Improved aircraft fuel tank water recovery system

Abstract

Improved Aircraft Fuel Tank Water Recovery Device A device (10) for recovering water from the main aircraft fuel tank (41) disposed on a fuel supply line (51) connecting a fuel pump (30) and an aircraft engine (1), the device (10) being disposed downstream of the fuel pump (30) and upstream of the engine (1), and comprising an inlet (10a) connected to the fuel pump (30) and an outlet (10b) connected to the engine (1), a main conduit (12) on which is disposed an activation valve (13) movable between a closed position preventing fuel flow through the main conduit (12), and an open position allowing such flow, a secondary conduit (14) parallel to the main conduit (12) and having a convergent, a divergent and a throat,the device (10) further comprising at least one recovery tube (15) connecting an orifice in the neck of the secondary conduit (14) and the bottom of at least one secondary fuel tank (42) of the aircraft. Fig. 4.

Description

Title of the invention: Improved aircraft fuel tank water recovery device. Technical field

[0001] The present invention relates to the field of aircraft engine fuel supply systems, and in particular to the drainage of water from the bottom of the fuel tank, and more specifically to a fuel tank water recovery device for aircraft turbomachinery. The present invention can notably be used for an aircraft turbojet engine or a turboprop engine. Prior art

[0002] Aircraft engines conventionally include fuel supply systems, drawing fuel from tanks located, in particular, in the wings of the aircraft. During its journey from these tanks to the aircraft's engine (or jet engine), the fuel first flows through various sub-compartments of these tanks belonging to the aircraft and then through a pipe belonging to the engine's fuel supply line.

[0003] During the operation of aircraft engines, water naturally present in the air inside the fuel tanks can condense on their walls and fall back into the tanks by gravity. Since water is denser than fuel, it tends to settle at the bottom of the tank. Similarly, water naturally present in the fuel that remains stagnant in the tanks can also settle at the bottom of the tank.

[0004] Fuel pumps, and more specifically lift pumps drawing fuel from the tanks, may therefore draw fuel / water mixtures with high water concentrations, these mixtures then being sent to the engines. This can cause problems during certain phases of flight, particularly at startup when the engine temperature is still low, as the water can then freeze and clog the filters in the engines / jets, thus hindering their proper operation. Consequently, water injection into the jets is only possible above a certain operating temperature.

[0005] To limit the aforementioned risks, devices and methods exist for draining / recovering water from fuel tanks. One solution involves periodically performing manual drainage operations to recover the water that has settled at the bottom of the tank when the aircraft is at rest.

[0006] However, this sedimentation method presents implementation difficulties. Indeed, the drains, generally located in the bottom wall of the tanks, are subject to freezing, which can prevent their use, or even damage them if they They can become stuck in the open or closed position. Furthermore, this operation can be complex (drains can be difficult to access on large aircraft) and time-consuming, as the drainage process can take longer when frozen water is present at the bottom of the tank.

[0007] Another solution is to agitate the fuel in the tank to homogenize the amount of water in the fuel. Thus, during reactor operation, the fuel delivered by the feed pump is a water / fuel mixture with a low water concentration that does not affect combustion. This agitation can be achieved by a dedicated ejector-type device. Alternatively, a water recovery ejector can be located at the lowest point of the tank, which is supplied with driving pressure and engine flow by the feed pump and which provides its flow at the pump's inlet.

[0008] However, the mixing method requires the integration of dedicated pressure lines for this equipment, increasing the system's mass and complexity. It also requires an additional flow rate to be supplied by the feed pump. Furthermore, in addition to the need for dedicated pressure lines, a higher concentration of water is sent to the reactor via the feed pump when the water recovery ejector is located at the lowest point of the tank. These high water concentrations can compromise the electrically non-conductive nature of the fuel in which the pump is immersed, which can damage the feed pump since it is equipped with an electric motor immersed in the fuel.

