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

Abstract

The invention relates to a method for determining a first heat quantity (ThA) and a second heat quantity (ThM) to be provided in a first heating module (31) and a second heating module (31) of a fuel conditioning system (SC), the method comprising a step of determining an overall range of second enthalpy (HC2) of a fuel flow (Q) from the formulas γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073 and γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778; and a step of determining, in a database (BdD) of elementary ranges of second enthalpies (HC2), a pair of heat quantities (ThA; ThM), the elementary range of which is entirely within the overall range so as to determine the first heat quantity (ThA) and the second heat quantity (ThM).

Description

System and method for determining at least two quantities 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, conditions (for example in pressure and temperature) which are known and 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 has, at the inlet of the turbomachine M, conditions which are known and 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 P3 of the fuel circuit 1, outlet conditions C3 (corresponding to the conditions at the inlet of the turbomachine M imposed by an engine manufacturer).

[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 C3, as is known, the conditioning system SC comprises a mechanical pump 2 and one or more heating modules 31, 32. The mechanical pump 2 is configured to circulate the fuel flow Q in the fuel circuit 1 and raise its pressure. Each heating module 31, 32 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, a first heating module 31 belonging to the aircraft reference system REF-A can take heat from the aircraft (air from the cabin, heat from electrical or electronic systems, etc.) while a second heating module 32 belonging to the turbomachine reference system REF-M can take heat from the turbomachine M (heat from the lubricating oil, calories at the turbine outlet, heat from the nozzle, etc.).

[0013] When designing a conditioning system SC, it is necessary to choose the heating modules 31, 32 in order to allow the output conditions C3 to be reached from the inlet conditions C1 for any flight phase of the aircraft A. This is complex since the heating modules 31, 32 do not provide the same amount of heat for all flight phases. For example, 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 modules 31, 32 but this would increase the mass and the size.

[0014] The design of a conditioning system SC must also take into account the constraints related to the connecting pipes L which connect the aircraft reference frame REF-M to the turbomachine reference frame REF-M for the circulation of fuel. In particular, the connecting pipes L connect the first heating module 31 to the second heating module 32. It was initially proposed to use cryogenic connecting pipes L to conduct a flow of liquid fuel Q but the latter are complex, heavy, bulky and expensive to conduct fuel in the liquid state without impacting the other fluids circulating in the vicinity, in particular when the connecting pipe L extends into a wing of the aircraft A as illustrated in. To eliminate this drawback, it was proposed to use non-cryogenic connecting pipes L to conduct a flow of gaseous fuel.This imposes further constraints on the design of a conditioning system SC since it is necessary to ensure the inlet conditions of the connecting lines L, hereinafter transfer conditions, so that the fuel is gaseous and sufficiently cold to avoid any oversizing of the connecting lines L. In addition, it is important that the fuel flow Q is 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 heating modules 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 heating modules that meet the requirements of a fuel conditioning system, in particular meeting the inlet conditions, the outlet conditions but also the transfer conditions, for any phase of flight of the aircraft. This advantageously makes it possible to limit the variation in the conditions of use of the connecting pipes, which makes it possible to limit their weight and their complexity of implementation. PRESENTATION OF THE INVENTION

[0017] The invention relates to a method for determining at least a first quantity of heat and at least a second quantity of heat to be supplied respectively in at least a first heating module and at least a second 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 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 aircraft being configured to evolve according to a plurality of flight phases,the first heating module being mounted in the aircraft reference frame and configured to heat the fuel flow to a circulation temperature from the first quantity of heat,the second heating module being mounted in the turbomachine reference frame downstream of the first heating module, the second heating module being configured to heat the fuel flow to an injection temperature from the second quantity of heat,at least one connecting pipe connecting the first heating module to the second heating module, the connecting pipe being configured to circulate a fuel flow in a gaseous single-phase state,the fuel flow having:at least a first predetermined enthalpy at an entry point positioned on the fuel circuit at the outlet of the cryogenic tank,at least a second enthalpy at a transfer point positioned on the fuel circuit at the inlet of the connecting pipe,at least a third predetermined enthalpy at an outlet point positioned on the fuel circuit at the inlet of the turbomachine,a database of pairs of a first quantity of heat and a second quantity of heat being accessible, each pair of quantities of heat being associated with an elementary range of second enthalpy, each elementary range being sized for the plurality of flight phases of the aircraft.,

[0018] The method comprises the steps of:determining an overall second enthalpy range defined between a first second enthalpy terminal and a second second enthalpy terminal, from:the first predetermined enthalpy at the entry point,the third predetermined enthalpy at the exit point, andthe following formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073,γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778,determining, in the database, at least one pair of heat quantities whose elementary range is entirely included in the overall range so as to determine the first heat quantity and the second heat quantity.

[0019] The method according to the invention makes it possible 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 to enable optimal distribution of the heating of the fuel flow in the aircraft reference frame and in the turbomachine reference frame. Such a distribution between the first quantity of heat and the second quantity of heat makes it possible to reach an injection temperature in the turbomachine while maintaining the heating power in the first heating module, which makes it possible to ensure, whatever the speed of the turbomachine, transport of the fuel flow between the aircraft reference frame and the turbomachine reference frame at optimal temperatures.In particular, the temperatures of the fuel flow between the aircraft reference frame and the turbomachine reference frame are both high enough to eliminate heavy and bulky cryogenic lines and low enough to avoid any oversizing of the elements of the conditioning system. In addition, the temperatures of the fuel flow in the fuel circuit can advantageously vary depending on the flight phases and the turbomachine speed without this penalizing the conditioning system and its operation.

