Apparatus for heating cryogenic fuel

Abstract

The invention relates to a system (2) for heating cryogenic fuel intended to supply a combustion chamber (15) of a turbine engine (1) of an aircraft, comprising: • a storage tank (20) for cryogenic fuel in liquid state, • a closed circuit (3) for the circulation of a heat transfer fluid, • a cryogenic fuel / heat transfer fluid heat exchanger (22) and at least one heat transfer fluid / work fluid heat exchanger (4, 5, 6, 7) installed in series in the closed circuit (3) for heating the cryogenic fuel. The system (2) is characterized in that the closed circuit (3) comprises a pump (30) for circulating the heat transfer fluid and a closed expansion vessel (31) connected to the circuit (3) upstream of the pump (30) and in that the pump (30) is controlled by a central unit (8) to ensure the heat transfer fluid flow, which makes it possible to cool one or more work fluids and to heat the cryogenic fuel, and in that the closed expansion vessel (31) maintains the heat transfer fluid under pressure, so that the heat transfer fluid remains in liquid state and does not undergo a phase change.

Description

Technical Field

[0001] This invention relates to the field of aircraft turbine engines.

[0002] More specifically, the present invention relates to a system for heating cryogenic fuel supplied to the combustion chamber of an aircraft turbine engine, and also to an aircraft turbine engine that operates with the cryogenic fuel and is equipped with such a system. Background Technology

[0003] The use of cryogenic fuels for supplying combustion chambers in aircraft turbine engines is known. Such cryogenic fuels are, for example, liquefied natural gas (GNL). This fuel is less expensive than conventional kerosene and further limits CO2 emissions. More specifically, GNL emits 25% less CO2 per unit of energy.

[0004] However, GNL must be transported in liquid form to ensure that the volume to be transported in the aircraft is acceptable. GNL in one bar (10 5 The liquefaction temperature of GNL at Pascals is -161°C. However, it is preferable to inject GNL in gaseous form into the combustion chamber because this has many advantages. Therefore, it is necessary to vaporize, pressurize, and heat the GNL until it reaches an injection temperature of about 50°C.

[0005] In the prior art, it has been suggested to heat cryogenic fuel by introducing it into a heat exchanger, within which exhaust air from the turbine engine nozzles circulates. However, there is a risk that water vapor present in these exhaust gases may convert into ice on the walls of the heat exchanger.

[0006] It is also suggested that cryogenic fuel be heated by allowing it to enter a heat exchanger, where oil (such as lubricating oil present in the accessories of the turbine engine) circulates. However, due to the extremely low temperature of the cryogenic fuel, there is a risk that the lubricating oil will be overcooled and become too viscous to be reused for lubricating these accessories.

[0007] Furthermore, rapidly injecting liquid cryogenic fuel onto the hot walls of the oil / cryofluid exchanger causes a portion of the liquid to vaporize. The resulting gas accumulates in the cryogenic fuel flow loop, increasing the pressure within the loop. This prevents the liquid cryogenic fuel from flowing. Therefore, it is necessary to start feeding the liquid cryogenic fuel into the oil / cryofluid exchanger at a low flow rate so that the formed gas can be vented until the exchanger walls cool to the temperature of the cryogenic fuel. Thus, only the two-phase oil / cryofluid exchanger operates.

[0008] For any liquid, the temperature at which it vaporizes and the energy required for vaporization depend on the liquid's pressure. In the current situation, if the cryogenic fuel is pressurized only upstream of the exchanger (in other words, while the fuel is still in the liquid phase) to eliminate the compressor downstream of the exchanger, then the change in the fuel's state temperature and the energy required for this change are no longer freely controllable, but are instead imposed by the pressure of the liquid fuel upstream of the exchanger. If multiple two-phase exchangers are present, system management becomes complex.

[0009] Finally, it may be desirable to use fuels that are no longer in a gaseous state, but rather in a supercritical fluid state, in the combustion chamber, which is not provided in the cited prior art. As a reminder, fuel reaches a supercritical fluid state when it reaches a temperature higher than its critical temperature and a pressure higher than its critical pressure. Summary of the Invention

[0010] The first objective of this invention is to develop a system for heating cryogenic fuel supplied to the combustion chamber of an aircraft turbine engine, a system that does not have the aforementioned disadvantages.

[0011] Another objective of this invention is that the system operates in a manner that is both ecologically and economically advantageous.

