Fuel conditioning system for supplying an aircraft turbomachine and associated method

Abstract

A fuel conditioning system (SC) configured to supply a main combustion chamber (CC) of an aircraft propulsion turbomachine (M), the conditioning system (SC) comprising: a fuel circuit (1) connected inlet to a cryogenic tank (R) and outlet to the propulsion turbomachine (M), a fuel stream (Q) circulating in the fuel circuit (1); a separator device (4) configured to separate an incoming air stream (A1) into a main air stream (A2), comprising at least 30% dioxygen, and at least one oxygen-depleted auxiliary air stream (B); and a secondary combustion chamber (5) configured to receive inlet a feed fuel stream (Qa) and the main air stream (A2), so as to effect combustion of the feed fuel stream (Qa) and the main air stream (A2) to heat the fuel stream (Q) from the cryogenic tank (R). Figure from the summary: Figure 3

Description

Title of the invention: Fuel conditioning system for supplying an aircraft turbomachine and associated method. 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] 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, in particular through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and, as essential complements to technological progress, aviation biofuels.

[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 an aircraft propulsion system powered by fuel from 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 -258 to -251°C (15 to 22 Kelvins) in a cryogenic tank on the aircraft.

[0008] In order to be injected into the combustion chamber of a turbomachine, for example, 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 fuel injectors.

[0009] It is shown in [Fig. 1], an SCAA conditioning system which includes a fuel circuit 101 mounted between a cryogenic tank R and a propulsion turbomachine M. A fuel flow Q circulates in the fuel circuit 101 and passes successively through a mechanical pump 102 and a heat exchanger 103.

[0010] In practice, the fuel heating stage is energy-intensive and requires extracting heat from hot sources. Among the various technologies for heating liquid fuel, it is known to generate a heating airflow AC by means of a heating turbomachine 104 powered, on the one hand, by a diverted fuel flow Qd from the cryogenic tank R and, on the other hand, by an ambient air flow A. The heat exchanger 103 transfers heat from the heating airflow AC to the fuel flow Q circulating in the fuel circuit 101 in order to make it gaseous. It is thus suitable for use in a combustion chamber CC of the aircraft engine M.

[0011] However, it is complex for the heating turbomachine 104 to generate a significant amount of heat with good efficiency. Indeed, by its very nature, a turbomachine is optimized to maximize its mechanical energy and limit its heat losses. In fact, the thermal energy generated is proportional to the mechanical energy generated. Therefore, to generate enough thermal energy to heat the fuel flow Q, it is necessary to use a large heating turbomachine 104, which generates a large amount of unwanted mechanical energy. The turbomachine is then heavy and bulky, which represents a significant drawback, particularly in the aeronautical field, which aims to limit aircraft mass to reduce their greenhouse gas emissions and their environmental impact.

[0012] To overcome this drawback, a prior art architecture is known, as shown in [Fig. 2], in which the airflow A entering the chamber of The combustion 142 of the heating turbomachine 104 corresponds to a mixture between a first compressed air flow AA from the ambient air and a second air flow AB from a high-pressure compressor CHP of the propulsion turbomachine M. The mixture of the air flows AA, AB thus has a higher pressure and temperature, which allows the compressor 141 of the heating turbomachine 104 to be sized less, allowing a lighter and less bulky heating turbomachine.

[0013] However, the heating turbomachine 104 requires a high-volume airflow. Therefore, even if the efficiency of the heating turbomachine is improved, the compressor 141 used to compress the ambient airflow AA remains oversized. In other words, it retains a significant mass and size, which can negatively impact the aircraft's energy and environmental performance.

[0014] The invention thus aims to eliminate at least some of these drawbacks, by proposing a new conditioning system which allows both optimal heating and compression of the fuel flow from a cryogenic tank while having limited mass and size. PRESENTATION OF THE INVENTION

[0015] The invention relates to a fuel conditioning system configured to supply a main combustion chamber of an aircraft propulsion turbomachine with fuel from a cryogenic tank, the conditioning system comprising at least: • a fuel circuit connected at the inlet to the cryogenic tank and at the outlet to the propulsion turbomachine, with a fuel flow circulating from upstream to downstream in the fuel circuit, • a separator device configured to divide an incoming airflow into a main airflow, comprising at least 30% oxygen, and at least one auxiliary airflow depleted of oxygen, and • a secondary combustion chamber configured to receive on the one hand a feed fuel stream, and, on the other hand, the main air stream from the separator device, so as to achieve combustion of the feed fuel stream and the main air stream to heat the fuel stream from the cryogenic tank to at least a vaporization temperature to supply the propulsion turbomachine.

