Fuel conditioning system and method for powering an aircraft turbine engine with cryotank fuel

Abstract

The present invention discloses a fuel conditioning system (SC) for supplying fuel (Q) from a cryogenic chamber (RC) to a turbine engine of an aircraft. The conditioning system (SC) includes: at least one pumping turbine (1); a first heat exchanger (31) configured to heat the fuel (Q) in a fuel circuit (CQ) by circulating an airflow (A); and at least one heating turbine (2) configured to supply airflow (A) to the first heat exchanger (31). The heating turbine (2) includes an intake compressor (21), a combustion chamber (24), and an exhaust turbine (22) configured to drive the intake compressor (21). The combustion chamber (24) is supplied with air obtained from the airflow (A) and fuel (Q) from the fuel circuit (CQ).

Description

Technical Field

[0001] This invention relates to the field of aircraft including turbine engines, which are supplied with fuel stored in cryogenic chambers. Background Technology

[0002] Storing fuels (especially hydrogen) in liquid form to reduce the size and mass of aircraft tanks is well-known. For example, fuel is stored in cryogenic chambers in aircraft at temperatures of approximately 20K to 22K (-253°C to -251°C). To inject fuel into the combustion chamber of a turbine engine, the fuel needs to be delivered and heated for optimal combustion. This heating step is necessary, for example, to reduce the risk of water vapor contained in the airflow circulating in the turbine engine icing, particularly at the turbine engine fuel injectors.

[0003] In practice, the fuel heating process is energy-intensive and requires heat to be drawn from the aircraft's heat sources. For example, heat generated by the turbine engine (heat from the lubricating oil, heat from the turbine outlet, heat from the nozzles, etc.) can be used. Heat from the aircraft itself (air from the cabin, heat from electrical or electronic systems, etc.) can also be used.

[0004] One challenge is optimizing fuel heating while utilizing its cooling properties as it is transported from the cryogenic chamber to the turbine engine combustion chamber, while simultaneously ensuring high safety.

[0005] Reference Figure 1 and Figure 2 In the prior art, several structures have been proposed to guide fuel Q from the cryogenic chamber RC to the combustion chamber CC of the turbine engine T. It is known that the cryogenic chamber RC belongs to the aircraft baseline REF-A, and the combustion chamber CC belongs to the turbine engine baseline REF-T.

[0006] Subsequently, the terms "upstream" and "downstream" are defined relative to the circulation direction of fuel Q from the cryogenic chamber RC to the combustion chamber CC.

[0007] Reference Figure 1 The diagram illustrates a first structure comprising, from upstream to downstream, a pump 101 and a heat exchanger 102. Pump 101 is associated with turbine engine baseline REF-T, and heat exchanger 102 obtains heat from turbine engine T, particularly at the nozzles of turbine engine T before fuel Q is injected into combustion chamber CC. Pump 101 is connected to cryogenic chamber RC via delivery pipe 100, which has an interface located between the two baselines REF-A and REF-T.

[0008] This first structure has several disadvantages. First, since the cryogenic, low-pressure fuel Q is delivered to the pump 101 via the delivery pipe 100, the pump 101 must be thermally adiabatic, which is disadvantageous. Additionally, the presence of the heat exchanger 102 in the nozzle of the turbine engine T affects the performance of the turbine engine T. Furthermore, during the transition phase (start-up / stop), the length of the delivery pipe 100 introduces a significant risk of thermofluid instability, resulting in prolonged cooling or recooling, which is detrimental to the safety and operability of the turbine engine T.

[0009] Reference Figure 2 The diagram illustrates a second configuration, proposing to offset pump 101 and position it as close as possible to the cryogenic chamber RC, i.e., within the aircraft baseline REF-A, in order to reduce its size within the turbine engine baseline REF-T. Fuel Q is compressed under high pressure and cryogenic pressure in a delivery pipe 100' that connects pump 101 to exchanger 102. Similar to the aforementioned cases, delivery pipe 100' must be thermally insulated and mechanically reinforced to withstand the high pressure of fuel Q.

