Fuel conditioning system for an aircraft turbomachine and method of use

Abstract

A fuel (H) conditioning system (100) for an aircraft turbomachine, comprising a first injection line (1) configured to supply a first flow (Q1) of fuel (H) to the combustion chamber (7), the first injection line (1) comprising a first pressurization device (10) configured to deliver the fuel (H) according to said first flow (Q1), said conditioning system (100) comprising a second injection line (2) configured to supply a second flow (Q2) of fuel (H) to the combustion chamber (7), the second injection line (2) comprising a second pressurization system (20) supplying at a supply pressure (P20) a sonic neck metering device (21), the sonic neck metering device (22) being configured, from a metering setpoint (C2) and the supply pressure (P20), to deliver the fuel (H) according to the second flow (Q2). Figure from the summary: Figure 1

Description

Title of the invention: Fuel conditioning system for an aircraft turbomachine and method of use Technical field

[0001] The present invention relates to the field of aircraft turbomachines, and aims at a fuel conditioning system, in particular hydrogen, for supplying a combustion chamber of 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] In this context, the Applicant's research and development work relates to new generations of aircraft engines, the lightening 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 aeronautical biofuels.

[0004] One of the Applicant's challenges is the development of an aircraft turbomachine running on hydrogen. In a known manner, such an aircraft turbomachine comprises a combustion chamber configured to be supplied with gaseous hydrogen which is burned to participate in the propulsion of the aircraft.

[0005] In practice, to limit the space required on board the aircraft, it is necessary to store the hydrogen in liquid form at a temperature of approximately -253°C and a pressure of less than 3 bars in cryogenic tanks. To enable its injection in gaseous form at a temperature between -40°C and 0°C and at a pressure between 30 bars and 80 bars into the combustion chamber, the aircraft turbomachine conventionally comprises a fuel conditioning system configured to pressurize, heat and meter the hydrogen. In a known manner, the conditioning system comprises a fuel circuit with one or more pumps to enable pressurization, as well as one or more heat exchangers, connected to one or more heat sources of the aircraft, to ensure heating.

[0006] To ensure metering in the fuel circuit, it is known to use a variable section metering valve. However, in the case of a fuel in the form of a Compressible fluid such as hydrogen gas, valve control is complex because the outlet mass flow rate depends on many physical parameters, including pressure, temperature, and diameter. One solution would be to control the valve using data from many sensors. However, such a solution is not satisfactory because it increases the footprint and would make control complex, expensive, and imprecise.

[0007] The invention thus aims to improve the precision of the dosage in a fuel conditioning system, in particular hydrogen, while limiting the related energy expenditure. PRESENTATION OF THE INVENTION

[0008] The invention relates to a fuel conditioning system for an aircraft turbomachine, said conditioning system comprising a fuel circuit configured to be fluidically connected upstream to at least one fuel tank and to supply downstream a combustion chamber of the aircraft turbomachine, said fuel circuit comprising: • At least one first injection line configured to supply a first flow rate of fuel to the combustion chamber, the first injection line comprising at least one first pressurization device configured to deliver the fuel according to said first flow rate.

[0009] The invention is remarkable in that the fuel circuit comprises at least one second injection line configured to supply a second fuel flow rate to the combustion chamber, the second injection line comprising at least one second pressurization system supplying a sonic throat metering device at a supply pressure, the sonic throat metering device being configured, from a metering setpoint and the supply pressure, to deliver the fuel at the second flow rate.

[0010] The conditioning system according to the invention advantageously makes it possible to supply a flow rate to the combustion chamber in a precise and rapid manner while limiting the energy consumption of pressurizing the fuel thanks to two injection lines of different operation. The first line, free of metering members, makes it possible to supply a high fuel flow rate for a low energy cost. The second line makes it possible to supply a precise flow rate with a short response time. Controlling the two injection lines according to the speed of the aircraft turbomachine advantageously makes it possible to combine the advantages of each line.

[0011] In one aspect, the pressurization system is configured to supply a predetermined priming pressure to the sonic neck metering device, at which the flow in the sonic neck metering device is sonic. advantageously, at the priming pressure, the flow rate provided by the metering device depends solely on the upstream pressure and temperature conditions of the fuel. The metering device thus advantageously makes it possible to provide a precise flow rate, particularly for a compressible fluid such as hydrogen in the gaseous state.

