Heating turbomachine for a fuel conditioning system configured to supply an aircraft engine with fuel from a cryogenic tank

Abstract

A heating turbomachine (1), for a fuel conditioning system (SC), comprising: a compressor (11) configured to be supplied by a first ambient air flow (A1), a combustion chamber (2) configured to be supplied by a feed air flow (A) and by a fuel flow (Q2), the combustion chamber (2) being configured to discharge an exhaust air flow (AE), a fluid circuit (CF) in which a heat transfer fluid (F) circulates, and a post-combustion heat exchanger (3), configured to take calories from the exhaust air flow (AE) so as to heat the heat transfer fluid (F), the heating turbomachine (1) is remarkable in that it comprises at least one manifold (4) configured to mix the first ambient air flow (A1) from the compressor (11) and the second air flow (A2) from the engine (M) in order to distribute the flow of supply air (A) to the combustion chamber (2).Abstract Figure: Figure 2.

Description

Title of the invention: Heating turbomachine for a fuel conditioning system configured to supply an aircraft engine with fuel from a cryogenic tank Technical field

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

[0002] It is known to store fuel, in particular hydrogen, in liquid form to limit the size and mass of the aircraft tanks. For example, hydrogen is stored at a temperature of the order of 20 to 22 Kelvins (-253 to -251°C) in a cryogenic tank of the aircraft.

[0003] In order to be injected into a combustion chamber of an aircraft engine, the fuel must be heated to allow optimal combustion. More precisely, it is necessary for the fuel initially in the liquid state to be heated to be in a gaseous state before entering the combustion chamber.

[0004] In practice, the fuel heating step is energy-intensive and requires taking calories from heat sources. Among the various technologies for heating liquid fuel, it is known, with reference to [Fig.l], to generate a heating air flow AC by means of a heating turbomachine 102 supplied, on the one hand, by fuel Q from a cryogenic tank RC and, on the other hand, by an air flow A. A heat exchanger 101 advantageously makes it possible to transfer calories from the heating air flow AC to the liquid fuel flow Q in order to make it gaseous. It is thus capable of being consumed in a combustion chamber CC of the aircraft engine M.

[0005] In practice, it is complex for the heating turbomachine 102 to generate a large quantity of heat with good efficiency. Indeed, by nature, a turbomachine is optimized to maximize its mechanical energy and limit its thermal losses. In fact, the thermal energy generated is proportional to the mechanical energy generated. Also, to generate enough thermal energy to heat the fuel flow, it is necessary to use a large turbomachine which generates a large quantity of unwanted mechanical energy.

[0006] The invention thus aims to eliminate at least some of these drawbacks by proposing a turbomachine which makes it possible to achieve optimal heating while having a small footprint and reduced mass. The invention also aims at a turbomachine which makes it possible to facilitate the regulation of the fuel system. PRESENTATION OF THE INVENTION

[0007] The invention relates to a heating turbomachine for a fuel conditioning system configured to supply an aircraft engine from a fuel flow from a cryogenic tank, the heating turbomachine comprising: • a compressor configured to be supplied by a first flow of ambient air from an air inlet, • a turbine connected to the compressor by a turbomachine shaft, the turbine being configured to drive the compressor, • a combustion chamber configured to be supplied, on the one hand, by a flow of supply air and, on the other hand, by a flow of fuel, the combustion chamber being configured to evacuate a flow of exhaust air loaded with calories and drive the turbine in rotation, • a fluid circuit in which a heat transfer fluid circulates from upstream to downstream, and • a post-combustion heat exchanger, mounted on the fluid circuit, configured to draw calories from the exhaust air flow between the combustion chamber and the turbine, so as to heat the heat transfer fluid.

[0008] The heating turbomachine is remarkable in that it comprises at least one manifold, mounted between the compressor and the combustion chamber, configured to mix the first ambient air flow from the compressor and the second air flow from the aircraft engine in order to form the supply air flow configured to supply the combustion chamber.

[0009] The mixing between an ambient air flow and an air flow from the aircraft engine allows a supply air flow having a higher flow rate, while limiting both the flow rate of the second air flow from the aircraft engine and the flow rate of the first air flow, which allows the use of a compressor having reduced characteristics compared to the compressor of the prior art. The heating turbomachine thus requires a less expensive, less heavy and less bulky compressor, allowing a heating turbomachine whose size and mass are also limited.

