A cooling system and method for aircraft wing fuel tanks

Abstract

This invention relates to a heat dissipation system and method for an aircraft wing fuel tank, belonging to the field of aircraft system design technology. It enables the cooling of fuel in the aircraft wing fuel tank, reducing the fuel temperature and increasing the stored cold energy in the fuel. The heat dissipation system includes a wing fuel tank, fuel lines, a skin heat exchanger, an electrically driven pump, a level gauge, a first switching valve, a second switching valve, and a controller. The skin heat exchanger is disposed on the lower skin of the wing and is fused to the lower skin. The wing fuel tank is connected to the skin heat exchanger via the fuel lines to form a loop. Following the fuel flow direction, the wing fuel tank, the first switching valve, the skin heat exchanger, the electrically driven pump, and the second switching valve are arranged sequentially. The level gauge is disposed inside the wing fuel tank. The first switching valve, the second switching valve, the electrically driven pump, and the level gauge are all electrically connected to the controller.

Description

Technical Field

[0001] This invention relates to the field of aircraft system design technology, and in particular to an aircraft wing fuel tank heat dissipation system and method. Background Technology

[0002] With the increasing electrification of civil aircraft, the heat output of onboard equipment is growing, leading to a greater demand for heat dissipation. The cooling function in civil aircraft is primarily handled by the air conditioning subsystem within the environmental control system. This subsystem introduces ram air through an opening in the fuselage, allowing it to transfer heat in the cold flow channels of the heat exchanger and ultimately providing cooling for the aircraft. However, relying solely on ram air for cooling the onboard equipment of future electrified civil aircraft would inevitably increase reliance on ram air and result in greater fuel compensation losses. Therefore, it is necessary to design new cooling methods or find usable heat sinks on the aircraft. Fuel itself has a limited cooling capacity, and most fuel tanks in civil aircraft are located in the wings. During flight, the wings exchange heat with the cold air at high altitudes. Fully utilizing the cooling capacity generated during this heat exchange process can significantly improve the aircraft's cooling capacity. This invention is designed with this in mind.

[0003] In the aviation field, especially in military aircraft, there are several technical solutions or structures that utilize skin heat exchangers to dissipate fuel heat. For example, patent CN108100273A provides a fuel cooling solution that uses the "structural heat dissipation effect" of the aircraft wing skin itself to cool the fuel in the wing fuel tank. To utilize the heat dissipation effect of the wing skin, at least one nozzle is installed to spray fuel from the fuel tank onto the upper surface of the wing fuel tank for cooling. This patent uses the wing skin for heat dissipation but does not specifically use a corresponding skin heat exchanger. The aircraft skin will be exposed to solar radiation, causing its temperature to rise and affecting the skin heat exchange effect. Furthermore, in the civil aircraft field, to prevent excessively high fuel temperatures in the fuel tanks and the increased risk of fuel explosion, fuel cooling in the fuel tanks utilizes the cooling capacity of the air conditioning system. For example, in the solution of patent CN 105644794 A, an air circulation cooling system or a liquid cooling system is used to cool the fuel in the center wing fuel tank; however, this solution does not affect the cooling capacity of the air conditioning system.

[0004] Therefore, it is necessary to study an aircraft wing fuel tank heat dissipation system and method using skin heat exchangers to address the shortcomings of existing technologies and solve or mitigate one or more of the aforementioned problems. Summary of the Invention

[0005] In view of this, the present invention provides a cooling system and method for an aircraft wing fuel tank using a skin heat exchanger, which can realize the function of cooling fuel in the aircraft wing fuel tank, reduce the fuel temperature in the fuel tank, and increase the cold energy stored in the fuel.