[0009] There is therefore a real need for a fuel tank water recovery device that is free, at least in part, from the disadvantages inherent in the aforementioned known solutions. Description of the invention

[0010] The present description relates to a device for recovering water from an aircraft's main fuel tank, the device being suitable for being placed on a fuel supply line connecting a fuel pump and an aircraft engine to be supplied, the device being placed downstream of the fuel pump and upstream of the engine, and comprising: - an inlet suitable for connection to the fuel pump and an outlet suitable for connection to the engine, - a main conduit on which is arranged a movable activation valve between a closed position preventing fuel flow through the main conduit, and an open position allowing this flow, - a secondary conduit parallel to the main conduit and having a convergent, a divergent and a neck, the device further comprising at least one recovery tube connecting an orifice of the neck of the secondary conduit and the bottom of at least one secondary fuel tank of the aircraft.

[0011] In the present exposition, the terms "upstream" and "downstream" are defined with respect to the direction of fuel flow in the various conduits of the device, from the fuel pump to the engine.

[0012] It is understood that the main fuel tank is the tank in which the fuel pump, or lift pump, is located, and the secondary tank(s) are tanks adjacent to the main tank, communicating with it, and also containing fuel.

[0013] It is further understood that, since the main conduit and the secondary conduit are in parallel with each other, the inlet and outlet of the device are common to the main conduit and the secondary conduit, the inlet allowing both the main conduit and the secondary conduit to be supplied.

[0014] Thus, when the activation valve is in the open position, a large proportion of the fuel taken by the fuel pump passes through the main line towards the engine, and a smaller proportion of the fuel passes through the secondary line, which generates little pressure loss.

[0015] Conversely, when the activation valve is in the closed position, all the fuel drawn by the fuel pump passes through the secondary line. However, the presence of the convergent section causes an acceleration of the fuel flow, and therefore a decrease in static pressure at the neck. The resulting vacuum causes the water present at the bottom of the secondary tank to be drawn through the recovery tube connecting the neck to the bottom of said tank.

[0016] The water recovery device according to the invention has the advantage of being able to be activated on demand, at any time during flight, by opening or closing the activation valve. In particular, when the fuel temperature is sufficient to allow water to be sent to the engine without risking icing of the engine's main filter, the activation valve can be closed to activate water drainage through the secondary channel. Conversely, when the temperature is too low, for example during startup, the activation valve is preferably open to prevent such drainage.

[0017] Furthermore, the water recovery device according to the invention does not require the addition of dedicated drive pressure lines and allows for localized action in the areas of the secondary reservoir where the water is located, enabling more efficient extraction thanks to the remote suction of the recovery tube. In addition, the fact that the strong Water concentrations are taken downstream of the fuel pump to limit damage to its electric motor.

[0018] Thus, the water recovery device according to the invention has the advantage of being efficient, simple to implement, while also improving the lifespan of engine and aircraft equipment.

[0019] In certain embodiments, the main conduit is able to be arranged in continuity with the fuel supply line, the secondary conduit being offset from the main conduit and the supply line, and in fluidic communication with an upstream end and a downstream end of the main conduit.

[0020] In some embodiments, the main conduit has a constant cross-section, the secondary conduit being a venturi tube.

[0021] In some embodiments, the activation valve is a ball valve.

[0022] In some embodiments, a longitudinal length of the convergent is shorter than a longitudinal length of the divergent.

[0023] In some embodiments, an angle between a main longitudinal axis of the secondary conduit and a wall of the divergent is between 2 and 4°, preferably between 3 and 3.5°.

[0024] In some embodiments, the neck of the secondary tube comprises a plurality of orifices, a plurality of recovery tubes being able to connect each orifice of the neck with the bottom of at least one secondary fuel tank of the aircraft.

[0025] The present exposition also relates to an aircraft fuel supply system, the system comprising a main fuel tank, a fuel pump disposed in said main tank, a supply line suitable for connecting the fuel pump to an aircraft engine, and a water recovery device according to any of the preceding embodiments, disposed on the supply line between the fuel pump and the engine.

[0026] In some embodiments, the fuel supply system is a first system suitable for being arranged in a first wing of the aircraft and for being connected to a second fuel supply system arranged in a second wing of the aircraft via an interconnecting valve arranged between the first wing and the second wing on an interconnecting channel, the supply line comprising a sampling channel connected to the fuel pump and a main channel on which the device is disposed, the sampling channel supplying both the main channel and the interconnecting channel.

