Inerting system for aircraft composite fuel tank

Abstract

The aircraft composite fuel tank inerting system comprises a gas source distribution pipeline; the central wing inerting branch is connected with the central wing oil tank; the central wing inerting branch is provided with a first inert gas generator and a multi-flow switching valve, and the multi-flow switching valve is located on the downstream side of the first inert gas generator and can be switched into a first flow mode and a second flow mode. The flow of inert gas provided in the first flow mode is smaller than that in the second flow mode, the outer wing inerting branch is provided with a second inert gas generator and a first flow limiting piece located on the downstream side of the second inert gas generator, and a single flow mode is achieved through the first flow limiting piece. Inert gas is injected into the center wing fuel tank such that the initial value of the oxygen concentration in the center wing fuel tank in the descent phase is equal to or less than a first upper limit threshold value, and inert gas is injected into the outer wing fuel tank such that the oxygen concentration in the outer wing fuel tank in the cruise phase is equal to or less than a second upper limit threshold value higher than the first upper limit threshold value.

Description

Technical Field

[0001] The present invention relates to an inerting system for a composite fuel tank of an aircraft, belonging to the technical field of fuel tank explosion protection. Background Art

[0002] For the purpose of weight reduction, the structural materials of fuel tanks of new-generation civil aircraft are mostly composite materials. The fuel tanks of such models do not belong to traditional non-heated aluminum wing fuel tanks. To meet the requirements of the FAR / CCAR25-R4 25.981 fuel tank explosion protection clause, inerting systems need to be installed in both the inboard and outboard fuel tanks of such models for reduced combustion design, that is, full fuel tank inerting is required. The larger the target fuel tank volume, the greater the flow rate of inert gas that the inerting system needs to provide, the larger the volume of the inerting system, the more air is bled, and the greater the weight. Especially in wide-body models, the need to reduce the weight of the inerting system, reduce system air bleeding, and improve the overall efficiency of the aircraft is becoming increasingly urgent.

[0003] A known inerting system for a wide-body aircraft is designed with a common main pipeline for distributing nitrogen-enriched air (NEA) to each fuel tank. Through the flow mode switching valves set in the main pipeline, the flow control of all fuel tanks is completed. The outer wing and the central wing both adopt a large-flow mode during the descent stage and a small-flow mode in other stages. The NEA demand in the large-flow mode determines the number of air separation modules (ASM). The existing architecture design of the inerting system results in a relatively large NEA demand, an increase in the number of core equipment ASMs, an increase in engine air bleeding, a large system weight, and a reduction in aircraft efficiency, etc. Summary of the Invention

[0004] The present invention is completed in view of the above problems, and the purpose is to provide an inerting system for a composite fuel tank of an aircraft that can reduce the weight of the inerting system, reduce system air bleeding, and improve the overall efficiency of the aircraft.

[0005] To achieve the above object, the present invention provides an aircraft composite fuel tank inerting system, comprising: a gas source distribution pipeline; a central wing inerting branch that bifurcates from the gas source distribution pipeline and is connected to the central wing fuel tank; and an outer wing inerting branch that bifurcates from the gas source distribution pipeline and is connected to the outer wing fuel tank. The central wing inerting branch is provided with a first inert gas generator and a multi-flow switching valve. The first inert gas generator receives the gas to be processed from the gas source and generates inert gas. The multi-flow switching valve is located on the downstream side of the first inert gas generator and can be switched to a first flow mode and a second flow mode. The flow rate of the inert gas provided in the first flow mode is less than the flow rate of the inert gas provided in the second flow mode. The outer wing inerting branch is provided with a second inert gas generator and a first flow restrictor. The second inert gas generator receives the gas to be processed from the gas source and generates inert gas. The first flow restrictor is located on the downstream side of the second inert gas generator, and a single flow mode is achieved through the first flow restrictor. The aircraft composite fuel tank inerting system is configured to inject inert gas into the central wing fuel tank in such a way that the initial value of the oxygen concentration in the central wing fuel tank during the descent phase is below a first upper threshold, and to inject inert gas into the outer wing fuel tank in such a way that the oxygen concentration in the outer wing fuel tank during the cruise phase is below a second upper threshold higher than the first upper threshold.

[0006] According to the aircraft composite fuel tank inerting system of the present invention, different fuel tanks of aircraft models that require full fuel tank inerting adopt a differential inert gas generation and inert gas distribution subsystem design. The central wing inerting branch leading to the central wing fuel tank is provided with a first inert gas generator and a multi-flow switching valve, and the outer wing inerting branch leading to the outer wing fuel tank is provided with a second inert gas generator and a first flow restrictor, forming a multi-branch inerting system design scheme.