[0012] Therefore, the present invention relates to a system for heating cryogenic fuel supplied to the combustion chamber of an aircraft turbine engine, the system comprising:

[0013] - Storage tanks, used for liquid cryogenic fuels.

[0014] - Supply lines, configured to connect the tank to the combustion chamber of the aircraft's turbine engine.

[0015] - A closed loop, used for the flow of heat transfer fluids.

[0016] - A cryogenic fuel / heat transfer fluid heat exchanger and at least one heat transfer fluid / working fluid heat exchanger, wherein the cryogenic fuel / heat transfer fluid heat exchanger and at least one heat transfer fluid / working fluid heat exchanger are connected in series in a closed loop, the cryogenic fuel / heat transfer fluid heat exchanger enabling heating of cryogenic fuel by an input of heat transferred by the heat transfer fluid.

[0017] According to the present invention, the closed loop for the flow of heat transfer fluid includes a pump for the flow of heat transfer fluid and a closed expansion container. The closed expansion container is connected to the closed loop upstream of the pump relative to the flow direction of the heat transfer fluid in the closed loop, and the pump is controlled by a central unit to ensure the flow rate of heat transfer fluid in the closed loop. This enables the cooling of one or more working fluids and the heating of cryogenic fuel according to the operating phase of the aircraft, and ensures that the closed expansion container maintains the heat transfer fluid flowing in the closed loop under pressure, so that the heat transfer fluid remains in a liquid state and does not undergo a phase change.

[0018] Through these features of the present invention, on the one hand, heat from various heat sources present in the aircraft can be recovered and transferred to a heat transfer fluid flowing in a closed loop; on the other hand, the heat transfer fluid can be kept in a single-phase liquid state with only temperature fluctuations; and finally, for cryogenic fuels, only a single heat exchanger is needed to heat the fuel.

[0019] According to a first alternative embodiment of the invention, the system includes a pump disposed in the supply line between the tank and the heat exchanger. The pump is a high-pressure pump capable of compressing the liquid cryogenic fuel to a pressure greater than the critical pressure of the cryogenic fuel. The cryogenic fuel / heat transfer fluid heat exchanger is a supercritical heat exchanger capable of heating the cryogenic fuel to a temperature higher than the critical temperature of the cryogenic fuel, so as to bring the cryogenic fuel into a supercritical fluid state before the cryogenic fuel is introduced into the combustion chamber.

[0020] According to a second alternative embodiment of the invention, the system includes a pump disposed in the supply line between the tank and the heat exchanger, the pump being capable of compressing liquid cryogenic fuel to a pressure lower than the critical pressure of the cryogenic fuel, the cryogenic fuel / heat transfer fluid heat exchanger being a two-phase heat exchanger capable of heating the cryogenic fuel by an input of heat transferred by the heat transfer fluid to change the cryogenic fuel from a liquid state to a gaseous state before the cryogenic fuel is introduced into the combustion chamber.

[0021] Other advantageous and non-limiting features, either individually or in combination, according to the invention:

[0022] - The working fluid of the heat exchanger is selected from exhaust air obtained at the outlet of the turbine engine compressor, lubricating oil, cooling air of the turbine blades of the turbine engine, and exhaust gas at the outlet of the turbine engine nozzle.

[0023] - The system includes at least one heat exchanger between the heat transfer fluid and the exhaust air obtained at the outlet of the turbine engine compressor;

[0024] - The system includes at least two heat transfer fluid / working fluid heat exchangers, each including a heat exchanger for the heat transfer fluid and exhaust air obtained at the outlet of the turbine engine compressor, and the heat exchanger for the heat transfer fluid and exhaust air obtained at the outlet of the turbine engine compressor is arranged upstream of one or more other heat transfer fluid / working fluid heat exchangers relative to the flow direction of the heat transfer fluid in the closed loop.

[0025] - The system includes a heat exchanger for heat transfer fluid and exhaust gas at the outlet of a turbine engine nozzle, and the heat exchanger is installed in series with a heat exchanger for heat transfer fluid and exhaust air obtained at the outlet of a turbine engine compressor, and is arranged downstream of the heat exchanger for heat transfer fluid and exhaust air obtained at the outlet of a turbine engine compressor relative to the flow direction of the heat transfer fluid in the closed loop.