[0012] According to one aspect, the sonic neck dosing device comprises a conduit of fixed section. Such a dosing device is advantageously simple to control.

[0013] According to one aspect, the first line is free of a metering device. In other words, the first flow rate is determined solely by the first pressurizing device, which is not suitable for metering the fuel. The first line advantageously makes it possible to provide a first flow rate at low energy cost in comparison with the second line. Indeed, the sonic neck metering device of the second line requires a high upstream pressure to operate. The second line, precise and fast, makes it possible to compensate for the response time and the lack of precision of the first pressurizing device.

[0014] According to one aspect, the second line comprises an electric heater mounted upstream of the sonic neck metering device. Such an electric heater makes it possible to heat the fuel with a low response time. The second line thus has a low response time for both heating and metering the fuel.

[0015] According to one aspect, the sonic neck metering device comprises a sonic neck conduit and at least one controllable pressure reducer mounted upstream of the sonic neck conduit. The pressure reducer advantageously makes it possible to control the pressure and temperature upstream of the sonic neck, to deliver a precise flow rate. In addition, the pressure reducer has a low response time.

[0016] According to a preferred aspect, the metering pressure P2 supplied by the pressurization system verifies: P2 > 1.5 * PI, with PI the fuel pressure in the first line, and preferably: P2 > 2 * PL Such a metering pressure is necessary to ensure the operation of the metering device.

[0017] According to a preferred aspect, the pressurizing device is in the form of at least one cryogenic pump, adapted to pressurize a cryogenic fluid, such as hydrogen.

[0018] The invention also relates to an assembly of a conditioning system and an aircraft turbomachine computer, said computer being configured to provide at least one metering instruction to the sonic neck metering device, so as to control the second fuel flow rate. Preferably, the computer is configured to provide at least one pressure instruction to the second pressurizing device, so as to control the second fuel flow rate.

[0019] Preferably, the computer is configured to provide at least one speed instruction to the first pressurizing device, so as to control the first fuel flow. The computer advantageously allows the distribution of the first flow and the second flow to be controlled in the total flow supplied to the combustion chamber, so as to take advantage of the benefits of each injection line depending on the speed of the aircraft turbomachine.

[0020] The invention also relates to an aircraft turbomachine comprising a combustion chamber, at least one fuel tank and at least one conditioning system as described above, the fuel circuit of which is fluidically connected upstream to the fuel tank and downstream to the combustion chamber.

[0021] The invention also relates to a method of using the conditioning system as described above, in which, during a stabilized regime of the aircraft turbomachine: • The sonic neck dosing device receives at least one predetermined dosing instruction, • The sonic neck dosing device provides, from the dosing setpoint and the supply pressure, a second flow to supply the combustion chamber.

[0022] According to one aspect, during an ignition phase during the stabilized regime of the aircraft turbomachine, the second flow rate is equal to at least 80% of the total flow rate provided by the conditioning system, preferably at least 95%. The second line makes it possible to provide a precise and rapid flow rate for the ignition phase of the aircraft turbomachine.

[0023] According to one aspect, during a cruise phase during the stabilized regime of the aircraft turbomachine: • The first line provides an initial flow to feed the combustion chamber, • The first flow rate being higher than the second flow rate provided by the second line, preferably at least twice as high.

[0024] The share of the first flow rate in the total flow rate supplied to the combustion chamber is advantageously maximized to reduce the energy consumption linked to the metering device of the second line.

[0025] According to one aspect, during a transient regime of the aircraft turbomachine: • The first line provides a first flow rate to feed the combustion chamber, the first flow rate varying during a response time until reaching a predetermined target flow rate, • The sonic neck dosing device receives at least one predetermined dosing instruction, • The sonic neck dosing device provides, from the dosing setpoint and the supply pressure, a second flow rate to supply the chamber of combustion, • The second flow rate being at least equal to the difference between the first flow rate and the predetermined target flow rate during the response time.

[0026] During a transient regime, the second line advantageously makes it possible to compensate for the response time of the first line, by providing a second flow rate precisely and quickly, taking into account the lack of flow rate of the first line. PRESENTATION OF FIGURES

[0027] 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.