[0007] Actually, according to the CCAR / FAR 25.981 composite fuel tank flammability analysis requirements, the outer wing fuel tank has no inerting requirement during climb and descent phases, and the injection of inert gas (such as NEA) into the outer wing fuel tank can be stopped during climb and descent phases. The existing inerting system architecture design does not support the differential inerting function for the outer wing and the central wing, resulting in a large demand for inert gas, an increase in engine bleed air, a large system weight, and a reduction in aircraft efficiency.

[0008] The present invention adopts a differential inert gas generation and inert gas distribution subsystem design for the central wing and the outer wing. Among them, a multi-flow mode is adopted for the central wing fuel tank, and a single flow mode is adopted for the outer wing fuel tank to achieve different oxygen concentration control targets, which can reduce the demand for inert gas in the outer wing fuel tank, reduce the overall weight of the inerting system, reduce system bleed air, and improve aircraft economy.

[0009] In addition, in the aircraft composite fuel tank inerting system of the present invention, it is also possible that the gas supply pipeline is provided with an air preparation subsystem, and the air preparation subsystem can adjust the pressure and temperature of the gas source. According to the above structure, the pressure and temperature of the inlet bleed air of the first inert gas generator and the second inert gas generator can be adjusted, which helps to improve the inerting performance.

[0010] In addition, in the aircraft composite fuel tank inerting system of the present invention, it is also possible that a pressure and temperature regulating device is provided at a position upstream of the first inert gas generator in the central wing inerting branch, and the pressure and temperature regulating device can adjust the pressure and temperature of the gas to be processed from the gas source. A temperature regulating device is provided at a position upstream of the second inert gas generator in the outer wing inerting branch, and the temperature regulating device can adjust the temperature of the gas to be processed from the gas source. According to the above structure, for the differentiated inerting requirements of the central wing fuel tank and the outer wing fuel tank, the central wing inerting branch adopts a pressure regulating design. The pressure regulating part works, for example, during the descent stage, for example, to increase the inlet bleed air pressure of the first inert gas generator and improve the inerting performance. The outer wing inerting branch does not require a pressure regulating design.

[0011] For example, the outer wing inerting branch bifurcates at a position downstream of the first flow limiting member to form a left outer wing inerting branch connected to the left outer wing fuel tank and a right outer wing inerting branch connected to the right outer wing fuel tank.

[0012] In addition, in the aircraft composite fuel tank inerting system of the present invention, it is preferred that two or more anti-backflow check valves are connected in series at a position downstream of the multi-flow switching valve in the central wing inerting branch. Thereby, it is possible to further reliably prevent the fuel from flowing back and contacting the upstream equipment. It is preferred that an anti-backflow check valve is provided at a position downstream of the first flow limiting member and upstream of the bifurcation position of the left outer wing inerting branch and the right outer wing inerting branch in the outer wing inerting branch, and anti-cross-fueling check valves are provided in the left outer wing inerting branch and the right outer wing inerting branch respectively. Thereby, it is possible to reliably prevent the fuel from flowing back and contacting the upstream equipment and prevent cross-fueling between the left and right outer wing fuel tanks.

[0013] In addition, in the aircraft composite fuel tank inerting system of the present invention, it is also possible that the central wing inerting branch is connected to a plurality of nozzles at the downstream end, and at least a part of the left outer wing inerting branch and the right outer wing inerting branch is located outside the central wing fuel tank. Thereby, the installation of equipment inside the fuel tank can be reduced.

[0014] In addition, in the aircraft composite fuel tank inerting system of the present invention, it is also possible that an ejector inlet air pipeline is further provided, and the central wing inerting branch is connected to an ejector pump at the downstream end, so that the gas in the central wing fuel tank is mixed with the inert gas entering the central wing fuel tank through the ejector inlet air pipeline and then filled into the central wing fuel tank. Thereby, the inerting uniformity inside the central wing fuel tank can be improved.

[0015] In addition, in the aircraft composite fuel tank inerting system of the present invention, it may also include a controller that can control the operation of the multi-flow switching valve.

[0016] In addition, in the aircraft composite fuel tank inerting system of the present invention, for example, the first upper limit threshold is 4% or less, and the second upper limit threshold is 14.5% or less.