[0010] Furthermore, thanks to the invention, the thermal energy of the turbomachine is captured directly at the outlet of the combustion chamber before it is converted into mechanical energy by the turbine. The thermal efficiency of the heating turbomachine is then improved, which is advantageous for extracting calories intended for heating a fuel flow from a cryogenic tank.

[0011] In one embodiment, the heat transfer fluid is the fuel flow from the cryogenic tank. In other words, the fuel flow is used directly as a heat transfer fluid in the heating turbomachine.

[0012] In one embodiment, the heating turbomachine comprises at least one exhaust heat exchanger, mounted on the fluid circuit, configured to heat the heat transfer fluid from calories taken from the exhaust air flow at the outlet of the turbine, the post-combustion heat exchanger being mounted downstream of the exhaust heat exchanger in the fluid circuit. The heat transfer fluid is thus gradually heated with increasingly hot sources

[0013] Preferably, the first air flow has at the compressor inlet a temperature of between -70°C and 45°C and a pressure of between 0.015 and 0.1 MPa (0.15 and 1 bar). The first air flow can thus easily be taken from any ambient air inlet of the aircraft.

[0014] In one embodiment, the first air flow has a temperature of between 50 and 400°C and a pressure of between 0.1 and 0.8 MPa (1 and 8 bar) at the compressor outlet. The compressor can thus have limited performance, which limits its mass and its size in the heating turbomachine.

[0015] Preferably, the second air flow from the aircraft engine has at the inlet of the manifold a temperature of between 50 and 400°C and a pressure of between 0.1 and 0.8 MPa (1 and 8 bar). The second air flow taken from the aircraft engine can thus be mixed with the first air flow to form a feed air flow that is sufficiently hot and has a pressure that is sufficiently high to burn the fuel in the combustion chamber of the heating turbomachine.

[0016] In one embodiment, the feed air flow has at the inlet of the combustion chamber a temperature of between 50 and 400°C and a pressure of between 0.1 and 0.8 MPa (1 and 8 bar).

[0017] In one embodiment, the exhaust air flow has at the outlet of the combustion chamber a temperature of between 300°C and 500°C and a pressure of between 0.1 and 0.8 MPa (1 and 8 bar).

[0018] In a first embodiment, the second air flow comes from a low pressure compressor of the aircraft engine, making it possible to limit the pressure in the heating turbomachine and to limit the energy cost of the extraction.

[0019] In one embodiment, for a triple-spool engine, the second airflow comes from an intermediate compressor of the aircraft engine.

[0020] In a second embodiment, the second air flow comes from a stage intermediate of a high-pressure compressor of the aircraft engine. This embodiment makes it possible to ensure a high pressure level in the combustion chamber of the heating turbomachine and therefore better combustion. In one embodiment, in which the exhaust air flow is routed to the ambient air, the size of the turbine is reduced because its pressure ratio is higher. Alternatively, thanks to the air flow from a high-pressure compressor of the aircraft engine, the exhaust air flow can advantageously be routed into a secondary flow of the aircraft engine since its pressure remains high after its expansion in the turbine.

[0021] In a second embodiment, the second air flow comes from a circuit supplying the aircraft with air from the aircraft engine. In this case, the supply circuit existing, it is then simple to form a bypass to take the second air flow directly from it.

[0022] Preferably, the engine comprises an air system configured to mix an air flow from an intermediate stage of a high pressure compressor of the aircraft engine and an air flow from a low pressure compressor of the aircraft engine.

[0023] In one embodiment, the heating turbomachine comprises a heat exchanger, mounted between the aircraft engine and the manifold, configured to heat the second air flow with a heat source from the aircraft engine, before it is mixed with the first air flow.

[0024] Preferably, in this embodiment, the heating turbomachine comprises a distribution valve, configured to eject the second air flow towards the ambient, so as to bypass the combustion chamber of the heating turbomachine. Such an embodiment makes it possible, for example, to cool an engine oil even when the combustion chamber is deactivated.

