An aircraft heat dissipation system, design method and control method

Abstract

The application provides an aircraft heat dissipation system, a design method and a control method, which mainly comprises a heat exchanger and a radiator. The radiator is installed on the surface of the aircraft and is provided with a first heat dissipation module and at least one second heat dissipation module. The first heat dissipation module is in a normally open state and is connected in parallel with the second heat dissipation module through a busbar structure. A heating module is installed on the surface of the second heat dissipation module and / or the corresponding busbar structure. According to the design result, the first and second heat dissipation modules are controlled to work in combination. The application reduces the occupation of the internal space of the aircraft by the external hanging mode of the aircraft, and meets the requirements of different working conditions by the combination of the modules of the radiator, so that the requirements of the power unit are reduced, and the control system is simplified. In combination with the application of the heating module, the system stability is improved, and the failure rate is reduced.

Description

Technical Field

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

[0002] Existing aircraft typically have one or more heat-generating components during flight, such as aircraft engines, fuel cells, lithium batteries, gasoline engines, and hybrid power systems. Although airflow can be used to dissipate heat during flight, a heat dissipation system of a certain scale is still needed to transfer heat to the corresponding heat dissipation areas or components for rapid heat dissipation. Furthermore, there are certain requirements for heat dissipation efficiency and results, such as the optimal operating temperature of the power system being within a relatively fixed range, neither too high nor too low.

[0003] In existing aircraft cooling designs, the common technical solution remains active cooling structures. These systems utilize components such as circulating pumps and flow valves to precisely regulate the flow of the cooling fluid, thereby controlling the operating temperature of the cooling system to meet the aforementioned cooling performance requirements and keep heat-generating components within their optimal operating temperature range. However, the addition of control components not only increases overall costs but also leads to a higher failure rate due to increased control complexity. Furthermore, a malfunction or crash in the control system can easily cause the entire cooling system to fail, jeopardizing flight safety. Additionally, the added power components also encroach on valuable space within the aircraft and reduce its effective payload capacity.

[0004] Although it is possible to save space and cost by linking the power unit of the cooling system to the aircraft's power system, the aircraft's power system is strongly correlated with the flight state and cannot be adjusted arbitrarily according to cooling requirements. Furthermore, the output of common aircraft power systems is also constant speed (such as rotorcraft, which mainly rely on changing the pitch of the rotor blades to adjust lift). Therefore, adding a controllable speed regulation mechanism will increase costs and occupy a lot of space.

[0005] In summary, improvements are needed to the existing cooling system to reduce the space and weight occupied by the power unit and to easily adjust the cooling capacity according to different operating conditions to meet system requirements. Summary of the Invention

[0006] In order to overcome the shortcomings of the prior art, the purpose of this invention is to provide an aircraft heat dissipation system, design method and control method, which reduces space occupation by being externally mounted on the aircraft and adapts to different operating conditions by using modular switch combination, thereby reducing the difficulty of control.

[0007] To achieve the above objectives, the present invention provides an aircraft heat dissipation system, including a heat exchanger and a radiator; the cold end of the heat exchanger is in direct or indirect contact with the heat-generating components of the aircraft and is connected to the radiator through a circulation pipeline; a cooling working fluid circulates between the heat exchanger and the radiator.

[0008] The radiator is an air-cooled radiator, which is installed on the surface of the aircraft's skin and / or gas flow channels;

[0009] The radiator is provided with a first heat dissipation module and at least one second heat dissipation module; the first heat dissipation module is normally open and is connected in parallel to the second heat dissipation module through a bus structure;

[0010] A switching valve is connected in series in the flow path of the second heat dissipation module, and a heating module is installed on the surface of the second heat dissipation module and / or the corresponding busbar structure.

[0011] Optionally, the heat exchanger is a cold plate, and the cold plate is provided with flow paths;

[0012] The temperature of the liquid inlet of the cold plate is within a fixed range.

[0013] Optionally, the radiator is a finned radiator, including an intermediate channel layer and a fin layer;

[0014] The intermediate channel layer is a harmonica tube-type liquid cooling plate; the fin layer is provided on the outer side of the intermediate channel layer.