[0028] [Fig.l] is a schematic representation of the fuel conditioning system for an aircraft turbomachine according to a first embodiment of the invention.

[0029] [Fig.2] is a schematic representation of the method of using the system of packaging of [Fig.l] according to an embodiment of the invention.

[0030] [Fig.3] is a schematic representation of the method of using [Fig.2] during a stabilized aircraft regime.

[0031] [Fig.4] is a schematic representation of the method of using [Fig.3] during a transient aircraft regime.

[0032] [Fig.5] is a schematic representation of the conditioning system of fuel for aircraft turbomachine according to a second embodiment of the invention.

[0033] [Fig.6] is a schematic representation of the conditioning system of fuel for aircraft turbomachine according to a third embodiment of the invention.

[0034] 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

[0035] The invention relates to a fuel conditioning system for an aircraft turbomachine and a method of using the same.

[0036] As illustrated in [Fig.l], an aircraft turbomachine comprises a combustion chamber 7 configured to be supplied with fuel H which is burned to participate in the propulsion of the aircraft. The fuel H is stored in tanks 4, 4L. The aircraft turbomachine comprises a conditioning system 100 configured to pressurize, heat and meter the fuel H from the tanks 4, 41, in order to allow their injection into the combustion chamber 7.

[0037] The invention is of particular interest for a fuel H configured to be stored in cryogenic form in the tanks 4, 41, such as hydrogen. As described previously, to limit the space required on board the aircraft, it is necessary to store the hydrogen in liquid form at a temperature of approximately -253°C and at a pressure of less than 7 bars, preferably less than 3 bars, in the cryogenic tanks 4, 41. The hydrogen must then be injected into the combustion chamber 7 in the gaseous state at a temperature between -40°C and 0°C and at a pressure between 30 bars and 80 bars to enable its combustion. Subsequently, the upstream and downstream terms are determined with respect to the direction of circulation of the fuel.

[0038] According to the invention and with reference to [Fig.l], the conditioning system 100 comprises a fuel circuit 3 configured to be fluidically connected upstream to at least one fuel tank 4 and to supply the combustion chamber 7 downstream. Still according to the invention, the fuel circuit 3 comprises: • at least one first injection line 1 configured to supply a first flow rate Q1 of fuel H to the combustion chamber 7, the first injection line 1 comprising at least one first pressurizing device 10 configured to discharge the fuel H according to said first flow rate Ql, • at least one second injection line 2 configured to supply a second flow rate Q2 of fuel H to the combustion chamber 7, the second injection line 2 comprising at least one second pressurization system 20 supplying a sonic throat metering device 21 at a supply pressure P20, the sonic throat metering device 21 being configured, from a metering setpoint C2 and the supply pressure P20, to deliver the fuel H according to the second flow rate Q2.

[0039] In the example of [Fig.l], the conditioning system 100 also comprises an upstream circuit 40 extending upstream of the fuel circuit 3. The upstream circuit 40 is configured to supply the tank 4 with fuel H, preferably cryogenic. The upstream circuit 40 comprises a tank 41 of fuel H, preferably cryogenic, as well as a device 42 for pressurizing the fuel H, such as a pump, in particular a cryogenic motor pump. The upstream circuit 40 also comprises a return loop mounted downstream of the pressurized device 42 and supplying the tank 4L. The return loop comprises a heater 43, preferably in the form of a heat exchanger, in order to control the temperature of the fuel H. A valve 44 ensures the opening and closing of the return loop.

[0040] With reference to [Fig.l], the first injection line 1 is configured to provide a first flow rate Q1 to the combustion chamber 7, more precisely to one or more first injectors 5 of the combustion chamber 7. The first flow rate Q1 is supplied by the first pressure device 10, preferably a pump, in particular a cryogenic motor pump.

[0041] In this example, the first injection line 1 further comprises a valve 11, a heater 12 and a flow rate measuring member 13. The valve 11 comprises an open state and a closed state in order to authorize or not the circulation of the fuel H in the first injection line 1. The valve 11 does not participate in modifying the first flow rate Ql. The heater 12 is configured to heat the fuel H, preferably to ensure the transition to the gaseous state of the hydrogen. The heater 12 is mounted downstream of the first pressure device 10. The heater 12 is preferably in the form of a heat exchanger, connected to a heat source of the aircraft turbomachine. The flow rate measuring member 13, such as a flow meter, is mounted downstream of the first pressure device 10 in order to measure the first flow rate Ql supplied.

