System and method for determining at least one heat quantity to be provided in a fuel conditioning system for supplying an aircraft turbine engine

Abstract

The invention relates to a method for determining a heat quantity (Th) to be provided in a heating module (3) of a fuel conditioning system (SC), the method comprising a step of determining an overall range of second specific enthalpy (HC2) of a fuel flow (Q) flowing through the conditioning system (SC) from the formulas γmin=min[(HC2a-HC1) / 106] in which γmin is a transfer criterion equal to 2.035 and γmax=max[(HC2b-HC1) / 106] in which γmax is a transfer criterion equal to 6.712; and a step of determining, in a database (BdD) of elementary ranges of second specific enthalpies (HC2), a heat quantity (Th), the elementary range of second specific enthalpies (HC2) of which is entirely within the overall range of second specific enthalpies (HC2) so as to determine the heat quantity (Th) to be supplied to the heating module (3).

Description

System and method for determining at least one quantity of heat to be supplied to a fuel conditioning system to power an aircraft turbomachine

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

[0002] Climate change is a major concern for many legislative and regulatory bodies around the world. Indeed, various carbon emission restrictions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies to both new aircraft types and those already in operation, requiring the implementation of technological solutions to ensure their compliance with current regulations. Civil aviation has been mobilizing 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 consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.

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

[0005] This sustained research and development work focuses on new generations of aircraft engines, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to ensure propulsion, and, as an essential complement to technological progress, aeronautical 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 turbomachines powered by fuel stored in a cryogenic tank.

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

[0008] In order to be injected into the combustion chamber of a turbomachine, the fuel must be conditioned, i.e. 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, particularly at the fuel injectors.

[0009] In a known manner, with reference to the, a conditioning system SC comprises a fuel circuit 1 connected at the inlet to a cryogenic tank R and at the outlet to the combustion chamber of a turbomachine M. In practice, the conditioning system SC is defined in an aircraft reference frame REF-A and in a turbomachine reference frame REF-M. The cryogenic tank R is positioned in the aircraft reference frame REF-A while the turbomachine M is positioned in the turbomachine reference frame REF-M. A fuel flow Q circulates from upstream to downstream in the fuel circuit 1.

[0010] In the conditioning system SC, the fuel flow Q has, at the outlet of the cryogenic tank R, known conditions (for example in pressure and temperature) which correspond to specific requirements to allow the fuel to be stored in the liquid state in the cryogenic tank R. Similarly, the fuel flow Q must have, at the inlet of the turbomachine M, specific conditions which are different from the conditions at the outlet of the cryogenic tank R. The conditions at the inlet of the turbomachine M allow the injection of the fuel Q into the combustion chamber of the turbomachine M, to allow optimal combustion as described previously.

[0011] With reference to the, the fuel flow Q has, at an inlet point P1 of the fuel circuit 1, inlet conditions C1 (corresponding to the conditions at the outlet of the cryogenic tank R imposed by an aircraft manufacturer), and, at an outlet point P2 of the fuel circuit 1, outlet conditions C2 (corresponding to the conditions at the inlet of the turbomachine M).

[0012] To enable the fuel flow Q to be conditioned, i.e. to enable its passage from the inlet conditions C1 to the outlet conditions C2, as is known, the conditioning system SC comprises a mechanical pump 2 and a heating module 3 (shown only on the). The mechanical pump 2 is configured to circulate the fuel flow Q in the fuel circuit 1 and raise its pressure. The heating module 3 is configured to provide calories to the fuel flow Q in order to heat it so that it can be injected into the turbomachine M. In practice, the heating module 3 may belong to the aircraft reference frame REF-A (as shown on the) and take heat from the aircraft (air from the cabin, heat from electrical or electronic systems, etc.) or belong to the REF-M turbomachine reference system to extract heat from the M turbomachine (heat from the lubricating oil, calories at the turbine outlet, heat from the nozzle, etc.). The SC conditioning system can also, in a known manner, comprise several heating modules 3 mounted in the REF-A aircraft reference system and / or in the REF-M turbomachine reference system.

[0013] When designing a conditioning system SC, it is necessary to choose the heating module 3 to achieve the outlet conditions C2 from the inlet conditions C1 for any flight phase of the aircraft A. This is complex since the different types of known heating modules 3 do not provide the same amount of heat for all flight phases. For example, in the case of a heating module mounted in the turbomachine reference REF-M, the turbomachine M generates less heat in the taxi phase than in the takeoff phase. An immediate solution to resolve this drawback would be to oversize the heating module 3 but this would increase the mass and the size.

[0014] The design of an SC conditioning system must also take into account constraints related to the fuel flow Q which must be in a single-phase state in the fuel circuit 1 to ensure flow stability and limit any damage to the SCAA conditioning system.

[0015] For each new aircraft, the selection of the heating module for an air conditioning system is currently carried out empirically to ensure that the output conditions are met. This generally leads to oversizing which affects the weight and size of the aircraft.

[0016] One of the objectives of the present invention is to enable the rigorous and reliable selection of one (or more) heating module(s) which meet(s) the requirements of a fuel conditioning system, in particular meeting the inlet conditions and the outlet conditions, for any phase of flight of the aircraft. The invention aims in particular to limit the mass and size of the conditioning system while limiting its implementation complexity. PRESENTATION OF THE INVENTION

[0017] The invention relates to a method for determining at least one quantity of heat to be supplied in at least one heating module of a fuel conditioning system configured to supply an aircraft turbomachine with fuel from a cryogenic tank, a fuel flow circulating between the cryogenic tank and the turbomachine via a fuel circuit, the aircraft being configured to operate according to a plurality of flight phases, the heating module being configured to heat the fuel flow from the quantity of heat, the fuel flow having: at least a first enthalpy specific to an entry point positioned on the fuel circuit at the outlet of the cryogenic tank, at least a second enthalpy specific to an exit point positioned on the fuel circuit at the inlet of the turbomachine, a database of quantities of heat being accessible,each quantity of heat being associated with an elementary range of second specific enthalpies, each elementary range being sized for the plurality of flight phases of the aircraft.,

[0018] A specific enthalpy is an enthalpy determined for a unit of mass.