[0010] In both configurations, because the temperature of the fuel Q in the delivery pipes 100 and 100' is very low, only helium can be used as a purging or conditioning gas when fuel Q leaks. For commercial aviation applications, helium must be avoided because this fluid is a high-cost rare gas. In the prior art of patent application EP3623604A1, a military propulsion turbine designed to operate at very high speeds is known, which includes a heat exchanger configured to extract heat from exhaust gas to supply it to the fuel stream.

[0011] Therefore, the present invention aims to at least partially eliminate the aforementioned disadvantages by proposing a new fuel regulation system. Summary of the Invention

[0012] This invention relates to a fuel regulation system for supplying fuel to an aircraft turbine engine via fuel from a cryogenic chamber, the regulation system comprising:

[0013] - At least one pumping turbine, including a pump and a turbine configured to drive the pump, the pump being configured to obtain fuel from a cryogenic tank and circulate the fuel from upstream to downstream in a fuel circuit, the fuel circuit including a supply outlet of the turbine engine, the turbine being mounted in the fuel circuit to drive the turbine to rotate by the circulation of the fuel;

[0014] - A first heat exchanger, located upstream of the turbine, is configured to heat the fuel in the fuel circuit by circulating airflow.

[0015] - At least one heating turbine configured to supply airflow to the first heat exchanger, the heating turbine including an intake compressor, a combustion chamber and an exhaust turbine configured to drive the intake compressor, the combustion chamber being supplied with air obtained from the airflow and fuel from the fuel circuit.

[0016] Advantageously, the use of a pumping turbine coupled to a heating turbine enables autonomous fuel heating while consuming only a small fraction of the fuel. The fuel supplied at the outlet has a temperature suitable for the turbine engine, preferably close to ambient temperature, and sufficient pressure to be delivered to the turbine engine. Therefore, the conditioning system can advantageously be located in the aircraft baseline, rather than in the turbine engine baseline, preferably near the cryogenic chamber. Advantageously, nitrogen can be used as a neutral gas for purification and purging, making the implementation of the conditioning system quite easy.

[0017] The use of a first heat exchanger allows heat to be transferred from the heating turbine to the fuel for optimal heating. Advantageously, the pumping turbine is driven by the circulation of fuel in the fuel circuit, which simplifies the design of the pumping turbine and improves safety by allowing for low fuel leakage between the pump and the turbine.

[0018] Preferably, the combustion chamber is supplied with air obtained from the airflow and fuel from the fuel circuit and downstream of the turbine. The fuel advantageously has an optimal temperature and pressure suitable for the combustion chamber.

[0019] Preferably, the system includes a second heat exchanger configured to heat the fuel in the fuel circuit by circulating the intake airflow of the heating turbine. Therefore, circulating the intake air at a lower temperature than the exhaust gas allows the fuel to be preheated before it is heated by the first heat exchanger. This preheating reduces the risk of condensation of hot gas passing through the first exchanger and also reduces the consumption of the heating turbine's intake compressor.

[0020] According to one aspect of the invention, the system includes an auxiliary heat exchanger configured to heat the fuel in the upstream portion of the fuel circuit by circulating fuel downstream of the turbine. Advantageously, gradual preheating is achieved through heat exchange involving the circulation of fuel at different temperatures. This preheating reduces the risk of air condensation in subsequent exchangers.

[0021] According to one aspect of the invention, the system includes a generator driven by the heating turbine. According to one aspect, the generator is driven by a free turbine configured to be driven by an airflow from the heating turbine. Therefore, the pumping turbine can generate an airflow for a first heat exchanger, which is then used by an auxiliary turbine to generate electricity. Thus, the regulating system generates electricity to promote its autonomy.

[0022] On the other hand, the generator is either directly driven by a heating turbine or driven through a gear system.

[0023] Preferably, the heating turbine is configured to implement a Brayton cycle in order to generate a large amount of heat by burning a small portion of the fuel.

