Fuel system for an aircraft turbomachine with a closed-loop heat transfer fluid system and its pressure control method

Abstract

A fuel system (20) for an aircraft turbomachine comprising a fuel circuit (1) including a heat exchanger (4) configured to heat the fuel (H) by heat exchange with a heat transfer fluid (F) circulating in a closed circuit (10), the fuel system (20) including a pressure control device (17) including a housing (11) in which is mounted a separator (14) separating a first compartment (12) and a second compartment (13), the first compartment (12) being supplied with fuel (H) by the fuel circuit (1) at a first pressure, the second compartment (13) being supplied with heat transfer fluid (F) by the closed circuit (10) at a second pressure, the separator (14) being adapted to be moved under the effect of variations in the first pressure and / or the second pressure, to modify the volume of the first compartment (12) and the second compartment (13).Figure from the summary: Figure 1.

Description

Title of the invention: Fuel system for an aircraft turbomachine with a closed-loop heat transfer fluid circuit and its pressure control method. Technical field

[0001] The present invention relates to the field of pressure control in a closed circuit of heat transfer fluid for an aircraft turbomachine fuel system.

[0002] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies to both new types of aircraft and those already in service, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively working 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 account the factors affecting all phases of design and development in order to obtain aeronautical components and products that are less energy-intensive, more environmentally friendly, and whose integration and use in civil aviation have moderate environmental consequences, with the aim of improving the energy efficiency of aircraft.

[0004] This sustained research and development work focuses in particular on new generations of aircraft turbomachinery powered by cryogenic fuels, such as hydrogen. Such an aircraft turbomachine conventionally comprises a cryogenic fuel circuit including, from upstream to downstream according to the direction of flow of the cryogenic fuel, a tank, a pressurization unit, a heating unit, a metering unit and a set of injectors.

[0005] As is known, the tank stores hydrogen in liquid form at a low temperature of approximately -253°C and a low pressure of less than 10 bar, in order to reduce its storage volume. The pressurization device, typically a pump, increases the pressure of the liquid hydrogen to maintain a pressure at the injectors higher than that in the combustion chamber. The heating device is typically in the form of a heat exchanger, in which the circulation of a heat transfer fluid, such as nitrogen or helium in a gaseous or supercritical state, heats the liquid hydrogen to make it gaseous for combustion. The metering device, typically a valve known by its English name "fuel metering valve," allows the mass flow rate of gaseous hydrogen to be metered and distributed to the injectors for injection into the combustion chamber.

[0006] In a known manner, the heat transfer fluid supplying the heat exchanger of the cryogenic fuel circuit circulates in a closed loop within the aircraft turbomachine. After heating the cryogenic fuel, the heat transfer fluid is guided, via a cold branch, to a second heat exchanger where it is heated by a hot air stream, such as that circulating in the nozzle of the aircraft turbomachine. The heat transfer fluid is then reinjected into the heat exchanger of the cryogenic fuel circuit via a hot branch.

[0007] In practice, the average temperature of the heat transfer fluid in the closed circuit is subject to significant variations depending on the operating points within the flight envelope. This generates significant variations in absolute pressure within the closed circuit, typically exceeding 100 bar, which severely constrain the mechanical strength of the equipment and operational safety, particularly by increasing the risk of leaks in the cryogenic fuel circuit.

[0008] The invention thus aims to eliminate at least some of these drawbacks. PRESENTATION OF THE INVENTION

[0009] The invention relates to a fuel system for an aircraft turbomachine comprising: • a fuel circuit including a heat exchanger configured to heat the fuel by heat exchange with a heat transfer fluid, and • a closed circuit of heat transfer fluid supplying the heat exchanger.

[0010] The invention is remarkable in that the fuel system includes a pressure control device comprising a chamber of determined total fluid volume and a separating element mounted in the chamber, the separating element separating in the chamber a first compartment and a second compartment, the first compartment being supplied with fuel by the fuel circuit at a first pressure, the second compartment being supplied with heat transfer fluid by the closed circuit at a second pressure, the separating element being adapted to be moved under the effect of variations in the first pressure and / or the second pressure, so as to modify the fluid volume of the first compartment and the second compartment.

[0011] The invention advantageously allows the fluid volume of the second compartment to be varied, so as to vary the pressure of the heat transfer fluid in the Closed circuit. This advantageously limits the absolute pressure variations occurring in the closed circuit during aircraft flight, typically exceeding 100 barA. The pressure control device according to the invention advantageously operates passively under the effect of the first fuel pressure, the second heat transfer fluid pressure, and optionally a third pressure from a return element. It advantageously takes advantage of the fact that the first pressure varies oppositely to the second pressure during aircraft flight. Furthermore, the pressure control device, according to a second function, filters high-frequency thermo-hydraulic disturbances in the fuel circuit, without requiring a dedicated accumulator. This further reduces the size and cost of the fuel system.