[0020] The distribution of heat quantities also allows the fuel flow to be in a single-phase or supercritical gaseous state, which helps ensure flow stability in the fuel system lines. This makes the conditioning system safe and secure.

[0021] The heat available in the aircraft frame is also advantageously used while being limited to avoid any risk of oversizing the connecting pipe. In addition, optimal heating in the aircraft frame ensures that the fuel flow temperature at the aircraft frame outlet is sufficient to avoid any risk of damage to materials exposed to the cold, such as the aircraft wings, through which the fuel flow generally passes.

[0022] The heat available in the turbomachine reference is also advantageously limited, which makes it possible to limit the volume of the heating modules and the associated losses (nozzle pressure drop for a nozzle exchanger for example).

[0023] The method according to the invention thus allows a significant saving of time when producing a conditioning system by automatically determining, from the inlet conditions and the outlet conditions of the fuel flow, all of the elements which allow an optimal distribution of the heating of the fuel flow between the aircraft reference frame and the turbomachine reference frame.

[0024] Preferably, the database associating for each pair of heat quantities, at least one pair of a first heating module and a second heating module, the method comprises a step of determining, in the database, at least one pair of heating modules. The method according to the invention thus makes it possible to choose from a plurality of possible heating modules, the heating modules to be used in the aircraft reference system and in the turbomachine reference system to heat the fuel flow optimally. Thanks to the invention, the choice of heating modules is no longer carried out empirically as was the case in the prior art, which allows a significant saving of time, while limiting costs.

[0025] Preferably, each 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.

[0026] In one embodiment, at least one of the first heating module and / or the second 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 heat exchanger 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 amount 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.

[0027] In one embodiment, the overall range of second enthalpies is determined from the formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.309, andγRmax=max[(HC2b-HC1) / (HC3-HC1)], in which the transfer criterion is equal to 0.681.

[0028] Such a range of transfer criteria allows a balanced distribution of the heating of the fuel flow between the first heating module in the aircraft reference frame and the second heating module in the turbomachine reference frame. Such a balanced distribution allows efficient maintenance of the heating power of the heating modules which ensures, regardless of the engine speed, transport of hydrogen in the connecting pipe, at free but sufficiently high temperatures up to the engine reference frame. A balanced distribution also allows efficient exploitation of the heat sources available in the aircraft reference frame.Furthermore, a balanced distribution advantageously allows the use of the heat sources present in the aircraft reference frame and in the turbomachine reference frame, while minimizing the use of the heat sources available in the aircraft reference frame and limiting the heat taken from the turbomachine reference frame, which makes it possible to limit the volume and the size of the heating modules.

[0029] Such transfer criteria values ​​ensure that the fuel flow is in a gaseous single-phase state or in a supercritical state, which helps to limit the use of cryogenic circulation lines. This also helps to limit the occurrence of hydraulic instabilities that would be due to two-phase or transcritical flow.

[0030] Such transfer criteria values ​​also make it possible to ensure a fuel flow temperature level sufficient to move away from the ductile-brittle transition temperature of the fuel, which advantageously makes it possible to limit the brittleness of the materials of the cold conditioning system, allowing its service life to be extended.

[0031] In a first alternative embodiment, the overall range of second enthalpies is determined from the formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.157, andγRmax=max[(HC2b-HC1) / (HC3-HC1)], in which the transfer criterion is equal to 0.778.

[0032] Such a range of transfer criteria allows efficient maintenance of the heating power of the heating modules which makes it possible to ensure, regardless of the engine speed, heating of the hydrogen to sufficient temperatures to allow the fuel flow to be transported safely to the turbomachine perimeter. Such a range of values ​​also makes it possible to efficiently exploit the heat sources available in the aircraft reference system. In addition, such a range of values ​​advantageously makes it possible to use the heat sources present in the aircraft reference system and in the turbomachine reference system, while minimizing the use of the heat sources available in the aircraft reference system and limiting the heat taken from the turbomachine reference system, which makes it possible to limit the volume and the size of the heating modules.

[0033] Such transfer criteria values ​​ensure that the fuel flow is in a gaseous single-phase state or in a supercritical state, which helps to limit the use of cryogenic circulation lines. This also helps to limit the occurrence of hydraulic instabilities that would be due to two-phase or transcritical flow.

[0034] Such transfer criteria values ​​also make it possible to ensure a fuel flow temperature level sufficient to move away from the ductile-brittle transition temperature of the fuel, which advantageously makes it possible to limit the brittleness of the materials of the cold conditioning system, allowing its service life to be extended.

[0035] In a second alternative embodiment, the overall range of second enthalpies is determined from the formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.073, andγRmax=max[(HC2b-HC1) / (HC3-HC1)], in which the transfer criterion is equal to 0.401.

[0036] Such a range of transfer criteria allows efficient maintenance of the heating power of the heating modules which makes it possible to ensure, regardless of the engine speed, heating of the hydrogen to minimum temperatures to allow the fuel flow to be transported safely to the turbomachine perimeter. Such a range of values ​​also makes it possible to advantageously exploit the hot sources available in the aircraft reference system. In addition, such a range of values ​​makes it possible to use the hot sources present in the aircraft reference system and in the turbomachine reference system, while minimizing the use of the hot sources available in the aircraft reference system and limiting the heat taken from the turbomachine reference system, which makes it possible to limit the volume and the size of the heating modules.