[0025] The invention also relates to a fuel conditioning system configured to supply an aircraft engine with fuel from a cryogenic tank, the conditioning system comprising: • a fuel system configured to connect the cryogenic tank to the aircraft engine, a fuel flow circulating in the fuel system and • a heating turbomachine as described previously for heating the fuel flow circulating in the fuel circuit from calories taken from the exhaust air flow, the combustion chamber being supplied by a fuel flow from the fuel circuit and by a supply air flow from the manifold.

[0026] In one embodiment, the fuel stream is configured to be heated in the post-combustion exchanger directly by the calories taken from the exhaust air stream.

[0027] In one embodiment, the conditioning system comprises a heat exchanger mounted on the fuel circuit configured to heat the fuel stream from calories taken from the heat transfer fluid, the heat transfer fluid being configured to be heated in the post-combustion exchanger from calories taken from the exhaust air stream.

[0028] According to one aspect, the fuel flow is used directly as a heat transfer fluid in the heating turbomachine.

[0029] The invention also relates to an assembly comprising a cryogenic tank, an aircraft engine and a fuel conditioning system, as presented previously, connecting the cryogenic tank to the aircraft engine.

[0030] The invention also relates to a method for heating a heat transfer fluid by means of a heating turbomachine as described above, the heat transfer fluid circulating from upstream to downstream in the fluid circuit, the method comprising the steps of: • Supply the compressor with a first flow of ambient air, • Mix in the collector the first air flow from the compressor and a second air flow from the aircraft engine, so as to form a supply air flow, • Supply the combustion chamber, on the one hand, with the feed air flow and, on the other hand, with a fuel flow, the combustion chamber evacuating a flow of exhaust air loaded with calories in order to drive the turbine in rotation, and • Heat the heat transfer fluid in the post-combustion heat exchanger using calories taken from the exhaust air flow leaving the combustion chamber. PRESENTATION OF FIGURES

[0031] 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.

[0032] [Fig.l] is a schematic representation of a fuel conditioning system according to the prior art.

[0033] [Fig.2] is a schematic representation of a fuel conditioning system according to a first embodiment of the invention.

[0034] [Fig.3] is a schematic representation of an aircraft engine.

[0035] [Fig.4] is a schematic representation of a conditioning system of fuel according to a second embodiment of the invention.

[0036] [Fig. 5] is a schematic representation of a fuel conditioning system according to a third embodiment of the invention.

[0037] [Fig.6] is a schematic representation of the operation of the heating turbomachine of the conditioning system of [Fig.5].

[0038] [Fig.7] is a schematic representation of a conditioning system of fuel according to a fourth embodiment of the invention.

[0039] 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

[0040] With reference to [Fig.2], there is shown a fuel conditioning system SC according to an embodiment of the invention for conducting fuel Q from a cryogenic tank RC to the combustion chamber of an engine of an aircraft M, in particular, a turbine engine.

[0041] In this example, the fuel is liquid hydrogen, however the invention applies to other types of fuel, for example, liquid methane or liquefied natural gas. The fuel is stored in the cryogenic tank RC at a temperature of the order of -253°C (20K) to -251°C (22K).

[0042] According to the invention, the conditioning system SC is configured to supply the combustion chamber of the aircraft engine M with liquid-phase fuel from the cryogenic tank RC. The conditioning system SC makes it possible to take a fuel flow Q from the cryogenic tank RC and to heat it to an optimal temperature so that it can be injected into the combustion chamber of the aircraft engine M, as will be described in more detail later.

[0043] In this example, the aircraft engine M is configured to provide propulsion for the aircraft, in particular, by driving at least one propulsion member OP ([Fig.3]).

[0044] As shown in [Fig. 3], in a known manner, the aircraft engine M comprises a high pressure compressor CHP and a low pressure compressor CBP, configured to suck in an external air flow and compress it so that the air reaches an optimum speed, pressure and temperature at the inlet of the combustion chamber CC of the aircraft engine M.

[0045] With reference to [Fig. 2], the conditioning system SC according to the invention comprises a fuel circuit CQ (in close dashed lines in [Fig. 2]) in which the fuel flow Q circulates from upstream to downstream. The fuel circuit CQ thus comprises an inlet configured to be fluidically connected to the cryogenic tank RC and an outlet configured to be fluidically connected to the aircraft engine M.