[0012] According to one aspect of the invention, the fuel is cryogenic and is configured to be heated from a liquid to a gaseous state in the heat exchanger. By limiting pressure variations in the closed circuit, the invention advantageously reduces the mechanical stresses on the equipment of the cryogenic fuel circuit, thereby increasing its lifespan and operational reliability. The risk of leakage from the cryogenic fuel circuit is advantageously reduced.

[0013] According to one aspect of the invention, the heat transfer fluid is in the form of a fluid in the gaseous or supercritical state, preferably nitrogen or helium, therefore compressible.

[0014] According to one aspect of the invention, the first compartment is mounted on the fuel circuit downstream of the heat exchanger in the direction of fuel flow. The cryogenic fuel is in a gaseous state. The initial pressure in the first compartment corresponds to that controlled by the pressurization device.

[0015] Preferably, the second compartment is mounted on the closed circuit downstream of the heat exchanger, in the direction of flow of the heat transfer fluid. The enclosure advantageously has a small footprint.

[0016] According to one aspect of the invention, the pressure control device comprises a return element mounted in the first compartment and configured to exert a third pressure on the separator element in addition to the initial fuel pressure. The return element reinforces the back pressure exerted on the separator element. The return element compensates for any potential pressure drop in the fuel circuit.

[0017] According to one aspect of the invention, the pressure control device includes a stop preventing the separation member from moving below a minimum volume of the first compartment. The stop ensures a minimum volume in the second compartment to filter high frequencies of thermo-hydraulic disturbances in the fuel circuit.

[0018] According to one aspect of the invention, the minimum volume of the first compartment is at least 1.1 times greater than the maximum volume of the second compartment in order to filter high-frequency thermo-hydraulic disturbances in the fuel circuit without requiring a dedicated accumulator. Preferably, the minimum volume of the first compartment is at most six times greater than the maximum volume of the second compartment to limit its size.

[0019] According to one aspect of the invention, the separating element is in the form of a membrane or a piston. The separating element hermetically seals the first compartment and the second compartment.

[0020] The invention also relates to an aircraft turbomachine comprising a fuel system as described above.

[0021] The invention also relates to a method of controlling the pressure in a closed circuit of heat transfer fluid of a fuel system as described above, in which, during a high speed of the aircraft turbomachine, the first pressure in the first compartment increases and the second pressure in the second compartment decreases, which moves the separating member to decrease the volume of the second compartment.

[0022] The invention also relates to a method of controlling the pressure in a closed circuit of heat transfer fluid of a fuel system as described above, in which, at a low speed of the aircraft turbomachine, the first pressure in the first compartment decreases and the second pressure in the second compartment increases, which moves the separating member to decrease the volume of the second compartment. PRESENTATION OF THE FIGURES

[0023] 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.

[0024] Fig. 1 is a schematic representation of an aircraft turbomachine fuel system according to one embodiment of the invention.

[0025] Fig. 2 is a schematic representation of the fuel system of Fig. 1 during high operating speed of the aircraft turbomachine.

[0026] Fig. 3 is a schematic representation of the fuel system of Fig. 1 during a low operating speed of the aircraft turbomachine.

[0027] Fig. 4 is a schematic representation of an aircraft turbomachine fuel system according to another embodiment of the invention.

[0028] It should be noted that the figures set out the invention in detail to implement the invention, said figures being of course able to serve to better define the invention where appropriate. DETAILED DESCRIPTION OF THE INVENTION

[0029] With reference to [Fig. 1], the invention relates to an aircraft turbomachine and more specifically to its fuel system 20 which supplies the combustion chamber with fuel H. The fuel system 20 comprises a fuel circuit 1 which typically includes, from upstream to downstream according to the direction of flow of the fuel H, one or more tanks 2, a pressurization unit 3, a heat exchanger 4, a metering unit 5 and a set of injectors 6. The fuel H is preferably of the cryogenic type and is, for example, in the form of hydrogen.

[0030] As illustrated in [Fig. 1], the tanks 2 allow the cryogenic fuel H to be stored in a liquid state, at a low temperature of approximately -253°C and at a low pressure of less than 10 bar, so as to reduce its storage volume. The pressurization device 3, such as a pump, increases the pressure of the cryogenic fuel H to a value higher than that in the combustion chamber. The heat exchanger 4 heats the cryogenic fuel H from a liquid to a gaseous state by heat exchange with a heat transfer fluid F. The metering device 5, typically a valve known by its English name "fuel metering valve," meters the mass flow rate of gaseous cryogenic fuel H distributed to the injectors 6 for injection into the combustion chamber.