[0019] The method comprises the steps of:determining a defined overall second specific enthalpy range, between a first second specific enthalpy boundary and a second second specific enthalpy boundary, from:the predetermined first specific enthalpy at the entry point, andthe following formulas:γmin = min[(HC2a-HC1) / 10 6 ] in which γmin is a transfer criterion equal to 2.035,γmax = max[(HC2b-HC1) / 10 6] in which γmax is a transfer criterion equal to 6.712.determine, in the database, the quantity of heat whose elementary range of second specific enthalpies is entirely included in the global range of second specific enthalpies so as to determine the quantity of heat to be supplied to the heating module.

[0020] The method according to the invention makes it possible to determine the quantity of heat to be supplied in the heating module to enable optimal heating of the fuel flow. The injection temperature is advantageously reached without the heating module being oversized, which makes it possible to limit both the mass and the size of the conditioning system. In particular, the fuel flow is thus injected at controlled thermo-fluidic conditions, such as density or injection speed, allowing stable and optimal combustion of the fuel in the combustion chamber of the turbomachine. Thanks to the determination system according to the invention, the transport of the fuel flow throughout the fuel circuit is ensured at optimal temperatures, regardless of the speed of the turbomachine.

[0021] The overall range of second specific enthalpies determined advantageously ensures that the fuel flow is in a single-phase state, thereby ensuring flow stability in the fuel circuit conduits. The conditioning system is thus safe and secure. In particular, this ensures that the fuel flow is in a gaseous single-phase state when injected into the combustion chamber of the turbomachine, regardless of the injection pressure of the fuel flow.

[0022] The determined range of second specific enthalpies also makes it possible to ensure both that the fuel flow is injected at low but sufficient temperatures (of the order of -123 to -73 °C (150 to 200K)) when the engine speed of the turbomachine is maximum and that the fuel flow is injected at high temperatures (of the order of 127 to 177 °C (400 to 450K)) when the engine speed of the turbomachine is minimum. The minimum and maximum injection temperatures in the turbomachine are thus advantageously optimal for all phases of flight of the aircraft.

[0023] In other words, thanks to the invention it is possible both to inject the fuel flow at so-called "cold" temperatures during maximum engine speeds, making it possible to minimize the sizing of the heating module, which leads to a minimization of the energy cost of heating the fuel during the most important flight phases, and to inject the fuel flow at so-called "hot" temperatures during minimum engine speeds, making it possible to maximize the thermodynamic efficiency of the engine without inducing oversizing of the heating module.

[0024] Preferably, the database associating for each quantity of heat, at least one heating module, the method comprises a step of determining, in the database, at least one heating module to be used in the conditioning system. The method according to the invention thus makes it possible to choose from a plurality of possible heating modules, the heating module to be used to heat the fuel flow optimally. Thanks to the invention, the choice of the heating module is no longer carried out empirically as was the case in the prior art, which allows a significant saving of time, while limiting costs. In addition, the heating module is not oversized as could be the case in the prior art, which makes it possible to limit both the mass and the size of the conditioning system.A limited mass also makes it possible to limit the aircraft's energy consumption, which advantageously limits greenhouse gas emissions and therefore the aircraft's impact on the environment.

[0025] In a first embodiment, the conditioning system being defined in an aircraft reference frame and a turbomachine reference frame distant from each other, the cryogenic tank being positioned in the aircraft reference frame and the turbomachine being positioned in the turbomachine reference frame, the heating module is mounted in the aircraft reference frame. Such an embodiment makes it possible to heat the fuel flow before it leaves the aircraft reference frame, which makes it possible to limit the use of heavy and bulky cryogenic lines.

[0026] In a second embodiment, the conditioning system being defined in an aircraft reference frame and a turbomachine reference frame distant from each other, the cryogenic tank being positioned in the aircraft reference frame and the turbomachine being positioned in the turbomachine reference frame, the heating module is mounted in the turbomachine reference frame. Such an embodiment makes it possible to heat the fuel flow efficiently by means of the heat sources present close to the turbomachine.

[0027] In one embodiment, the conditioning system being defined in an aircraft reference frame and a turbomachine reference frame distant from each other, the cryogenic tank being positioned in the aircraft reference frame and the turbomachine being positioned in the turbomachine reference frame, the conditioning system comprising at least a first heating module, mounted in the aircraft reference frame and configured to heat the fuel flow from a first quantity of heat, and at least a second heating module, mounted in the turbomachine reference frame and configured to heat the fuel flow from a second quantity of heat, the database associating for each pair of a plurality of pairs of quantities of heat an elementary range of second specific enthalpies, each elementary range being sized for the plurality of flight phases of the aircraft,the step of determining the method makes it possible to determine at least one pair of heat quantities whose elementary range is entirely included in the overall range, so as to determine the first quantity of heat to be supplied in the first heating module and the second quantity of heat to be supplied in the second heating module. The fuel flow can thus be gradually heated in the fuel circuit by means of two heating modules positioned in the two reference frames, which makes it possible to limit the sizing of each heating module while allowing the fuel flow to reach optimal conditions at the inlet of the combustion chamber of the turbomachine.,

[0028] Preferably, the heating module is chosen, alone or in combination, from: a heat exchanger, an electric heating system, a thermal storage system, a fuel cell, a gas turbine, a burner. The method thus makes it possible to choose from a plurality of modules capable of heating the fuel flow efficiently. The method also allows the use of types of heating modules whose operation and capacities are known.

[0029] In one embodiment, the heating module being a heat exchanger configured to heat the fuel flow from calories transferred by at least one heat source, the database associating for each heating module corresponding to a heat exchanger, a heat source to be used, the method comprises a step of determining, in the database, the heat source to be used in the heating module to heat the fuel flow. Thanks to the method according to the invention, when the heating module is a heat exchanger, the heat source to be used to achieve the desired quantity of heat is chosen simply and quickly from a plurality of possible heat sources, without requiring a significant number of iterations as was the case in the prior art.