[0024] Preferably, the pumping turbine is completely submerged in the fuel. This eliminates the need for a seal between the pump and the turbine. Advantageously, fluid bearing guidance can be implemented in a practical manner.

[0025] Preferably, the pumping turbine is an expansion hydrogen turbine pump so that it is driven by fuel circulation when the fuel is hydrogen.

[0026] The present invention also relates to a component comprising at least one cryogenic chamber, an aircraft turbine engine, and a conditioning system as described above, the conditioning system being fluidly connected to the cryogenic chamber and the aircraft turbine engine.

[0027] Preferably, the conditioning system is located near the cryogenic chamber and connected to the turbine engine via at least one circulation line that circulates the fuel at ambient temperature. This line advantageously has a simple structure and requires no special insulation or cooling.

[0028] According to one aspect, the regulating system is supplied with air by airflow from outside the turbine engine. According to another aspect, the regulating system is supplied with air by airflow from the turbine engine, particularly by airflow from the turbine engine's compressor.

[0029] Preferably, the component includes at least one fuel buffer storage unit configured to be supplied by the regulating system and configured to supply the aircraft turbine engine.

[0030] Preferably, the fuel buffer reservoir is configured to store fuel at high pressure and ambient temperature. Therefore, the fuel in the buffer reservoir can be directly used by the turbine engine and respond quickly to its demands.

[0031] The present invention also relates to a method for regulating fuel using the regulating system described above to supply fuel from a cryogenic chamber to an aircraft turbine engine, wherein:

[0032] The pump of the pumping turbine obtains fuel from the cryogenic tank and circulates the fuel from upstream to downstream in the fuel circuit.

[0033] - Circulate the airflow of the heating turbine so that the first heat exchanger heats the fuel in the fuel circuit.

[0034] The circulation of the fuel in the fuel circuit drives the rotation of the turbine of the pumping turbine.

[0035] The combustion chamber of the heated turbine is supplied with air obtained from the airflow and fuel from the fuel circuit. Attached Figure Description

[0036] The invention will be better understood by reading the following description given as an example and by referring to the following drawings given as a non-limiting example, wherein the same reference numerals are used to indicate similar objects.

[0037] Figure 1 This is a schematic diagram of a first structure in the prior art for regulating fuel from a cryogenic chamber to a turbine engine.

[0038] Figure 2 This is a schematic diagram of a prior art second structure for regulating fuel from a cryogenic chamber to a turbine engine.

[0039] Figure 3 This is a schematic diagram of the regulating system of the present invention for regulating the fuel from the cryogenic chamber to the turbine engine.

[0040] Figure 4 This is a schematic diagram of the first embodiment of the control system.

[0041] Figure 5 This is a schematic diagram of a second embodiment of the control system.

[0042] Figure 6 This is a schematic diagram of the third embodiment of the control system.

[0043] Figure 7 This is a schematic diagram of a regulating system that provides fuel to a buffer storage container.

[0044] It should be noted that the accompanying drawings illustrate the invention in detail in order to implement the invention, and the drawings can of course be used to better define the invention where applicable. Detailed Implementation

[0045] Reference Figure 3This diagram illustrates a structure according to an embodiment of the invention for delivering fuel Q from a cryogenic chamber RC to the combustion chamber CC of a turbine engine T in an aircraft. It is known that the cryogenic chamber RC belongs to the aircraft baseline REF-A, while the combustion chamber CC belongs to the turbine engine baseline REF-T.

[0046] In this example, the fuel is liquid hydrogen, but the invention is applicable to other types of fuels, such as liquid methane or liquefied natural gas.

[0047] According to the present invention, a fuel conditioning system SC is provided, configured to supply liquid fuel from a cryogenic chamber RC to the combustion chamber CC of a turbine engine T. The conditioning system SC is located at the aircraft baseline REF-A and pumps fuel Q as close as possible to the cryogenic chamber RC, thereby reducing the space required in the turbine engine baseline REF-T. As will be described later, the conditioning system SC heats the fuel to an optimal temperature by circulating the airflow from the airflow inlet EA. The fuel conditioning system SC is connected to the turbine engine T via a circulation pipe 6.