[0031] With reference to [Fig. 1], the fuel system 20 also includes a closed circuit 10 of heat transfer fluid F for supplying the heat exchanger 4 of the fuel circuit 1 with heat transfer fluid F. The closed circuit 10 is so described in that no exchange of matter takes place between the heat transfer fluid F and the outside. The heat transfer fluid F is typically nitrogen or helium, in a gaseous or supercritical state, and therefore compressible.

[0032] As illustrated in [Fig. 1], the closed circuit 10 typically includes a hot branch 8, configured to guide the heat transfer fluid F into the heat exchanger 4 of the fuel circuit 1, and a cold branch 7, configured to guide the fuel-cooled heat transfer fluid F out of the heat exchanger 4 of the fuel circuit 1. The closed circuit 10 also typically includes a drive pump 9 and a second heat exchanger 19. The second heat exchanger heat 19 is configured to heat the heat transfer fluid F supplied by the cold branch 7 and reinject it into the hot branch 8. The second heat exchanger 19 is typically mounted in contact with a hot air stream A from the aircraft turbomachine, for example located at the nozzle.

[0033] According to the invention and with reference to Figures 1 and 2, the fuel system 20 also includes a pressure control device 17 comprising a chamber 11 with a determined total fluid volume Vtot and a separating element 14 mounted in the chamber 11. In the chamber 11, the separating element 14 separates a first compartment 12 and a second compartment 13. The first compartment 12 is supplied with fuel H by the fuel circuit 1 at a first pressure PL. The second compartment 13 is supplied with heat transfer fluid F by the closed circuit 10 at a second pressure P2. Also according to the invention, the separating element 14 is adapted to be displaced under the effect of pressure variations PI, P2 of the first compartment 12 and / or the second compartment 13, so as to modify their fluid volume VI, V2.

[0034] As illustrated in Figures 2 and 3, the pressure control device 17 according to the invention allows the fluid volume V2 of the second compartment 13 to be decreased or increased, thereby respectively increasing or decreasing the second pressure P2 of the heat transfer fluid F in the closed circuit 10. This advantageously limits variations in the second pressure P2 in the closed circuit 10 that occur during aircraft flight. The second pressure P2 is directly related to the average temperature of the heat transfer fluid F in the closed circuit 10, which varies according to the operating conditions of the aircraft turbomachine and external conditions.

[0035] As illustrated in Figures 2 and 3, the pressure control device 17 advantageously operates passively. The separating element 14 is moved solely by the first pressure PI of the fuel H, the second pressure P2 of the heat transfer fluid F, and preferably a third pressure P3 from a return element 16 as described below. No manual or electronic control is required. Advantageously, the fact that the first pressure PI varies inversely with the second pressure P2 during aircraft flight is exploited: at a high R1 of the aircraft turbomachine (see [Fig. 2]), the first pressure PI controlled by the pressurization element 3 increases while the second pressure P2 decreases in the closed circuit 10 due to a drop in the average temperature of the heat transfer fluid F. At a low R2 (see [Fig. 3]), the first pressure PI controlled by the pressurization element 3 increases while the second pressure P2 decreases in the closed circuit 10 due to a decrease in the average temperature of the heat transfer fluid F.3]), the first pressure PI controlled by the pressurization device 3 decreases while the second pressure P2 increases in the closed circuit 10 under the effect of a rise in the average temperature of the heat transfer fluid F. .

[0036] Thus, the first pressure PI provides a significant back pressure relative to the second pressure P2, enabling or contributing to enabling the movement of the separation element 14 without the need to integrate or order dedicated, bulky, and expensive equipment. Furthermore, the pressure control device 17, according to a second function, filters high-frequency thermo-hydraulic disturbances in the fuel circuit 1, without requiring a dedicated accumulator. This further reduces the size and cost of the fuel system 20.

[0037] By limiting pressure variations in the closed circuit 10, the invention advantageously reduces the mechanical stresses on the equipment, particularly the cryogenic fuel circuit 1, and thus increases its lifespan and operational reliability. The risk of leakage from the cryogenic fuel circuit 1 is advantageously reduced.