[0046] As shown in [Fig.2], the packaging system SC comprises in in addition to a heating turbomachine 1 which comprises a compressor 11, configured to be supplied by a first ambient air flow Al from an air inlet EA, and a turbine 12 connected to the compressor 11 by a turbomachine shaft 13. The turbine 12 is configured to drive the compressor 11.

[0047] In one embodiment, the first air flow Al has, at the inlet of the compressor 11, a temperature Tl of between -70 and 45°C and a pressure PI of between 0.015 and 0.1 MPa. At the outlet of the compressor 11, the first air flow Al preferably has a temperature Tic of between 50 and 400°C and a pressure Pic of between 0.1 and 0.8 MPa.

[0048] Still with reference to [Fig. 2], the heating turbomachine 1 comprises a combustion chamber 2 configured to be supplied, on the one hand, by a feed air flow A and, on the other hand, by a fuel flow Q2 from the cryogenic tank RC. The combustion chamber 2 is further configured to evacuate an exhaust air flow AE loaded with calories in order to drive the turbine 12 in rotation.

[0049] According to the invention, the supply air flow A is a mixture between the first ambient air flow Al coming from the compressor 11 and a second air flow A2 coming from the aircraft engine M.

[0050] With reference to [Fig.3], according to several embodiments, the second air flow A2 comes from: • of the low pressure compressor CBP of the aircraft engine M, in the case of a twin-spool engine (solid arrow), • of the inter-pressure compressor of the aircraft engine M, in the case of a triple-body engine (not shown), • an intermediate stage of the high pressure compressor CHP of the aircraft engine M (dotted arrow), • an outlet of an air system S mounted in the aircraft engine M, the air system S being configured to mix at least two air sources of the aircraft engine M. In this example, the air system S is configured to mix an air flow from the intermediate stage of the high pressure compressor CHP and an air flow from the low pressure compressor CBP.

[0051] Preferably, the second air flow A2 has at the outlet of the aircraft engine M a temperature T2 of between 50 and 400°C, and a pressure P2 of between 0.1 and 0.8 MPa. The second air flow A2 has a high temperature T2 and pressure P2. Advantageously, it is not necessary for the first air flow A1 to be at a temperature T1 and a pressure P1 that are too high, since it will be mixed with the second air flow A2. The heating turbomachine 1 according to the invention thus does not require a high-performance compressor 11, which limits its cost, its mass and its size.

[0052] Still with reference to [Fig. 2], to allow the mixing of the first air flow Al and the second air flow A2, the heating turbomachine 1 comprises a manifold 4 mounted between the compressor 11 and the combustion chamber 2. The manifold 4 is configured to receive the first ambient air flow Al from the compressor 11 and the second air flow A2 from the aircraft engine M and to distribute the supply air flow A to the combustion chamber 2. In this example, the manifold 4 is in the form of a three-way valve.

[0053] Preferably, the feed air flow A has at the inlet of the combustion chamber 2 a temperature Ta of between 50 and 400°C, and a pressure Pa of between 0.1 and 0.8 MPa. Thanks to the second air flow A2 from the aircraft engine M, the feed air flow A has at the inlet of the combustion chamber 2 a temperature and a pressure sufficient to burn the fuel flow Q2.

[0054] Preferably, the exhaust air flow AE has at the outlet of the combustion chamber 2 a temperature Te of between 300 and 500°C and a pressure Pe of between 0.1 and 0.8 MPa.

[0055] In one embodiment, the heating turbomachine 1 comprises a cut-off valve 6, configured to cut off the supply of the second air flow A2 from the aircraft engine M, when the heating turbomachine 1 is not in use.

[0056] In this example, the combustion chamber 2 is supplied by a fuel flow Q2 taken from the fuel circuit CQ upstream of the aircraft engine M. For this purpose, the conditioning system SC preferably comprises a distribution member 5, in particular a three-way valve, in order to supply the aircraft engine M with a first fuel flow Q1 and the combustion chamber CC with a second fuel flow Q2, as shown in [Fig.2].

[0057] According to the invention, still with reference to [Fig. 2], the heating turbomachine 1 further comprises a fluid circuit CF in which a heat transfer fluid F circulates from upstream to downstream. In this embodiment, the heat transfer fluid F is nitrogen but it goes without saying that this could be different.