[0026] - The system includes at least one heat exchanger installed in series in a closed loop for the flow of heat transfer fluid, a heat transfer fluid / oil heat exchanger, a heat exchanger for the heat transfer fluid and air used to cool turbine blades, and a heat exchanger for the heat transfer fluid and exhaust gas at the outlet of the turbine engine nozzle.

[0027] - The system includes, in a closed loop for the flow of heat transfer fluid, a heat exchanger between the heat transfer fluid and exhaust air obtained at the outlet of the turbine engine compressor, a heat transfer fluid / oil heat exchanger, a heat exchanger between the heat transfer fluid and air for cooling the turbine blades, and a heat exchanger between the heat transfer fluid and exhaust gas at the outlet of the turbine engine nozzle, which are installed in series in the following order.

[0028] - The system includes at least two heat transfer fluid / working fluid heat exchangers. For at least one of the two heat transfer fluid / working fluid heat exchangers, a switching valve is installed in a closed loop. Upstream of the heat transfer fluid / working fluid heat exchanger, a bypass loop connects a point of the closed loop located upstream of the switching valve to a point of the closed loop located downstream of the heat transfer fluid / working fluid heat exchanger, and the bypass loop is provided with a switching valve.

[0029] - Cryogenic fuels are liquefied natural gas or liquid hydrogen.

[0030] - The two-phase cryogenic fuel / heat transfer fluid heat exchanger is a plate heat exchanger.

[0031] The present invention also relates to aircraft turbine engines.

[0032] According to the invention, the turbine engine includes a combustion chamber supplied with cryogenic fuel and a system as described above for heating the cryogenic fuel. Attached Figure Description

[0033] Other features, objects, and advantages of the invention will become apparent from the following description, which is given only in an illustrative and non-limiting manner and should be read with reference to the accompanying drawings, in which:

[0034] [ Figure 1 [Illustration] is a schematic diagram illustrating an exemplary embodiment of an aircraft turbine engine (here, a twin-shaft ducted turbine engine).

[0035] [ Figure 2 [Illustration] is a schematic diagram illustrating an exemplary embodiment of a system for heating cryogenic fuel according to the present invention, the fuel being intended to supply the combustion chamber of an aircraft turbine engine.

[0036] [ Figure 3 [This indicates the period during takeoff of the aircraft] Figure 2 A schematic diagram of the system's operation.

[0037] [ Figure 4 [This is an illustration of the aircraft during its climb.] Figure 2 A schematic diagram of the system's operation.

[0038] [ Figure 5 This indicates when the aircraft is in cruise, during descent, or on the ground with its engines idling. Figure 2 A schematic diagram of the system's operation.

[0039] [ Figure 6 [Illustration 1] is a plan view of the stage of the plate switch used in the above system.

[0040] [ Figure 7 ] is with Figure 6 A plan view of adjacent stages of a plate switch.

[0041] exist Figures 3 to 5 The system is shown in a simplified manner in the image. Detailed Implementation

[0042] As a reminder, the following will refer to Figure 1 Review the structure of a twin-shaft ducted turbine engine.

[0043] The figure shows a ducted turbine 1, which has an air inlet 10 and a fan 11 successively in the direction of airflow (in other words, from upstream (left side of the figure) to downstream (right side of the figure)). The fan delivers air to a main channel 12 and a secondary channel 13. The term "channel" refers to the volume through which airflow passes.

[0044] The airflow flowing in the main channel 12 passes successively through the low-pressure compressor 14a, the high-pressure compressor 14b, the combustion chamber 15, the high-pressure turbine 16a, and the low-pressure turbine 16b before being discharged through the main channel nozzle 120.

[0045] Furthermore, the secondary airflow flowing in the secondary channel 13 is individually sprayed through the secondary flow nozzle 130 after passing through a series of guide vanes 131.

[0046] Now we will combine Figure 2 To describe the system 2 for heating cryogenic fuel according to the present invention.

[0047] System 2 can be combined with the above. Figure 1 The system 2 is used with the described turbine engine 1. However, the system 2 can also be used with another type of turbine engine (e.g., a single-shaft turbine engine or a turbine engine that consists of only one compressor and a single turbine).

[0048] Therefore, in Figure 2 The figure schematically shows a turbine engine 1, wherein the compressor is generally indicated by reference numeral 14 and the turbine is generally indicated by reference numeral 16, and the compressor and turbine are arranged upstream and downstream of the combustion chamber 15, respectively.