[0006] On one hand, the present invention provides an aircraft wing fuel tank heat dissipation system, the heat dissipation system including a wing fuel tank, fuel lines, skin heat exchanger, electric drive pump, level gauge, first switching valve, second switching valve and controller;

[0007] The skin heat exchanger is disposed on the lower skin of the wing, and the skin heat exchanger is fused to the lower skin;

[0008] The wing fuel tank is connected to the skin heat exchanger through the fuel line to form a loop; following the flow direction of the fuel in the loop, the loop is sequentially provided with the wing fuel tank, the first switching valve, the skin heat exchanger, the electric drive pump, and the second switching valve;

[0009] The level gauge is installed inside the wing fuel tank;

[0010] The first switching valve, the second switching valve, the electric drive pump, and the level gauge are all electrically connected to the controller.

[0011] In addition to the aspects and any possible implementations described above, a further implementation is provided in which the fusion connection is specifically a connection in which the lower skin surface of the skin heat exchanger and the lower skin at the corresponding position are the same skin.

[0012] In addition to the aspects and any possible implementations described above, a further implementation is provided in which the skin heat exchanger is a counter-flow multi-channel gas-liquid heat exchanger; the heat exchange fins of the skin heat exchanger are tightly fitted to the lower skin.

[0013] In addition to the aspects and any possible implementations described above, a further implementation is provided in which a first temperature sensor is provided at the lower skin and a second temperature sensor is provided inside the wing fuel tank; both the first temperature sensor and the second temperature sensor are electrically connected to the controller.

[0014] In addition to the aspects and any possible implementations described above, a further implementation is provided in which a third temperature sensor is provided on the fuel line downstream of the skin heat exchanger, the third temperature sensor being electrically connected to the controller.

[0015] In addition to the aspects described above and any possible implementations, a further implementation is provided in which a flow meter and a pressure sensor are provided on the fuel line, both of which are electrically connected to the controller.

[0016] In addition to the aspects and any possible implementations described above, a further implementation is provided in which a check valve is also provided in the circuit, and the check valve, the second switching valve, and the inlet of the wing fuel tank are arranged sequentially.

[0017] In addition to the aspects and any possible implementations described above, a further implementation is provided in which the heat dissipation system further includes a partition disposed in the internal space of the aircraft wing and dividing the internal space into upper and lower chambers; the wing fuel tank is disposed in the upper chamber of the two chambers, and the skin heat exchanger is disposed in the lower chamber of the two chambers.

[0018] On the other hand, the present invention provides a method for heat dissipation of an aircraft wing fuel tank, wherein the method is implemented using any of the heat dissipation systems described above;

[0019] The steps of the method include:

[0020] S1. Collect the aircraft's flight status signal to determine whether the aircraft is in flight; if yes, proceed to the next step; otherwise, continue collecting and judging.

[0021] S2. Determine whether the heat dissipation system meets the heat dissipation conditions. If yes, proceed to the next step; otherwise, continue to determine.

[0022] The heat dissipation conditions include: the real-time fuel level in the wing fuel tank is higher than 1 / 3 of the full fuel level;

[0023] S3. The controller controls the opening of the first and second switching valves and controls the electric drive pump to start for fuel cooling.

[0024] S4. Monitor the operating parameters of the heat dissipation system to determine whether the conditions for stopping heat dissipation are met; if yes, proceed to the next step; otherwise, continue heat dissipation while continuing to monitor and determine the conditions.

[0025] The conditions for stopping heat dissipation are that the real-time fuel level in the wing fuel tank is not higher than 1 / 3 of the full fuel level or the aircraft is on the ground.

[0026] S5. The controller shuts down the electric drive pump and closes the first and second switching valves.

[0027] In addition to the aspects and any possible implementations described above, a further implementation is provided in which the heat dissipation conditions in step S2 further include: the real-time temperature of the fuel in the wing fuel tank is higher than the real-time temperature of the lower skin.

[0028] The cooling stop condition in step S4 is that the real-time temperature of the fuel after cooling by the skin radiator is lower than the real-time temperature of the fuel in the wing fuel tank, and the difference is greater than a preset threshold.