[0058] In a first embodiment, the heat transfer fluid F is distinct from the fuel flow Q.

[0059] In a second embodiment, shown in [Fig. 4], the heat transfer fluid F is the fuel flow Q from the cryogenic tank RC. In other words, the fuel flow Q is used directly as heat transfer fluid F in the heating turbomachine 1, as will be described in more detail later.

[0060] With reference to [Fig.2], the heating turbomachine 1 according to the invention comprises a post-combustion heat exchanger 3, mounted on the fluid circuit CF between the combustion chamber 2 and the turbine 12. The post-combustion heat exchanger 3 is configured to take calories from the exhaust air flow AE leaving the combustion chamber 2 so as to heat the heat transfer fluid F.

[0061] Preferably, the post-combustion heat exchanger 3 is mounted in the fluid circuit CF upstream of the turbine 12 so as to take calories from the exhaust air flow AE which is hot and has a high pressure. The post-combustion heat exchanger 3 is preferably made of a material suitable for high temperatures, preferably not very porous, for example, ceramic.

[0062] At the outlet of the post-combustion heat exchanger 3, the exhaust air flow AE is configured to pass through the turbine 12. The exhaust air flow AE makes it possible to generate mechanical energy via the turbine 12 which can drive the turbomachine shaft 13 and therefore the compressor 11. At the outlet of the turbine 12, the exhaust air flow AE is configured to be ejected from the heating turbomachine 2. In one embodiment, the heating turbomachine 2 comprises a nozzle Tu, configured to generate residual thrust for the aircraft.

[0063] In the second embodiment in which the heat transfer fluid F is the fuel flow Q, as shown in [Fig. 4], the heating turbomachine 1 is configured to directly heat the fuel flow Q circulating in the fuel circuit CQ, in the post-combustion heat exchanger 3, from calories taken from the exhaust air flow AE.

[0064] In one embodiment, shown in Figures 5 and 6, the heating turbomachine 1 comprises at least one exhaust heat exchanger 31, mounted on the fluid circuit CF, configured to heat the heat transfer fluid F from calories taken from the exhaust air flow AE at the outlet of the turbine 12.

[0065] Preferably, the exhaust heat exchanger 31 is mounted upstream of the post-combustion heat exchanger 3 in the fluid circuit CF. The heat transfer fluid F is thus heated progressively (less thermal gradient) by the exhaust heat exchanger 31 then the post-combustion heat exchanger 3, which makes it possible to distribute the heat flows. The service life of the heat exchangers 3, 31 is thus increased. It goes without saying that a reverse circulation, that is to say, from the post-combustion heat exchanger 3 to the exhaust heat exchanger 31 is also possible. Preferably, the circulation of the fluids in each of the heat exchangers 3, 31 can be organized in counter-current or cross-current mode which are effective. A co-current exchange is however possible.

[0066] In one embodiment, still with reference to [Fig. 5], the conditioning system SC comprises a heat exchanger 32 mounted on the fuel circuit CQ and on the fluid circuit CF. The heat exchanger 32 is configured to heat the fuel flow Q from calories taken from the heat transfer fluid F, the heat transfer fluid F being configured to have been previously heated in the post-combustion exchanger 3 from the calories taken from the exhaust air flow AE. Preferably, the heat exchanger 32 is positioned upstream of the distribution member 5 so as to supply a fuel flow Q1, Q2 in the gaseous state to the combustion chamber 2 of the heating turbomachine 1 and to the aircraft engine M. The heat transfer fluid F thus circulates in a closed loop. The heat exchanger 32 can be used without an exhaust heat exchanger 31.

[0067] Different types of technologies can be implemented in the heat exchangers 3, 31, 32, namely, tubular technology, plate technology, fin technology, etc.

[0068] In one embodiment, with reference to [Fig. 7], the heating turbomachine 1 comprises a heat exchanger 33, mounted between the aircraft engine M and the manifold 4. The heat exchanger 33 makes it possible to increase the temperature T2 of the second air flow A2 before it is mixed with the first air flow A1 with a heat source from the aircraft engine M, in particular, oil H. Such an architecture thus makes it possible to cool a hot engine oil flow H from the aircraft engine M by means of the second air flow A2.