[0048] Reference Figure 4 The regulating system SC includes a pumping turbine 1, a fuel circuit CQ, a heating turbine 2, a first heat exchanger 31, and a second heat exchanger 32.

[0049] like Figure 4 As shown, a pumping turbine 1 (also referred to as a turbopump) is illustrated, comprising a pump 11 and a turbine 12 configured to drive the pump 11. The pump 11 is configured to obtain fuel Q from a cryogenic chamber RC and circulate the fuel Q from upstream to downstream in a fuel circuit CQ. In this example, the pump 11 and the turbine 12 are connected as a single unit via a shaft 13 and rotate synchronously.

[0050] Fuel circuit CQ (in) Figure 4 (Outlined with dashed lines) Therefore, it includes the inlet that is fluidly connected to the cryogenic chamber RC and the supply outlet S of the turbine engine T.

[0051] Turbine 12 is mounted in fuel circuit CQ so that it is driven to rotate by the circulation of fuel Q. In other words, pump 11 enables the increase of fuel pressure Q in fuel circuit CQ downstream of pump 11. This pressure allows turbine 12 to be driven.

[0052] The pumping turbine 1 is preferably in the form of an expansion hydrogen turbopump. Advantageously, the turbine 12 and the pump 11 are completely immersed in fuel, especially pure hydrogen, thereby reducing dynamic sealing problems associated with the drive, particularly in the case of electric motor drive. Any leakage from the pump 11 can be reinjected into the turbine 12 upstream or downstream. Rotational guidance can be advantageously ensured by fluid bearings, thus guaranteeing optimal service life without the need for lubrication.

[0053] like Figure 4 As shown, a heating turbine 2 is illustrated, which is configured to acquire airflow A at an airflow inlet EA, preferably outside the aircraft or within the turbine engine T, and supply airflow A to a first heat exchanger 31. For example, the airflow inlet EA may correspond to airflow acquired at the compressor of the turbine engine T.

[0054] The heating turbine 2 includes an intake compressor 21, a combustion chamber 24, and an exhaust turbine 22, which drives the intake compressor 21.

[0055] In this example, the intake compressor 21 and the exhaust turbine 22 are integrally rotatably connected via shaft 23. The heating turbine 2 causes the airflow A obtained from the upstream to be heated and then discharged from the downstream.

[0056] In this embodiment, the first heat exchanger 31 is supplied with an exhaust flow A (at a high temperature) downstream, and the second heat exchanger 32 is supplied with an intake flow A (at a temperature lower than the exhaust temperature) upstream. Preferably, the second heat exchanger 32 is located upstream of the first heat exchanger 31 in the fuel circuit CQ. Therefore, the second heat exchanger 32 performs a preheating function relative to the first heat exchanger 31, which provides the main heating effect. This arrangement optimizes heating efficiency and prevents water condensation and / or icing in the air that comes into contact with the fuel Q in the first heat exchanger 31. Therefore, fuel at an optimal temperature is supplied to the outlet S of the fuel circuit CQ.

[0057] According to the present invention, the combustion chamber 24 is supplied with air obtained from the airflow A and fuel Q from the fuel circuit CQ and obtained from downstream of the turbine 12. Therefore, the combustion chamber 24 of the heated turbine 2 is supplied with fuel Q having optimal temperature and optimal pressure.

[0058] The heating turbine 2 is preferably in the form of a Brayton cycle turbine and allows the generation of hot airflow in order to provide heat to the fuel Q.

[0059] Still refer to Figure 4 The regulating system SC includes a flow divider 4, which includes, for example, a three-way valve configured to receive fuel Q from turbine 12 at the inlet and supply fuel to the combustion chamber 24 of the heated turbine 2 on the one hand and to the supply outlet S of the turbine engine T on the other.