[0037] Such transfer criteria values ​​ensure that the fuel flow is in a gaseous single-phase state or in a supercritical state, which helps to limit the use of cryogenic circulation lines. This also helps to limit the occurrence of hydraulic instabilities that would be due to two-phase or transcritical flow.

[0038] The invention also relates to a system for determining at least a first quantity of heat and at least a second quantity of heat to be supplied respectively in at least a first heating module and at least a second 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 previously, 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 aircraft being configured to evolve 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 first heating module mounted in the aircraft reference frame, the first heating module being configured to heat the fuel flow to a circulation temperature from the first quantity of heat,at least one second heating module mounted downstream of the first heating module in the turbomachine reference frame, the second heating module being configured to heat the fuel flow to an injection temperature from the second quantity of heat,at least one connecting pipe connecting the first heating module to the second heating module, the connecting pipe being configured to circulate a fuel flow in a gaseous single-phase state,the fuel flow having:at least a first predetermined enthalpy at an entry point positioned on the fuel circuit at the outlet of the cryogenic tank,at least a second enthalpy at a transfer point positioned on the fuel circuit at the inlet of the connecting pipe,at least a third predetermined enthalpy at an exit point positioned on the fuel circuit at the inlet of the turbomachine.the determination system comprising:a database of elementary ranges of second enthalpies, each elementary range being sized for the plurality of flight phases of the aircraft, the database being configured to associate for each elementary range, at least one pair of a first quantity of heat and a second quantity of heat,a calculator configured to:determine an overall range of second enthalpy between a first terminal of second enthalpy and a second terminal of second enthalpy,from:the first predetermined enthalpy at the entry point,the third predetermined enthalpy at the exit point, andthe following formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073,γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778,determine, in the database, at least one pair of heat quantities whose elementary range is entirely included in the global range.,

[0039] 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 distribution of the first quantity of heat to be supplied to the first heating module in the aircraft reference system and of the second quantity of heat to be supplied to the second heating module in the turbomachine reference system, to heat a flow of fuel from a cryogenic tank with a view to its injection into an aircraft turbomachine.

[0040] The invention also relates to a fuel conditioning system configured to supply an aircraft turbomachine with fuel from a cryogenic tank, 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 aircraft being configured to evolve 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 first heating module mounted in the aircraft reference frame, the first heating module being configured to heat the fuel flow to a circulation temperature from the first quantity of heat,at least one second heating module mounted downstream of the first heating module in the turbomachine reference system, the second heating module being configured to heat the fuel flow to an injection temperature from the second quantity of heat,at least one connecting pipe connecting the first heating module to the second heating module, the connecting pipe being configured to circulate a fuel flow in a gaseous single-phase state,the fuel flow having:at least one first predetermined enthalpy at an entry point positioned on the fuel circuit at the outlet of the cryogenic tank,at least one second enthalpy at a transfer point positioned on the fuel circuit at the inlet of the connecting pipe,at least one third predetermined enthalpy at an exit point positioned on the fuel circuit at the inlet of the turbomachine,conditioning system in which,the second enthalpy of the fuel flow is included in an interval [HC2a, HC2b] respecting the following formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073,γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778.,

[0041] In one embodiment, the second enthalpy of the fuel flow in the range [HC2a, HC2b] complies with the following formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion γRmin is equal to 0.309, andγRmax=max[(HC2b-HC1) / (HC3-HC1)] in which the transfer criterion γRmax is equal to 0.681.

[0042] In a first alternative embodiment, the second enthalpy of the fuel flow included in the interval [HC2a, HC2b] complies with the following formulas: γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion γRmin is equal to 0.157, and γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which the transfer criterion γRmax is equal to 0.778.

[0043] In a second alternative embodiment, the second enthalpy of the fuel flow included in the interval [HC2a, HC2b] complies with the following formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion γRmin is equal to 0.073, andγRmax=max[(HC2b-HC1) / (HC3-HC1)] in which the transfer criterion γRmax is equal to 0.401.

[0044] 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

[0045] 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.

[0046] This is a schematic representation of an aircraft including a fuel conditioning system.

[0047] This is a schematic representation of the packaging system of the.

[0048] This is a schematic representation of a determination system according to one embodiment of the invention.

[0049] This is a schematic representation of a database associating for each elementary range of second enthalpy, at least one pair of heat quantities.

[0050] This is a schematic representation of a global second enthalpy range and a plurality of elementary second enthalpy ranges from the database.

[0051] This is a diagram of the steps of a determination method according to an embodiment of the invention.

[0052] 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

[0053] 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.

[0054] 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.

[0055] The turbomachine M is configured to provide propulsion for the aircraft A, in particular, by driving at least one propulsion member (not shown in the). 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.

[0056] 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.

[0057] 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.

[0058] 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.

[0059] 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. Similarly, the fuel flow Q has, at the inlet of the turbomachine M, known conditions imposed by an engine manufacturer and 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.

[0060] 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 P3 of the fuel circuit 1, positioned directly before the inlet into the turbomachine M, predetermined outlet conditions C3 corresponding to the inlet conditions of the turbomachine M.