[0069] In this example, the heating turbomachine 1 further comprises a distribution valve 7, for example a three-way valve allowing in a particular configuration to bypass the combustion chamber by ejecting the second air flow A2 towards the ambient. Such an embodiment makes it possible for example to cool the engine oil H even when the combustion chamber is deactivated.

[0070] An example of fuel conditioning will now be described, with reference to [Fig. 5] and [Fig. 6]. According to one embodiment of the invention, the fuel conditioning comprises a plurality of steps of a method of heating the heat transfer fluid F according to the invention to enable the fuel to be conditioned.

[0071] A flow of fuel Q in the liquid state is taken from the cryogenic tank RC and is conducted from upstream to downstream in the fuel circuit CQ. The flow of fuel Q is heated in the heat exchanger 32 by circulation of the heat transfer fluid F. The fuel flows Q1, Q2 are then conveyed to the aircraft engine M and to the heating turbomachine 1 via the distribution member 5.

[0072] With reference to [Fig. 6], the circulation of the heat transfer fluid F in the exhaust heat exchanger 31 and then in the post-combustion heat exchanger 3 is shown in detail so that it takes calories from the air flow. AE exhaust.

[0073] More precisely, in an exemplary implementation, the method comprises a first step E1 in which a first ambient air flow Al coming from an air inlet EA, initially at a first temperature T1 and at a first pressure PI, is compressed in the compressor 11 to reach a second temperature Tic, higher than the first temperature T1, and a second pressure Pic, higher than the first pressure PI.

[0074] At the outlet of the compressor 11, in a step E2, the first air flow A1 is mixed in the manifold 4 with a second air flow A2 from the aircraft engine M, the second air flow A2 being at the temperature T2 and the pressure P2. At the outlet of the manifold 4, the mixture of the first air flow A1 and the second air flow A2 forms the feed air flow A, then at the temperature Ta and the pressure Pa. Thanks to the temperature T2 and the pressure P2 of the second air flow A2, the feed air flow A has a temperature Ta and a pressure Pa that are sufficiently high to allow combustion of the fuel in the combustion chamber 2, without the first ambient air flow Al having to be compressed too much in the compressor 11.

[0075] The combustion chamber 2 of the heating turbomachine 1 is then supplied, in a step E3, by a fuel flow Q2 and the feed air flow A at the pressure Pa in order to emit at the outlet an exhaust air flow AE having a higher pressure Pe and a higher temperature Te.

[0076] The exhaust air flow AE then passes through the post-combustion heat exchanger 3, in a step E4, in order to transfer calories to it, which makes it possible to lower its temperature and its pressure and to heat the heat transfer fluid F.

[0077] The exhaust air flow AE is then expanded in the turbine 12, in a step E5, which drives the compressor 11 into rotation via the turbomachine shaft 13. During its expansion, the pressure and temperature of the exhaust air flow AE drop again.

[0078] The exhaust air flow AE then passes through the exhaust heat exchanger 31, in a step E6, in order to transfer calories to it to heat the heat transfer fluid F, the temperature and pressure of the exhaust air flow AE is again lowered.

[0079] The exhaust air flow AE is finally discharged into the ambient environment. In this example, the exhaust air flow AE passes through a nozzle Tu, making it possible to generate residual thrust.

[0080] A method is presented in which the heat transfer fluid F and the fuel flow Q are distinct. Alternatively, the fuel flow Q can be used as the heat transfer fluid F. In this embodiment, the exhaust air flow AE transfers calories to the post-combustion exchanger 3 to directly heat the fuel flow Q, the temperature and pressure of the exhaust air flow AE are lowered in a manner analogous to the implementation method presented previously.