[0038] As illustrated in [Fig. 1], the enclosure 11 is typically mounted on the fuel circuit 1 downstream of the pressurization device 3, the initial pressure PI corresponding to the pressure controlled by the pressurization device 3 so as to be higher than that in the combustion chamber 3. Preferably, the enclosure 11 is mounted on the fuel circuit 1 downstream of the heat exchanger 3, so that the cryogenic fuel H is in a gaseous state, and upstream of the metering device 5, in order to filter out high-frequency thermo-hydraulic disturbances before metering and injection into the combustion chamber. Also preferably, the enclosure 11 is mounted on the closed circuit 10 downstream of the heat exchanger 3, as close as possible to the fuel circuit 1 to minimize its size.

[0039] With reference to Figures 2 and 3, the enclosure 11 comprises a rigid wall, in this example cylindrical, externally delimiting the first compartment 12 and the second compartment 13. The enclosure 11 typically includes an inlet and an outlet opening into the first compartment 12 for the intake and discharge of fuel H, as well as an inlet and an outlet opening into the second compartment 13 for the intake and discharge of the heat transfer fluid F. The total fluid volume Vtot of the enclosure 11 is constant and satisfies: Vtot = VI + V2, the first fluid volume VI and the second fluid volume V2 being variable depending on the position of the separating element 14. The total fluid volume Vtot is preferably greater than 25 L and less than 75 L to absorb the pressure variations P2 of the closed circuit 10.

[0040] In the example shown in Figures 1 to 3, the separating element 14 is in the form of a membrane. The separating element 14 could also be in the form of a piston, for example. The first compartment 12 and the second compartment 13 are fluidly separated in a sealed manner by the separating element 14.

[0041] With further reference to Figures 1 to 3, the first compartment 12 preferably extends vertically above the second compartment 13, depending on gravity. The back pressure applied to the separating element 14, formed by the first pressure PI of the fuel H and preferably the third pressure P3 of a return element 16, is thus reinforced by taking advantage of gravity.

[0042] As illustrated in figures 1 to 3, the pressure control device 17 preferably includes a stop 15 preventing the separation member 14 from moving below a minimum volume of the first compartment 12. The separation member 14 is free to move above said minimum volume of the first compartment 12 according to pressure variations PI, P2, P3.

[0043] In the presence of a second high pressure P2 in the closed circuit 10, this ensures a minimum volume in the first compartment 12, preferably at least 1.1 times greater than the maximum volume of the second compartment 13, to filter high-frequency thermo-hydraulic disturbances in the fuel circuit 1, without requiring a dedicated accumulator. Preferably, the minimum volume of the first compartment 12 is at most six times greater than the maximum volume of the second compartment 13 to limit bulk and effectively absorb pressure variations P2 in the closed circuit 10.

[0044] In the example shown in Figures 1 to 3, the stop 15 protrudes into the first compartment 12 from the wall of the enclosure 11, so as to prevent the separation element 14 from moving beyond the minimum volume. Also in this example, the stop 15 extends over a portion of the circumference of the first compartment 12, but it is understood that the stop 15 could be circumferential. Several stops 15 could also be distributed around the circumference of the first compartment 12.

[0045] As illustrated in Figures 1 to 3, the pressure control device 17 preferably comprises a return element 16 mounted in the first compartment 12 and configured to exert a third pressure P3 on the separating element 14. In this example, the return element 16 is in the form of a spring connecting the separating element 14 to an opposite face of the wall of the enclosure 11. The third pressure P3 and the first pressure PI together form a counter-pressure opposing the second pressure P2. This advantageously compensates for a potentially insufficient first pressure PI in the fuel circuit 1.

[0046] Preferably, the third pressure P3 exerted on the separating element 14 is lower than the first pressure PL. The first pressure PI of the fuel H in the fuel circuit 1 is advantageously sufficient to provide the majority of the back pressure. Such a return element 16 is advantageously passive and compact.

[0047] Alternatively, in the example of [Fig.4], the pressure control device 17 is free of a return member 16. The first pressure PI of the fuel H is advantageously sufficient to ensure the entire back pressure.

[0048] A method for controlling the second pressure P2 in the closed circuit 10 of heat transfer fluid F is described below.

[0049] As illustrated in [Fig. 2], during high R1 of the aircraft turbomachine, the first pressure PI controlled by the pressurization device 3 in the fuel circuit 1 increases. Conversely, the second pressure P2 in the closed circuit 10 decreases because the average temperature of the heat transfer fluid F decreases due to the increased cooling requirements of the fuel H. The fuel H and the return device 16 exert a first pressure PI and a third pressure P3, respectively, on the separation device 14, the sum of which is greater than that of the second pressure P2 of the heat transfer fluid F. This displaces the separation device 14 in a direction that reduces the fluid volume V2 of the second compartment 13 and increases the fluid volume VI of the first compartment 12. This results in an increase in the second pressure P2 in the closed circuit 10. The first pressure PI in the open-type fuel circuit 1 is not affected.The separating element 14 ceases to move when the equilibrium of the pressures PI, P2, P3 is restored.