[0049] System 2 includes a tank 20 for storing the liquid cryogenic fuel. The fuel is, for example, liquefied natural gas or liquid hydrogen. The tank 20 is connected to a combustion chamber 15 via a pipeline 21. A cryogenic fuel / heat transfer fluid heat exchanger 22 is installed in the pipeline 21 between the tank 20 and the combustion chamber 15. This heat exchanger 22 can heat the cryogenic fuel before it is injected into the combustion chamber 15. For this purpose, an injector is arranged in the pipeline 21, downstream of the heat exchanger 22, although this is not shown in the figures.

[0050] Finally, pump 23 is installed in pipeline 21, and preferably between storage tank 20 and heat exchanger 22. The operation of the pump is controlled by central unit 8 (e.g., computer).

[0051] According to the first alternative embodiment, the pump is a high-pressure pump that can bring the cryogenic fuel to a pressure above its critical pressure, and since the heat exchanger can keep the cryogenic fuel above its critical temperature, the heat exchanger 22 is a so-called "supercritical" heat exchanger. Therefore, at the outlet of the heat exchanger 22, the fuel is in a supercritical fluid state and is injected into the combustion chamber 15 in this state.

[0052] As a reminder, the critical point of hydrogen is at 32K (-241.15℃) and 12.8 bar (12.8·10⁻⁶ bar). 5 The critical point of natural gas is at 190 K (-83.15℃) and 46.8 bar (46.8·10). 5 (Pa).

[0053] According to the second alternative embodiment, pump 23 is a pump capable of pressurizing liquid cryogenic fuel to a pressure lower than the critical pressure of cryogenic fuel in order to deliver the required cryogenic fuel flow to the aircraft according to these flight phases (takeoff, cruise flight, landing, taxiing), and heat exchanger 22 is a so-called "two-phase" exchanger because the heat exchanger is capable of changing the cryogenic fuel from liquid to gas (phase change) before the cryogenic fuel is injected into the combustion chamber 15.

[0054] System 2 also includes a closed loop 3 for the flow of a liquid heat transfer fluid. The heat transfer fluid flowing in the closed loop is, for example, a haloalkane sold by 3M under the trade name "NOVEC 7500". However, any other heat transfer fluid that remains in the liquid phase and has sufficient thermal properties to ensure the heat exchange described below can be used.

[0055] The flow of heat transfer fluid within the closed loop 3 is ensured by pump 30, which is also controlled by the central unit 8.

[0056] A closed expansion vessel 31 is connected to the circuit 3 at a point upstream of the flow pump 30 in the flow direction of the heat transfer fluid within the closed circuit 3. This closed expansion vessel 31 is able to maintain the heat transfer fluid in the circuit 3 under pressure, so that the heat transfer fluid remains in a liquid state and does not undergo a phase change.

[0057] Heat exchanger 22 is also installed in closed loop 3 to exchange heat with heat transfer fluid.

[0058] In addition, at least one heat transfer fluid / working fluid heat exchanger is installed in series with exchanger 22 in loop 3. Various working fluids can be used, which will be described below.

[0059] Advantageously, the circuit 3 is provided with at least one heat exchanger 4 "for the heat transfer fluid and the exhaust air obtained at the outlet of the turbine engine compressor 14". The air obtained at the outlet of the compressor 14 is referred to as "exhaust air".

[0060] The heat exchanger 4 is connected at its inlet 41 to the hot air outlet of the compressor 14. Hot air flows within the heat exchanger 4 in a counter-current manner with the heat transfer fluid, which flows in a closed loop 3. After transferring heat to the heat transfer fluid, the hot air exits the heat exchanger 4 via its outlet 42. This cooled air can then be delivered to the aircraft's cabin 43.

[0061] Advantageously, system 2 may also include a heat transfer fluid / lubricating oil heat exchanger 5. The exchanger 5 is connected at its inlet 51 to a lubricating oil network 50, which is used for various mechanical components of the turbine engine.

[0062] Inside a turbine engine, a number of mechanical components must indeed be cooled by lubricating oil to maintain their mechanical integrity. These components include, for example, reduction gears in an accessory gearbox or bearings supporting rotating shafts. The oil, heated in contact with these components, enters heat exchanger 5, transferring heat to the heat transfer fluid flowing in loop 3, and flows out through outlet 52 of exchanger 5 to be returned to the lubricating oil network 50, allowing the oil to be reused and in contact with these mechanical components again.