[0061] In particular, the fuel flow Q has at the entry point P1, a first predetermined enthalpy HC1. More precisely, the fuel flow Q has at the entry point P1 a first enthalpy HC1 determined for example for a predetermined reference enthalpy value at liquid saturation and at the boiling point, i.e. for a pressure of 101325 Pa. The first enthalpy HC1 varies within a range of first enthalpies HC1 valid for all phases of flight of the aircraft.

[0062] As is known, the 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 enthalpy value to be determined as a function of the type of fluid and the static pressure and temperature measurements.

[0063] For example, it is possible to calculate a variation in HC enthalpy from the following formula: , in which: is the pressure variation between the entry point P1 and the exit point P3, is the temperature variation between the entry point P1 and the exit point P3, is the specific heat capacity at constant pressure, is the density, is the static temperature, and is the coefficient of thermal expansion.

[0064] The fuel flow Q also has a predetermined third enthalpy HC3 at the outlet point P3. Analogously to the inlet point P1, the fuel flow Q has a third enthalpy HC3 at the outlet point P3, which varies within a range of third enthalpies HC3, determined for example for a predetermined reference enthalpy value at liquid saturation and at the boiling point, i.e. for a pressure of 101325 Pa. The range of third enthalpies HC3 is valid for all flight phases of the aircraft.

[0065] For example, for a para-hydrogen, the first enthalpy HC1 of the fuel flow Q at the entry point P1 is in the interval [-29,942; 49,693] J / kg and the third enthalpy HC3 at the exit point P3 is in the interval [2,084,731; 6,681,802] J / kg for a zero reference enthalpy value HC=0 at liquid saturation and boiling point, i.e. for a pressure of 101325 Pa.

[0066] To enable the fuel flow Q to be conditioned, i.e. to enable its passage from the inlet conditions C1 to the outlet conditions C3, the conditioning system SC comprises a mechanical pump 2 and two heating modules 31, 32.

[0067] 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.

[0068] Each heating module 31, 32 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 two heating modules 31, 32, however the conditioning system SC could alternatively comprise a different number of heating modules 31, 32, in particular a number greater than two.

[0069] In this example, a first heating module 31 is mounted in the aircraft reference frame REF-A and is configured to heat the fuel flow Q to a circulation temperature Tc. The circulation temperature Tc corresponds to a temperature above a cryotemperature threshold, so as to avoid the installation of heavy and bulky cryogenic lines between the aircraft reference frame REF-A and the turbomachine reference frame REF-M. In this example, the circulation temperature Tc is between 44.5 and 299.7°C. A second heating module 32 is mounted in the turbomachine reference frame REF-M and is configured to heat the fuel flow Q to an injection temperature Ti. The injection temperature Ti is above the circulation temperature Tc and 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. This document describes a conditioning system SC comprising 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; it goes without saying that the conditioning system SC could comprise a different number of heating modules 31 mounted in the aircraft reference frame REF-A and / or a different number of heating modules 32 mounted in the turbomachine reference frame REF-M.

[0070] In this example, the first heating module 31 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. The second heating module 32 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.

[0071] Each heating module 31, 32 is preferably a heat exchanger, an electric heating system, a thermal storage system, a fuel cell, a gas turbine or a burner. Each heating module 31, 32 is configured to provide a quantity of its own heat.

[0072] In particular, the first heating module 31 is configured to provide a first quantity of heat ThA making it possible to heat the fuel flow Q up to the circulation temperature Tc. 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.

[0073] To connect the first heating module 31 and the second heating module 32, the conditioning system SC comprises a connecting pipe L, as shown in the. The connecting pipe L is configured to circulate a fuel flow Q in a gaseous single-phase state. In other words, the connecting pipe L is not a cryogenic pipe, which makes it possible to dispense with a heavy and bulky connecting pipe L, as described previously. In addition, a single-phase state makes it possible to limit flow instabilities in the connecting pipe L.

[0074] The fuel flow Q has, at a transfer point P2 of the fuel circuit 1, positioned directly at the inlet of the connecting pipe L, i.e. at the outlet of the first heating module 31, transfer conditions C2. In particular, the fuel flow Q has, at the transfer point P2, a second enthalpy HC2. More precisely, the fuel flow Q has at the transfer point P2 a second enthalpy HC2 which varies in a second enthalpy HC2 interval determined for example for a zero reference enthalpy value at liquid saturation and at the boiling point, i.e. for a pressure of 101325 Pa. The second enthalpy HC2 interval is valid for all flight phases of the aircraft.

[0075] The second enthalpy HC2 is configured to be determined from the first enthalpy HC1 at the inlet point P1, the third enthalpy HC3 at the outlet point P3 and a transfer criterion γR, according to which γR=(HC2-HC1) / (HC3-HC1). In practice, the second enthalpy HC2 depends on the first heat quantity ThA supplied by the first heating module 31 and the second heat quantity ThM supplied by the second heating module 32, as will be described in more detail later.

[0076] According to one aspect of the invention, the transfer criterion γR is defined according to a minimum transfer criterion γRmin and a maximum transfer criterion γRmax, so as to cover all the flight phases of the aircraft A. In practice, the transfer criterion γR depends on the definition intervals of each enthalpy HC1, HC2, HC3 which vary according to the flight phases of the aircraft. In other words, the transfer criterion γR is defined in a predetermined range [γRmin; γRmax]. According to one aspect of the invention, the transfer criterion γR for determining a range of second enthalpies HC2 is defined in a predetermined range [0.073; 0.778]. Such a range of transfer criterion γR advantageously allows an optimal distribution of the heating between the aircraft reference frame REF-A and the turbomachine reference frame REF-M.