[0016] The secondary combustion chamber provides heat to warm the fuel stream and change it from a liquid to a gaseous state. It is not advantageously, there is no need to use an external heat source that could freeze on contact with cryogenic fuel or a heavy and bulky heat transfer fluid loop.

[0017] The separator device allows the secondary combustion chamber to be supplied with a main airflow enriched in oxygen compared to the incoming airflow (and therefore depleted in non-reactive components). Furthermore, the main airflow has a significantly higher oxidant density. Since its reactive density is higher, its total mass flow rate can advantageously be limited while still ensuring the same performance. The dimensions of the secondary combustion chamber are advantageously controlled, which limits its mass and size. Fuel flow heating is thus achieved at a lower cost.

[0018] According to a preferred aspect, the secondary combustion chamber is supplied with oxidant solely by the main airflow. This allows for a low mass flow rate of the oxidant flow (i.e., the main airflow) and therefore an advantageously lightweight and compact secondary combustion chamber.

[0019] In a preferred embodiment, the separator device is configured to separate the incoming airflow into the main airflow and the oxygen-depleted auxiliary airflow, using cooling transferred by the fuel flow from the cryogenic tank. In other words, the very low cryogenic temperatures are advantageously used to separate the oxygen from the other components in a simple manner, avoiding the use of elements external to the system and, for example, the installation of additional piping.

[0020] Preferably, the separator device is mounted directly on the fuel circuit, making it possible to avoid a fuel flow bypass to separate the incoming airflow into the main airflow and the oxygen-depleted auxiliary airflow, from the cold transferred by the fuel flow from the cryogenic tank.

[0021] According to one aspect, the separator device is a distillation column, allowing the use of a known and simple-to-use device. Advantageously, the environment is taken advantage of to obtain, by distillation, an air stream rich in dioxygen and depleted of non-reactive components.

[0022] Alternatively, the separator device is chosen from: a device suitable for ensuring filtration by particle size, by distillation, by means of a separation column, etc.

[0023] In one embodiment, with the secondary combustion chamber configured to generate an exhaust gas flow, the conditioning system includes a heat exchanger mounted on the fuel circuit downstream of the separator device and configured to heat the fuel flow initially in a liquid state. and convert it into a gaseous state, using calories transferred by the exhaust gas flow. In other words, the amount of heat corresponds to the exhaust gas flow from the secondary combustion chamber, which allows the reuse of products from fuel combustion in the secondary combustion chamber and enables a compact conditioning system.

[0024] According to one aspect, the secondary combustion chamber is configured to be supplied by a gaseous fuel flow from the heat exchanger, which allows the assembly of a simple secondary combustion chamber supplied only by gaseous flows.

[0025] In one embodiment, the secondary combustion chamber is in the form of a staged combustion chamber. The fuel flow can thus be heated directly in the secondary combustion chamber by the heat generated during combustion of the fuel fraction. This eliminates the need for a heat exchanger.

[0026] According to one aspect, the secondary combustion chamber is mounted directly on the fuel circuit. Thus, all the fuel circulating in the fuel circuit can pass through the staged secondary combustion chamber and only a fraction reacts with the main airflow to generate heat and warm the rest of the fuel.

[0027] According to one aspect, the conditioning system includes at least one heat exchanger mounted on the fuel circuit between the separator device and the secondary combustion chamber and configured to preheat the feed fuel stream before its admission into the secondary combustion chamber, allowing the fuel stream to be preheated to make it pass into the gaseous state and limiting the sizing of the secondary combustion chamber.

[0028] Preferably, the feed fuel stream is heated by a fuel stream from the secondary combustion chamber and heated by the partial combustion of the feed fuel stream, which allows the optimization of the use of heat sources already present in the conditioning system and limits its size.