[0063] Advantageously, system 2 may also include a heat exchanger 6 between the heat transfer fluid and the air used to cool the turbine blades. This exchanger 6 is connected at its inlet 61 to the cooling air circuit of the turbine 16 blades, such that once the air comes into contact with and is heated by the air, it is recovered. This hot air transfers heat to the heat transfer fluid flowing in circuit 3 and exits via the outlet 62 of the exchanger 6, where it is cooled. The air can then be returned to the turbine 16 blades to cool them.

[0064] Finally, advantageously, system 2 also includes a heat exchanger 7 "for the heat transfer fluid and the exhaust gas at the outlet of the turbine engine nozzle". The exhaust gas is acquired at the outlet of nozzle 120 (in other words, at the nozzle of the main channel).

[0065] The heat exchanger 7 is connected to the nozzle 120 at its inlet 71. Hot air flows in the heat exchanger 7 in a countercurrent manner with the heat transfer fluid, which flows in a closed loop 3. After transferring heat to the heat transfer fluid, the hot air exits through the outlet 72 and is then discharged into the atmosphere via the downstream portion of the nozzle.

[0066] Therefore, it can be seen that the cooled heat transfer fluid leaving the two-phase heat exchanger 22 is gradually heated after each journey through one of the exchangers 4, 5, 6, 7, until it reaches the highest temperature at the inlet 223 of the exchanger 22.

[0067] The intensity of heat transfer from the heat source (in other words, the various working fluids flowing in the various exchangers 4, 5, 6, and 7) to the cold source (in other words, the liquid heat transfer fluid) is controlled by the flow rate of the heat transfer fluid flowing in the closed loop 3. This flow rate is itself controlled by the rotational speed of the pump 30. The greater the flow rate of the heat transfer fluid, the more heat the heat transfer fluid acquires from the heat source, and the greater the temperature rise of the heat transfer fluid.

[0068] The enclosed expansion container 31 comprises two sections separated by an elastic membrane. The first section is connected to the loop 3 and receives the heat transfer fluid, while the second section is filled with compressed air (or gas). When the heat transfer fluid heats up after heat exchange with at least one of the heat exchangers 4, 5, 6, or 7, the fluid expands, and the first section of the expansion container 31 receives excess heat transfer fluid, thereby preventing an increase in pressure in the loop 3. Conversely, if the temperature of the heat transfer fluid decreases, the fluid is compressed, and the excess heat transfer fluid present in the expansion container 31 returns to the loop 3.

[0069] The expansion vessel 31 can maintain the heat transfer fluid at a pressure sufficient to raise its boiling point and prevent it from entering the gaseous state. Therefore, the heat transfer fluid flowing in loop 3 only changes in temperature, but the pressure of the heat transfer fluid remains constant, and the heat transfer fluid remains in the liquid phase.

[0070] Advantageously, the two-phase heat exchanger 22 is a plate heat exchanger. An exemplary embodiment of such a plate heat exchanger is described in... Figure 6 and Figure 7 As shown in the figure. This type of exchanger comprises multiple stacked plates made of a material with good thermal conductivity (usually a metal, such as aluminum or stainless steel), which are assembled together by welding, brazing or using seals.

[0071] One of the main advantages of this type of exchanger compared to other types is that it maximizes convection exchange. Finally, advantageously, in the embodiment shown in the accompanying drawings, and in order to maximize heat transfer, each stage of the plate includes multiple channels arranged to form a V-shaped pattern.

[0072] In a given stage of the two-phase heat exchanger 22 (i.e., in Figure 6 In the stage (marked by reference numeral 22a, and arranged between two adjacent plates), liquid cryogenic fuel enters through inlet 221 of exchanger 22, passes through the stage, and exits as a gas through outlet 222. In the adjacent stage (i.e., in...) Figure 7In the stage (with reference numeral 22b), the heat transfer fluid enters through inlet 223 of exchanger 22, passes through the stage, and exits through outlet 224 (see also...). Figure 2 (Schematic diagram).

[0073] Preferably, as in Figure 2 As shown, the circuit 3 is equipped with the four types of heat exchangers 4, 5, 6 and 7 described above, which are installed in series in the closed circuit 3. More preferably, the heat exchangers are installed in this order from the outlet 224 to the inlet 223 of the heat exchanger 22, relative to the flow direction of the heat transfer fluid in the closed circuit 3.