[0029] Compared with the prior art, one of the above technical solutions has the following advantages or beneficial effects: The solution of the present invention realizes the function of cooling the fuel in the wing fuel tank by setting a skin heat exchanger integrated with the lower wing skin in the wing of a civil aircraft. It has high heat dissipation efficiency, convenient and reasonable structural setting, and safe and stable heat dissipation process.

[0030] Of course, any product implementing this invention does not necessarily need to achieve all of the technical effects described above at the same time. Attached Figure Description

[0031] To more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0032] Figure 1 This is a schematic diagram of a civil aircraft wing fuel tank heat dissipation system using a skin heat exchanger, provided in one embodiment of the present invention.

[0033] Figure 2 This is a top view of the flow channel structure of a skin heat exchanger provided in one embodiment of the present invention;

[0034] Figure 3 This is a flowchart of the wing fuel tank cooling system provided in one embodiment of the present invention.

[0035] In the figure:

[0036] 1. Controller; 2. First temperature sensor; 3. Flow meter; 4. Skin heat exchanger; 5. Third temperature sensor; 6. Electric drive pump; 7. Lower skin; 8. Upper skin; 9. Pressure sensor; 10. Wing fuel tank; 11. Second temperature sensor; 12. Fuel line; 13. Baffle; 14. Level gauge; 15. Check valve; 16. First switching valve; 17. Second switching valve. Detailed Implementation

[0037] To better understand the technical solution of the present invention, the embodiments of the present invention will be described in detail below with reference to the accompanying drawings.

[0038] It should be understood that the described embodiments are merely some, not all, of the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without inventive effort are within the scope of protection of the present invention.

[0039] To address the shortcomings of existing technologies, this invention provides a cooling system for the wing fuel tanks of a civil aircraft using a skin heat exchanger. The structure of the cooling system of this invention is as follows: Figure 1 As shown, it includes a flow meter 3, temperature sensors (first temperature sensor 2, second temperature sensor 11, third temperature sensor 5), pressure sensor 9, electric drive pump 6, skin heat exchanger 4, wing fuel tank 10, level gauge 14, controller 1 for wing fuel tank cooling system, check valve 15, first switching valve 16, second switching valve 17, fuel line 12, and baffle 13.

[0040] A partition 13 is installed inside the wing, dividing the wing's internal space into upper and lower chambers. The wing fuel tank 10 is located in the upper chamber and contains fuel that needs to be cooled. One end of the wing fuel tank 10 has a first cooling vent, and the other end has a second cooling vent. The first and second cooling vents are connected to the inlet and outlet of the skin heat exchanger 4, respectively, via fuel lines. A first switching valve 16 and a flow meter 3 are sequentially installed on the fuel line between the first cooling vent and the inlet of the skin heat exchanger 4. A third temperature sensor 5, an electric pump 6, a pressure sensor 9, a check valve 15, and a second switching valve 17 are sequentially installed on the fuel line between the skin heat exchanger outlet and the second cooling vent. The first switching valve 16, the second switching valve 17, the flow meter 3, the third temperature sensor 5, the electric pump 6, the pressure sensor 9, and the check valve 15 are all electrically connected to the controller 1.

[0041] When the fuel circuit of the entire heat dissipation structure is working, the working force of the electric pump 6 draws fuel out of the wing fuel tank 10 and flows sequentially through the first switching valve 16 at the beginning of the heat dissipation channel, the flow meter 3, the skin heat exchanger 4, the third temperature sensor 5, the electric pump 6, the pressure sensor 9, the check valve 15, and the second switching valve 17 at the end of the heat dissipation channel before entering the wing fuel tank 10. The fuel dissipates heat during this cycle. The electric pump 6 can be located at any point in the fuel circuit. The reason for choosing to place it after the skin heat exchanger 4 is that, in this invention, the wing fuel tank 10 and the skin heat exchanger 4 are arranged vertically. Since the cooled fuel requires greater power to flow upwards, placing the electric pump 6 at the bottom of the upward-flowing pipeline better ensures the circulation of the fuel.