[0029] According to one aspect, the secondary combustion chamber is configured to provide an amount of heat suitable for warming the fuel flow to an injection temperature to directly supply the main combustion chamber of the propulsion turbomachine, thus eliminating the need to mount additional components on the fuel circuit.

[0030] In one embodiment, the conditioning system being defined in an aircraft frame of reference and a turbomachine frame of reference, the cryogenic tank extending in the aircraft frame of reference and the propulsion turbomachine extending In the turbomachine frame of reference, at least the separator device and the auxiliary combustion chamber are mounted in the aircraft frame of reference. Thus, the fuel flow is vaporized directly at the outlet of the cryogenic tank, which limits the use of heavy and bulky cryogenic piping, particularly between the two frames of reference.

[0031] The invention also relates to an aircraft comprising a cryogenic tank, a propulsion system comprising at least one propulsion turbomachine, and a conditioning system as described above, the propulsion turbomachine being supplied by a fuel flow circulating in the fuel circuit of the conditioning system.

[0032] Finally, the invention relates to a method of supplying fuel to an aircraft propulsion system from fuel from a cryogenic tank by means of a conditioning system as described above, a fuel flow circulating from upstream to downstream in the fuel circuit connecting the cryogenic tank at the inlet and the propulsion turbomachine at the outlet, the method comprising steps consisting of: • separating, in the separator device, the incoming airflow into a main airflow, comprising at least 30% oxygen, and at least one auxiliary airflow depleted of oxygen, and • to achieve combustion of the feed fuel stream and the main air stream in the secondary combustion chamber so as to heat the fuel stream from the cryogenic tank to at least a vaporization temperature to power the propulsion turbomachine. PRESENTATION OF THE FIGURES

[0033] 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.

[0034] Fig. 1 is a schematic representation of a first conditioning system according to the prior art.

[0035] Fig. 2 is a schematic representation of a second conditioning system according to the prior art.

[0036] Fig. 3 is a schematic representation of a conditioning system according to a first embodiment of the invention.

[0037] Fig. 4 is a schematic representation of a conditioning system according to a second embodiment of the invention.

[0038] Fig. 5 is a schematic representation of a conditioning system according to a third embodiment of the invention.

[0039] 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

[0040] With reference to [Fig. 3], a fuel conditioning system SC is shown, configured to supply an aircraft propulsion turbomachine M with fuel Q from a cryogenic tank R. The propulsion turbomachine M is configured to provide propulsion for the aircraft, in particular by driving at least one propulsion unit (not shown in [Fig. 3]). In this example, the fuel Q is dihydrogen, but the invention applies to other types of fuel, in particular fuels such as methane or liquefied natural gas.

[0041] In practice, the fuel Q is stored in the cryogenic tank R at cryogenic temperatures. For example, the fuel Q is stored in the cryogenic tank R at a temperature of approximately -253 to -251°C (20 to 22 Kelvin). At this temperature, the fuel Q is liquid. In order to be introduced into a combustion chamber of the turbomachine M (hereinafter referred to as the main combustion chamber CC), the fuel Q must be heated.

[0042] For this purpose, the aircraft includes an SC conditioning system configured to heat and pressurize the Q fuel.

[0043] The conditioning system SC is defined, in this example, in an aircraft reference frame REF-A and a turbomachine reference frame REF-M distant from each other, the cryogenic tank R being positioned in the aircraft reference frame REF-A and the propulsion turbomachine M being positioned in the turbomachine reference frame REF-M.

[0044] With reference to [Fig.3], the conditioning system SC includes a fuel circuit 1 connected inlet to the cryogenic tank R and outlet to the main combustion chamber CC of the propulsion turbomachine M. The conditioning system SC also includes a mechanical pump 2, preferably high pressure, configured to circulate a fuel flow Q from upstream to downstream in the fuel circuit 1.

[0045] According to one aspect of the invention, the conditioning system SC comprises a separator device 4 and a secondary combustion chamber 5 for generating a quantity of heat configured to warm the fuel stream Q. The device separator 4 is configured to generate an oxidant flow intended to be injected into the secondary combustion chamber 5.