[0017] In addition, in the aircraft composite fuel tank inerting system of the present invention, it may also be that the inert gas is nitrogen-rich gas, and the first inert gas generator and the second inert gas generator respectively use hollow fiber membrane type air separation modules. When the first inert gas generator and the second inert gas generator use hollow fiber membrane type air separation modules, the number of air separation modules can be reduced or the diameter of the hollow fiber membrane of the air separation module can be decreased, reducing the weight of the air separation module and the air intake volume. The first flow limiting member may be a flow limiting orifice. Description of the Drawings

[0018] Figure 1 Shows a schematic structural diagram of an aircraft composite fuel tank inerting system according to an embodiment of the present invention.

[0019] Figure 2 Shows a schematic structural diagram of an aircraft composite fuel tank inerting system according to another embodiment of the present invention.

[0020] Figure 3 Shows a schematic structural diagram of an aircraft composite fuel tank inerting system according to yet another embodiment of the present invention.

[0021] Symbol Description

[0022] 1 Air preparation subsystem

[0023] 1a Pressure and temperature regulating device

[0024] 1b Temperature regulating device

[0025] 2 First inert gas generator

[0026] 3 Second inert gas generator

[0027] 4 First flow limiting member

[0028] 5 Multi-flow switching valve

[0029] 6a, 6b, 6c Anti-backflow check valves

[0030] 7a, 7b Distribution flow limiting members

[0031] 8a, 8b Inert gas nozzles

[0032] 9 Left outer wing fuel tank

[0033] 10 Right outer wing fuel tank

[0034] 11 Center wing fuel tank

[0035] 12 Left ventilation fuel tank

[0036] 13 Right ventilation fuel tank

[0037] 15a Center wing fuel tank ventilation pipeline

[0038] 15b Left outer wing fuel tank ventilation pipeline

[0039] 15c Right outer wing fuel tank ventilation pipeline

[0040] 16 Ejector pump

[0041] 17 Ejector flow inlet pipeline

[0042] 18a, 18b Anti - cross - flow check valve

[0043] 19 Center wing inert gas distribution branch

[0044] L1 Gas source distribution pipeline

[0045] L2 Center wing inerting branch

[0046] L3 Outer wing inerting branch

[0047] L4 Left outer wing inerting branch

[0048] L5 Right outer wing inerting branch Specific embodiments

[0049] Hereinafter, with reference to the drawings, the technical solutions of the embodiments and variants of the present invention will be described. In addition, the scope of the present invention is not limited to the following embodiments and variants, and can be arbitrarily changed within the scope of the technical idea of the present invention. In addition, in the following drawings, in order to facilitate the understanding of each structure, sometimes the actual structure is different from the scale, quantity, etc. in each structure.

[0050] <First embodiment>

[0051] Figure 1 A schematic structural diagram of an aircraft composite material fuel tank inerting system according to the first embodiment of the present invention is shown. The aircraft composite material fuel tank inerting system includes a gas source distribution pipeline L1, a center wing inerting branch L2, and an outer wing inerting branch L3.

[0052] The gas source distribution pipeline L1 supplies the gas source for circulation. The gas source distribution pipeline L1 may be provided with an air preparation subsystem 1. The air preparation subsystem 1 can adjust the pressure and temperature of the gas source. The air preparation subsystem 1 can use any existing means of pressure and temperature regulation for air preparation, and the specific treatment will not be elaborated here. Additionally, the gas source can be, for example, engine bleed air from an engine, etc. The air preparation subsystem 1 can process the air with a relatively high temperature, such as engine bleed air, etc., for example, adjust the gas by cooling, pressurizing, etc., and then input the gas into the subsequent inert gas generator for gas separation, etc.

[0053] The central wing inerting branch L2 and the outer wing inerting branch L3 branch off from the gas source distribution pipeline L1 respectively. That is, one end of the central wing inerting branch L2 and one end of the outer wing inerting branch L3 are connected to the gas source distribution pipeline L1 in a manner that allows gas to flow through. In Figure 1 the example, the central wing inerting branch L2 and the outer wing inerting branch L3 branch off from the same position of the gas source distribution pipeline L1.

[0054] The other end (downstream end) of the central wing inerting branch L2 is connected to the central wing fuel tank 11 (the oil-free space). The central wing inerting branch L2 is provided with a first inert gas generator 2 and a multi-flow switching valve 5. The first inert gas generator 2 receives the gas to be processed from the gas source and generates nitrogen-enriched air (NEA) as an inert gas. In this embodiment, the first inert gas generator 2 uses a hollow fiber membrane type air separation module (ASM). Multiple first inert gas generators 2 can be arranged in parallel. In Figure 1 the example, two first inert gas generators 2 are arranged in parallel. In addition, one first inert gas generator 2 can also be provided.