[0015] Optionally, the cross-section of the fin layer has a closely spaced zigzag structure, and the protrusions face outwards from the aircraft;

[0016] Preferably, the long side direction of the fin layer protrusions is perpendicular to the flow direction of the cooling medium in the intermediate channel layer.

[0017] Optionally, the radiator may further include an upper panel and a lower panel;

[0018] The upper panel is attached to the skin of the aircraft; the intermediate channel layer is located between the upper panel and the lower panel; the fin layer is covered by the lower panel.

[0019] Optionally, the heating module includes a heating jacket;

[0020] The manifold structure is a collector pipe, and the heating sleeve is fitted onto the collector pipe;

[0021] A temperature sensor is also provided on the busbar structure;

[0022] Preferably, the heating module includes a heating film;

[0023] The heating film is attached to one side surface of the second heat dissipation module, and the second heat dissipation module is also equipped with a temperature sensor;

[0024] The system is equipped with at least two second heat dissipation modules, and the heat dissipation areas of the second heat dissipation modules are not all the same.

[0025] On the other hand, the present invention also provides a design method for an aircraft heat dissipation system, which mainly includes:

[0026] First, based on the atmospheric parameters at different altitudes and the corresponding flight states under different aircraft operating conditions, the heat dissipation power requirements for each aircraft operating condition are determined; a simulation model of the radiator is established, and then, based on the heat dissipation power requirements, the required heat dissipation area for each aircraft operating condition is calculated using the simulation model.

[0027] The smallest heat dissipation area is selected as the heat dissipation area of ​​the first heat dissipation module, and the difference between the heat dissipation area required under each aircraft operating condition and the heat dissipation area of ​​the first heat dissipation module is calculated to obtain the heat dissipation area to be increased.

[0028] The dataset consisting of the heat dissipation area to be increased is subjected to ordered clustering, and the resulting categories are used as the working stages of the heat sink.

[0029] The preset heat dissipation system includes several second heat dissipation modules, and the second heat dissipation modules include low-power second heat dissipation modules and high-power second heat dissipation modules;

[0030] Calculate the average heat dissipation area to be increased in each category, and remove the categories whose average heat dissipation area to be increased is less than a threshold. Select the category with the smallest average heat dissipation area as the low-power category, and take the maximum heat dissipation area to be increased in the low-power category as the total heat dissipation area of ​​the low-power second heat dissipation module. Then, take the difference between the maximum value of all heat dissipation areas to be increased and the heat dissipation area of ​​the low-power second heat dissipation module. The difference obtained is the total heat dissipation area of ​​the high-power second heat dissipation module.

[0031] Next, the rated power of the required heating module is calculated based on the aircraft's operating conditions, and then the corresponding selection process is carried out.

[0032] Optionally, a mapping relationship is established between the heat dissipation area to be increased and the corresponding aircraft operating conditions, and the dataset is formed.

[0033] Furthermore, the present invention also provides a corresponding control method based on the above-mentioned aircraft heat dissipation system and design method; the working stages of the radiator are divided according to the aircraft's operating conditions;

[0034] The control methods based on the aforementioned work stages include:

[0035] The first heat dissipation module remains in a normally open working state during each working stage, and the second heat dissipation module is selectively switched on and off during the corresponding working stage.

[0036] The heating module regulates the temperature of the cooling medium entering the heat exchanger to a fixed range.

[0037] Optionally, the heating module is installed on the second heat dissipation module, and during the period when the second heat dissipation module is turned off, the temperature of the second heat dissipation module is maintained within a fixed range by the corresponding heating module.

[0038] The present invention employs the above-described aircraft heat dissipation system, design method, and control method, which, compared with the prior art, has the following beneficial effects:

[0039] This invention reduces the space occupied by the aircraft by attaching the heat sink externally, effectively dissipates heat by utilizing airflow during flight, and modularizes the heat sink to meet the needs of different operating conditions through combination. This reduces the requirements of the power unit and simplifies the control system, thereby reducing costs while meeting operating requirements. In addition, the application of heating modules can assist in temperature regulation without adjusting the flow rate. Furthermore, it can ensure that each heat dissipation module is within its operating temperature range in low-temperature environments, improving system stability and reducing the failure rate. Attached Figure Description

[0040] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, 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.