[0042] The skin heat exchanger of this invention is integrally installed at the bottom of the fuel tank, and is fused with the lower surface skin of the wing (i.e., the lower skin 7). Specifically, the skin heat exchanger is rectangular in shape, with the bottom surface of the rectangular prism being the lower skin 7, thus achieving fusion between the skin heat exchanger and the lower skin. This structural arrangement allows the lower skin 7 to serve as the heat dissipation surface of the skin heat exchanger, enabling energy exchange with the external environment. It is important to note that the area occupied by the skin heat exchanger is much smaller than the entire lower wing skin; that is, only a portion of the lower skin is used, which can be referred to as the design section, serving as the heat dissipation surface of the skin heat exchanger. The upper part of the skin heat exchanger is physically isolated from the fuel tank by a partition 13, i.e., the aforementioned separation creates two chambers. The reason why the lower skin was chosen as part of the skin heat exchanger instead of the upper skin 8 is that the upper surface skin of the wing is located on the upper surface of the aircraft and is in direct contact with sunlight, resulting in solar heat radiation that affects the heat exchange efficiency. In addition, the upper skin 8 is not flat, but has different curvatures at different positions. When integrated with the skin heat exchanger, the geometric curvature of the skin heat exchanger needs to be consistent with that of the upper skin 8 of the wing, which results in higher design, manufacturing and installation costs. The lower skin with a planar structure is more practical.

[0043] In one specific embodiment of the present invention, the skin heat exchanger 4 is a gas / liquid heat exchanger, specifically a counter-flow single-layer multi-channel gas-liquid heat exchanger, wherein the liquid flow path is fuel, and the gas is air outside the lower surface of the wing. The fuel flow direction is opposite to the flight direction. During aircraft flight, the outside cold air and fuel exchange heat through the fins of the skin heat exchanger, and the fuel temperature decreases after the heat exchange. Figure 2 The diagram shows a top view of the simplified flow channel structure of the skin heat exchanger in this embodiment. The heat exchanger fins of the skin heat exchanger 4 are tightly connected to the lower skin 7. Multi-layer flow channel heat exchangers are not recommended in this invention because their heat exchange efficiency will decrease. The fuel flow direction can also be the same as the flight direction; however, this embodiment only describes a flow direction opposite to the flight direction because counter-flow has higher heat exchange efficiency. The inlet and outlet of the skin heat exchanger are rigidly connected to the fuel line 12, which should be a rigid pipe commonly used for fuel transfer in conventional aircraft.

[0044] In one specific embodiment of the present invention, the wing fuel tank 10 is provided with a level gauge 14 for measuring the fuel level. The level gauge 14 is electrically connected to the controller 1, which receives the fuel level data measured by the level gauge and uses it as the basis for controlling the operation of the electrically driven pump 6, the first switching valve 16, and the second switching valve 17.

[0045] The sensor devices in the fuel circuit are fixedly connected to the fuel line 12 (structure is not limited). The second temperature sensor 11 is installed in the wing fuel tank to measure the fuel temperature T2 in the tank. The third temperature sensor 5 is installed in the fuel circuit after the skin heat exchanger 4 to measure the fuel temperature T3 in the circuit after the skin heat exchanger has dissipated heat. The first temperature sensor 2 is installed on the lower surface of the wing to measure the skin temperature T1 of the lower wing surface. The probe of the first temperature sensor 2 is preferably in direct contact with the outside air. All three temperature sensors are electrically connected to the controller 1 to control the operation of the electric drive pump 6, the first switching valve 16, and the second switching valve 17.