[0046] With further reference to [Fig. 3], the separator device 4 is configured to receive an incoming airflow Al and to divide this incoming airflow Al into a main airflow A2 and at least one oxygen-depleted auxiliary airflow B. The incoming airflow Al is preferably atmospheric air. As such, the incoming airflow Al is an outside airflow and contains oxygen, nitrogen, water particles, etc.

[0047] In practice, the main airflow A2 is intended to supply the secondary combustion chamber 5. According to one feature of the invention, this airflow, at the outlet of the separator device 4, consists of at least 30% oxygen. The main airflow A2 is thus also depleted in nitrogen compared to the incoming airflow AL. Therefore, although the main airflow A2 is referred to as the "airflow," the latter, enriched in oxygen compared to the incoming airflow Al and depleted in nitrogen, is, in practice, in the form of a gaseous flow. The main airflow A2 can be in a liquid or gaseous state.

[0048] The oxygen-depleted auxiliary airflow B comprises a plurality of components generally present in ambient air, such as nitrogen, water, oxygen, carbon dioxide, and rare gases.

[0049] In other words, the separator device 4 is configured to separate oxygen from the other components present in the ambient air and to supply the secondary combustion chamber 5 with a flow enriched in oxygen relative to the air. The separator device 4 increases the oxidant density in the airflow to supply the auxiliary combustion chamber 5. In this example, the oxygen-depleted auxiliary airflow(s) B is / are vented into the atmosphere. It is understood that the oxygen-depleted auxiliary airflow B, and therefore richer in nitrogen than the incoming airflow A1, could alternatively be reused in the aircraft, for example, to perform inerting functions.

[0050] Preferably, the separator device 4 is mounted on the fuel circuit 1 and is configured to separate the main airflow A2 from the oxygen-depleted auxiliary airflow B by means of cooling transferred by the liquid fuel flow Q from the cryogenic reservoir R. In other words, the cryogenic fuel flow Q is advantageously used in the separator device 4 to separate the oxygen from the other components contained in the incoming airflow AL

[0051] In a preferred embodiment, the separator device 4 is a distillation column that allows liquid fluids to be separated into individual components according to their liquefaction point. A distillation column can be conveniently integrated by taking advantage of the cooling capacity available in the flow of Fuel Q. Such a distillation column is known to those skilled in the art, and its operation will not be described in further detail in this document. It goes without saying that the separator device 4 could alternatively take a different form, for example, any known device suitable for particle size filtration, distillation, separation column, etc.

[0052] According to one aspect of the invention, as shown in [Fig. 3], the conditioning system SC comprises a compressor 6 configured to compress the main airflow A2 exiting the separator device 4 before it is injected into the secondary combustion chamber 5. The compressor 6 is configured to increase the pressure of the main airflow A2 to at least a combustion pressure Pc corresponding to a pressure at which the main airflow A2 can be consumed in the secondary combustion chamber 5, as will be described in more detail later. In this example, the combustion pressure Pc is between 0.1 and 10 MPa.

[0053] Thanks to the separator device 4, the main airflow A2 contains a significant proportion of oxygen. Its oxidant density during combustion is thus increased, which makes it possible to significantly reduce its flow rate for the same service provided. Given the limited flow rate, the compressor 6 can advantageously have limited performance, and therefore limited mass and size, while still ensuring optimal compression to supply the secondary combustion chamber 5.

[0054] In this example, the SC conditioning system includes a first control valve 91 mounted between the separator device 4 and the compressor 6, so as to control the circulation and flow rate of the main airflow A2.

[0055] With further reference to [Fig. 3], the secondary combustion chamber 5 is configured to receive, on the one hand, the main air flow A2 from the separator device 4, which has been compressed in the compressor 6, and, on the other hand, a feed fuel flow Qa. The feed fuel flow Qa preferably corresponds to a fraction of the fuel flow Q circulating in the fuel circuit 1.

[0056] Preferably, the secondary combustion chamber 5 is supplied with air (in other words, oxidant) only with the main air flow A2 from the separator device 4. In other words, the secondary combustion chamber 5 is preferably supplied with air only with a gaseous flow enriched in dioxygen and depleted in nitrogen, which allows efficient combustion of the fuel Q while limiting the size of the secondary combustion chamber 5.