[0074] Preferably, when at least two of the heat exchangers 4, 5, 6 and 7 are present, in order to selectively select only some of these heat exchangers, a switching valve and a bypass circuit can be provided for at least one of these heat exchangers. The switching valve is installed upstream of the heat exchanger, and the bypass circuit connects a point of closed loop 3 located upstream of the switching valve and a point of closed loop 3 located downstream of the heat exchanger.

[0075] Therefore, for example, exchangers 4, 5, 6, and 7 are provided with switching valves (with reference numerals 44, 54, 64, and 74, respectively) and bypass circuits (with reference numerals 45, 55, 65, and 75, respectively). In addition, each of the above-mentioned bypass circuits 45, 55, 65, and 75 is provided with switching valves 450, 550, 650, and 750.

[0076] The switching valves 44, 54, 64, 74, 450, 550, 650, and 750 are controlled by the central unit 8.

[0077] Figure 3 , Figure 4 as well as Figure 5 The diagrams illustrate the various operational scenarios of System 2 based on the flight phases of the aircraft. In each of these diagrams, the flow of the heat transfer fluid is indicated by a bold line.

[0078] Figure 3 The diagram illustrates the situation during aircraft takeoff. In this scenario, valves 450, 54, 64, and 74 are open, while valves 44, 550, 650, and 750 are closed.

[0079] This makes it possible to eliminate the need for a heat exchanger 4 between the heat transfer fluid and the exhaust air obtained at the outlet of the turbine engine compressor. In fact, during takeoff, the exhaust air is cut off to obtain maximum propulsive power. Therefore, there is no airflow to be cooled.

[0080] Figure 4The diagram illustrates the situation during the climb after takeoff. In this scenario, valves 450, 54, 650, and 74 are open, while valves 44, 550, 64, and 750 are closed. This allows the heat exchanger 4 between the heat transfer fluid and the exhaust air obtained at the turbine engine compressor outlet to be omitted, and the heat exchanger 6 between the heat transfer fluid and the cooling air for the turbine blades is also omitted. Therefore, after takeoff, when the engine reaches its maximum temperature, the cooling function of the air cooling the blades is no longer necessary, and thus the heat exchanger 6 can be shut off to limit losses.

[0081] at last, Figure 5 The diagram illustrates the situation during flight or descent, or when the aircraft is taxiing on the ground, with the aircraft's turbine engines idling. In this case, valves 44, 54, 650, and 750 are open, while valves 450, 550, 64, and 74 are closed. This allows the absence of heat exchanger 6 between the heat transfer fluid and the cooling air for the turbine blades, and also eliminates the need for heat exchanger 7 between the heat transfer fluid and the exhaust gas at the turbine engine nozzle outlet. In practice, in the aforementioned scenario, the cryogenic fuel flow rate is low, and heat exchangers 4 and 5 alone are sufficient to heat the cryogenic fuel. Therefore, heat exchanger 7 can be closed, and air from the nozzles bypasses it to limit airflow losses due to its presence.

[0082] According to only Figures 3 to 5 In the alternative embodiment shown, a bypass line 32 may be provided in the closed loop 3, connecting a point in loop 3 downstream of the flow pump 30 and upstream of the heat exchanger 22 to a point downstream of the heat exchanger 22 and upstream of the heat exchanger 4 between the heat transfer fluid and the exhaust air obtained at the outlet of the turbine engine compressor. A thermostatic valve 33 equipped with a temperature sensor 34 is arranged on the bypass line. The thermostatic valve 33 is controlled based on temperature and optional factors such as flight phase or start-up / restart. The bypass line 32 may divert a portion of the flow into the heat exchanger 22 during start-up or restart and during some flight phases.

[0083] Below a certain temperature, valve 33 is controlled to allow a portion of the heat transfer fluid flow to pass through bypass line 32 instead of solely through exchanger 22, and this is to control the phase change or supercritical state of the fuel. It can also be configured such that, within a given temperature range, an additional portion of the heat transfer fluid is directed into bypass line 32 as the temperature detected by sensor 34 increases.

[0084] The heating system 2 according to the present invention has the following advantages:

[0085] - The heating system includes only a single cryogenic fuel / heat transfer fluid heat exchanger 22 and is capable of providing the heat required to heat the cryogenic fuel through a single closed loop 3, within which the heat transfer fluid, which remains in a liquid state and only changes in temperature, flows.