[0050] As illustrated in [Fig. 3], during a low R2 operating speed of the aircraft turbomachine, the first pressure PI controlled by the pressurization device 3 in the fuel circuit 1 decreases. Conversely, the second pressure P2 in the closed circuit 10 increases because the average temperature of the heat transfer fluid F increases due to the reduced cooling requirements of the fuel H. The second pressure P2 becomes greater than the sum of the first and third pressures PI, P3, which displaces the separator 14 in the opposite direction, increasing the fluid volume V2 of the second compartment 13 and reducing the fluid volume VI of the first compartment 12. This results in a decrease in the second pressure P2 in the closed circuit 10. The first pressure PI in the open-type fuel circuit 1 is not affected. The separator 14 ceases to move when the pressure equilibrium of PI, P2, and P3 is restored or when it comes into contact with the stop 15..

[0051] In the example of [Fig.4], the process described above differs in that no third pressure P3 is exerted by a return member 16, the first pressure PI ensuring the entire back pressure.

[0052] The invention advantageously allows passive and compact control of the pressure P2 in a closed circuit 10 of heat transfer fluid F for an aircraft turbomachine fuel system 1, by exploiting the pressure PI in the circuit fuel 1. The invention avoids having to use an accumulator in the fuel circuit 1.

Claims

Demands

1. A fuel system (20) for an aircraft turbomachine comprising: • a fuel circuit (1) including a heat exchanger (4) configured to heat the fuel (H) by heat exchange with a heat transfer fluid (F), • a closed circuit (10) of heat transfer fluid (F) supplying the heat exchanger (4), • the fuel system (20) being characterized in that it comprises a pressure control device (17) including a chamber (11) of determined total fluid volume (Vtot) and a separator (14) mounted in the chamber (11), the separator (14) separating in the chamber (11) a first compartment (12) and a second compartment (13), the first compartment (12) being supplied with fuel (H) by the fuel circuit (1) at a first pressure (PI), the second compartment (13) being supplied with heat transfer fluid (F) by the closed circuit (10) at a second pressure (P2),the separating element (14) being adapted to be displaced under the effect of variations in the first pressure (PI) and / or the second pressure (P2), so as to modify the fluidic volume (VI, V2) of the first compartment (12) and the second compartment (13).

2. Fuel system (20) according to claim 1, wherein the fuel (H) is of cryogenic type and is configured to be heated from a liquid state to a gaseous state in the heat exchanger (4).

3. Fuel system (20) according to any one of claims 1 and 2, wherein the heat transfer fluid (F) is in the form of a fluid in the gaseous or supercritical state, preferably nitrogen or helium.

4. Fuel system (20) according to any one of claims 1 to 3, wherein the first compartment (12) is mounted on the fuel circuit (1) downstream of the heat exchanger (4) in the direction of fuel flow (H).

5. Fuel system (20) according to any one of claims 1 to 4, wherein the pressure control device (17) includes a return member (16) mounted in the first compartment (12) and configured to exert a third pressure (P3) on the separator member (14) in addition to the first pressure (PI) of the fuel (H).

6. Fuel system (20) according to any one of claims 1 to 5, wherein the pressure control device (17) includes a stop (15) prohibiting the movement of the separator member (14) below a minimum volume of the first compartment (12).

7. Fuel system (20) according to any one of claims 1 to 6, wherein the minimum volume of the first compartment (12) is at least 1.1 times greater than the maximum volume of the second compartment (13), preferably at most six times greater.

8. Fuel system (20) according to any one of claims 1 to 7, wherein the separation member (14) is in the form of a membrane or a piston.

9. A method for controlling the pressure in a closed circuit (10) of heat transfer fluid (F) of a fuel system (20) according to any one of claims 1 to 8, wherein, during a high speed (RI) of the aircraft turbomachine, the first pressure (PI) in the first compartment (12) increases and the second pressure (P2) in the second compartment (13) decreases, which moves the separating member (14) to decrease the volume (V2) of the second compartment (13).

10. A pressure control method according to claim 9, wherein, during a low engine speed (R2) of the aircraft turbomachine, the first pressure (PI) in the first compartment (12) decreases and the second pressure (P2) in the second compartment (13) increases, which moves the separating member (14) to decrease the volume (V2) of the second compartment (13).

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