[0041] Figure 1 This is a simplified diagram of the heat dissipation system of the present invention;

[0042] Figure 2 This is a perspective view of the first heat dissipation module of the present invention;

[0043] Figure 3 This is an exploded view of the first heat dissipation module of the present invention;

[0044] Figure 4 for Figure 3 Enlarged view of point A in the middle;

[0045] Figure 5 for Figure 3 Enlarged diagram of point B in the middle.

[0046] In the diagram, 1 is the radiator; 2 is the circulating pump; 3 is the cold plate; 4 is the power system; 101 is the first heat dissipation module; 102 is the second heat dissipation module; 103 is the heating jacket; 1011 is the upper panel; 1012 is the intermediate channel layer; 1013 is the lower panel; and 1014 is the fin layer. Detailed Implementation

[0047] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0048] It should be noted that similar labels and letters in the following figures indicate similar items. Therefore, once an item is defined in one figure, it does not need to be further defined and explained in subsequent figures.

[0049] In the description of this invention, it should be understood that the relative relationships indicated by terms such as "upper", "lower", and "back side" are based on the results of parameter comparisons in actual applications and are only for the convenience of describing this invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific position, and therefore should not be construed as a limitation of this invention.

[0050] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified.

[0051] In this invention, unless otherwise explicitly specified and limited, the terms "series connection" and "connection" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral part; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0052] It should also be noted that, unless otherwise specified, the methods used in this invention are conventional methods; and the raw materials and apparatus used are, unless otherwise specified, conventional commercially available products.

[0053] like Figure 1As shown in the figure, this embodiment provides an aircraft heat dissipation system, which mainly includes a heat exchanger and a radiator.

[0054] In this embodiment, the cold end of the heat exchanger is in direct or indirect contact with the heat-generating components of the aircraft. Specifically, in this embodiment, the core component of the heat exchanger is the cold plate 3, and the cold plate 3 is provided with a flow path for the cooling working fluid. It transfers heat with the heat-generating components of the aircraft that need to dissipate heat. In this embodiment, the main heat-generating component of the aircraft is the power system 4.

[0055] The cold plate 3 is connected to the radiator 1 via a circulation pipe, and a circulation pump 2 is connected in series on the pipe. The pump's input end is connected to the aircraft's power system via a transmission mechanism, thereby obtaining the power required for operation. The cooling medium flows along... Figure 1 The direction of the middle arrow circulates between the cold plate 3 and the radiator 1, continuously transferring the heat of the power system 4 to the radiator 1.

[0056] In this embodiment, the radiator 1 is an air-cooled radiator, which is specifically mounted on the surface of the aircraft skin by means of a hanging mount. After contouring processing, the radiator 1 is made to fit the skin of the mounting area, thereby reducing the wind resistance caused by the external mount.

[0057] The radiator 1 is provided with a first heat dissipation module 101 and at least one second heat dissipation module 102. In this embodiment, the radiator 1 consisting of one first heat dissipation module 101 and two second heat dissipation modules 102 is used as an example for system description. The first heat dissipation module 101 is in a normally open state and always participates in the operation of the heat dissipation system after power-on. The two second heat dissipation modules 102 are connected in parallel through a bus structure.

[0058] In this embodiment, both the first and second heat dissipation modules adopt the same structural design; therefore, the following description focuses only on the structure of the first heat dissipation module 101. Figure 2 , 3 As shown (the piping on both sides of the heat dissipation module is omitted in the figure), the first heat dissipation module 101, due to the use of contouring processing, has an overall arc-shaped structure that is approximately flat (the curvature of the heat dissipation module is specifically set according to the curvature of the skin). The first heat dissipation module 101 is a finned heat sink, and along the direction of the skin to the outside, as shown... Figure 3 The structure includes, in sequence, an upper panel 1011, a middle channel layer 1012, a lower panel 1013, and a fin layer 1014.

[0059] Among them, the upper and lower panels serve as transitional connection structures between each layer. The upper panel 1011 connects upward to the skin at the bottom of the aircraft, and can be fixedly connected to the skin or other airframe structural components by bolts or adapters.