[0077] In a first alternative embodiment, the transfer criterion γR for determining a range of second enthalpies HC2 is defined in the predetermined range [0.309; 0.681]. Such a range makes it possible to ensure that the fuel flow Q circulates in the fuel circuit 1 at satisfactory temperatures both in the aircraft reference frame REF-A and in the turbomachine reference frame REF-M.

[0078] In a second alternative embodiment, the transfer criterion γR for determining a range of second enthalpies HC2 is defined in the predetermined range [0.157; 0.778]. Such a range makes it possible to ensure that the fuel flow Q circulates between the aircraft reference frame and the turbomachine reference frame at free but sufficiently high temperatures to allow optimal distribution of the heating between the first heating module 31 and the second heating module 32.

[0079] In a third alternative embodiment, the transfer criterion γR for determining a range of second enthalpies HC2 is defined in the predetermined range [0.073; 0.401]. Such a range makes it possible to ensure that the fuel flow Q circulates between the aircraft reference frame and the turbomachine reference frame at a minimum temperature to allow the need for cryogenic circulation lines to be dispensed with.

[0080] There is also shown on the, a system for determining SD, according to an embodiment of the invention, the first quantity of heat ThA and the second quantity of heat ThM to be supplied respectively in the first heating module 31 and in the second heating module 32, to make it possible to obtain at the transfer point P2, a second enthalpy HC2 which is sized to form an optimal conditioning system SC.

[0081] 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 enthalpies HC2, and a calculator 9, configured to determine the first quantity of heat ThA and the second quantity of heat ThM to be supplied respectively in the first heating module 31 and the second heating module 32.

[0082] In the BdD database, each elementary range HC2-1, HC2-2, … HC2-X of second HC2 enthalpies 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 HC2 enthalpies is configured to allow optimal operation of the SC conditioning system regardless of the flight phase of the aircraft A.

[0083] As shown in the, the BdD database associates for each elementary range HC2-1, HC2-2, … HC2-X of second enthalpies HC2 one or more pairs of a first quantity of heat ThA and a second quantity of heat ThM.

[0084] The database BdD also associates for each pair of a first quantity of heat ThA and a second quantity of heat ThM, a pair of heating modules 31, 32 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 enthalpies HC2, the type of heating modules 31, 32 to be mounted in the conditioning system SC, and in particular to be mounted in the aircraft reference frame REFA and in the turbomachine reference frame REF-M, to allow the fuel flow Q to be in a gaseous single-phase state in the connecting pipe L, which makes it possible to ensure a conditioning system SC which is defined optimally according to the inlet conditions C1 and the outlet conditions C3.Alternatively, the database BdD could also be configured to associate with each pair of heat quantities ThA, ThM a plurality of possible pairs of heating modules 31, 32.

[0085] For each heating module 31, 32 corresponding to a heat exchanger, configured to heat the fuel flow Q from the calories supplied by a hot source SA, SM 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 31, 32, a hot source SA, SM to be used. Alternatively, the database BdD could also be configured to associate with each pair of heating modules 31, 32, a plurality of possible pairs of hot sources SA, SM.

[0086] By way of example, with reference to the, for an elementary range HC2-2 of second enthalpies HC2, the database BdD is configured to associate three possible pairs (ThA-21; ThM-21), (ThA-22; ThM-22), (ThA-23; ThM-23) of a first quantity of heat ThA and a second quantity of heat ThM to be supplied respectively in the first heating module 31 and in the second heating module 32. The database BdD is also configured to associate with each pair of quantities of heat (ThA; ThM) a pair of heating modules 31, 32 making it possible to supply such quantities of heat ThA, ThM.In one example, the first pair of heating modules (31-21; 32-21) corresponds to an electric heating system mounted in the aircraft reference frame REF-A and to a fuel cell mounted in the turbomachine reference frame REF-M, the second pair of heating modules (31-22; 32-22) corresponds to a heat exchanger mounted in the aircraft reference frame REF-A and to a fuel cell mounted in the turbomachine reference frame REF-M, and the third pair of heating modules (31-23; 32-23) corresponds to an electric heating system mounted in the aircraft reference frame REF-A and to a heating gas turbine mounted in the turbomachine reference frame REF-M. In this example, in which the second pair of heating modules (31-22; 32-22) comprises a heat exchanger mounted in the aircraft reference REF-A, the database BdD is also configured to associate with the heating module 31-22 several heat sources SA1, SA2 to be used.

[0087] 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 enthalpies HC2 could alternatively be associated with a different number of pairs of quantities of heat ThA, ThM. Likewise, each pair of heat quantities ThA, ThM could alternatively be associated with a different number of pairs of heating modules 31, 32. Furthermore, there are shown on the figure two examples of pairs of heating modules 31, 32 in which a single heating module 31-22, 32-3 is a heat exchanger with which several heat sources SA1, SA2, SM1, SM2 are associated, however it goes without saying that the pair of heating modules 31, 32 could just as well correspond to two heat exchangers, one mounted in the aircraft reference frame REF-A and the other mounted in the turbomachine reference frame REF-M.In this case, the database BdD would be configured to associate with the pair of heating modules 31, 32 one or more pairs of hot sources SA, SM. Similarly, the database BdD could alternatively associate with the pair of heating modules 31, 32 comprising a heat exchanger, a single hot source SA, SM to be used.