[0055] The multi-flow switching valve 5 is located on the downstream side of the first inert gas generator 2 and can be switched to a first flow mode (small flow mode) and a second flow mode (large flow mode). The flow rate of the inert gas (in this embodiment, NEA) provided in the first flow mode is less than the flow rate of the inert gas (in this embodiment, NEA) provided in the second flow mode. In this embodiment, the multi-flow switching valve 5 is configured as a dual-flow switching valve.

[0056] Preferably, a check valve for preventing fuel backflow, i.e., a backflow prevention check valve, is provided at a position on the central wing inerting branch L2 downstream of the multi-flow switching valve 5. More preferably, two or more ( Figure 1 in the example, two) backflow prevention check valves 6b, 6c are provided in series connection, whereby the reliability of preventing fuel backflow can be further improved and single-point failure can be avoided.

[0057] In addition, the aircraft composite fuel tank inerting system of this embodiment may include an ejector flow inlet pipeline 17. The ejector flow inlet pipeline 17 is, for example, arranged in the center wing fuel tank 11. The center wing inerting branch L2 may be connected to the ejector pump 16 at the downstream end, so that the gas in the center wing fuel tank 11 is mixed with the inert gas (NEA in this embodiment) entering the center wing fuel tank 11 via the ejector flow inlet pipeline 17 and the center wing inerting branch L2, and then filled into the center wing fuel tank.

[0058] The other end (downstream end) of the outer wing inerting branch L3 is connected to the outer wing fuel tank (the oil-free space). The outer wing inerting branch L3 is provided with a second inert gas generator 3 and a first flow limiter 4. The second inert gas generator 3 receives the gas to be processed from the gas source and generates inert gas (NEA in this embodiment). In this embodiment, the second inert gas generator 3 adopts a hollow fiber membrane type air separation module (ASM).

[0059] The first flow limiter 4 is located on the downstream side of the second inert gas generator 3. In the outer wing inerting branch L3, a single flow mode is realized through the first flow limiter 4. That is, the first flow limiter 4 is installed on the outer wing inerting branch L3, and a single flow mode is adopted throughout the whole flight. The setting of the first flow limiter helps to reduce the flow rate of the inert gas (NEA in this embodiment) and prevent the fuel tank from overpressure caused by excessive gas. The first flow limiter is, for example, a flow limiting device with a fixed aperture such as a flow limiting hole (single flow mechanical flow limiting hole).

[0060] In addition, the outer wing inerting branch L3 bifurcates at a position downstream of the first flow limiter 4 to form a left outer wing inerting branch L4 connected to the left outer wing fuel tank 9 and a right outer wing inerting branch L5 connected to the right outer wing fuel tank 10. That is, one end of the left outer wing inerting branch L4 and one end of the right outer wing inerting branch L5 are respectively connected to the outer wing inerting branch L3 in a manner that allows gas to flow through. The outer wing inerting branch L3 is connected to the left outer wing fuel tank 9 (the oil-free space) via the left outer wing inerting branch L4 and to the right outer wing fuel tank 10 (the oil-free space) via the right outer wing inerting branch L5. In Figure 1 the example, the left outer wing inerting branch L4 and the right outer wing inerting branch L5 bifurcate from the same position of the outer wing inerting branch L3.

[0061] In addition, an inert gas nozzle 8a may be connected to the other end of the left outer wing inerting branch L4. An inert gas nozzle 8b may be connected to the other end of the right outer wing inerting branch L5.

[0062] Preferably, a check valve 6a for preventing backflow is provided at a position on the outer wing inerting branch L3 downstream of the first flow-limiting member 4 and upstream of the bifurcation points of the left outer wing inerting branch L4 and the right outer wing inerting branch L5. Preferably, a check valve for preventing oil mixing between the left and right outer wing fuel tanks, i.e., an anti-oil-mixing check valve 18a, is provided in the left outer wing inerting branch L4. Preferably, an anti-oil-mixing check valve 18b is provided in the right outer wing inerting branch L5.

[0063] In addition, a distribution flow-limiting member 7a (such as a flow-limiting orifice) may be provided in the left outer wing inerting branch L4, and a distribution flow-limiting member 7b (such as a flow-limiting orifice) may be provided in the right outer wing inerting branch L5. The distribution flow-limiting member 7a may be provided downstream of the anti-oil-mixing check valve 18a or upstream of the anti-oil-mixing check valve 18a. The distribution flow-limiting member 7b may be provided downstream of the anti-oil-mixing check valve 18b or upstream of the anti-oil-mixing check valve 18b.