[0027] The present description also relates to an aircraft comprising at least one water recovery device according to any one of the preceding embodiments. Brief description of the drawings

[0028] The invention and its advantages will be better understood upon reading the following detailed description of various embodiments of the invention, given by way of non-limiting examples. This description refers to the accompanying figure pages, on which:

[0029] [Fig-1] Fig. 1 represents an aircraft comprising a power supply system according to the invention;

[0030] [Fig.2] Fig.2 schematically represents a fuel supply network in the two wings of the aircraft, comprising a supply system according to the invention;

[0031] [Fig.3] The [Fig.3] schematically represents a detailed view of the fuel supply system of one of the aircraft's engines;

[0032] [Fig.4] Fig.4 schematically represents a supply system comprising a water recovery device according to the invention;

[0033] [Fig.5] Fig.5 schematically and in isolation represents a water recovery device according to the invention. Description of the implementation methods

[0034] To make the explanation more concrete, an example of a water recovery device 10 and a fuel supply system 50 comprising such a device is described in detail below, with reference to the accompanying drawings. It should be noted that the invention is not limited to this example.

[0035] Fig. 1 represents an aircraft 100 comprising a first engine 1 carried by a first wing 40, and a second engine carried by a second wing 60. The rest of the description is based on the first wing 40 and the first engine 1, but the characteristics described also apply to the second wing 60 and the second engine 2.

[0036] An internal enclosure within the first wing 40 includes fuel storage tanks. In [Fig. 1], only the main tank 41 is shown. The engine 1 is supplied with fuel drawn from the main tank 41, via a fuel supply system 50.

[0037] Fig. 2 schematically represents a top view of the wings 40, 60 of the aircraft 100, illustrating the fuel supply network of the engines 1, 2. This network includes in particular fuel tanks and fuel supply systems according to the invention.

[0038] It should be noted that for the sake of clarity and conciseness, only the engine fuel supply network is shown; the other networks and conduits (fuel mixing / transfer between wings, fuel filling / draining, power unit supply) are not shown. auxiliary power...) also usually present in an aircraft being intentionally masked.

[0039] A fuel supply system 50 is disposed in each wing 40, 60 of the aircraft 100. The system 50 of the first wing 40, for example, includes a main fuel tank 41, a fuel pump 30 (or lift pump) disposed in the main tank 41, a supply line 51 connected to the fuel pump 30 and the engine 1, and a water recovery device 10 described below and disposed on the supply line 51 between the fuel pump 30 and the engine 1.

[0040] The "main tank 41" is the tank in which the fuel pump 30, which pumps fuel to the engine 1, is located and immersed. In this respect, the system is configured so that the main tank 41 is always filled with fuel, ensuring that the fuel pump 30 is constantly immersed in fuel and therefore draws only fuel, and not air.

[0041] To do this, the main tank 41 can be delimited on the one hand by a solid and hermetically sealed wall 43, and on the other hand, by a communicating wall 44 allowing communication with at least one secondary tank 42. In this example, each wing 40, 60 comprises two secondary tanks 42, communicating with each other via a communicating wall 44.

[0042] It is understood that the communicating walls 44, symbolized by dashed lines on [Fig.2], allow fuel to be transferred from one tank to another, for example from secondary tanks 42 to the main tank 41, or vice versa.

[0043] Furthermore, the fuel drawn by the fuel pump 30 flows to the engine 1 via a supply line 51, comprising a draw channel 52 connected to the fuel pump 30, and a main channel 53 connecting the draw channel 52 to the engine 1. It should also be noted that the draw channel 52 branches off on one side towards the main channel 53 in the direction of the engine 1, and on the other side towards an interconnecting channel 54 in the direction of the opposite wing 60. Thus, the fuel supply systems 50 of each wing 40, 60 are interconnected via an interconnecting channel 54, on which an interconnecting valve 70 is located, allowing, when this valve is open, the transfer of fuel from one wing to the other, for example when only one of the fuel pumps 30 needs to supply fuel to both engines 1, 2.