[0060] According to the optional aspects, refer to Figure 5The regulating system SC also includes an auxiliary heat exchanger 33, which is configured to heat the fuel Q upstream of the fuel circuit CQ by circulating the fuel Q that circulates downstream of the turbine 12. Preferably, the auxiliary heat exchanger 33 is located near the pumping turbine 1. The fuel / fuel type of the auxiliary heat exchanger 33 allows for gradual heating of the fuel at the outlet of the cryogenic chamber RC. This prevents water in the air in contact with the fuel Q in the second heat exchanger 32 and / or the first heat exchanger 31 from condensing and / or freezing.

[0061] Additionally, at least one heat exchanger may be located upstream of turbine 12 or upstream of the first heat exchanger 31 to provide heat from a heat source from the aircraft and / or turbine engine T.

[0062] Depending on the optional aspect, refer to Figure 6 The regulating system SC also includes a generator 51 driven by a free turbine 52, which is driven by airflow A from the heated turbine 2. In this example, the generator 51 is driven by the free turbine 52 via a shaft 53. The rotation of the free turbine 52 advantageously enables the generation of electrical energy to supply non-propulsion functions, such as those of an aircraft, thereby reducing the need to obtain mechanical power from the turbine engine.

[0063] According to another aspect of the invention, the generator 51 may be connected to the heating turbine 2, in particular directly or via a gear set (not shown).

[0064] Now refer to Figure 4 A method for regulating fuel using a regulating system SC according to the present invention is described.

[0065] During the method, pump 11 of pumping turbine 1 obtains fuel from cryogenic tank RC and circulates it from upstream to downstream in fuel system CQ. During pumping, the pressure of fuel Q increases.

[0066] Advantageously, the circulation of fuel Q in the fuel circuit CQ drives the turbine 12 of the pump turbine 1 to rotate, so that motion can be generated by using the enthalpy of fuel Q. Furthermore, after the drive, the pressure of fuel Q is optimal for supplying the combustion chamber, i.e., the pressure of the turbine engine T and the pressure of the heating turbine 2.

[0067] In this example, fuel Q is first heated in the fuel circuit CQ by a second heat exchanger 32 by circulating airflow A obtained from upstream of the heating turbine 2. Airflow A ensures preheating. Then, fuel Q is heated in the fuel circuit CQ by a first heat exchanger 31 by circulating airflow A downstream of the heating turbine 2. The combustion chamber 24 of the heating turbine 2 is supplied with air obtained from airflow A and fuel Q obtained from the fuel circuit CQ and downstream of the turbine 12. The heating turbine 2 can perform a Brayton cycle to bring heat directly to the fuel circuit CQ. Therefore, the temperature and pressure of fuel Q are increased by the regulation system SC for direct supply to the turbine engine T.

[0068] Reference Figure 3 Positioning the regulating system SC within the aircraft baseline REF-A reduces the size of the turbine engine baseline REF-T. This regulating system SC achieves a balance between simplicity and implementation safety in connecting it to the recirculation pipe 6 of the turbine engine T. In fact, since the regulating system SC can increase the pressure and temperature of fuel Q, there is no need to isolate the recirculation pipe 6 connecting the regulating system SC and the turbine engine T. Fuel Q is compressed and heated to a supercritical state, where its implementation will not present any difficulties related to the inherent thermal instabilities of cooling. The pressurized gas can then be advantageously distributed using common techniques, replacing the more complex solutions required for cryogenic fluids.

[0069] Advantageously, thanks to the regulating system SC, fuel regulation in the cryogenic chamber RC is separated from fuel use in the turbine engine T, simplifying management. Advantageously, the speeds of the pumping turbine 1 and the heating turbine 2 can be independently adjusted to optimize the heating of fuel Q.

[0070] Furthermore, advantageously, it eliminates the need for an exchanger at the nozzle of the turbine engine T as is present in the prior art. The performance of the turbine engine T, especially its performance related to the flow of the main airflow, remains unaffected.

[0071] Advantageously, nitrogen can be used as a neutral gas for the purification and purging of circulation line 6, which is more convenient than using helium. Finally, the regulation system SC achieves autonomous operation by consuming a small portion of fuel Q pumped into the cryogenic chamber RC.