[0060] The intermediate channel layer 1012 is a harmonica tube type liquid cooling plate; please refer to the details. Figure 4The structure is shown in the middle, with a large number of parallel microchannels machined in the layer plate, serving as flow paths for the cooling working fluid, and along... Figure 3 The flow is in the direction of the upper arrow in the middle.

[0061] The lower surface of the intermediate channel layer 1012 is bonded to the lower panel 1013, and the lower surface of the lower panel 1013 is bonded to the fin layer 1014. Optionally, combined with... Figure 5 As shown, the fin layer 1014 has a closely spaced "Z"-shaped cross-section, with the protrusions facing outwards and the roots fixed to the lower panel 1013. Within the fin layer 1014, the "Z"-shaped protrusions form elongated ridges. The long side of these ridges is aligned with the aircraft's axis and projects along the normal direction of the skin. The long side of the ridges is perpendicular to the flow direction of the cooling medium in the intermediate channel layer 1012. Figure 3 As shown, in common flight scenarios, the external air flows relative to the aircraft in the direction of the downward arrow, thus smoothly passing through the fin layer 1014 and making more full contact with the convex structure. Furthermore, since the flow direction of the cooling medium is perpendicular to the long side of the convex ridge, the heat exchange efficiency of the cooling medium in the flow path is higher, and the simulation conditions are more simplified in the design stage. In addition, the same-direction setting has higher requirements for machining accuracy, requiring the bottom of the middle of the two convex ridges to be strictly aligned with the flow path. Therefore, the vertical setting can also reduce the machining difficulty.

[0062] In this embodiment, to ensure that the temperature of the cooling medium entering the inlet of the power system 4 is always at its optimal temperature, an electrically controlled switching valve is connected in series in the flow path of each second heat dissipation module 102. This valve controls the temperature of the cooling medium entering the cold plate 3 by switching on and off the corresponding number of second heat dissipation modules 102 under different operating conditions. Optionally, this embodiment also considers the low temperature conditions at high altitudes. If the second heat dissipation modules 102 are not properly heated, the following problems may occur, specifically:

[0063] a. When the ambient temperature is lower than the rated operating temperature of the cooling medium, some types of cooling medium may appear as flocculent or lumpy, causing flow problems. This can lead to the failure of the second heat dissipation module 102 when it restarts, resulting in a heat dissipation system failure and endangering flight safety.

[0064] b. When the second heat dissipation module 102 restarts normally, the temperature of the internal cooling medium is too low, which leads to an overall low temperature after the current converges. This causes the operating temperature of the power system 4 to drop sharply, the battery efficiency to decrease significantly, and affects the operation of the aircraft in many ways, even endangering flight safety.

[0065] Unlike existing technologies that use flow-controlled heat dissipation power, this embodiment avoids the aforementioned problems because it does not have the issue of intermittent operation at the working end. However, this embodiment needs to address these issues. Therefore, while adopting a multi-heat dissipation module combined switch control design, this embodiment also requires the addition of different temperature control methods to ensure temperature controllability and system reliability. Consequently, this embodiment also installs heating modules on the surface of the second heat dissipation module 102 and on the corresponding busbar structure.

[0066] Optionally, the heating module also includes a first heating module, specifically a PI heating film (polyimide film electrothermal film) in this embodiment. The auxiliary heating PI heating film, along with the heat insulation layer and PSA adhesive (pressure-sensitive adhesive), is integrated and glued to the back of the second heat dissipation module 102, optionally at the position between the top plate and the skin. Correspondingly, each second heat dissipation module 102 is also equipped with a temperature sensor, which, in conjunction with the PI heating film, monitors the temperature of the cooling medium and performs necessary first-stage heating. Optionally, the radiator 1 in this embodiment has two second heat dissipation modules 102, and the two second heat dissipation modules 102 have different heat dissipation areas. Based on the size of the heat dissipation area, they are respectively a low-power second heat dissipation module and a high-power second heat dissipation module, so as to adapt to more heat dissipation requirements by changing the combination.