[0030] In one embodiment, the overall range of second specific enthalpies is determined from the formulas: γmin = min[(HC2a-HC1) / 10 6 ] in which γmin is a transfer criterion equal to 2.035, and γmax = max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 4.507.

[0031] Such a range of transfer criteria makes it possible to inject the fuel flow into the combustion chamber of the turbomachine at low temperatures for all engine speeds, and therefore for all phases of flight of the aircraft, while respecting the minimum and maximum temperature values ​​that the fuel flow must have in order to be injected. The sizing of the heating module can thus be minimized, which makes it possible to limit the mass and size of the conditioning system. A limited mass also makes it possible to limit the energy consumption dedicated to heating the fuel while limiting the fuel consumption of the aircraft, which makes it possible to limit greenhouse gas emissions and therefore to limit the impact of the aircraft on the environment.

[0032] Alternatively, the overall range of second specific enthalpies is determined from the formulas:γmin = min[(HC2a-HC1) / 10 6] in which γmin is a transfer criterion equal to 4.034, and γmax = max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 6.712.

[0033] Such a range of transfer criteria makes it possible to inject the fuel flow into the combustion chamber of the turbomachine at high temperatures for all engine speeds, and therefore for all flight phases of the aircraft, while respecting the minimum and maximum temperature values ​​that the fuel flow must have in order to be injected. The thermodynamic efficiency of the turbomachine is thus advantageously maximized while mounting an optimally sized heating module in the fuel circuit, unlike the prior art in which the heating module was oversized to ensure that it could operate even at high speeds.

[0034] The invention also relates to a system for determining at least one quantity of heat to be supplied in at least one heating module of a fuel conditioning system configured to supply an aircraft turbomachine with fuel from a cryogenic tank, the determination system being configured to implement the determination method as described above, the aircraft being configured to operate according to a plurality of flight phases, the conditioning system comprising: a fuel circuit connected at the inlet to the cryogenic tank and at the outlet to the turbomachine, a fuel flow circulating in the fuel circuit, at least one heating module configured to heat the fuel flow from the quantity of heat, the fuel flow having: at least one first specific enthalpy at an inlet point positioned on the fuel circuit at the outlet of the cryogenic tank,at least one second specific enthalpy at an outlet point positioned on the fuel circuit at the inlet of the turbomachine, the determination system comprising: a database of heat quantities, each heat quantity being associated with an elementary range of second specific enthalpies, each elementary range being sized for the plurality of flight phases of the aircraft, a computer configured to determine a global range of second specific enthalpies defined, between a first terminal of second specific enthalpy and a second terminal of second specific enthalpy, from: the first specific enthalpy predetermined at the inlet point, and the following formulas: γmin = min[(HC2a-HC1) / 10, 6 ] in which γmin is a transfer criterion equal to 2.035,γmax = max[(HC2b-HC1) / 10 6] in which γmax is a transfer criterion equal to 6.712.determine, in the database, the quantity of heat whose elementary range of second specific enthalpies is entirely included in the global range of second specific enthalpies so as to determine the quantity of heat to be supplied to the heating module.

[0035] The determination system according to the invention advantageously makes it possible to determine simply and quickly, by means of a database and a calculator, the optimal quantity of heat to be supplied to the heating module, to heat a flow of fuel from a cryogenic tank with a view to its injection into an aircraft turbomachine.

[0036] The invention also relates to a fuel conditioning system configured to supply an aircraft turbomachine with fuel from a cryogenic tank, the aircraft being configured to operate in a plurality of flight phases, the conditioning system comprising: a fuel circuit connected at the inlet to the cryogenic tank and at the outlet to the turbomachine, a fuel flow circulating in the fuel circuit, at least one heating module configured to heat the fuel flow from the quantity of heat, the fuel flow having: at least a first specific enthalpy at an inlet point positioned on the fuel circuit at the outlet of the cryogenic tank, at least a second specific enthalpy at an outlet point positioned on the fuel circuit at the inlet of the turbomachine,conditioning system in which the second specific enthalpy of the fuel flow is included in an interval [HC2a; HC2b] respecting the following formulas:γmin = min[(HC2a-HC1) / 10, 6 ] in which γmin is a transfer criterion equal to 2.035,γmax = max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 6.712.

[0037] In a first alternative embodiment, the second specific enthalpy of the fuel flow in the interval [HC2a, HC2b] complies with the following formulas: γmin = min[(HC2a-HC1) / 10 6 ] in which γmin is a transfer criterion equal to 2.035,γmax = max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 4.507.

[0038] In a second alternative embodiment, the second specific enthalpy of the fuel flow in the interval [HC2a, HC2b] complies with the following formulas: γmin = min[(HC2a-HC1) / 10 6 ] in which γmin is a transfer criterion equal to 4.034,γmax = max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 6.712.

[0039] Finally, the invention relates to an aircraft comprising a cryogenic tank, a turbomachine and a conditioning system as described previously for supplying the turbomachine with fuel from the cryogenic tank. PRESENTATION OF FIGURES

[0040] 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.

[0041] This is a schematic representation of an aircraft including a fuel conditioning system.

[0042] This is a schematic representation of the packaging system of the.

[0043] This is a schematic representation of a determination system according to a first embodiment of the invention.

[0044] It is a schematic representation of a database associating for each elementary range of second specific enthalpies, at least one quantity of heat.

[0045] This is a schematic representation of an overall range of second specific enthalpies and a plurality of elementary ranges of second specific enthalpies from the database of the.

[0046] This is a schematic representation of a determination system according to a second embodiment of the invention.

[0047] This is a diagram of the steps of a determination method according to an embodiment of the invention.

[0048] 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 if necessary. DETAILED DESCRIPTION OF THE INVENTION

[0049] With reference to the, there is shown a fuel conditioning system SC configured to supply a turbomachine M of an aircraft A from fuel Q from a cryogenic tank R.