[0042] With reference to [Fig.l], the first injection line 1 is free of a fuel H metering member. The first flow rate Ql is provided solely by the first pressurizing device 10 which is not suitable for metering the fuel H. The first injection line 1 thus makes it possible to provide a first flow rate Ql which is not very precise and whose reactivity is low. As will be seen later, only the second injection line 2 comprises a metering member. It is the second injection line 2 which makes it possible to compensate for the lack of precision of the first injection line with a high reactivity 1 so that the conditioning system 100 as a whole provides a precise total flow rate which dynamically follows the needs of the turbomachine.

[0043] With reference to [Fig.l], the second injection line 2 is configured to supply a second flow rate Q2 to the combustion chamber 7, more precisely to one or more second injectors 6 of the combustion chamber 7. A sonic throat metering device 21 is advantageously adapted to ensure precise metering of a fuel H in the form of a compressible fluid, such as hydrogen in the gaseous state, unlike a metering valve which is only suitable for incompressible fluids.

[0044] As illustrated in [Fig.l], the sonic neck metering device 21 comprises a pressure reducer 25 mounted upstream of a conduit 26 comprising a sonic neck. The conduit 26 is configured to meter the second flow rate Q2 as a function of the pressure P2 and the temperature T2 upstream of the sonic neck. The pressure reducer 25 is configured to control the pressure P2 reactively upstream of the sonic neck. The pressure reducer 25 advantageously has a rapid response time, less than 500ms, preferably less than 200ms and preferentially less than 100ms. In addition to its precision, the sonic neck metering device 21 thus makes it possible to meter a second flow rate Q2 of quickly following receipt of a dosing instruction C2. Such a sonic neck dosing device 21 advantageously makes it possible to compensate for the slower response time (typically between 500ms and 2s) of the first pressurizing device 10 of the first line 1.

[0045] With reference to [Fig.l], the conduit 26 preferably comprises a fixed, i.e. non-movable, section. The conduit 26 is preferably in the form of a Venturi tube comprising a convergent upstream section and a divergent downstream section on either side of the sonic throat. Such a conduit 26 makes it possible to avoid detachment of the fuel H at the outlet of the sonic throat, and thus to avoid pressure losses.

[0046] As illustrated in [Fig.l], the sonic throat metering device 21 is supplied by a second pressurizing device 20 configured to provide a supply pressure P20. The second pressurizing device 20 and the sonic throat metering device 21 make it possible to provide a priming pressure for which the flow at the sonic throat is sonic and known with great precision, in particular independently of the downstream conditions. Preferably, the second pressurizing device 20 comprises several pumps, preferably two, in particular in the form of cryogenic motor pumps, to provide sufficient pressure to the sonic throat metering device 21.

[0047] Preferably, to enable the operation of the sonic neck metering device 21, the following equation is verified: P2 > 1.5 * PI, and preferably P2 > 2 * PI, with PI the pressure of the first injection line 1. Due to the absence of a metering member, the first injection line 1 has the advantage of being energy-efficient. This makes it possible to compensate for the energy required to pressurize the fuel H in the second injection line 2. The conditioning system 100 advantageously combines the advantages specific to each injection line 1, 2: namely a low energy cost for the first line 1, and a precise and rapid response for the second line 2.

[0048] Still with reference to [Fig.l], the second injection line 2 also comprises a valve 29, a heating device 22 as well as pressure 30 and temperature 31 measuring members. The valve 29 comprises an open state and a closed state in order to authorize or not the circulation of the fuel H in the second injection line 2. The valve 29 does not participate in adjusting the second flow rate Q2. The pressure measuring member 30, such as a pressure sensor, is mounted downstream of the first metering device 21 in order to measure the downstream pressure. The temperature measuring member 31, such as a temperature sensor, is mounted downstream of the first metering device 21 in order to measure the downstream temperature. This makes it possible to know the characteristics of the fuel in the second injection line 2 prior to its injection.