[0044] The fuel supply system 50 further includes a water recovery device 10, disposed here in a secondary tank 42, downstream of the fuel pump 30 and upstream of the engine 1. In this respect, it is understood that the terms "upstream" and "downstream" are defined with respect to the direction of fuel flow in the fuel supply system 50, from its extraction by the fuel pump 30 to its injection into the engine 1. Arrows also symbolize the direction of flow on figures 4 and 5 described later.

[0045] The water recovery device 10 makes it possible to recover the water that has settled and stagnated at the bottom of the tank 42, and to send it to the motor 1 when necessary.

[0046] Figure 3 schematically represents only the first wing 40, illustrating in more detail the water recovery device 10. In addition, Figure 4 schematically represents the fuel supply system 50 including the device 10, and Figure 5 schematically represents a detailed view of the device 10 alone.

[0047] The device 10 is arranged on the supply line 51, in particular on the main channel 53. It comprises an inlet 10a connected to the fuel pump 30, and an outlet 10b connected to the engine 1. It further comprises a main conduit 12 and a secondary conduit 14, in parallel with each other. More precisely, the main conduit 12 is preferably arranged in line with the supply line 51, and in particular with the main channel 53. This makes it possible to limit pressure losses when the fuel flows to the engine 1 via the main conduit 12. Indeed, the pressure loss in the device 10 is then equivalent to that of a straight conduit. Typically, the main conduit 12 is a tube of constant cross-section, for example circular, having a diameter equal to the diameter of the main channel 53.

[0048] The main conduit 12 includes an activation valve 13, movable between a closed position preventing fuel flow through the main conduit 12, and an open position allowing this flow. The activation valve 13 is preferably a ball valve, which helps to limit pressure losses when it is in the open position.

[0049] The secondary conduit 14 is arranged in parallel with the main conduit 12, and offset from it and from the main channel 53. Thus, when the activation valve 13 is open, for the majority of the flight time of the aircraft 100, the majority of the fuel passes through the main conduit 12, with a smaller proportion passing through the secondary conduit 14. As the main conduit 12 is in continuity with the main channel 53, and the majority of the fuel passes through the main conduit 12, pressure losses are consequently minimized.

[0050] Furthermore, the secondary conduit 14 communicates with an upstream end 12a of the main conduit 12, and a downstream end 12b thereof. It is therefore understood that when the activation valve 13 is in the open position and fuel flows from upstream to downstream, the majority of the fuel flows via the main conduit 12 from the upstream end 12a, with a small proportion of fuel being diverted to the conduit secondary 14, this small proportion joining downstream the main conduit 12 at the level of the downstream end 12b.

[0051] Conversely, when the activation valve 13 is in the closed position, all of the fuel flowing from the upstream end 12a of the main conduit 12 is diverted into the secondary conduit 14, then rejoins the main conduit 12 at the downstream end 12b, and then the supply line 51 to the engine 1.

[0052] The secondary conduit 14 is typically a venturi tube, and comprises, from upstream to downstream, a convergent 141, a throat 143 and a divergent 142. Preferably, a longitudinal length of the convergent 141, along a principal longitudinal axis X of the secondary conduit 14, is less than a longitudinal length of the divergent 142. In a venturi tube, the head losses occur mainly in the divergent due to flow separation; maximizing the length of the divergent 142 limits the risk of flow separation, and therefore limits head losses. In other words, with the same overall dimensions, i.e. for a given total length of the secondary conduit 14, increasing the length of the divergent 142, and decreasing that of the convergent 141, the latter having little impact on pressure losses, makes it possible to limit the pressure losses, while limiting the overall dimensions.

[0053] Preferably, in an axial cross-sectional view of the device 10 ([Fig. 5]), an angle [3] between the principal axis X, which is also the central axis of revolution of the venturi tube, which is typically axisymmetric, and the wall of the divergent 142, is between 2 and 4°, preferably between 3 and 3.5°. These values ​​optimize the ratio between the length of the divergent 142, and therefore the size of the secondary conduit 14, and the pressure losses.