[0050] As is known, aircraft A is configured to operate in a plurality of flight phases, for example a takeoff phase, a cruise phase, a landing phase and a taxi phase.

[0051] The turbomachine M is configured to provide propulsion for the aircraft A, in particular, by driving at least one propulsion unit (not shown).

[0052] In this example, the fuel Q in the cryogenic tank R is stored at a temperature of the order of -253 to -251°C (20 to 22 Kelvins). At this temperature, the fuel stream Q is liquid. In this example, the fuel Q is liquid hydrogen, but the invention applies to other types of fuel, for example, liquid methane or liquefied natural gas.

[0053] The conditioning system SC is defined in an aircraft reference frame REF-A and in a turbomachine reference frame REF-M distant from each other. The cryogenic tank R extends in the aircraft reference frame REF-A while the turbomachine M extends in the turbomachine reference frame REF-M.

[0054] Still with reference to the, the conditioning system SC comprises a fuel circuit 1 connected at the inlet to the cryogenic tank R and at the outlet to the combustion chamber of a turbomachine M. A fuel flow Q circulates from upstream to downstream in the fuel circuit 1.

[0055] The fuel flow Q has, at the outlet of the cryogenic tank R, known conditions (for example in pressure and temperature) which correspond to specific requirements imposed by an aircraft manufacturer to allow the fuel Q to be stored in the liquid state in the cryogenic tank R. The fuel flow Q must also have, at the inlet of the turbomachine M, conditions which allow the injection of the fuel Q into the combustion chamber of the turbomachine M, to allow optimal combustion as described previously. In practice, the conditions at the inlet of the turbomachine M are different from the conditions at the outlet of the cryogenic tank R.

[0056] More precisely, as shown in the, the fuel flow Q has, at an inlet point P1 of the fuel circuit 1, positioned directly at the outlet of the cryogenic tank R, predetermined inlet conditions C1 corresponding to the conditions at the outlet of the cryogenic tank R. Similarly, the fuel flow Q has, at an outlet point P2 of the fuel circuit 1, positioned directly before the inlet into the turbomachine M, outlet conditions C2 corresponding to the inlet conditions of the turbomachine M.

[0057] In particular, the fuel flow Q has at the entry point P1, a first predetermined specific enthalpy HC1. More precisely, the fuel flow Q has at the entry point P1 a first specific enthalpy HC1 determined for example for a predetermined reference specific enthalpy value at liquid saturation and at the boiling point, i.e. for a pressure of 101325 Pa. The first specific enthalpy HC1 varies within a range of first specific enthalpies HC1 valid for all flight phases of the aircraft.

[0058] As is known, the specific enthalpy of a substance corresponds to the value of its enthalpy per unit mass. In practice, the specific enthalpy is determined from the measurement of the static pressure and the measurement of the static temperature of the fluid (here the fuel flow Q), preferably in a single-phase state. Enthalpy diagrams or data tables known in the literature allow the specific enthalpy value to be determined depending on the type of fluid and the static pressure and temperature measurements.

[0059] As an example, it is possible to calculate a specific enthalpy variation HC from the following formula: , in which: is the pressure variation between the entry point P1 and the exit point P2, is the temperature variation between the entry point P1 and the exit point P2, is the specific heat capacity at constant pressure, is the density, is the static temperature, and is the coefficient of thermal expansion.

[0060] For example, for a para-hydrogen, the first specific enthalpy HC1 of the fuel flow Q at the entry point P1 is in the interval [-29,942; 49,693] J / kg for a zero reference specific enthalpy value HC=0 at liquid saturation and boiling point, i.e. for a pressure of 101325 Pa.

[0061] The fuel flow Q also has a second specific enthalpy HC2 at the outlet point P2. Analogously to the inlet point P1, the fuel flow Q has a second specific enthalpy HC2 at the outlet point P2, which varies within a range of second specific enthalpies HC2, which corresponds to a specific enthalpy value, for example, for a predetermined reference specific enthalpy value at liquid saturation and boiling point, i.e., for a pressure of 101325 Pa. The range of second specific enthalpies HC2 is valid for all flight phases of the aircraft.

[0062] To enable the fuel flow Q to be conditioned, i.e. to enable its passage from the inlet conditions C1 to the outlet conditions C2, the conditioning system SC comprises a mechanical pump 2 and a heating module 3 mounted in the fuel circuit 1.

[0063] The mechanical pump 2 is preferably a high-pressure pump and is configured to circulate fuel flow Q from upstream to downstream in the fuel circuit 1 and to raise its pressure. The mechanical pump 2 is preferably mounted in the aircraft reference frame REF-A, i.e. as close as possible to the cryogenic tank R.

[0064] The heating module 3 is configured to provide calories to the fuel flow Q in order to heat it so that it can be injected into the turbomachine M. In this example, the conditioning system SC comprises a single heating module 3, however the conditioning system SC could alternatively comprise a different number of heating modules 3, in particular a number greater than one, as will be described in more detail later.

[0065] In this example, still with reference to the, the heating module 3 is mounted in the aircraft reference frame REF-A and is configured to heat the fuel flow Q up to an injection temperature Ti. The injection temperature Ti corresponds to a temperature at which the fuel flow Q can be injected into the combustion chamber of the turbomachine M. In this example, the injection temperature Ti is between 150 and 450°C. A conditioning system SC is described in which the heating module 3 is mounted in the aircraft reference frame REF-A, it goes without saying that the heating module 3 could alternatively be mounted in the turbomachine reference frame REF-M.Similarly, in the case of a conditioning system SC comprising a plurality of heating modules 31, 32 (as shown in the), each heating module 31, 32 could just as easily be mounted in a different reference frame, for example a first heating module 31 in the aircraft reference frame REF-A and a second heating module 32 in the turbomachine reference frame REF-M.