[0064] In addition, it may further include a controller (not shown), which can communicate with the multi-flow switching valve 5 and the like by wired or wireless means to control the operation of the aircraft composite fuel tank inerting system. For example, the controller can control the operation of the multi-flow switching valve 5 according to the flight phase of the aircraft and the like.

[0065] For example, through the flammability analysis of the composite fuel tanks of wide-body aircraft models, it is recognized that the flammability requirement of the outer wing composite fuel tank is lower than that of the center wing fuel tank, and it is not necessary to meet the requirement that the flammability is lower than 3% during the ground / climb phase in warm weather (when the temperature is greater than 26.7°C). The design of the outer wing fuel tank fuel reduction system can be different from that of the center wing fuel tank. The stage with the largest contribution to the basic flammability of the outer wing is the cruise stage. Approximately 90% of the combustible time in the whole profile is in the cruise stage. Therefore, it is only necessary to ensure the inerting of the outer wing fuel tank during the cruise stage, and no requirements are made for the other stages, so as to ensure that the flammability of the outer wing fuel tank is lower than 3%. That is, it is not necessary to impose the same restrictions on the oxygen concentration in the outer wing fuel tank during the descent stage as on the center wing fuel tank.

[0066] In view of the above situation, the present inventor designs a differential air separation and inert gas distribution subsystem for the center wing and the outer wing to achieve global optimization, improve the economic efficiency of the whole aircraft, reduce the demand for inert gas in the outer wing fuel tank, reduce the requirements for the system (such as reducing the number of ASMs or reducing the size of ASMs, reducing engine bleed air), reduce the overall weight of the inerting system, and improve the economy of the aircraft.

[0067] The inerting system for the composite fuel tank of the aircraft according to the present invention is configured to inject inert gas into the center wing fuel tank 11 so that the initial value of the oxygen concentration in the center wing fuel tank 11 during the descent phase is below the first upper limit threshold, and to inject inert gas into the outer wing fuel tanks 9 and 10 so that the oxygen concentration in the outer wing fuel tanks 9 and 10 during the cruise phase is below the second upper limit threshold, where the second upper limit threshold is higher than the first upper limit threshold. Here, the "oxygen concentration" refers to the oxygen concentration in the oil-free space inside the fuel tank (also referred to as the "fuel tank"). For example, the first upper limit threshold is 4% or less, and the second upper limit threshold is 14.5% or less.

[0068] In addition, a ventilation system for connecting the aircraft fuel tank to the outside can be realized by the left ventilation fuel tank 12 and the right ventilation fuel tank 13 located near the wing tip. A center wing fuel tank ventilation pipeline 15a can be provided on the wing to connect the center wing fuel tank 11 and the right ventilation fuel tank 13. In addition, a left outer wing fuel tank ventilation pipeline 15b can be provided to connect the left outer wing fuel tank 9 and the left ventilation fuel tank 12; a right outer wing fuel tank ventilation pipeline 15c can be provided to connect the right outer wing fuel tank 10 and the right ventilation fuel tank 13. Multiple ( Figure 1 in the example, two) left outer wing fuel tank ventilation pipelines 15b can be respectively connected to different positions of the left outer wing fuel tank 9, and multiple ( Figure 1 in the example, two) right outer wing fuel tank ventilation pipelines 15c can be respectively connected to different positions of the right outer wing fuel tank 10

[0069] In addition, pressure sensors, temperature sensors, etc. (not shown) can be installed in the center wing fuel tank 11, the outer wing fuel tanks 9 and 10, and other components such as cut-off valves can be installed in several pipelines / branches.

[0070] The operation process of the above-mentioned inerting system for the composite fuel tank of the aircraft will be described below.