[0057] The secondary combustion chamber 5 is configured to carry out combustion of the feed fuel stream Qa and to generate a quantity of heat configured to heat the fuel stream Q from the cryogenic tank R to at least a vaporization temperature Tv, so as to allow its delivery to the propulsion turbomachine M without requiring the use of cryogenic piping (i.e., piping generally incorporating a vacuum insulation system), particularly between the aircraft reference frame REF-A and the turbomachine reference frame REF-M, for example, in the aircraft wings. Preferably, the secondary combustion chamber 5 is configured to provide sufficient heat to raise the fuel stream Q to an injection temperature to supply the main combustion chamber CC of the propulsion turbomachine M. Since the feed fuel stream Qa corresponds to a fraction of the fuel stream Q circulating in the fuel circuit 1, this is referred to as partial combustion of the fuel stream Q.

[0058] In a first embodiment, with reference to [Fig.3], the secondary combustion chamber 5 is configured to generate an exhaust gas flow AE.

[0059] In this embodiment, the conditioning system SC preferably includes a heat exchanger 3 mounted on the fuel circuit 1 downstream of the separator device 4. The heat exchanger 3 is configured to heat the fuel flow Q circulating in the fuel circuit 1 initially in a liquid state and change it into a gaseous state.

[0060] Preferably, the fuel flow Q is heated from calories transferred by the exhaust gas flow AE. It is understood that the fuel flow Q could alternatively be heated by an external heat source.

[0061] At the outlet of the heat exchanger 3, the exhaust gas flow AE is, in this example, vented into the atmosphere. It goes without saying that the exhaust gas flow AE could alternatively be reused in the aircraft, for example at the propulsion turbomachine M.

[0062] In practice, in this embodiment, the secondary combustion chamber 5 is mounted in parallel with the fuel circuit 1.

[0063] According to a preferred aspect, in this embodiment, the secondary combustion chamber 5 is configured to be supplied by a feed fuel flow Qa exiting the heat exchanger 3. A fraction of the fuel flow Q is thus diverted to supply the secondary combustion chamber 5 with a gaseous fuel flow Qg.

[0064] In this example, the SC conditioning system includes a second control valve 92 mounted between the heat exchanger 3 and the secondary combustion chamber 5, so as to control the circulation and flow rate of the supply fuel flow Qa.

[0065] A single heat exchanger 3 is described; however, it is understood that the conditioning system SC could alternatively comprise several. Similarly, it is understood that the heat exchanger 3 could be mounted in parallel with the fuel circuit 1 to heat only the supply fuel flow Qa.

[0066] In a second embodiment, with reference to [Fig. 4], the secondary combustion chamber 5 is in the form of a staged combustion chamber. As is known, a staged combustion chamber comprises a plurality of successive combustion stages, so as to successively burn a predetermined quantity of fuel Q. The heat generated during this partial combustion allows the remaining fuel flow Q to be vaporized and heated. Such a staged combustion chamber is known to those skilled in the art, and its operation will not be described in further detail in this document.

[0067] In this embodiment, the secondary combustion chamber 5 is mounted directly on the fuel circuit 1. The feed fuel flow Qa corresponds directly to the fuel flow Q from the cryogenic tank R. In other words, the entire fuel flow Q from the cryogenic tank R is configured to pass through the secondary combustion chamber 5, and only a fraction of the fuel flow Q is burned. In practice, the fraction of fuel Q burned, corresponding to the feed fuel flow Qa, is a function of the main air flow A2 supplied to the secondary combustion chamber 5.

[0068] In a third embodiment, with reference to [Fig.5], the secondary combustion chamber 5 is presented, in a manner analogous to the second embodiment, in the form of a staged combustion chamber configured to generate gases from the partial combustion between the feed fuel flow Qa and the main air flow A2. These gases mainly comprise fuel Q in liquid form, water, etc.

[0069] As shown in [Fig.5], the SC conditioning system also includes a heat exchanger 7 mounted on the fuel circuit 1 between the separator device 4 and the secondary combustion chamber 5.

[0070] The heat exchanger 7 is configured to preheat the fuel stream Q before it enters the secondary combustion chamber 5. In practice, in this example, the fuel stream Q is heated by the fuel stream Q exiting the secondary combustion chamber 5, which has been heated by the partial combustion of the feed fuel stream Qa. The heated fuel stream Q exiting the heat exchanger 3 is then configured to be routed to the propulsion turbomachine M.