[0066] In this example, the heating module 3 is configured to heat the fuel flow Q from the heat coming from the aircraft A, such as the air coming from the cabin, the heat coming from electrical or electronic systems, etc. Alternatively, when the heating module 3 is mounted in the turbomachine reference frame REF-M, it is configured to heat the fuel flow Q from the heat coming from the turbomachine M, such as the heat coming from the lubricating oil, the calories at the turbine outlet, the heat from the nozzle, etc.

[0067] The heating module 3 is preferably a heat exchanger, an electric heating system, a thermal storage system, a fuel cell, a gas turbine or a burner.

[0068] According to one aspect of the invention, with reference to the, the heating module 3 is configured to provide a quantity of heat Th making it possible to heat the fuel flow Q up to the injection temperature Ti.In the embodiment in which the conditioning system SC comprises a first heating module 31 mounted in the aircraft reference frame REF-A and a second heating module 32 mounted in the turbomachine reference frame REF-M (as shown in the), the first heating module 31 is configured to provide a first quantity of heat ThA making it possible to heat the fuel flow Q, for example up to a circulation temperature Tc, lower than the injection temperature Ti and allowing the circulation of the fuel flow Q in a gaseous state between the aircraft reference frame REF-A and the turbomachine reference frame REF-M, and the second heating module 32 is configured to provide a second quantity of heat ThM making it possible to heat the fuel flow Q up to the injection temperature Ti.

[0069] The fuel flow Q is configured to flow in the fuel circuit 1 in a single-phase state, so as to limit flow instabilities. In this example, the fuel flow Q is configured to flow in a liquid single-phase state upstream of the heating module 3 and in a gaseous single-phase state downstream of the heating module 3. In this example, in which the heating module 3 is mounted in the aircraft reference frame REF-A, this makes it possible to avoid heavy and bulky cryogenic lines between the aircraft reference frame REF-A and the turbomachine reference frame REF-M.

[0070] As described above, the fuel stream Q has at the outlet point P2 a second specific enthalpy HC2. The second specific enthalpy HC2 is configured to be determined from the first specific enthalpy HC1 at the inlet point P1 and a transfer criterion γ, according to which γ=(HC2-HC1) / 10 6. In practice, the second specific enthalpy HC2 depends on the amount of heat Th supplied by the heating module 3, as will be described in more detail later.

[0071] According to one aspect of the invention, the transfer criterion γ is defined according to a minimum transfer criterion γmin and a maximum transfer criterion γmax, so as to cover all the flight phases of the aircraft A. In practice, the transfer criterion γ depends on the definition intervals of each specific enthalpy HC1, HC2 which vary according to the flight phases of the aircraft. In other words, the transfer criterion γ is defined in a predetermined range [γmin; γmax]. According to one aspect of the invention, the transfer criterion γ making it possible to determine a range of second specific enthalpies HC2 is defined in a predetermined range [2.035; 6.712]. Such a range of transfer criteria γ advantageously allows optimal heating of the fuel flow Q in the heating module 3 to allow it to be injected optimally into the combustion chamber of the turbomachine M.

[0072] In a first alternative embodiment, the transfer criterion γ for determining a range of second specific enthalpies HC2 is defined in the predetermined range [2.035; 4.507]. Such a range makes it possible to inject the fuel flow Q into the turbomachine M at low temperatures regardless of the engine speed, thus making it possible to limit the mass and size of the heating module 3.

[0073] In a second alternative embodiment, the transfer criterion γ for determining a range of second specific enthalpies HC2 is defined in the predetermined range [4.034; 6.712]. Such a range makes it possible to inject the fuel flow Q into the turbomachine M at high temperatures regardless of the engine speed, thus making it possible to maximize the energy efficiency of the turbomachine M.

[0074] There is also shown on the, a system for determining SD, according to an embodiment of the invention, the quantity of heat Th to be supplied in the heating module 3, to make it possible to obtain at the outlet point P2, a second specific enthalpy HC2 which is sized to form an optimal conditioning system SC.

[0075] The determination system SD comprises a database BdD, represented on the, comprising a plurality of elementary ranges HC2-1, HC2-2, … HC2-X of second specific enthalpies HC2, and a calculator 9 (represented on the), configured to determine the quantity of heat Th to be supplied in the heating module 3.

[0076] In the BdD database, each elementary range HC2-1, HC2-2, … HC2-X of second specific enthalpies HC2 is sized for the plurality of flight phases of the aircraft A. In other words, each elementary range HC2-1, HC2-2, … HC2-X of second specific enthalpies HC2 is configured to allow optimal operation of the SC conditioning system regardless of the flight phase of the aircraft A.

[0077] As shown in the, the BdD database associates for each elementary range HC2-1, HC2-2, … HC2-X of second specific enthalpies HC2 one or more quantities of heat Th.

[0078] The database BdD also associates for each heat quantity Th, a heating module 3 to be used in the conditioning system SC. In other words, the database BdD is configured to determine for a given elementary range HC2-1, HC2-2, … HC2-X of second specific enthalpies HC2, the type of heating module 3 to be mounted in the conditioning system SC to ensure that the conditioning system SC is optimally defined according to the inlet conditions C1 and to thus heat the fuel flow Q optimally to allow its injection into the combustion chamber of the turbomachine M. Alternatively, the database BdD could also be configured to associate with each heat quantity Th, a plurality of possible heating modules 3.

[0079] For each heating module 3 corresponding to a heat exchanger, configured to heat the fuel flow Q from the calories provided by a hot source SA present in the aircraft reference frame REF-A (respectively in the turbomachine reference frame REF-M), the database BdD is also configured to associate, for each heating module 3, a hot source SA to be used. Alternatively, the database BdD could also associate with each heating module 3, a plurality of possible hot sources SA.