Claims

Claims

1. Heating turbomachine (1) for a fuel conditioning system (SC) configured to supply an aircraft engine (M) from a fuel flow (Q) from a cryogenic tank (RC), the heating turbomachine (1) comprising: • a compressor (11) configured to be supplied by a first air flow (Al) from an air inlet (EA), • a turbine (12) connected to the compressor (11) by a turbomachine shaft (13), the turbine (12) being configured to drive the compressor (11), • a combustion chamber (2) configured to be supplied, on the one hand, by a feed air flow (A) and, on the other hand, by a fuel flow (Q2), the combustion chamber (2) being configured to evacuate an exhaust air flow (AE) loaded with calories and to rotate the turbine (12), • a fluid circuit (CF) in which circulates, from upstream to downstream,a heat transfer fluid (F) and • a post-combustion heat exchanger (3), mounted on the fluid circuit (CF), configured to take calories from the exhaust air flow (AE) between the combustion chamber (2) and the turbine (12), so as to heat the heat transfer fluid (F), • heating turbomachine (1) characterized in that it comprises at least one manifold (4), mounted between the compressor (11) and the combustion chamber (2), configured to mix the first air flow (Al) from the compressor (11) and a second air flow (A2) from the aircraft engine (M) in order to distribute the supply air flow (A) to the combustion chamber (2).,

2. Heating turbomachine (1) according to claim 1, in which the heat transfer fluid (F) is the fuel flow (Q) from the cryogenic tank (RC).

3. Heating turbomachine (1) according to one of claims 1 and 2, in which the first air flow (Al) present at the inlet of the compressor (11) a temperature (Tl) between -70°C and 45°C and a pressure (PI) between 0.015 and 0.1 MPa (0.15 and 1 bar).

4. Heating turbomachine (1) according to one of claims 1 to 3, in which the second air flow (A2) from the aircraft engine (M) has at the inlet of the manifold (4) a temperature (T2) of between 50 and 400°C and a pressure (P2) of between 0.1 and 0.8 MPa (between 1 and 8 bar).

5. Heating turbomachine (1) according to one of claims 1 to 4, in which the second air flow (A2) comes from a low pressure compressor of the aircraft engine (M).

6. Heating turbomachine (1) according to one of claims 1 to 5, in which the second air flow (A2) comes from an intermediate stage of a high pressure compressor of the aircraft engine (M).

7. Heating turbomachine (1) according to one of claims 1 to 6, in which the second air flow (A2) comes from the aircraft air supply circuit of the aircraft engine (M).

8. Heating turbomachine (1) according to one of claims 1 to 7, comprising a heat exchanger (33), mounted between the aircraft engine (M) and the manifold (4), configured to heat the second air flow (A2) with a heat source from the aircraft engine (M), before it is mixed with the first air flow (Al).

9. Fuel conditioning system (SC) configured to supply an aircraft engine (M) of an aircraft from fuel (Q) from a cryogenic tank (RC), the conditioning system (SC) comprising: • a fuel circuit (CQ) configured to connect the cryogenic tank (RC) to the aircraft engine (M) of an aircraft, a fuel flow (Q) circulating in the fuel circuit (CQ) and • a heating turbomachine (1) according to one of claims 1 to 8 for heating the fuel flow (Q) circulating in the fuel circuit (CQ) from calories taken from the exhaust air flow (AE), the combustion chamber (2) being supplied by a fuel flow (Q2) from the fuel circuit (CQ) and by a feed air flow (A) from the manifold (4).

10. A fuel conditioning system (SC) according to claim 9, in which the fuel flow (Q) is configured to be heated in the post-combustion exchanger (3) directly by the calories taken from the exhaust air flow (AE).

11. Fuel conditioning system (SC) according to claim 9, comprising a heat exchanger (32) mounted on the fuel circuit (CQ) configured to heat the fuel flow (Q) from calories taken from the heat transfer fluid (F), the heat transfer fluid (F) being configured to be heated in the post-combustion exchanger (3) from calories taken from the exhaust air flow (AE).

12. Method for heating a heat transfer fluid (F) by means of a heating turbomachine (2) according to one of claims 1 to 8, the heat transfer fluid (F) circulating from upstream to downstream in the fluid circuit (CF), the method comprising the steps of: • Supply the compressor (11) with a first air flow (Al), • Mix in the manifold (4) the first air flow (Al) from the compressor (11) and a second air flow (A2) from the aircraft engine (M), so as to form a supply air flow (A), • Supply the combustion chamber (2), on the one hand, with the feed air flow (A) and, on the other hand, with a fuel flow (Q2), the combustion chamber (2) discharging an exhaust air flow (AE) loaded with calories in order to drive the turbine (12) in rotation, and • Heat the heat transfer fluid (F) in the post-combustion heat exchanger (3) using calories taken from the exhaust air flow (AE) leaving the combustion chamber (2).

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