[0067] Optionally, the heating module also includes a second heating module. The second heat dissipation module 102 is connected to the main flow path via a manifold, and a heating sleeve 103 is fitted on the manifold. A temperature sensor is installed on the manifold or its upstream position to monitor the temperature of the cooling medium of the flow and perform necessary second-stage heating.

[0068] The aforementioned aircraft cooling system simplifies the power unit structure and control system in conventional designs. This embodiment uses a combination of radiator module switches to create different operating states, meeting the needs of multiple operating conditions. It avoids the complexity of flow regulation systems and, through pre-designing the radiator, achieves direct connection of the circulating pump to the aircraft's power system. This eliminates the need to consider whether the power system's output to the cooling system is at a constant speed; even with variable speed output, no special design is required because the variable speed output is strongly correlated with the overall aircraft operating conditions, thus naturally creating a correlation. In summary, this embodiment reduces the requirements for the power unit and transmission mechanism, thereby reducing the space and payload requirements of the aircraft, lowering costs, and improving reliability.

[0069] Based on the above embodiments, the present invention also provides a design method for an aircraft heat dissipation system, the main contents of which are as follows:

[0070] First, in this embodiment, the aircraft's power system (including fuel cells, lithium batteries, gasoline engines, and hybrid power) operates at different altitudes during takeoff, climb, and cruise, resulting in varying heat generation and external heat dissipation environments. Therefore, the radiator's operation is divided into multiple phases. Specifically, during climb, cruise, and takeoff, the heat dissipation power requirements are determined based on atmospheric parameters at different altitudes. A simulation model of the radiator is then established. Based on the heat dissipation power requirements, the required heat dissipation area for each operating condition is calculated using the simulation model and used for preliminary engineering design and calculations.

[0071] Furthermore, the heat dissipation performance of the radiator is analyzed, including the microchannels of the intermediate channel layer and the zigzag protrusions of the fin layer. Considering the inlet temperature, outlet temperature, and structural design of the intermediate channel layer, the heat dissipation capacity per unit area of ​​the radiator related to flow rate is obtained. Simulation calculations are performed based on external conditions (including flight speed and ambient temperature) at different altitudes and internal operating conditions (including the flow rate pushed by the circulating pump and the required cooling power of the power system). To simplify the radiator simulation model calculation, this embodiment only selects typical flight altitudes and corresponding flight states for simulation calculations, thereby calculating the required heat dissipation area of ​​the radiator under the aforementioned aircraft operating conditions. The smallest heat dissipation area is selected as the area of ​​the first heat dissipation module. The difference between the required heat dissipation area for each aircraft operating condition and the heat dissipation area of ​​the first heat dissipation module is calculated to obtain several different additional heat dissipation areas. Furthermore, based on design requirements, it is preliminarily determined that the heat dissipation system includes K second heat dissipation modules. In this embodiment, K=2 is set for ease of explanation. Furthermore, although in reality, the ambient temperature changes almost linearly with flight altitude (e.g., the temperature decreases by 6 degrees Celsius for every 1000 meters increase in altitude), the required heat dissipation area varies in stages due to changes in airflow and heat generation from the power system. Therefore, a mapping relationship can be established between the several heat dissipation areas to be increased obtained above and the aircraft's operating conditions. Then, the resulting dataset is subjected to ordered clustering, and the resulting categories correspond to the operating stages of the radiator. In this embodiment, the two second heat dissipation modules have different heat dissipation areas: one is a low-power heat dissipation module, and the other is a high-power heat dissipation module. Calculate the average heat dissipation area to be increased in each category. Since at least one category corresponds to the working stage where only the first heat dissipation module is running, its average heat dissipation area to be increased is close to 0. Therefore, a threshold is designed. Categories with an average heat dissipation area to be increased below the threshold do not belong to the working stage where only the first heat dissipation module is running. Therefore, remove the categories with an average heat dissipation area to be increased below the threshold. Among the remaining categories, select the category with the smallest average heat dissipation area to be increased as the low-power category. Take the maximum value in this category as the heat dissipation area of ​​the low-power second heat dissipation module. Then, select the maximum value among all heat dissipation areas to be increased and take the difference from it. The difference is taken as the heat dissipation area of ​​the high-power second heat dissipation module.