[0046] The first switching valve 16 and the second switching valve 17 in this invention are designed to prevent fuel from flowing into the cooling circuit when the heat dissipation system is not activated. To ensure a reasonable and safe fuel circuit, the inlet of the wing fuel tank should be located at the top of the tank, and the outlet should be located at the bottom. The design of the number and installation position of the level gauges 14 in the fuel tank should consider the influence of the fuel tank attitude on the fuel level gauge measurement during takeoff, climb, descent, approach, and landing phases. In one specific embodiment, there is one second temperature sensor 11 and one level gauge 14 in the wing fuel tank to avoid an overly complex system. The level gauge 14 is preferably installed near the trailing edge of the wing in the fuel tank. This ensures that the fuel level can be measured to meet the heat dissipation requirements when the fuel tank is tilted forward and the fuel level is higher at the front and lower at the rear during descent, approach, and landing, thus ensuring the safe and stable operation of the fuel tank and the heat dissipation system. The flow meter 3 is connected to the fuel line flange and can be a volumetric flow meter or a mass flow meter. The purpose of the partition 13 is to ensure physical isolation between the components of the heat dissipation system and to ensure the safety and stability of the fuel tank in terms of physical structure.

[0047] The heat dissipation system of this invention is suitable for cooling the fuel in the wing fuel tanks of an aircraft during flight. The operating logic of this heat dissipation system is as follows: Figure 3 As shown. The steps include:

[0048] Step 1: After the aircraft engine starts, power is supplied to the measurement and control components in the wing fuel tank cooling system;

[0049] Step 2: The aircraft sends a landing gear grounding signal to the fuel tank cooling system as a signal to determine system activation;

[0050] Step 3: The fuel level gauge in the wing fuel tank detects the fuel level H in the fuel tank, the second temperature sensor 11 measures the fuel temperature T2 in the fuel tank, the first temperature sensor 2 detects the temperature T1 of the lower surface skin of the wing, and all the above signals are transmitted to the controller 1 for the controller 1 to make judgments.

[0051] Step 4: When T1 < T2 and H is higher than 1 / 3 of the full fuel tank level, the controller 1 controls the first switching valve 16 and the second switching valve 17 to open, and at the same time controls the electric drive pump 6 to start, and starts fuel heat dissipation; during the heat dissipation process, the third temperature sensor 5 measures the fuel temperature T3 after heat dissipation by the skin heat exchanger 4 and transmits it to the controller 1. The controller 1 can judge data such as heat dissipation efficiency based on T2, T3 and T1, which can be used when needed;

[0052] The determination condition of 1 / 3 of the aforementioned full fuel tank level is based on the reading information of the liquid level gauge 14; if multiple liquid level gauges are configured at the same time, the lowest fuel liquid level reading among all readings is used as the determination object, that is, among the multiple liquid level values that can be measured in the fuel tank, if one does not meet 1 / 3 of the full fuel tank level, the heat dissipation operation will not be carried out;

[0053] Step 5: When the fuel liquid level is lower than 1 / 3 of the full fuel tank level, or it is determined that the aircraft has landed safely according to the landing gear grounding signal, or the temperature difference between T1 and T2 is lower than the preset first threshold, or the temperature difference between T3 and T2 is lower than the preset second threshold, the controller 1 controls the electric drive pump 6 to stop working, and the system stops the heat dissipation operation. The first threshold is 5°C and the second threshold is 3°C.

[0054] The above has introduced in detail a civil aircraft wing fuel tank heat dissipation system and method using a skin heat exchanger provided by the embodiments of the present application. The description of the above embodiments is only used to help understand the method and its core idea of the present application; at the same time, for ordinary technical personnel in the field, according to the idea of the present application, there will be changes in the specific implementation manner and application scope. In summary, the content of this specification should not be construed as a limitation to the present application.

[0055] The terminology used in the embodiments of this invention is for the purpose of describing particular embodiments only and is not intended to limit the invention. The singular forms “a,” “the,” and “the” used in the embodiments of this invention and the appended claims are also intended to include the plural forms unless the context clearly indicates otherwise. In this application, the terms “upper,” “lower,” “left,” “right,” “inner,” “outer,” and “middle,” etc., indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings. Some of the above terms may also be used to indicate other meanings besides orientations or positional relationships; for example, the term “upper” may in some cases indicate a dependency or connection relationship. Those skilled in the art can understand the specific meaning of these terms in this application according to the specific circumstances. The term “and / or” used herein is merely a description of the association relationship between related objects, indicating that three relationships can exist. For example, A and / or B can represent: A alone, A and B simultaneously, and B alone. Additionally, the character “ / ” in this document generally indicates that the preceding and following related objects have an “or” relationship.