[0086] - In the prior art, cryogenic fuel is heated by traveling through a heat exchanger, and air from the secondary channel circulates within the heat exchanger; compared to the prior art, the risk of icing on the heat exchanger 22 is reduced because the temperature of the heat transfer fluid circulating within the closed loop 3 is easier to control.

[0087] - Because only a single heat exchanger 22 is required between tank 20 and combustion chamber 15, instead of multiple heat exchangers (which would be necessary when heat needs to be recovered from different working fluids), pipeline 21 is shorter. This allows for the restriction of the flow of cryogenic fuels (especially liquefied natural gas) in the hot and therefore flammable areas of the aircraft.

[0088] Furthermore, two-phase heat exchangers where a phase change occurs internally are more difficult to control. Therefore, it is preferable to reduce the number of two-phase heat exchangers.

[0089] - In addition, within pipeline 21, the pressure of the cryogenic fuel only needs to be managed by pump 23, and the temperature is managed by the heat supplied by the heat transfer fluid flowing in the closed loop 3.

[0090] -exist Figure 2 In a preferred alternative embodiment of system 2 shown, the heat transfer fluid is heated by passing through the exchanger 4 before reaching the exchanger 5. Therefore, the oil brought through the exchanger 5 is cooled, but only to a temperature high enough to prevent the oil from freezing or solidifying.

[0091] Heat exchangers 4, 5, and 6 can cool the various working fluids supplied to these exchangers according to their cooling requirements.

[0092] - The heat exchanger 7 can use a heat source (in this case, the exhaust gas at the outlet of the turbine engine nozzle 120) to regulate the temperature of the heat transfer fluid flowing in the loop 3. This heat source is continuous as long as the aircraft is running, and it does not need to be cooled for the operation of the aircraft. Therefore, the heat exchanger 7 constitutes a regulating variable for the operation of the heating system 2.

Claims

1. A system (2) for heating cryogenic fuel supplied to the combustion chamber (15) of an aircraft turbine engine (1), the system (2) comprising: - Storage tank (20), the storage tank being used for the liquid cryogenic fuel, - A supply line (21) configured to connect the storage tank (20) to the combustion chamber (15) of the aircraft turbine engine. - Closed loop (3), the closed loop being used for the flow of heat transfer fluid, - A cryogenic fuel / heat transfer fluid heat exchanger (22) and at least one heat transfer fluid / working fluid heat exchanger (4, 5, 6, 7), wherein the cryogenic fuel / heat transfer fluid heat exchanger and the at least one heat transfer fluid / working fluid heat exchanger are installed in series in the closed loop (3), wherein the cryogenic fuel / heat transfer fluid heat exchanger (22) enables the cryogenic fuel to be heated by the input of heat transferred by the heat transfer fluid. The system (2) is characterized in that the closed loop (3) for the flow of the heat transfer fluid includes a pump (30) for the flow of the heat transfer fluid and a closed expansion container (31), the closed expansion container being connected to the closed loop (3) upstream of the pump (30) relative to the flow direction of the heat transfer fluid in the closed loop (3), the pump (30) being controlled by a central unit (8) to ensure the flow rate of the heat transfer fluid within the closed loop (3), which enables the cooling of one or more of the working fluids and the heating of the cryogenic fuel according to the operating phase of the aircraft, and causes the closed expansion container (31) to maintain the heat transfer fluid flowing in the closed loop (3) under pressure, so that the heat transfer fluid remains in a liquid state and does not undergo a phase change, the system (2) including at least two heat transfer fluids / The working fluid heat exchangers (4, 5, 6, 7) include at least two heat transfer fluid / working fluid heat exchangers, each comprising a heat exchanger (4) between the heat transfer fluid and exhaust air obtained at the outlet of the turbine engine compressor and a heat exchanger (6) between the heat transfer fluid and air used to cool the turbine blades. The heat exchanger (4) between the heat transfer fluid and exhaust air obtained at the outlet of the turbine engine compressor is arranged upstream of the heat exchanger (6) between the heat transfer fluid and the air used to cool the turbine blades, or upstream of the heat exchanger (6) between the heat transfer fluid and the air used to cool the turbine blades and a plurality of other heat transfer fluid / working fluid heat exchangers (5, 7), relative to the flow direction of the heat transfer fluid in the closed loop (3).