[0071] The bleed air (air source) of the air source system enters the first inert gas generator 2 and the second inert gas generator 3 through the air preparation subsystem 1;

[0072] The inert gas (in this embodiment, nitrogen-rich gas) generated by the first inert gas generator 2 enters the center wing fuel tank 11 through the multi-flow switching valve 5 and two check valves 6b and 6c to prevent backflow;

[0073] The inert gas pipeline entering the center wing fuel tank 11 (the downstream end part of the center wing inerting branch L2) is connected to the ejector pump 16, so that the gas in the center wing fuel tank 11 is mixed with the inert gas (in this embodiment, nitrogen-rich gas) entering the center wing fuel tank through the ejector inlet pipeline 17 and then recharged into the center wing fuel tank 11, improving the inerting uniformity in the center wing fuel tank 11;

[0074] The inert gas (in this embodiment, nitrogen-rich gas) generated by the second inert gas generator 3 enters the fuel tank through a single-flow mechanical flow-limiting orifice (the first flow-limiting member 4) and a non-return check valve 6a and is divided into two streams: the left side (refer to Figure 1 ) enters the left outer wing fuel tank 9 through a non-cross-flow check valve 18a, a distribution flow-limiting member 7a, and an inert gas nozzle 8a; the right side (refer to Figure 1 ) enters the right outer wing fuel tank 10 through a non-cross-flow check valve 18b, a distribution flow-limiting member 7b, and an inert gas nozzle 8b;

[0075] During the climb and cruise phases of the aircraft, the multi-flow switching valve 5 is controlled in the first flow mode (small flow mode). During the descent phase, the multi-flow switching valve 5 switches to the second flow mode (large flow mode). When the aircraft descends to a certain altitude (e.g., 4000 ft), the multi-flow switching valve 5 switches to the small flow mode until the system shuts down after the aircraft lands.

[0076] In the inerting system for the composite material fuel tank of the aircraft of the present invention, there are two branches for inert gas generation and inert gas distribution, injecting inert gas (nitrogen-rich gas, NEA) into the fuel tank. One branch injects into the central wing fuel tank 11, and the other branch injects into the outer wing fuel tanks 9 and 10. The two branches correspond to different inert gas generators;

[0077] The inert gas distribution branch corresponding to the central wing fuel tank 11 (i.e., the central wing inerting branch L2 leading to the central wing fuel tank 11) adopts a multi-flow mode, which is realized by a dual-flow switching valve. For example, during the descent phase, a large flow mode is adopted for the central wing fuel tank 11 to reduce the entry of outside air into the fuel tank to meet the requirements for controlling the oxygen concentration in the central wing fuel tank 11 when landing. For wide-body aircraft, the oxygen concentration after flight is lower than 11%, and for narrow-body aircraft, it is lower than 12%. During the cruise phase, a small flow mode is adopted to reduce the oxygen concentration in the fuel tank to the lowest level, such as 2% - 4%, providing an ideal initial value for the descent phase;

[0078] The single-flow mode is adopted for the outer wing fuel tanks, which is realized by a flow-limiting device with a fixed aperture. The inert gas (in this embodiment, nitrogen-rich gas) is evenly distributed to the left and right outer wing fuel tanks through the distribution branch (the outer wing inerting branch L3 and the left outer wing inerting branch L4 and the right outer wing inerting branch L5 bifurcated from the outer wing inerting branch L3). It only needs to maintain the oxygen concentration in the fuel tank lower than the limit value during the cruise phase, such as 12% - 14.5%.

[0079] According to the aircraft composite fuel tank inerting system of the present invention, by taking advantage of the characteristic that the outer wing fuel tank only needs to be inerted during the cruise phase to achieve the design goal of reducing flammability, the outer wing inerting branch can meet the requirements by using low-purity inert gas (in this embodiment, NEA). When the flow rate of the inert gas (in this embodiment, NEA) remains unchanged, the demand for the inert gas (in this embodiment, NEA) by the outer wing fuel tank can be reduced. In particular, when the first inert gas generator 2 and the second inert gas generator 3 use the hollow fiber membrane type ASM, the number of ASMs can be reduced or the diameter of the ASM (hollow fiber membrane) can be decreased, reducing the weight of the ASM and the engine bleed air volume (lower purity, improving nitrogen production efficiency, and reducing the engine bleed air volume).

[0080] In one example, after adopting the aircraft composite fuel tank inerting system of this embodiment, the weight of the ASM is reduced by 26.6%, and the engine bleed air volume is reduced by 13.3%.

[0081] <Second Embodiment>

[0082] Figure 2 The schematic structural diagram of the aircraft composite fuel tank inerting system according to the second embodiment of the present invention is shown. Hereinafter, the differences from the above first embodiment will be mainly described. Sometimes, the same reference numerals are used for the constituent elements that are the same as those in the above first embodiment, and repeated descriptions are omitted. Unless there are contrary descriptions or conflicts with other technical features in the following content, the features described in the first embodiment also apply to this second embodiment, and no detailed description will be given here.