[0054] The wall of the neck 143, at which the flow velocity is greatest, and therefore the pressure is lowest, includes at least one orifice 17, to which a recovery tube 15 is attached. The recovery tube 15 may be a flexible tube, connecting the neck 143 to a desired area of ​​the secondary reservoir 42, in particular the bottom of the secondary reservoir 42 where settled water may be found. The use of such a long and flexible recovery tube 15 thus makes it easy to choose the sampling point, in particular a location where the water is most likely to settle.

[0055] Furthermore, although only one recovery tube 15 and one orifice 17 are shown in [Fig. 5], the neck 143 can have two or more orifices 17 and as many recovery tubes 15 as there are in [Fig. 3]. The device 10 comprises two recovery tubes 15 connecting the neck 143 to the bottom of two separate secondary reservoirs 42. Alternatively, the two tubes 15 could connect the neck 143 to two separate points in the same reservoir 42. This device 10 thus offers a large flexibility in choosing the points where stagnant water from the bottom of the tanks 42 is to be drawn. This multi-tube configuration 15 also allows for anticipating a possible blockage of one of the recovery tubes 15.

[0056] In addition, the recovery tube(s) 15 may include at their end corresponding to the point of water intake at the bottom of the tanks 42, a filter 16, allowing to limit the risk of insertion of impurities into the flow of fuel intended to be injected into the engine 1 via the device 10 and the supply line 51.

[0057] It is thus understood that when the activation valve 13 is closed, all the fuel passes through the secondary tube 14, and the vacuum at the throat 143 is then at its maximum. This vacuum, proportional to the consumption rate of the engine 1, causes the water present at the bottom of the secondary tank(s) 42 to be drawn via the recovery tube(s) 15, the water being further filtered by the filter 16. The water drawn is then mixed with the fuel flowing into the device 10 via the secondary conduit 14. The fuel / water mixture thus formed is then directed downstream to the engine 1 into which it is injected.

[0058] Conversely, when the activation valve 13 is open, a small proportion of the fuel passes through the secondary tube 14. The pressure difference generated between the two ends of the recovery tube 15 is then too small to draw the water present at the bottom of the secondary tank 42. On the contrary, the pressure at the neck 143 can be positive and a backflow of fuel can thus occur in the recovery tube 15, from the neck 143 towards the filter 16, which allows the filter 16 to be cleaned and limits the risk of icing of the latter.

[0059] The water sampling device 10 according to the invention is particularly advantageous in that it can be activated on demand when necessary at any time during the flight, for limited periods, without risking damage to the engine equipment, in particular without risking icing of the main filter of the engine 1.

[0060] For example, the activation valve 13 should not be closed when starting the engine 1, as the fuel and engine 1 temperatures are still low and the risk of water freezing is highest. Conversely, when a certain temperature threshold is reached, the activation valve 13 can be activated, either manually or automatically via a control unit (not shown), to close it. The fuel flow then passes entirely through the secondary conduit 14, which causes the water to be drawn through the recovery tubes 15. The water thus extracted can then be injected into the engine 1 without incurring the aforementioned risks.

[0061] Device 10 further reduces the impact on integration since it is a simple and compact mechanical part, without moving parts, and it integrates directly onto the supply line 51, without requiring the addition of dedicated drive pressure lines or drawing flow from the fuel pump 30, as pressure losses are low and the fuel flows through the venturi tube which itself pumps the water. Furthermore, as mentioned above, the device 10 is located downstream of the fuel pump 30, and therefore draws water downstream of it. Consequently, since the water does not pass through the fuel pump 30, this limits the risk of damage to its electric motor.

[0062] Furthermore, the device 10 provides a thermal expansion function via the secondary conduit 14 and the recovery tube 15, such thermal expansion being usually achieved by means of drillings on ball valves of the supply system. This thermal expansion helps to limit overpressures in the pipes.