[0072] Next, the rated power of the required heating module is calculated based on the operating conditions, and then the corresponding selection is carried out.

[0073] The present invention also provides a corresponding heat dissipation system control method based on the above design, the specific method being as follows:

[0074] 1. Based on the working stages corresponding to the aforementioned categories (including the working stage in which only the first heat dissipation module operates), there are four working stages in this embodiment as the working stages of the heat dissipation system. Optionally, in this embodiment, each working stage is ordered according to the time sequence of the corresponding aircraft operating conditions, such as: first working stage (takeoff stage); second working stage (climb stage); third working stage (cruise stage); fourth working stage (landing stage).

[0075] 2. The first heat dissipation module remains constantly on during each working stage, while the second heat dissipation module is selectively switched on and off during the corresponding working stage; specifically as follows:

[0076] In the first working stage, the first heat dissipation module and the high-power second heat dissipation module are turned on; the low-power second heat dissipation module is turned off. In this stage, the heat dissipation power generated by the total heat dissipation area of ​​the first heat dissipation module and the high-power second heat dissipation module meets the system requirements. If there is any heat dissipation power exceeding the requirements in this stage, it will be balanced by the heating module.

[0077] In the second working phase, the first heat dissipation module, the low-power second heat dissipation module, and the high-power second heat dissipation module are activated. In this phase, the heat dissipation power generated by the total heat dissipation area of ​​the first and second heat dissipation modules meets the system requirements. If the heat dissipation power exceeds the requirements in this phase, the heating module will balance it.

[0078] In the third working stage, the first heat dissipation module and the low-power second heat dissipation module are turned on; the high-power second heat dissipation module is turned off. In this stage, the heat dissipation power generated by the total heat dissipation area of ​​the first heat dissipation module and the low-power second heat dissipation module meets the system requirements. If there is any heat dissipation power exceeding the requirements in this stage, it will be balanced by the heating module.

[0079] In the fourth working phase, the first heat dissipation module is turned on, and the low-power second heat dissipation module and the high-power second heat dissipation module are turned off. During this phase, the total heat dissipation area of ​​the first heat dissipation module generates enough heat dissipation power to meet the system requirements.

[0080] During each working stage, the heating module adjusts the temperature of the cooling medium entering the heat exchanger to be between 50℃ and 70℃ according to the working conditions.

[0081] Optionally, the heating module can operate in the following states:

[0082] In the first working stage, the PI heating film on the high-power second heat dissipation module and / or the heating sleeve on the high-power second heat dissipation module manifold are turned on.

[0083] In the second working stage, the PI heating film and heating jacket are turned off or adjusted.

[0084] In the third working stage, the PI heating film on the low-power second heat dissipation module and / or the heating sleeve on the manifold of the low-power second heat dissipation module are turned on; the PI heating film on the high-power second heat dissipation module is turned on.

[0085] In the fourth working stage, the PI heating film on the low-power second heat dissipation module and the high-power second heat dissipation module is turned on, and the heating sleeve is turned off.

[0086] This document uses specific examples to illustrate the principles and implementation methods of the present invention. The descriptions of the above embodiments are only for the purpose of helping to understand the method and core ideas of the present invention. Furthermore, those skilled in the art will recognize that, based on the ideas of the present invention, there will be changes in the specific implementation methods and application scope. Therefore, the content of this specification should not be construed as a limitation of the present invention.

Claims

1. An aircraft cooling system, comprising a heat exchanger and a radiator; the cold end of the heat exchanger is in direct or indirect contact with the heat-generating components of the aircraft and is connected to the radiator via a circulation pipeline; a cooling working fluid circulates between the heat exchanger and the radiator; Its features are, The radiator is an air-cooled radiator, which is installed on the surface of the aircraft's skin. The radiator is provided with a first heat dissipation module and at least one second heat dissipation module; the first heat dissipation module is normally open and is connected in parallel to the second heat dissipation module through a bus structure; A switching valve is connected in series in the flow path of the second heat dissipation module, and a heating module is installed on the surface of the second heat dissipation module and the corresponding busbar structure.