[0083] In this embodiment, a pressure and temperature regulating device 1a is provided at a position upstream of the first inert gas generator 2 in the central wing inerting branch L2. The pressure and temperature regulating device 1a can regulate the pressure (for example, boosting) and temperature (for example, cooling) of the gas to be processed from the gas source. The pressure and temperature regulating device 1a includes, for example, a pressure regulating part and a temperature regulating part. In addition, one first inert gas generator 2 is provided in the central wing inerting branch L2.

[0084] A temperature regulating device 1b is provided at a position upstream of the second inert gas generator 3 in the outer wing inerting branch L3. The temperature regulating device 1b can regulate the temperature (for example, cooling) of the gas to be processed from the gas source.

[0085] In this embodiment, the first inert gas generator 2 and the second inert gas generator 3 respectively adopt hollow fiber membrane type air separation modules (ASMs).

[0086] The bleed air pressure regulation design, specifically the bleed air pressurization design, can adopt, for example, electric supercharging or turbocharging, which is used to improve the ASM separation performance and reduce the number of ASMs during the descent phase. In this embodiment, for the differentiated inerting requirements of the center wing tank and the outer wing tank, the center wing inerting branch L2 adopts a pressure regulation design, specifically a pressurization design, while the outer wing inerting branch L3 does not require a pressure regulation design. Thus, the main difference between this embodiment and the first embodiment is that independent upstream bleed air regulation subsystems are respectively provided for the first inert gas generator 2 and the second inert gas generator 3. For the first inert gas generator 2, its bleed air regulation subsystem has functions of temperature regulation and pressure regulation. The pressure regulation part works during the descent phase to increase the inlet bleed air pressure of the first inert gas generator 2, improve the separation performance, and reduce the number of ASM components in the first inert gas generator 2. For the second inert gas generator 3, its bleed air regulation subsystem does not require a pressurization function and only needs to have a temperature regulation function. In addition, for example, since the working temperature of the air separation module is lower than the gas source temperature, the temperature regulation part (temperature regulation device) works throughout the flight.

[0087] <Third Embodiment>

[0088] Figure 3 Fig. shows a schematic structural diagram of the aircraft composite material fuel tank inerting system according to the third embodiment of the present invention. Hereinafter, the differences from the above-mentioned first embodiment will be mainly described. Sometimes, the same reference numerals are used for the same constituent elements as those in the above-mentioned first embodiment, and repeated descriptions are omitted. Unless there are contrary descriptions or conflicts with other technical features in the following content, the features described in the first embodiment are also applicable to this third embodiment, and no detailed description will be made here.

[0089] In this embodiment, the downstream end of the center wing inerting branch L2 is connected to a plurality of nozzles 16a, 16b, 16c. The downstream end of the center wing inerting branch L2 can be respectively connected to the plurality of nozzles via a multi-branched center wing inert gas distribution branch 19.

[0090] In addition, at least a part of the left outer wing inerting branch L4 and the right outer wing inerting branch L6 is located outside the center wing tank 11.

[0091] There are already a large number of fuel systems and hydraulic system pipe networks in the wing tank, and the installation space in the tank is limited. In the first embodiment, a part of the inert gas distribution pipeline (including the left outer wing inerting branch L4 and the right outer wing inerting branch L5 branched from the outer wing inerting branch L3) is concentrated in the center wing tank 11 (refer to Figure 1), there is a great difficulty in installation and layout. To address this problem, this embodiment is designed. The main differences between this embodiment and the first embodiment are as follows: The configuration of the internal and external distribution pipelines of the fuel tank has been changed. On the one hand, instead of the ejector pump design at the downstream end of the central wing inerting branch L2 in the first embodiment, a multi-nozzle design is adopted, reducing the installation of equipment inside the fuel tank. On the other hand, a part of the pipeline network of the external wing fuel tank inert gas distribution branch (the external wing inerting branch L3 and the left external wing inerting branch L4 and the right external wing inerting branch L5 bifurcated from the external wing inerting branch L3) located in the central wing fuel tank 11 is transferred outside the fuel tank, reducing the installation of the pipeline network inside the fuel tank.

[0092] The present invention has been described exemplarily above in conjunction with the accompanying drawings. Obviously, the specific implementation of the present invention is not limited by the above embodiments.

[0093] For example, in the above embodiments, the first inert gas generator and the second inert gas generator respectively use hollow fiber membrane type air separation modules, but not limited thereto. The first inert gas generator and the second inert gas generator may also use other air separation modules or other inert gas generation means such as pressure swing adsorption (PSA). The inert gas can be nitrogen, nitrogen-rich gas, carbon dioxide, etc.