[0049] As illustrated in [Fig.l], the heating device 22 makes it possible to heat the hydrogen in the gaseous state. The heating device 22 is mounted between the second pressurizing device 20 and the metering device 21. The heating device 22 preferably comprises an electric heater 27 and a heat exchanger 28. The electric heater 27 advantageously has a faster response time than the heat exchanger 28. The heat exchanger 28 is less energy-intensive than the electric heater 27. The alternative or combined use of the electric heater 27 and the heat exchanger 28 allows precise, rapid and energy-efficient control of the temperature of the fuel H.

[0050] In the embodiment of [Fig. 5], the heating device 22 comprises only an electric heater 27 configured to ensure the heating of the fuel H before its injection into the combustion chamber 7. This makes it possible to limit the size of the second fuel line 2.

[0051] In the embodiment of [Fig. 6], the first injection line 1 and the second injection line 2 comprise a common portion 9, on which the first pressurizing device 10 is mounted. In other words, the first pressurizing device 10 is mounted on the first injection line 1 and the second injection line 2. Preferably, the second pressurizing device 20 comprises a single pump, sufficient with the first pressurizing device 10 to supply the metering device 21. The sharing of the first pressurizing device 10 advantageously makes it possible to limit the size of the packaging system 100.

[0052] With reference to [Fig. 2], the aircraft turbomachine also comprises a computer 50 configured to control the first flow rate Q1 and the second flow rate Q2 provided by the injection lines 1, 2. As illustrated in [Fig. 2], the computer 50 is configured to provide a first setpoint C1 to the first pressurizing device 10 in order to control the first flow rate Q1. The first setpoint C1 is predetermined as a function of a speed setpoint of the aircraft turbomachine, for example, imposed by a pilot of the aircraft. Preferably, the computer 50 is also configured to receive a flow rate measurement DI from the flow meter 13 of the first injection line 1 and to correct the first setpoint C1 as a function of the flow rate measurement DI (regulation loop). The first setpoint C1 is preferably a speed setpoint. It is specified that [Fig. 2] illustrates the operation of the computer 50 and that the elements of [Fig.l] to which the computer 50 is not connected are not shown for clarity purposes only.

[0053] As illustrated in [Fig.2], the computer 50 is configured to provide a dosing instruction C2 to the dosing device 21, and more precisely to the pressure reducer 25, in order to control the pressure P2 and the temperature T2 upstream of the sonic neck. The cal The computer 50 is also configured to provide a second setpoint C2 to the second pressurizing device 20 in order to control the supply pressure P20 supplied to the metering device 21. The second setpoint C2 is preferably a pressure setpoint. The metering setpoint C2 and the second setpoint C20 are predetermined as a function of a speed setpoint of the aircraft turbomachine, for example, imposed by a pilot of the aircraft. Preferably, the computer 50 is also configured to receive a pressure and temperature measurement D2 from the pressure 30 and temperature 31 sensors of the second injection line 2. The computer 50 is configured to correct the second setpoint C20 and the metering setpoint C2 as a function of the pressure and temperature measurement D2, and preferably the flow rate measurement DI (regulation loop).

[0054] Preferably and as illustrated in [Fig.2], the computer 50 is also configured to provide a setpoint C4 to the pump 42 of the upstream circuit 40.

[0055] The method of using the conditioning system 100 is described below during a stabilized regime RS of the aircraft turbomachine, during which the sum of the first flow rate Q1 and the second flow rate Q2 supplied to the combustion chamber varies by less than 10%, as well as during a transient regime RT of the aircraft turbomachine, namely unstabilized.

[0056] With reference to [Fig.2], during a stabilized regime RS: • The sonic neck dosing device 21 receives El (reception step) at least one predetermined dosing instruction C2, • The sonic neck dosing device 21 provides E2 (supply step), from the dosing setpoint C2 and the supply pressure P20, a second flow rate Q2 to supply the combustion chamber 7, and• In certain embodiments, the first line 1 provides E3 (supply step) a first flow rate Q1 to supply the combustion chamber 7.