[0080] By way of example, with reference to the, for an elementary range HC2-2 of second specific enthalpies HC2, the database BdD is configured to associate three possible quantities of heat Th-21, Th-22, Th-23 to be supplied in the heating module 3. The database BdD is also configured to associate with each quantity of heat Th, one (or more) heating module(s) 3 making it possible to supply such a quantity of heat Th. In one example, a first possible quantity of heat Th-21 is associated with a first heating module 3-21 corresponding to an electric heating system mounted in the aircraft reference system REF-A. A second possible quantity of heat Th-22 is associated with two possible heating modules 3-221, 3-222 corresponding for example respectively to a heat exchanger and a fuel cell mounted in the aircraft reference system REF-A.In this example, a third heat quantity Th-23 is associated with a third heating module 3-23 corresponding, for example, to an electric heating system mounted in the aircraft reference frame REF-A. In this example, in which the second heating module 3-221 corresponds to a heat exchanger mounted in the aircraft reference frame REF-A, the database BdD is also configured to associate with the heating module 3-221 several heat sources SA1, SA2 to be used. In the example in which the heating module 3 is mounted in the turbomachine reference frame REF-M, the database BdD is configured in a similar manner to determine the heat quantity Th and the optimal type of heating module 3 to be mounted in the turbomachine reference frame REF-M, for example a gas turbine, a burner or a fuel cell.

[0081] The database BdD represented on the figure is given only as an example, it goes without saying that each elementary range HC2-1, HC2-2, HC2-3, …, HC2-X of second specific enthalpies HC2 could alternatively be associated with a different number of possible heat quantities Th. Similarly, each heat quantity Th could alternatively be associated with a different number of possible heating modules 3. Similarly, the database BdD could alternatively associate with the heating module 3 corresponding to a heat exchanger, a single hot source SA to be used or a different number of two possible hot sources SA.

[0082] In the embodiment in which the conditioning system SC comprises a first heating module 31 mounted in the aircraft reference frame REF-A and a second heating module 32 mounted in the turbomachine reference frame REF-M, the database BdD is configured to associate with each pair of heating modules (31, 32) one or more pairs of heat quantities (ThA, ThM) and for each pair of heat quantities (ThA, ThM), one or more pairs of heating modules (31, 32) to be mounted respectively in the aircraft reference frame REF-A and in the turbomachine reference frame REF-M to allow optimal combustion of the fuel flow Q in the turbomachine M.

[0083] As described above, the determination system SD also comprises a computer 9. The computer 9 is configured to determine the heating module 3 to be mounted in the conditioning system SC from the inlet conditions C1 of the fuel flow Q. For this, the computer 9 is configured to determine a range of second specific enthalpies [HC2a-HC2b] required at the outlet point P2 to ensure optimal heating of the fuel flow Q in the conditioning system SC.

[0084] More precisely, the calculator 9 is configured to: determine a first terminal of second specific enthalpy HC2a from the first specific enthalpy HC1 predetermined at the entry point P1 and the formula γmin=min[(HC2a-HC1) / 10 6] in which γmin is a transfer criterion equal to 2.035, determine a second terminal of second specific enthalpy HC2b from the first specific enthalpy HC1 predetermined at the entry point P1 and the formula γmax=max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 6.712, anddetermine an overall range HC2-G of second specific enthalpies HC2 between the first terminal of second specific enthalpy HC2a and the second terminal of second specific enthalpy HC2b determined.

[0085] In particular, in the transfer criterion γmin=min[(HC2a-HC1) / 10 6], the minimum γmin is sought for all values ​​of first specific enthalpies HC1 in the interval of first specific enthalpies HC1 valid for all phases of flight of the aircraft and for all values ​​of second specific enthalpies HC2 in the interval of second specific enthalpies HC2 valid for all phases of flight of the aircraft. Similarly, in the transfer criterion γmax=max[(HC2b-HC1) / 10 6 ], the maximum γmax is sought for all values ​​of first specific enthalpies HC1 in the interval of first specific enthalpies HC1 valid for all phases of flight of the aircraft and for all values ​​of second specific enthalpies HC2 in the interval of second specific enthalpies HC2 valid for all phases of flight of the aircraft.

[0086] In a first alternative embodiment, the calculator 9 is configured to determine the terminals HC2a, HC2b of second specific enthalpy HC2 from the preceding formulas, in which the transfer criterion γmin is equal to 2.035 and the transfer criterion γmax is equal to 4.507.

[0087] In a second alternative embodiment, the calculator 9 is configured to determine the terminals HC2a, HC2b of second specific enthalpy HC2 from the preceding formulas, in which the transfer criterion γmin is equal to 4.034 and the transfer criterion γmax is equal to 6.712.

[0088] The calculator 9 is also configured to determine, among the plurality of elementary ranges HC2-1, HC2-2, …, HC2-X of second specific enthalpies HC2 of the database BdD, the elementary range HC2-X which is entirely included in the determined global range HC2-G, as shown in the. In this example, only the second specific enthalpy HC2-2 is concerned. It goes without saying that the calculator 9 could alternatively determine a plurality of elementary ranges HC2-1, HC2-2, …, HC2-X of second specific enthalpies HC2 according to the determined global range HC2-G of second specific enthalpies HC2.

[0089] According to one aspect of the invention, the computer 9 is configured to determine, from the database BdD: the quantity(ies) of heat Th corresponding to the determined elementary range HC2-X, the heating module(s) 3 to be mounted in the fuel circuit 1, and when the heating module 3 is a heat exchanger, the heat source(s) SA, SM to be used in the heating module 3.

[0090] The determination system SD according to the invention advantageously makes it possible to determine the heating module 3 to be used in the conditioning system SC so as to allow optimal heating of the fuel flow Q for injection into the combustion chamber of the turbomachine M while optimizing the mass and size of the conditioning system SC.

[0091] A method for determining a quantity of heat Th to be supplied in a heating module 3 of a fuel conditioning system SC of an aircraft A will now be described, with reference to, according to an embodiment of the invention.

[0092] In this example, in a preliminary step E0, the inlet conditions C1 at the inlet point P1 of the fuel circuit 1 are predetermined and known. In particular, the first specific enthalpy HC1 at the inlet point P1 is known for the different flight phases of the aircraft. The inlet point P1 and the outlet point P2 are determined so that the fluid is single-phase in the entire fuel circuit Q between the inlet point P1 and the outlet point P2.