[0071] It goes without saying that the fuel flow Q could be heated in the heat exchanger 7 from calories transferred by the exhaust gas flow AE generated by the secondary combustion chamber 5, in a manner analogous to the heat exchanger 3 described previously. In practice, the SC conditioning system could comprise a single heat exchanger 3, 7, or several heat exchangers.

[0072] A preheating allows the supply fuel stream Qa to be vaporized before it enters the secondary combustion chamber 5, which allows the use of a simpler staged combustion chamber configured to receive a gaseous air stream and a gaseous fuel stream.

[0073] In a preferred embodiment, as shown in the figures, the separator device 4, the compressor 6, and the secondary combustion chamber 5 are mounted in the aircraft frame of reference REF-A so as to condition the fuel flow Q directly at the outlet of the cryogenic tank R. This limits the use of cryogenic piping, thereby reducing both the size of the conditioning system SC and the mass of the aircraft. It is understood that the separator device 4, the compressor 6, and / or the secondary combustion chamber 5 could alternatively be mounted in the turbomachine frame of reference REF-M.

[0074] According to one aspect, the secondary combustion chamber 5 is connected to a turbine (not shown), which is itself connected to the compressor 6 or to a gas generator, for example. The turbine is configured to receive the exhaust gas flow AE and generate mechanical energy, so as to drive the compressor 6 or the gas generator, for example.

[0075] A method for supplying fuel to an aircraft propulsion turbomachine M, using fuel Q from a cryogenic tank R, will now be described. The supply method is carried out using the conditioning system SC as described previously. A flow of fuel Q circulates from upstream to downstream in the fuel circuit 1 which connects the cryogenic tank R to the propulsion turbomachine M. In this example, the fuel Q is hydrogen stored in liquid form in the cryogenic tank R.

[0076] The process will be described subsequently for the first embodiment of the SC conditioning system, shown in [Fig. 3]. In other words, in this example, the SC conditioning system comprises a heat exchanger 3 mounted on the fuel circuit 1 to heat the fuel flow Q, and the secondary combustion chamber 5 is mounted in parallel with the fuel circuit 1. It is understood that the process operates in a substantially analogous manner for the other embodiments.

[0077] In a first step El, an incoming airflow Al is introduced into the separator device 4 in which it is separated into a main airflow A2, comprising at less than 30% oxygen, and at least one oxygen-depleted auxiliary air stream B, including nitrogen, water, etc. In this example, the separator device 4 is in the form of a distillation chamber and is mounted on the fuel circuit 1. The separation of the components of the incoming air stream Al is achieved by means of the cooling transferred by the cryogenic fuel stream Q.

[0078] At the outlet of the separator device 4, the main airflow A2, enriched in dioxygen compared to the incoming airflow Al, passes through the compressor 6 to reach a combustion pressure Pc and is routed, in a second stage E2, to the secondary combustion chamber 5.

[0079] In parallel, the fuel flow Q circulating in the fuel circuit 1 is routed from the separator device 4 to the heat exchanger 3 where it is heated. A supply fuel flow Qa in a gaseous state from the outlet of the heat exchanger 3 is then routed to the secondary combustion chamber 5.

[0080] In a third stage E3, the feed fuel stream Qa mixed with the main air stream A2, enriched in dioxygen relative to the incoming air stream Al, is consumed in the secondary combustion chamber 5. An exhaust gas stream AE is generated at the outlet of the secondary combustion chamber 5 and is routed to the heat exchanger 3 to heat the fuel stream Q circulating in the fuel circuit 1. At the outlet of the heat exchanger 3, a fraction of the fuel stream Q forms the feed fuel stream Qa destined for the secondary combustion chamber 5. The remainder of the heated fuel stream Q, in a gaseous state, is routed to the main combustion chamber CC of the propulsion turbomachine M.

[0081] Thanks to the conditioning system according to the invention, the fuel flow, initially in a liquid state, is heated simply and at a lower cost by limiting the mass and size of each component, which is particularly advantageous in the aeronautical field, which aims to limit the mass of aircraft to limit their fuel consumption and thus limit greenhouse gas emissions. tight.