[0057] [Fig. 3] illustrates three phases RS1, RS2, RS3 during the stabilized regime RS of the aircraft turbomachine. During an ignition phase RS1 of the aircraft turbomachine, only the receiving step E1 and the supply step E2 of a second flow rate Q2 are preferably implemented. The valve 11 closes the first injection line 1. The flow rate provided by the conditioning system 100 comes only from the second injection line 2. During the receiving step E1, the computer 50 provides a second setpoint C20 to the second pressurization device 20 and a metering setpoint C2 to the metering device 21, which are preferably predetermined for the ignition phase RS1.According to a preferred aspect, during the ignition phase RS1, the dosing device 21 provides a second flow rate Q2 at least equal to 50% of the maximum flow rate that the dosing device 21 is adapted to provide. The setpoints C2, C20 are preferably corrected using the measurements D2 from the sensors. of pressure 30 and temperature 31.

[0058] The method described for the ignition phase RS1 is preferably identical for an idling phase of the aircraft turbomachine, except for the value of the setpoints C2, C20 which are predetermined for the idling phase.

[0059] During a takeoff phase RS2 of the aircraft turbomachine, the supply step E3 is also implemented. During the supply step E3, the computer 50 supplies a first setpoint C1 to the first pressurizing device 10. The setpoints C2, C20 transmitted by the computer 50 also take into account the flow rate measurement DI of the flow meter 13 so that the second flow rate Q2 compensates for the inaccuracy of the first flow rate Q1 supplied. According to a preferred aspect, during the ignition phase RS1, the metering device 21 supplies a second flow rate Q2 at least equal to 80%, preferably at least 90%, of the maximum flow rate that the metering device 21 is adapted to supply. The same applies to the first pressurizing device 10 to supply a sufficient total flow rate. According to a preferred aspect, the first flow rate Q1 represents more than 50% of the total flow rate supplied, preferably between 60% and 75%.

[0060] During a cruise phase RS3, known by the English term “cruise”, steps E1, E2 and E3 are implemented. The first pressurizing device 10 operates at a higher speed than that of the metering device 21 in order to maximize the share of the first flow rate Q1 in the total flow rate supplied to the combustion chamber 7. The first flow rate Q1 supplied by the first line 1 is higher than the second flow rate Q2 supplied by the second line 2, preferably at least twice higher. This advantageously makes it possible to reduce the energy consumption of the conditioning system 100.

[0061] The method described for the RS3 cruise phase is preferably identical for a climb, descent and approach phase of the aircraft.

[0062] According to a preferred aspect, during a stabilized regime RS, a small variation in flow rate is reflected in the setpoints C2, C20 because the response time of the second line 2, and in particular of the pressure reducer 25, is less than that of the first line 1, and in particular of the first pressurizing device 10. The second line 2 advantageously makes it possible to adjust the second flow rate Q2 with precision and speed during a fluctuation in the stabilized regime RS.

[0063] With reference to figures 2 and 4, during a transient RT regime: • The first line 1 provides E4 a first flow rate Q1 to supply the combustion chamber 7, the first flow rate Q1 varying during a response time t until reaching a predetermined target flow rate Qtl, • The sonic neck dosing device 21 receives E5 at least one predetermined dosing instruction C2, • The sonic neck dosing device 21 provides E6, from the dosing setpoint C2 and the supply pressure P20, a second flow Q2 to supply the combustion chamber 7, • The second flow rate Q2 being at least equal to the difference between the first flow rate Q1 and the predetermined target flow rate Qtl during the response time t.

[0064] The second line 2, with a low response time t0, advantageously makes it possible to compensate for the higher response time t of the first line 1. As illustrated in [Fig.4], during the response time t of the first line 1, the first flow rate Q1 supplied is lower than the target flow rate Qtl requested by the computer 50 via the first instruction CL. Preferably, the difference between the first flow rate Q1 and the target flow rate Qtl is determined by the computer 50 thanks to the flow rate measurement DI by the flow rate measuring member 13.

[0065] Preferably, the difference between the first flow rate Q1 and the target flow rate Qtl is taken into account to determine the dosing instruction C2, as well as the second instruction C20 transmitted to the second line 2. During the response time t of the first line 1, the instructions C2, C20 thus comprise a target flow rate Qt2 desired for the second line 2, but also an excess flow rate corresponding to that not provided by the first line 1. At the end of the response time t, the first flow rate Q1 is equal to the target flow rate Qtl. The instructions C2, C20 transmitted to the second line 2 only comprise the target flow rate Qt2 desired for the second line 2. At the end of the response time t, the second flow rate Q2 is equal to the target flow rate Qt2.