2. The system according to claim 1, characterized in that, The system includes a pump (23) arranged in the supply line (21) between the storage tank (20) and the cryogenic fuel / heat transfer fluid heat exchanger (22). The pump (23) is a high-pressure pump capable of compressing the liquid cryogenic fuel to a pressure greater than the critical pressure of the cryogenic fuel. The cryogenic fuel / heat transfer fluid heat exchanger (22) is a supercritical heat exchanger capable of heating the cryogenic fuel to a temperature higher than the critical temperature of the cryogenic fuel, so as to bring the cryogenic fuel into a supercritical fluid state before the cryogenic fuel is introduced into the combustion chamber (15).

3. The system according to claim 1, characterized in that, The system includes a pump (23) arranged in the supply line (21) between the storage tank (20) and the cryogenic fuel / heat transfer fluid heat exchanger (22), the pump (23) being capable of compressing the liquid cryogenic fuel to a pressure lower than the critical pressure of the cryogenic fuel, and the cryogenic fuel / heat transfer fluid heat exchanger (22) being a two-phase heat exchanger capable of heating the cryogenic fuel by the input of heat transferred by the heat transfer fluid to change the cryogenic fuel from a liquid state to a gaseous state before the cryogenic fuel is introduced into the combustion chamber (15).

4. The system according to claim 1, characterized in that, The working fluid of the heat transfer fluid / working fluid heat exchanger (4, 5, 6, 7) is selected from exhaust air obtained at the outlet of the turbine engine compressor, lubricating oil, cooling air of the turbine blades of the turbine engine, and exhaust gas at the outlet of the turbine engine nozzle.

5. The system according to claim 4, characterized in that, The system includes a heat exchanger (7) for the heat transfer fluid and the exhaust gas at the outlet of the turbine engine nozzle, and the heat exchanger is installed in series with the heat exchanger (4) for the heat transfer fluid and the exhaust air obtained at the outlet of the turbine engine compressor, and is arranged downstream of the heat exchanger for the heat transfer fluid and the exhaust air obtained at the outlet of the turbine engine compressor relative to the flow direction of the heat transfer fluid in the closed loop (3).

6. The system according to claim 1, characterized in that, The system includes at least the following heat exchangers installed in series in the closed loop (3) for the flow of the heat transfer fluid: a heat exchanger (4) for the heat transfer fluid to the exhaust air obtained at the outlet of the turbine engine compressor, a heat transfer fluid / lubricating oil heat exchanger (5), a heat exchanger (6) for the heat transfer fluid to the air for cooling the turbine blades, and a heat exchanger (7) for the heat transfer fluid to the exhaust gas at the outlet of the turbine engine nozzle.

7. The system according to claim 6, characterized in that, The system includes, in the closed loop (3) for the flow of the heat transfer fluid, a heat exchanger (4) for the heat transfer fluid to discharge air obtained at the outlet of the turbine engine compressor, a heat transfer fluid / lubricating oil heat exchanger (5), a heat exchanger (6) for the heat transfer fluid to air for cooling the turbine blades, and a heat exchanger (7) for the heat transfer fluid to exhaust gas at the outlet of the turbine engine nozzle, which are installed in series in the following order.

8. The system according to claim 1, characterized in that, The system includes at least two heat transfer fluid / working fluid heat exchangers (4, 5, 6, 7). For at least one of the two heat transfer fluid / working fluid heat exchangers (4, 5, 6, 7), a switching valve (44, 54, 64, 74) is installed in the closed loop (3). Upstream of the heat transfer fluid / working fluid heat exchanger (4, 5, 6, 7), a bypass loop (45, 55, 65, 75) connects a point of the closed loop (3) located upstream of the switching valve (44, 54, 64, 74) to a point of the closed loop (3) located downstream of the heat transfer fluid / working fluid heat exchanger (4, 5, 6, 7). The bypass loop (45, 55, 65, 75) is provided with switching valves (450, 550, 650, 750).

9. The system according to claim 1, characterized in that, The cryogenic fuel is liquefied natural gas or liquid hydrogen.

10. The system according to claim 3, characterized in that, The cryogenic fuel / heat transfer fluid heat exchanger (22) is a plate heat exchanger.

11. An aircraft turbine engine (1), characterized in that, The aircraft turbine engine includes a combustion chamber (15) supplied with cryogenic fuel and a system (2) for heating the cryogenic fuel, according to any one of claims 1 to 10.

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