[0093] The method comprises a first step E1 of determining, by the computer 9, a global range HC2-G of second specific enthalpies HC2 defined between a first terminal of second specific enthalpy HC2a and a second terminal of second specific enthalpy HC2b. For this, the computer 9 determines, in this step E1, the first terminal of second specific enthalpy HC2a from the first specific enthalpy HC1 predetermined at the entry point P1 and the formula γmin=min[(HC2a-HC1) / 10 6 ] in which γmin is a transfer criterion equal to 2.035. In this step E1, the computer 9 also determines the second terminal of second specific enthalpy HC2b from the first specific enthalpy HC1 predetermined at the entry point P1 and from the formula γmax=max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 6.712.

[0094] In a first alternative implementation mode, in this step E1, the calculator 9 determines the first terminal of second specific enthalpy HC2a and the second terminal of second specific enthalpy HC2b, such that the transfer criteria γmin and γmax are respectively equal to 2.035 and 4.507.

[0095] In a second alternative implementation mode, in this step E1, the calculator 9 determines the first terminal of second specific enthalpy HC2a and the second terminal of second specific enthalpy HC2b, such that the transfer criteria γmin and γmax are respectively equal to 4.034 and 6.712.

[0096] In a second step E2, the computer 9 compares the overall range HC2-G of second specific enthalpies HC2 with a plurality of elementary ranges HC2-1, HC2-2, …, HC2-X of second specific enthalpies HC2 listed in the database BdD. More precisely, in this step E2, the computer 9 determines which elementary range HC2-1, HC2-2, …, HC2-X of second specific enthalpies HC2 is entirely included in the overall range HC2-G of second specific enthalpies HC2, so as to ensure the operation of the conditioning system SC in all the flight phases of the aircraft A. It goes without saying that the computer 9 could alternatively, in this step E2, determine several elementary ranges HC2-1, HC2-2, …, HC2-X entirely included in the overall range HC2-G.

[0097] The method then comprises a step E3 of determination, by the calculator 9, in the database BdD, of a quantity of heat Th to be supplied to the heating module 3 from the elementary range HC2-1, HC2-2, …, HC2-X of second specific enthalpies HC2 determined. It goes without saying that the calculator 9 could associate, in the database BdD, a plurality of quantities of heat Th with the elementary range HC2-1, HC2-2, …, HC2-X determined.

[0098] In a fourth step E4, the computer 9 determines, in the database BdD, the heating module 3 to be used in the conditioning system SC, from the quantity of heat Th to be supplied to said heating module 3, so as to allow optimal heating of the fuel flow Q in the heating module 3 to allow its injection into the combustion chamber of the turbomachine M. In this step, the computer 9 could alternatively determine several possible heating modules 3

[0099] In the case where the heating module 3 is a heat exchanger, the method comprises a fifth step E5 of determining the heat source SA to be used in the heating module 3.

[0100] When the computer 9 determines a plurality of possible heating modules 3 (as is the case in the example of the elementary range HC2-2 of the database BdD represented on the), the person skilled in the art then chooses according to other criteria such as for example the space available in the aircraft.

[0101] In practice, to ensure that the SC conditioning system is correctly sized and meets the γ transfer criteria defined previously: a difference between the second specific enthalpy value HC2 and the first specific enthalpy value HC1, determined for each flight phase, is measured, and the measured difference is divided by 10 6 .

[0102] The first specific enthalpy HC1 and the second specific enthalpy HC2 are determined from a local measurement of the static temperature and static pressure of the fuel stream Q and the use of the appropriate diagrams for the fluid.

[0103] When the transfer criterion γ is included in the calculated interval [γmin; γmax], then the conditioning system SC is optimally sized for all phases of aircraft flight, so as to optimally heat the fuel flow Q for its injection into the gas turbomachine M.

Claims

Method for determining at least one quantity of heat (Th) to be supplied in at least one heating module (3) of a fuel conditioning system (SC) configured to supply an aircraft turbomachine (M) with fuel from a cryogenic tank (R), a fuel flow (Q) circulating between the cryogenic tank (R) and the turbomachine (M) via a fuel circuit (1), the aircraft (A) being configured to operate according to a plurality of flight phases, the heating module (3) being configured to heat the fuel flow (Q) from the quantity of heat (Th), the fuel flow (Q) having: at least a first specific enthalpy (HC1) at an inlet point (P1) positioned on the fuel circuit (1) at the outlet of the cryogenic tank (R), at least a second specific enthalpy (HC2) at an outlet point (P2) positioned on the fuel circuit (1) at the inlet of the turbomachine (M),a database (BdD) of quantities of heat (Th) being accessible, each quantity of heat (Th) being associated with an elementary range (HC2-1, HC2-2, …, HC2-X) of second specific enthalpies (HC2), each elementary range (HC2-X) being sized for the plurality of flight phases of the aircraft (A), the method comprises the steps of: determining a global range (HC2-G) of second specific enthalpies (HC2) defined, between a first terminal of second specific enthalpy (HC2a) and a second terminal of second specific enthalpy (HC2b), from: the first predetermined specific enthalpy (HC1) at the entry point (P1), and the following formulas: γmin = min[(HC2a-HC1) / 10, 6 ] in which γmin is a transfer criterion equal to 2.035,γmax = max[(HC2b-HC1) / 10 6] in which γmax is a transfer criterion equal to 6.712.determine, in the database, the quantity of heat (Th) whose elementary range (HC2-x) of second specific enthalpies (HC2) is entirely included in the global range (HC2-G) of second specific enthalpies (HC2) so as to determine the quantity of heat (Th) to be supplied to the heating module (3). Determination method according to claim 1, in which the database associating for each quantity of heat (Th), at least one heating module (3), the method comprises a step of determining (E1), in the database (BdD), at least one heating module (3) to be used in the conditioning system (SC). Determination method according to one of claims 1 to 2, in which the conditioning system (SC) is defined 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 turbomachine (M) being positioned in the turbomachine reference frame (REF-M), the conditioning system (SC) comprising at least one first heating module (31), configured to be mounted in the aircraft reference frame (REF-A) and to heat the fuel flow (Q) from a first quantity of heat (ThA), and at least one second heating module (32), configured to be mounted in the turbomachine reference frame (REF-M) and to heat the fuel flow (Q) from a second quantity of heat (ThM), the database (BdD) associating for each pair of a plurality of pairs of quantities of heat (ThA, ThM) an elementary range (HC2-1, HC2-2, …,HC2-X) of second specific enthalpies (HC2), each elementary range (HC2-X) being sized for the plurality of flight phases of the aircraft (A), the determination step (E3) of the method makes it possible to determine at least one pair of quantities of heat (ThA, ThM) whose elementary range (HC2-X) is entirely included in the overall range (HC2-G), so as to determine the first quantity of heat (ThA) to be supplied to the first heating module (31) and the second quantity of heat (ThM) to be supplied to the second heating module (32)., Determination method according to one of claims 1 to 3, in which the heating module (3) is chosen, alone or in combination, from: a heat exchanger, an electric heating system, a thermal storage system, a fuel cell, a gas turbine, a burner. Determination method according to one of claims 1 to 4, in which the heating module (3) is a heat exchanger configured to heat the fuel flow (Q) from calories transferred by at least one heat source (SA, SM), the database (BdD) associating for each heating module (3) corresponding to a heat exchanger, a heat source (SA, SM) to be used, the method comprises a step of determining (E5), in the database (BdD), the heat source (SA) to be used in the heating module (3) to heat the fuel flow (Q). Determination method according to one of claims 1 to 5, in which the overall range (HC2-G) of second specific enthalpy (HC2) is determined (E1) from the formulas: γmin = min[(HC2a-HC1) / 10 6 ] in which γmin is a transfer criterion equal to 2.035, and γmax = max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 4.