[0066] The conditioning system 100 thus makes it possible to supply a flow rate to the combustion chamber 7 in a precise and rapid manner and by limiting the energy consumption of pressurization thanks to two injection lines 1, 2 of different operation. The first line 1, free of metering member, makes it possible to supply a high flow rate for a low energy cost. The second line 2 makes it possible to supply a precise flow rate with a short response time. The control of the two injection lines 1, 2 according to the speed of the aircraft turbomachine advantageously makes it possible to combine the advantages of each line 1, 2.

Claims

Claims

1. Fuel (H) conditioning system (100) for an aircraft turbomachine, said conditioning system (100) comprising a fuel circuit (3) configured to be fluidically connected upstream to at least one fuel tank (4) and to supply downstream a combustion chamber (7) of the aircraft turbomachine, said fuel circuit (3) comprising: • At least one first injection line (1) configured to supply a first flow rate (Q1) of fuel (H) to the combustion chamber (7), the first injection line (1) comprising at least one first pressurizing device (10) configured to discharge the fuel (H) according to said first flow rate (Q1), the first injection line (1) being free of a metering device, • At least one second injection line (2) configured to supply a second flow rate (Q2) of fuel (H) to the combustion chamber (7),the second injection line (2) comprising at least one second pressurizing system (20) supplying a sonic throat metering device (21) with a supply pressure (P20), the sonic throat metering device (21) being configured, from a metering setpoint (C2) and the supply pressure (P20), to deliver the fuel (H) according to the second flow rate (Q2).,

2. The packaging system (100) of claim 1, wherein the pressurizing system (20) is configured to supply a predetermined priming pressure to the sonic neck metering device (21), at which the flow in the sonic neck metering device is sonic.

3. Packaging system (100) according to one of claims 1 and 2, in which the sonic neck dosing device (21) comprises a conduit (26) of fixed section.

4. Packaging system (100) according to one of claims 1 to 3, in which the second line (2) comprises an electric heater (27) mounted upstream of the sonic neck dosing device (21).

5. Packaging system (100) according to one of claims 1 to 4, in which the sonic neck dosing device (21) comprises a sonic neck conduit (26) and at least one controllable pressure reducer (25) mounted upstream of the sonic neck conduit (26).

6. Aircraft turbomachine comprising a combustion chamber (7), at least one fuel tank (4) and at least one conditioning system (100) according to one of claims 1 to 5, the fuel circuit (3) of which is fluidically connected upstream to the fuel tank (4) and downstream to the combustion chamber (7).

7. Method of using the conditioning system (100) according to one of claims 1 to 5, in which, during a stabilized regime (RS) of the aircraft turbomachine: • The sonic throat metering device (21) receives (El) at least one predetermined metering instruction (C2), • The sonic throat metering device (21) provides (E2), from the metering instruction (C2) and the supply pressure (P20), a second flow rate (Q2) to supply the combustion chamber (7).

8. Method of use according to claim 7, in which, during an ignition phase (RS1) during the stabilized regime (RS) of the aircraft turbomachine, the second flow rate (Q2) is equal to at least 80% of the total flow rate provided by the conditioning system (100), preferably at least 95%.

9. Method of use according to one of claims 7 and 8, in which, during a cruise phase (RS3) during the stabilized regime (RS) of the aircraft turbomachine: • The first line (1) provides (E3) a first flow rate (Ql) to supply the combustion chamber (7), • The first flow rate (Ql) being greater than the second flow rate (Q2) provided by the second line (2), preferably at least twice greater.

10. Method of use according to one of claims 7 to 9, in which, during a transient regime (RT) of the aircraft turbomachine: • The first line (1) provides (E4) a first flow rate (Ql) for supplying the combustion chamber (7), the first flow rate (Ql) varying during a response time (t) until reaching a predetermined target flow rate (Qtl), The sonic neck dosing device (21) receives (E5) at least one predetermined dosing instruction (C2), The sonic neck dosing device (21) provides (E6), from the dosing setpoint (C2) and the supply pressure (P20), a second flow rate (Q2) to supply the combustion chamber (7), The second flow rate (Q2) being at least equal to the difference between the first flow rate (Ql) and the predetermined target flow rate (Qtl) during the response time (t).

Images