Claims

1. A heat dissipation system for an aircraft wing fuel tank, characterized in that, The cooling system includes a wing fuel tank, fuel lines, a skin heat exchanger, an electric pump, a level gauge, a first switching valve, a second switching valve, and a controller. The skin heat exchanger is disposed on the lower skin of the wing, and the skin heat exchanger is fused to the lower skin; The wing fuel tank is connected to the skin heat exchanger through the fuel line to form a loop; following the flow direction of the fuel in the loop, the loop is sequentially provided with the wing fuel tank, the first switching valve, the skin heat exchanger, the electric drive pump, and the second switching valve; The level gauge is installed inside the wing fuel tank; The first switching valve, the second switching valve, the electric drive pump, and the level gauge are all electrically connected to the controller. Specifically, the fusion connection is achieved by having the lower skin surface of the skin heat exchanger and the corresponding lower skin surface be the same skin; the skin heat exchanger is a counter-flow multi-channel gas-liquid heat exchanger; and the heat exchange fins of the skin heat exchanger are tightly fitted to the lower skin surface. The lower skin is equipped with a first temperature sensor, and the wing fuel tank is equipped with a second temperature sensor; both the first and second temperature sensors are electrically connected to the controller.

2. The aircraft wing fuel tank cooling system according to claim 1, characterized in that, A third temperature sensor is installed on the fuel line behind the skin heat exchanger, and the third temperature sensor is electrically connected to the controller.

3. The aircraft wing fuel tank cooling system according to claim 1, characterized in that, The fuel line is equipped with a flow meter and a pressure sensor, both of which are electrically connected to the controller.

4. The aircraft wing fuel tank cooling system according to claim 1, characterized in that, The circuit is also equipped with a check valve, and the check valve, the second switching valve, and the inlet of the wing fuel tank are arranged in sequence.

5. The aircraft wing fuel tank cooling system according to claim 1, characterized in that, It also includes a partition, which is located in the internal space of the aircraft wing and divides the internal space into upper and lower chambers; the wing fuel tank is located in the upper chamber of the two chambers, and the skin heat exchanger is located in the lower chamber of the two chambers.

6. A method for heat dissipation of fuel tanks in aircraft wings, characterized in that, The method is implemented using the heat dissipation system described in any one of claims 1-5; The steps of the method include: S1. Collect the aircraft's flight status signal to determine whether the aircraft is in flight; if yes, proceed to the next step; otherwise, continue collecting and judging. S2. Determine whether the heat dissipation system meets the heat dissipation conditions. If yes, proceed to the next step; otherwise, continue to determine. The heat dissipation conditions include: the real-time fuel level in the wing fuel tank is higher than 1 / 3 of the full fuel level; S3. The controller controls the opening of the first and second switching valves and controls the electric drive pump to start for fuel cooling. S4. Monitor the operating parameters of the heat dissipation system to determine whether the conditions for stopping heat dissipation are met; if yes, proceed to the next step; otherwise, continue heat dissipation while continuing to monitor and determine the conditions. The conditions for stopping heat dissipation are that the real-time fuel level in the wing fuel tank is not higher than 1 / 3 of the full fuel level or the aircraft is on the ground. S5. The controller shuts down the electric drive pump and closes the first and second switching valves.

7. The method for heat dissipation of aircraft wing fuel tanks according to claim 6, characterized in that, The heat dissipation conditions in step S2 also include: the real-time temperature of the fuel in the wing fuel tank is higher than the real-time temperature of the lower skin. The cooling stop condition in step S4 is that the real-time temperature of the fuel after cooling by the skin radiator is lower than the real-time temperature of the fuel in the wing fuel tank, and the difference is greater than a preset threshold.

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