[0063] Furthermore, placing the device 10 on the main channel 53 of the supply line 51, rather than on the sampling channel 52, is advantageous because it simplifies its sizing. Indeed, when the interconnecting valve 70 is open, the fuel pump 30 located in the first wing 40 must provide sufficient flow to supply both engines 1 and 2. However, if the device were placed on the sampling channel 52, the sizing of the secondary conduit 14, i.e., the venturi tube, would have to take this scenario into account to avoid restricting the flow supplied by the fuel pump 30 and thus avoid pressure losses. Placing the device 10 on the main channel 53 therefore simplifies its sizing and consequently its design.

[0064] It should be noted that the fuel supply system 50 may also include a fuel shut-off valve 80 (known by the acronym "FSOV" for "Fuel Shut-Off Valve" in English) located downstream of the device 10 and upstream of the engines 1, 2, allowing the engines 1, 2 to be decoupled from the rest of the system and the supply network when necessary, for example in the event of a fire in the engine 1, 2.

[0065] Although the present invention has been described with reference to specific embodiments, it is evident that modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, individual features of the various embodiments illustrated / mentioned can be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than a restrictive sense.

Claims

Demands

1. Device (10) for recovering water from the main aircraft fuel tank (41), the device (10) being adapted to be disposed on a fuel supply line (51) connecting a fuel pump (30) and an engine (1) of the aircraft to be supplied, the device (10) being disposed downstream of the fuel pump (30) and upstream of the engine (1), and comprising: - an inlet (10a) adapted to be connected to the fuel pump (30) and an outlet (10b) adapted to be connected to the engine (1), - a main conduit (12) on which is disposed an activation valve (13) movable between a closed position preventing fuel flow via the main conduit (12), and an open position allowing such flow, - a secondary conduit (14) parallel to the main conduit (12) and having a convergent (141), a divergent (142) and a throat (143),the device (10) further comprising at least one recovery tube (15) connecting an orifice (17) of the neck (143) of the secondary conduit (14) and the bottom of at least one secondary fuel tank (42) of the aircraft.

2. Device (10) according to claim 1, wherein the main conduit (12) is able to be arranged in continuity with the fuel supply line (51), the secondary conduit (14) being offset from the main conduit (12) and the supply line (51), and in fluidic communication with an upstream end (12a) and a downstream end (12b) of the main conduit (12).

3. Device (10) according to claim 1 or 2, wherein the main conduit (12) has a constant cross-section, the secondary conduit (14) being a venturi tube.

4. Device (10) according to any one of claims 1 to 3, wherein the activation valve (13) is a ball valve.

5. Device (10) according to any one of claims 1 to 4, wherein a longitudinal length of the convergent (141) is shorter than a longitudinal length of the divergent (142).

6. Device (10) according to any one of claims 1 to 5, wherein an angle (|3) between a longitudinal principal axis (X) of the conduit secondary (14) and a divergent wall (142) is between 2 and 4°, preferably between 3 and 3.5°.

7. Device (10) according to any one of claims 1 to 6, wherein the neck (143) of the secondary tube (14) comprises a plurality of orifices (17), a plurality of recovery tubes (15) being able to connect each orifice (17) of the neck (143) with the bottom of at least one secondary fuel tank (42) of the aircraft.

8. Aircraft fuel supply system (50), the system (50) comprising a main fuel tank (41), a fuel pump (30) disposed in said main tank (41), a supply line (51) capable of connecting the fuel pump (30) to an engine (1) of the aircraft (100), and a water recovery device (10) according to any one of the preceding claims, disposed on the supply line (51) between the fuel pump (30) and the engine (1).

9. Fuel supply system (50) according to claim 8, wherein said fuel supply system (50) is a first system suitable for being arranged in a first wing (40) of the aircraft (100) and for being connected to a second fuel supply system (50) arranged in a second wing (60) of the aircraft (100) via an interconnecting valve (70) arranged between the first wing (40) and the second wing (60) on an interconnecting channel (54), the supply line (51) comprising a sampling channel (52) connected to the fuel pump (30) and a main channel (53) on which the device (10) is disposed, the sampling channel (52) supplying both the main channel (53) and the interconnecting channel (54).

10. Aircraft (100) comprising at least one water recovery device (10) according to any one of claims 1 to 7.

Images