[0088] As described previously, the determination system SD also comprises a computer 9. The computer 9 is configured to determine for the conditioning system SC the heating modules 31, 32 to be mounted respectively in the aircraft reference system REF-A and in the turbomachine reference system REF-M from the inlet conditions C1 and the outlet conditions C3 of the fuel flow Q. For this, the computer 9 is configured to determine a range of second enthalpies [HC2a-HC2b] necessary at the transfer point P2 to ensure optimal heating of the fuel flow Q in the conditioning system SC.

[0089] More specifically, the computer 9 is configured to:determine a first terminal of second enthalpy HC2a from the first enthalpy HC1 predetermined at the input point P1, the third enthalpy HC3 predetermined at the output point P3 and the formula γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073,determine a second terminal of second enthalpy HC2b from the first enthalpy HC1 predetermined at the input point P1, the third enthalpy HC3 predetermined at the output point P3 and the formula γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778, anddetermine an overall range HC2-G of second enthalpy HC2 between the first terminal of second enthalpy HC2a and the second terminal of second enthalpy HC2b determined.

[0090] In particular, in the transfer criterion γRmin=min[(HC2a-HC1) / (HC3-HC1)] the minimum γRmin is sought for all the values ​​of first enthalpies HC1 in the interval of first enthalpies HC1 valid for all the flight phases of the aircraft, for all the values ​​of second enthalpies HC2 in the interval of second enthalpies HC2 valid for all the flight phases of the aircraft and for all the values ​​of third enthalpies HC3 in the interval of third enthalpies HC3 valid for all the flight phases of the aircraft.Similarly, in the transfer criterion γRmax=max[(HC2b-HC1) / (HC3-HC1)] the maximum γRmax is sought for all the values ​​of first enthalpies HC1 in the interval of first enthalpies HC1 valid for all the flight phases of the aircraft, for all the values ​​of second enthalpies HC2 in the interval of second enthalpies HC2 valid for all the flight phases of the aircraft and for all the values ​​of third enthalpies HC3 in the interval of third enthalpies HC3 valid for all the flight phases of the aircraft.

[0091] In one embodiment, the calculator 9 is configured to determine the second enthalpy limits HC2 from the preceding formulas, in which the transfer criterion γRmin is equal to 0.309 and the transfer criterion γRmax is equal to 0.681.

[0092] In a first alternative embodiment, the calculator 9 is configured to determine the second enthalpy limits HC2 from the preceding formulas, in which the transfer criterion γRmin is equal to 0.157 and the transfer criterion γRmax is equal to 0.778.

[0093] In a second alternative embodiment, the calculator 9 is configured to determine the second enthalpy limits HC2 from the preceding formulas, in which the transfer criterion γRmin is equal to 0.073 and the transfer criterion γRmax is equal to 0.401.

[0094] The calculator 9 is also configured to determine, among the plurality of elementary ranges HC2-1, HC2-2, …, HC2-X of second HC2 enthalpies 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 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 HC2 enthalpies according to the determined global range HC2-G of second HC2 enthalpies.

[0095] According to one aspect of the invention, the computer 9 is configured to determine, from the database BdD: the pair(s) of heat quantities ThA, ThM corresponding to the determined elementary range HC2-X, the pair(s) of heating modules 31, 32 to be mounted in the aircraft reference frame REF-A and in the turbomachine reference frame REF-M, and when one or more heating modules 31, 32 is a heat exchanger, the heat source(s) SA, SM to be used in the heat module(s) 31, 32.

[0096] The determination system SD according to the invention advantageously makes it possible to determine the heating modules 31, 32 to be used in the conditioning system SC so as to best distribute the quantity of heat to be supplied to the fuel flow Q in the aircraft reference frame REF-A and in the turbomachine reference frame REF-M to allow optimal heating of the fuel flow Q while avoiding the use of cryogenic lines and optimizing the mass and size of the conditioning system SC.

[0097] A method for determining a first quantity of heat ThA and a second quantity of heat ThM to be supplied respectively in a first heating module 31 and a second heating module 32 of a fuel conditioning system SC of an aircraft A will now be described, with reference to, according to an embodiment of the invention.

[0098] In this example, in a preliminary step E0, the inlet conditions C1 at the inlet point P1 of the fuel circuit 1 and the outlet conditions C3 at the outlet point P3 of the fuel circuit 1 are predetermined and known. In particular, the first enthalpy HC1 at the inlet point P1 and the third enthalpy HC3 at the outlet point P3 are known for the different flight phases of the aircraft.

[0099] The method comprises a first step E1 of determining, by the computer 9, a global range HC2-G of second enthalpies HC2 defined between a first terminal of second enthalpy HC2a and a second terminal of second enthalpy HC2b. For this, the computer 9 determines, in this step E1, the first terminal of second enthalpy HC2a from the first enthalpy HC1 predetermined at the entry point P1, from the third enthalpy HC3 predetermined at the exit point P3 and from the formula γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073. In this step E1, the calculator 9 also determines the second terminal of second enthalpy HC2b from the first enthalpy HC1 predetermined at the entry point P1, from the third enthalpy HC3 predetermined at the exit point P3 and from the formula γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778.