[0094] The specific embodiments of the present invention have been specifically described above in conjunction with the accompanying drawings. However, it can be understood that the above description does not limit the present invention in any form. The technical features in each embodiment can be combined with each other in any way to form new embodiments. In addition, after understanding the above specific embodiments, those skilled in the art can make various other modifications and changes to the present invention as needed. These do not deviate from the essence of the present invention.

Claims

1. An aircraft composite fuel tank inerting system, characterized in that: have: Gas source distribution pipeline; a center wing inerting branch, the center wing inerting branch branching from the gas source distribution pipeline and connected to the center wing oil tank; and an outer wing inerting branch, the outer wing inerting branch is branched from the gas source distribution pipeline and connected to the outer wing oil tank, The central wing inerting branch is provided with a first inert gas generator and a multi-flow switching valve. The first inert gas generator receives the gas to be treated from the gas source and generates the inert gas. The multi-flow switching valve is located at the downstream side of the first inert gas generator and can be switched to a first flow mode and a second flow mode. The flow rate of the inert gas provided in the first flow mode is less than the flow rate of the inert gas provided in the second flow mode. The outer wing inerting branch is provided with a second inert gas generator and a first flow limiter. The second inert gas generator receives the gas to be treated from the gas source and generates the inert gas. The first flow limiter is located at the downstream side of the second inert gas generator. The single flow mode is realized by the first flow limiter. The aircraft composite fuel tank inerting system is configured to inject inert gas into the center wing tank so that an initial value of oxygen concentration in the center wing tank during a descent phase becomes equal to or less than a first upper limit threshold, and to inject inert gas into the outer wing tanks so that an oxygen concentration in the outer wing tanks during a cruise phase becomes equal to or less than a second upper limit threshold that is higher than the first upper limit threshold.

2. The aircraft composite fuel tank inerting system according to claim 1, characterized in that: The gas source distribution pipeline is provided with an air preparation subsystem, and the air preparation subsystem can adjust the pressure and temperature of the gas source.

3. The aircraft composite fuel tank inerting system according to claim 1, characterized in that: A pressure and temperature regulating device is provided at a position upstream of the first inert gas generator in the central wing inerting branch, and the pressure and temperature regulating device can regulate the pressure and temperature of the gas to be treated from the gas source. A temperature control device is provided at a position upstream of the second inert gas generator in the outer wing inerting branch, and the temperature control device is capable of adjusting the temperature of the gas to be treated from the gas source.

4. The aircraft composite fuel tank inerting system according to claim 1, characterized in that: The outer wing inerting branch is bifurcated into a left outer wing inerting branch connected to the left outer wing oil tank and a right outer wing inerting branch connected to the right outer wing oil tank at a position downstream of the first flow restrictor.

5. The aircraft composite fuel tank inerting system according to claim 4, characterized in that: Two or more anti-backflow check valves connected in series are provided at a position downstream of the multi-flow switching valve in the center wing inerting branch. A backflow prevention one-way valve is provided in the outer wing inerting branch at a position downstream of the first flow limiting member and upstream of the bifurcation of the left outer wing inerting branch and the right outer wing inerting branch, and an oil cross-talk prevention one-way valve is provided in the left outer wing inerting branch and the right outer wing inerting branch, respectively.

6. The aircraft composite fuel tank inerting system according to claim 4, characterized in that: The central wing inerting branch is connected to a plurality of nozzles at the downstream end. At least a portion of the left outer wing inerting branch and the right outer wing inerting branch are located outside the center wing fuel tank.

7. The aircraft composite fuel tank inerting system according to claim 1, characterized in that: It also has a jet flow intake pipe. The center wing inerting branch is connected to the ejector pump at the downstream end, so that the gas in the center wing fuel tank is mixed with the inert gas entering the center wing fuel tank through the ejector flow intake pipeline and then filled into the center wing fuel tank.

8. The aircraft composite fuel tank inerting system according to claim 1, characterized in that: It also includes a controller capable of controlling the operation of the multi-flow switching valve.

9. The aircraft composite fuel tank inerting system according to claim 1, characterized in that: The first upper limit threshold is below 4%, The second upper limit threshold is less than 14.5%.

10. The aircraft composite fuel tank inerting system according to claim 1, characterized in that: The inert gas is nitrogen-rich gas, The first inert gas generator and the second inert gas generator respectively use hollow fiber membrane air separation modules. The first flow limiting member is a flow limiting hole.

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