main portion 61A of the closed loop 61 to the heating module 63. The cold heat transfer fluid FC-0 then circulates internally to the return circulation channel 65 and the walls of the forward circulation channel 64 are not in contact with the secondary air flow F2 which circulates in the secondary vein V2. The forward circulation channel 64 thus advantageously presents no risk of frost forming on its external walls.

[0072] In a second step E2, the cold heat transfer fluid FC-0 passes through the heating module 63 in the primary vein VI, in which it is reheated from the calories transferred by the exhaust air flow AE which circulates in the exhaust nozzle 14 at the outlet of the turbine stage 13 of the turbomachine 1. At the outlet of the heating module 63, the heat transfer fluid FC is hot.

[0073] The hot heat transfer fluid FC-1 is then conveyed, in a third step E3, in the return circulation channel 65 from the heating module 63 to the main portion 61A of the closed loop 61 mounted in the outer casing 5. The hot heat transfer fluid FC-1 is conveyed in this step externally to the forward circulation channel 64. The hot heat transfer fluid FC-1 thus circulates between the cold heat transfer fluid FC-0 and the secondary air flow F2.

[0074] In a preferred embodiment, the hot heat transfer fluid FC-1 is then conveyed, in a fourth step E4, to the heat exchanger EC, mounted both on the fuel circuit CQ and on the closed loop 61, in which it transfers calories to the fuel flow Q to heat it. At the outlet of the heat exchanger EC, the heat transfer fluid FC is cold and circulates in the closed loop 61 to be heated in the heating module 63 by the exhaust air flow AE.

Claims

Claims

1. Gas turbomachine (1) for aircraft configured to be powered by a fuel (Q) from a cryogenic tank (R), the gas turbomachine (1) comprising an inner primary stream (VI) and an outer secondary stream (V2) delimited radially by an intermediate casing (4), the secondary stream (V2) being delimited externally by an outer casing (5), the gas turbomachine (1) comprising at least one compression stage (11), a combustion chamber (12), a turbine stage (13) and an exhaust nozzle (14) forming the primary stream (VI), an exhaust air flow (AE) from the combustion chamber (12) circulating in the exhaust nozzle (14), the gas turbomachine (1) comprising a heating system (6) for a heat transfer fluid (FC) configured to heat the fuel (Q) from the cryogenic tank (R) so as to allow its injection in the combustion chamber (12), the heating system (6) comprising: • a closed loop (61) for circulating the heat transfer fluid (FC), the closed loop (61) comprising at least one main portion (61 A) mounted in the outer casing (5), • a mechanical pump (62) configured to circulate the heat transfer fluid (FC) in the circulation loop (61), • a heating module (63) mounted on the closed loop (61) in the primary vein (VI), the heating module (63) being configured to heat the heat transfer fluid (FC) from calories transferred by the exhaust air flow (AE), • the closed loop (61) comprising a forward circulation channel (64), configured to convey the heat transfer fluid (FC) from the main portion (61 A) of the closed loop (61) to the heating module (63), and a return circulation channel (65), configured to convey the heat transfer fluid (FC) from the heating module (63) to the main portion (61 A), the forward circulation channel (64) and the return circulation channel (65) extending in the secondary vein (V2), • the forward circulation channel (64) and the return circulation channel (65) being concentric.

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10. Gas turbomachine (1) according to claim 1, in which the forward circulation channel (64) extends radially internally to the peripheral return circulation channel (65). Gas turbomachine (1) according to one of claims 1 to 2, in which the heat transfer fluid (FC) is an inert gas. Gas turbomachine (1) according to claim 3, in which the heat transfer fluid (FC) is chosen from: nitrogen, helium, carbon dioxide, argon, neon or a synthetic oil. Gas turbomachine (1) according to one of claims 1 to 4, in which the heat transfer fluid (FC) has at the inlet of the flow channel (64) a temperature less than or equal to -25°C. Gas turbomachine (1) according to one of claims 1 to 5, in which the heating system (6) comprises a heat exchanger (EC) configured to transfer calories taken from the heat transfer fluid (FC) to the fuel flow (Q), so as to heat the fuel flow (Q) to at least a vaporization temperature. Aircraft comprising at least one gas turbomachine (1) according to one of claims 1 to 6 and a cryogenic tank (R) in which a fuel (Q) is stored to supply the gas turbomachine (1). Method for heating a heat transfer fluid (FC) in a gas turbomachine (1) according to one of claims 1 to 6, the method comprising the steps of: • convey (El) the heat transfer fluid (FC) in the forward circulation channel (64) from the main portion (61 A) of the closed loop (61) to the heating module (63), • heating (E2) the heat transfer fluid (FC) in the heating module (63) from the calories transferred by the exhaust air flow (AE), and • convey (E3) the heat transfer fluid (FC) in the return circulation channel (65) from the heating module (63) to the main portion (61A) of the closed loop (61). Heating method according to claim 8, wherein the heat transfer fluid (FC) heated in the heating module (63) is conveyed in the return circulation channel (65) which extends externally to the supply circulation channel (64). Heating method according to one of claims 8 to 9, comprising, after the heat transfer fluid (FC) has been conveyed into the return circulation channel (65), a step (E4) of reheating the fuel flow (Q) in a heat exchanger (EC) by the heat transfer fluid (FC).

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