2. The aircraft heat dissipation system according to claim 1, characterized in that, The heat exchanger is a cold plate, and the cold plate is provided with flow paths; The temperature of the liquid inlet of the cold plate is within a fixed range.

3. The aircraft heat dissipation system according to claim 1, characterized in that, The radiator is a finned radiator, including an intermediate channel layer and a fin layer; The intermediate channel layer is a harmonica tube-type liquid cooling plate; the fin layer is provided on the outer side of the intermediate channel layer.

4. The aircraft heat dissipation system according to claim 3, characterized in that, The cross-section of the fin layer has a closely spaced zigzag structure, with the protrusions facing outwards from the aircraft.

5. The aircraft heat dissipation system according to claim 4, characterized in that, The long side of the finned layer protrusions is perpendicular to the flow direction of the cooling medium in the intermediate channel layer.

6. The aircraft heat dissipation system according to claim 3 or 5, characterized in that, The radiator also includes an upper panel and a lower panel; The upper panel is attached to the skin of the aircraft; the intermediate channel layer is located between the upper panel and the lower panel; the fin layer is covered by the lower panel.

7. The aircraft heat dissipation system according to claim 1, characterized in that, The heating module includes a heating jacket; The manifold structure is a collector pipe, and the heating sleeve is fitted onto the collector pipe; A temperature sensor is also installed on the busbar structure.

8. The aircraft heat dissipation system according to claim 7, characterized in that, The heating module includes a heating film; The heating film is attached to one side surface of the second heat dissipation module, and the second heat dissipation module is also equipped with a temperature sensor; The system is equipped with at least two second heat dissipation modules, and the heat dissipation areas of the second heat dissipation modules are not all the same.

9. The aircraft heat dissipation system design method according to any one of claims 1 to 8, characterized in that, Design methods include: First, based on the atmospheric parameters at different altitudes and the corresponding flight states under different aircraft operating conditions, the heat dissipation power requirements for each aircraft operating condition are determined; a simulation model of the radiator is established, and then, based on the heat dissipation power requirements, the required heat dissipation area for each aircraft operating condition is calculated using the simulation model. The smallest heat dissipation area is selected as the heat dissipation area of ​​the first heat dissipation module, and the difference between the heat dissipation area required under each aircraft operating condition and the heat dissipation area of ​​the first heat dissipation module is calculated to obtain the heat dissipation area to be increased. The dataset consisting of the heat dissipation area to be increased is subjected to ordered clustering, and the resulting categories are used as the working stages of the heat sink. The preset heat dissipation system includes several second heat dissipation modules, and the second heat dissipation modules include low-power second heat dissipation modules and high-power second heat dissipation modules; Calculate the average heat dissipation area to be increased in each category, and remove the categories whose average heat dissipation area to be increased is less than a threshold. Select the category with the smallest average heat dissipation area as the low-power category, and take the maximum heat dissipation area to be increased in the low-power category as the total heat dissipation area of ​​the low-power second heat dissipation module. Then, take the difference between the maximum value of all heat dissipation areas to be increased and the heat dissipation area of ​​the low-power second heat dissipation module. The difference obtained is the total heat dissipation area of ​​the high-power second heat dissipation module. Next, the rated power of the required heating module is calculated based on the aircraft's operating conditions, and then the corresponding selection process is carried out.

10. The aircraft heat dissipation system design method according to claim 9, characterized in that, A mapping relationship is established between the heat dissipation area to be increased and the corresponding aircraft operating conditions, and the dataset is formed.

11. The aircraft heat dissipation system control method according to any one of claims 1 to 8, characterized in that, The working stages of the radiator are divided according to the operating conditions of the aircraft. The control methods based on the aforementioned work stages include: The first heat dissipation module remains in a normally open working state during each working stage, and the second heat dissipation module is selectively switched on and off during the corresponding working stage. The heating module regulates the temperature of the cooling medium entering the heat exchanger to a fixed range.

12. The aircraft heat dissipation system control method according to claim 11, characterized in that, The heating module is installed on the second heat dissipation module, and when the second heat dissipation module is turned off, the temperature of the second heat dissipation module is maintained within a fixed range by the corresponding heating module.

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