[0072] According to another aspect of the invention, referring to Figure 7The regulating system SC is configured to provide one or more fuel buffer reservoirs 7, hereinafter referred to as storage containers 7. Storage containers 7 ensure sufficient fuel flow Q during the transition phase of the turbine engine T or when the regulating system SC does not provide sufficient flow. When fuel flow demand is low, the regulating system SC supplies fuel to the storage containers 7. Therefore, even under deteriorating conditions, the turbine engine T always has fuel at high pressure and ambient temperature for its use.

Claims

1. A fuel regulation system (SC) configured to supply fuel (Q) from a cryogenic chamber (RC) to a turbine engine (T) of an aircraft, characterized in that, The regulation system (SC) includes: - At least one pumping turbine (1), which is separate from the aircraft turbine engine (T), includes a pump (11) and a turbine (12) for driving the pump (11), the pump (11) being configured to obtain the fuel (Q) from the cryogenic chamber (RC) and circulate the fuel from upstream to downstream in a fuel circuit (CQ), the fuel circuit (CQ) including a supply outlet (S) of the turbine engine (T), the turbine (12) being mounted in the fuel circuit (CQ) to drive the turbine (12) to rotate by the circulation of the fuel (Q); - A first heat exchanger (31), located upstream of the turbine (12), is configured to heat the fuel (Q) in the fuel circuit (CQ) by circulating the airflow (A); - At least one heating turbine (2), which is separate from the aircraft turbine engine (T), is configured to supply the airflow (A) to the first heat exchanger (31). The heating turbine (2) includes an intake compressor (21), a combustion chamber (24), and an exhaust turbine (22) configured to drive the intake compressor (21). The combustion chamber (24) is supplied with air in the airflow (A) and fuel (Q) in the fuel circuit (CQ).

2. The fuel regulation system as claimed in claim 1, characterized in that, Includes a second heat exchanger (32) configured to heat fuel (Q) in the fuel circuit (CQ) by circulating the intake airflow of the heating turbine (2).

3. The fuel regulation system as described in claim 1, characterized in that, It includes an auxiliary heat exchanger (33) configured to heat the fuel (Q) in the upstream portion of the fuel circuit (CQ) by means of fuel (Q) circulating downstream of the turbine (12).

4. The fuel regulation system as claimed in claim 1, characterized in that, This includes a generator (51) configured to be driven by the heating turbine (2).

5. The fuel regulation system as claimed in claim 1, characterized in that, The heating turbine (2) is configured to perform a Brayton cycle.

6. The fuel regulation system as claimed in claim 1, characterized in that, The pumping turbine (1) is completely submerged in the fuel (Q).

7. The fuel regulation system as claimed in claim 1, characterized in that, The pumping turbine (1) is an expansion hydrogen turbine pump.

8. A component, characterized in that, It includes at least one cryogenic chamber (RC), an aircraft turbine engine (T), and a conditioning system (SC) as described in any one of claims 1-7, the conditioning system connecting the cryogenic chamber (RC) to the aircraft turbine engine (T).

9. The component as claimed in claim 8, characterized in that, It includes at least one fuel buffer storage (7), which is configured to be supplied by the regulation system (SC) and configured to supply the aircraft turbine engine (T).

10. A method for regulating fuel (Q) using a regulating system (SC) according to any one of claims 1-7 to supply fuel (Q) from a cryogenic chamber (RC) to an aircraft turbine engine (T), characterized in that, Includes the following steps: - The pump (11) of the pumping turbine (1) obtains fuel from the cryogenic chamber (RC) and circulates the fuel from upstream to downstream in the fuel circuit (CQ). - The airflow (A) of the heating turbine (2) is circulated so that the first heat exchanger (31) heats the fuel (Q) in the fuel circuit (CQ). - The circulation of fuel (Q) in the fuel circuit (CQ) drives the turbine (12) of the pumping turbine (1) to rotate. - The combustion chamber (24) of the heated turbine (2) is supplied with air obtained from the airflow (A) and fuel (Q) from the fuel circuit (CQ).

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