[0100] In a first alternative implementation mode, in this step E1, the calculator 9 determines the first terminal of second enthalpy HC2a and the second terminal of second enthalpy HC2b, such that the transfer criteria γRmin and γRmax are respectively equal to 0.309 and 0.681.

[0101] In a second alternative implementation mode, in this step E1, the calculator 9 determines the first terminal of second enthalpy HC2a and the second terminal of second enthalpy HC2b, such that the transfer criteria γRmin and γRmax are respectively equal to 0.157 and 0.778.

[0102] In a third alternative implementation mode, in this step E1, the calculator 9 determines the first terminal of second enthalpy HC2a and the second terminal of second enthalpy HC2b, such that the transfer criteria γRmin and γRmax are respectively equal to 0.073 and 0.401.

[0103] In a second step E2, the computer 9 compares the overall range HC2-G of second enthalpies HC2 with a plurality of elementary ranges HC2-1, HC2-2, …, HC2-X of second 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 enthalpies HC2 is entirely included in the overall range HC2-G of second 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.

[0104] The method then comprises a step E3 of determination, by the computer 9, in the database BdD, of a pair of quantities of heat ThA, ThM to be supplied to the first heating module 31 and to the second heating module 32 from the elementary range HC2-1, HC2-2, …, HC2-X of second enthalpies HC2 determined. It goes without saying that the computer 9 could associate, in the database BdD, a plurality of pairs of quantities of heat ThA, ThM with the elementary range HC2-1, HC2-2, …, HC2-X determined.

[0105] In a fourth step E4, the computer 9 determines, in the database BdD, the first heating module 31 and the second heating module 32 to be used in the conditioning system SC, from the pair of heat quantities ThA, ThM to be provided in said first heating module 31 and in said second heating module 32, so as to distribute the heating of the fuel flow Q optimally between the aircraft reference frame REF-A and the turbomachine reference frame REF-M in the conditioning system SC. In this step, the computer 9 could alternatively determine different possible pairs of heating modules 31, 32.

[0106] In the case in which one or more heating module(s) 31, 32 is a heat exchanger, the method comprises a fifth step E5 of determining the heat source(s) to be used in each heating module 31, 32 of the heat exchanger type.

[0107] When the calculator 9 determines a plurality of possible pairs of heating modules 31, 32 (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.

[0108] In practice, to ensure that the SC conditioning system is correctly sized and meets the γR transfer criteria defined previously: a first difference between the second enthalpy value HC2 and the first enthalpy value HC1, determined for each flight phase, is measured, a second difference between a third enthalpy value HC3 and a first enthalpy value HC1, determined for each flight phase, is measured, and the ratio between the first difference (HC2-HC1) and the second difference (HC3-HC1) is calculated.

[0109] 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.

[0110] When the transfer criterion γR is included in the calculated interval [γRmin; γRmax], then the conditioning system SC is optimally sized for all phases of aircraft flight, so as to optimally heat the fuel flow Q between the cryogenic tank R and the gas turbine engine M.

Claims

Method for determining at least a first quantity of heat (ThA) and at least a second quantity of heat (ThM) to be supplied respectively in at least a first heating module (31) and at least a second heating module (32) of a fuel conditioning system (SC) configured to supply a turbomachine (M) of an aircraft (A) 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 conditioning system (SC) being 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 aircraft (A) being configured to evolve according to a plurality of flight phases,the first heating module (31) being configured to be mounted in the aircraft reference frame (REF-A) and to heat the fuel flow (Q) to a circulation temperature (Tc) from the first quantity of heat (ThA),the second heating module (32) being configured to be mounted in the turbomachine reference frame (REF-M) downstream of the first heating module (31), the second heating module (32) being configured to heat the fuel flow (Q) to an injection temperature (Ti) from the second quantity of heat (ThM),at least one connecting pipe (L) connecting the first heating module (31) to the second heating module (32), the connecting pipe (L) being configured to circulate a fuel flow (Q) in a gaseous single-phase state,the fuel flow (Q) having:at least a first predetermined 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 enthalpy (HC2) at a transfer point (P2) positioned on the fuel circuit (1) at the inlet of the connecting pipe (L),at least a third predetermined enthalpy (HC3) at an outlet point (P3) positioned on the fuel circuit (1) at the inlet of the turbomachine (M),a database (BdD) of pairs of a first quantity of heat (ThA) and a second quantity of heat (ThM) being accessible, each pair of quantities of heat (ThA, ThM) being associated with an elementary range (HC2-1, HC2-2, …, HC2-X) of second enthalpy (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 (E1) a global range (HC2-G) of second enthalpy (HC2) defined between a first second enthalpy terminal (HC2a) and a second second enthalpy terminal (HC2b), from:the first predetermined enthalpy (HC1) at the entry point (P1),the third predetermined enthalpy (HC3) at the exit point (P3), andthe following formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073,γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778,determining (E3), in the database (BdD), at least one pair of heat quantities (ThA, ThM) whose elementary range (HC2-X) is entirely within the global range (HC2-G) so as to determine the first quantity of heat (ThA) and the second quantity of heat (ThM)., Determination method according to claim 1, in which the database (BdD) associating for each pair of heat quantities (ThA, ThM), at least one pair of a first heating module (31) and a second heating module (32), the method comprises a step of determining (E4), in the database (BdD), at least one pair of heating modules (31, 32). Determination method according to one of claims 1 to 2, in which each heating module (31, 32) 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 claim 3, wherein at least one of the first heating module (31) and / or the second heating module (32) is a heat exchanger configured to heat the fuel flow (Q) from calories transferred by at least one hot source (SA, SM), the database (BdD) associating for each heating module (31, 32) corresponding to a heat exchanger, a hot source (SA, SM) to be used, the method comprises a step of determining (E5), in the database (BdD), the hot source (SA, SM) to be used in the heat exchanger to heat the fuel flow (Q). Determination method according to one of claims 1 to 4, in which the overall range (HC2-G) of second enthalpy (HC2) is determined (E1) from the formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.309, andγRmax=max[(HC2b-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.