Claims

Demands

1. Fuel conditioning system (SC) configured to supply a main combustion chamber (CC) of an aircraft propulsion turbomachine (M) with fuel (Q) from a cryogenic tank (R), the conditioning system (SC) comprising at least: • a fuel circuit (1) connected inlet to the cryogenic tank (R) and outlet to the propulsion turbomachine (M), a fuel flow (Q) circulating upstream to downstream in the fuel circuit (1), • a separator device (4) configured to divide an incoming air stream (A1) into a main air stream (A2), comprising at least 30% dioxygen, and at least one oxygen-depleted auxiliary air stream (B), and • a secondary combustion chamber (5) configured to receive inlet, on the one hand, a feed fuel stream (Qa), and, on the other hand, the main air stream (A2) from the separator device (4),in order to achieve combustion of the feed fuel stream (Qa) and the main air stream (A2) to heat the fuel stream (Q) from the cryogenic tank (R) to at least a vaporization temperature (Tv) to power the propulsion turbomachine (M).

2. Conditioning system (SC) according to claim 1, wherein the separator device (4) is configured to separate the incoming airflow (A1) into the main airflow (A2) and the oxygen-depleted auxiliary airflow (B), from cooling transferred by the fuel flow (Q) from the cryogenic tank (R).

3. Conditioning system (CS) according to any one of claims 1 to 2, wherein the separator device (4) is a distillation column.

4. Conditioning system (SC) according to any one of claims 1 to 3, wherein the secondary combustion chamber (5) being configured to generate an exhaust gas flow (AE), the conditioning system (SC) comprises a heat exchanger (3) mounted on the fuel circuit (1) downstream of the separator device (4) and configured to heat the fuel flow (Q) initially in a liquid state and changing it into a gaseous state, from calories transferred by the exhaust gas flow (AE).

5. Conditioning system (SC) according to claim 4, wherein the secondary combustion chamber (5) is configured to be supplied by a gaseous fuel flow (Qg) from the heat exchanger (3).

6. Conditioning system (CS) according to any one of claims 1 to 5, wherein the secondary combustion chamber (5) is in the form of a staged combustion chamber.

7. Conditioning system (CS) according to claim 6, wherein the secondary combustion chamber (5) is mounted directly on the fuel circuit (1).

8. Conditioning system (SC) according to any one of claims 1 to 7, the conditioning system (SC) comprising at least one heat exchanger (7) mounted on the fuel circuit (1) between the separator device (4) and the secondary combustion chamber (5) and configured to preheat the feed fuel stream (Qa) before its admission into the secondary combustion chamber (5).

9. Conditioning system (CS) according to any one of claims 1 to 8, the conditioning system (CS) being defined in an aircraft frame of reference (REF-A) and a turbomachine frame of reference (REF-M), the cryogenic tank (R) extending in the aircraft frame of reference (REF-A) and the propulsion turbomachine (M) extending in the turbomachine frame of reference (REF-M), at least the separator device (4) and the auxiliary combustion chamber (5) are mounted in the aircraft frame of reference (REF-A).

10. Aircraft comprising a cryogenic tank (R), a propulsion system comprising at least one propulsion turbomachine (M), and a conditioning system (SC) according to any one of claims 1 to 9, the propulsion turbomachine (M) being supplied by a fuel flow (Q) circulating in the fuel circuit (1) of the conditioning system (SC).

11. A method for supplying fuel to an aircraft propulsion system from fuel (Q) from a cryogenic tank (R) by means of a conditioning system (SC) according to any one of claims 1 to 9, a fuel flow (Q) circulating from upstream to downstream in the fuel circuit (1) connecting at the inlet to the tank cryogenic (R) and at the output the propulsion turbomachine (M), the process comprising steps consisting of: • separate, in the separator device (4), the incoming airflow (A1) into a main airflow (A2), comprising at least 30% dioxygen, and at least one auxiliary airflow depleted in dioxygen (B), and • to carry out combustion of the feed fuel stream (Qa) and the main air stream (A2) in the secondary combustion chamber (5) so as to heat the fuel stream (Q) from the cryogenic tank (R) to at least a vaporization temperature (Tv) to supply the propulsion turbomachine (M).

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