507. Determination method according to one of claims 1 to 5, in which the overall range (HC2-G) of second specific enthalpy (HC2) is determined (E1) from the formulas: γmin = min[(HC2a-HC1) / 10 6 ] in which γmin is a transfer criterion equal to 4.034, and γmax = max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 6.

712. System (SD) for determining at least one quantity of heat (Th) to be supplied in at least one heating module (3) of a fuel conditioning system (SC) configured to supply an aircraft turbomachine (M) with fuel from a cryogenic tank (R), the determination system (SD) being configured to implement the determination method according to one of claims 1 to 7, the aircraft (A) being configured to evolve according to a plurality of flight phases, the conditioning system (SC) comprising: a fuel circuit (1) connected at the inlet to the cryogenic tank (R) and at the outlet to the turbomachine (M), a fuel flow (Q) circulating in the fuel circuit (1), at least one heating module (3) configured to heat the fuel flow (Q) from the quantity of heat (Th),the fuel flow (Q) having:at least a first specific enthalpy (HC1) at an inlet point (P1) positioned on the fuel circuit (1) at the outlet of the cryogenic tank (R),at least a second specific enthalpy (HC2) at an outlet point (P2) positioned on the fuel circuit (1) at the inlet of the turbomachine (M),the determination system (SD) comprising:a database (BdD) of quantities of heat (Th) being accessible, each quantity of heat (Th) being associated with an elementary range (HC2-1, HC2-2, …, HC2-X) of second specific enthalpies (HC2), each elementary range (HC2-X) being sized for the plurality of flight phases of the aircraft (A),a computer (9) configured to determine a global range (HC2-G) of second specific enthalpies (HC2) defined, between a first terminal of second specific enthalpy (HC2a) and a second terminal of second specific enthalpy (HC2b),from: the first predetermined specific enthalpy (HC1) at the entry point (P1), and the following formulas: γmin = min[(HC2a-HC1) / 10, 6 ] in which γmin is a transfer criterion equal to 2.035,γmax = max[(HC2b-HC1) / 10 6 ] in which γmax is a transfer criterion equal to 6.712.determine, in the database, the quantity of heat (Th) whose elementary range (HC2-x) of second specific enthalpies (HC2) is entirely included in the global range (HC2-G) of second specific enthalpies (HC2) so as to determine the quantity of heat (Th) to be supplied to the heating module (3). Fuel conditioning system (SC) configured to supply a turbomachine (M) of an aircraft (A) with fuel from a cryogenic tank (R), the aircraft (A) being configured to operate according to a plurality of flight phases, the conditioning system (SC) comprising: a fuel circuit (1) configured to be connected at the inlet to the cryogenic tank (R) and at the outlet to the turbomachine (M), a fuel flow (Q) circulating in the fuel circuit (1), at least one heating module (3) configured to heat the fuel flow (Q) from a quantity of heat (Th), the heating module (3) being configured to heat the fuel flow (Q) between an inlet point (P1) positioned on the fuel circuit (1) at the outlet of the cryogenic tank (R) and an outlet point (P2) positioned on the fuel circuit (1) at the inlet of the turbomachine (M),the fuel flow (Q) having a first specific enthalpy (HC1) at the inlet point (P1) and a second specific enthalpy (HC2) at the outlet point (P2), the heating module (3) is configured to heat the fuel flow (Q) between the inlet point (P1) and the outlet point (P2) while respecting the condition that the second specific enthalpy (HC2) is determined from the first specific enthalpy (HC1), the second specific enthalpy (HC2) being included in an interval [HC2a, HC2b] respecting the following formulas: γmin = min[(HC2a-HC1) / 106] in which γmin is a transfer criterion equal to 2.035, γmax = max[(HC2b-HC1) / 106] in which γmax is a transfer criterion equal to 6.712., Aircraft (A) comprising a cryogenic tank (R), a turbomachine (M) and a conditioning system (SC) according to claim 9 for supplying the turbomachine (M) from fuel (Q) from the cryogenic tank (R).

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