681. Determination method according to one of claims 1 to 4, in which the overall range (HC2-G) of second enthalpy (HC2) is determined (E1) from the formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.157, andγRmax=max[(HC2b-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.

778. Determination method according to one of claims 1 to 4, in which the overall range (HC2-G) of second enthalpy (HC2) is determined (E1) from the formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.073, andγRmax=max[(HC2b-HC1) / (HC3-HC1)] in which the transfer criterion is equal to 0.

401. System (SD) for determining at least a first quantity of heat (ThA) and at least a second quantity of heat (ThM) to be supplied respectively in at least a first heating module (31) and at least a second heating module (32) of a fuel conditioning system (SC) configured to supply a turbomachine (M) of an aircraft (A) from 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 conditioning system (SC) being 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 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 first heating module (31) mounted in the aircraft reference frame (REF-A), the first heating module (31) being configured to heat the fuel flow (Q) to a circulation temperature (Tc) from the first quantity of heat (ThA),at least one second heating module (32) mounted downstream of the first heating module (31) in the turbomachine reference frame (REF-M), the second heating module (32) being configured to heat the fuel flow (Q) to an injection temperature (Ti) from the second quantity of heat (ThM),at least one connecting pipe (L) connecting the first heating module (31) to the second heating module (32),the connecting pipe (L) being configured to circulate a fuel flow (Q) in a gaseous single-phase state, the fuel flow (Q) having: at least a first predetermined 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 enthalpy (HC2) at a transfer point (P2) positioned on the fuel circuit (1) at the inlet of the connecting pipe (L), at least a third predetermined enthalpy (HC3) at an outlet point (P3) positioned on the fuel circuit (1) at the inlet of the turbomachine (M), the determination system (SD) comprising: a database (BdD) of elementary ranges (HC2-1, HC2-2, …, HC2-X) of second enthalpies (HC2), each elementary range (HC2-X) being sized for the plurality of flight phases of the aircraft (A), the database (BdD) being configured to associate for each elementary range (HC2-X),at least one pair of a first quantity of heat (ThA) and a second quantity of heat (ThM),a calculator (9) configured to:determine a global range (HC2-G) of second enthalpy (HC2) between a first terminal of second enthalpy (HC2a) and a second terminal of second enthalpy (HC2b), from:the first predetermined enthalpy (HC1) at the entry point (P1),the third predetermined enthalpy (HC3) at the exit point (P3), andthe following formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073,γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778,determine, in the database (BdD), at least one pair of quantities of heat (ThA, ThM) whose elementary range (HC2-X) is entirely included in the global range (HC2-G)., Fuel conditioning system (SC) configured to supply a turbomachine (M) of an aircraft (A) with fuel from a cryogenic tank (R), the conditioning system (SC) being 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 aircraft (A) being configured to evolve 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 first heating module (31) configured to be mounted in the aircraft reference frame (REF-A),the first heating module (31) being configured to heat the fuel flow (Q) to a circulation temperature (Tc) from a first quantity of heat (ThA),at least one second heating module (32) configured to be mounted downstream of the first heating module (31) in the turbomachine reference frame (REF-M), the second heating module (32) being configured to heat the fuel flow (Q) to an injection temperature (Ti) from a second quantity of heat (ThM),at least one connecting pipe (L) connecting the first heating module (31) to the second heating module (32), the connecting pipe (L) being configured to circulate a fuel flow (Q) in a gaseous single-phase state,the first heating module (31) is 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 a transfer point (P2) positioned on the fuel circuit (1) at the inlet of the connecting pipe (L), the fuel flow (Q) having a first predetermined enthalpy (HC1) at the inlet point (P1) and a second enthalpy (HC2) at the transfer point (P2), the second heating module (32) is configured to heat the fuel flow (Q) up to an outlet point (P3) positioned on the fuel circuit (1) at the inlet of the turbomachine (M), the fuel flow (Q) having a third predetermined enthalpy (HC3) at the outlet point (P3),the first heating module (31) and the second heating module (32) are configured together to heat the fuel flow (Q) between the inlet point (P1) and the outlet point (P3) respecting the condition that the second enthalpy (HC2) is determined from the first predetermined enthalpy (HC1) and the third predetermined enthalpy (HC3), the second enthalpy (HC2) being comprised in an interval [HC2a, HC2b] respecting the following formulas:γRmin=min[(HC2a-HC1) / (HC3-HC1)] in which γRmin is a transfer criterion equal to 0.073,γRmax=max[(HC2b-HC1) / (HC3-HC1)] in which γRmax is a transfer criterion equal to 0.778, 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).

Images