A cooling system for long-endurance high-power airborne electronic equipment

Abstract

The application relates to a cooling system of long-time high-power airborne electronic equipment, and belongs to the technical field of aircraft equipment heat dissipation, and solves the problem of insufficient heat dissipation efficiency of high-power electronic equipment of a subsonic aircraft. The cooling system comprises a heat dissipation structural member, a driving pump, a heat storage pool and an aircraft skin; the electronic equipment is in contact with the heat dissipation structural member, the heat dissipation structural member is heated and the electronic equipment is cooled at the same time through heat conduction; the heat dissipation structural member is internally provided with a cooling liquid internal flow channel for circulating cooling liquid; the driving pump is used for driving the cooling liquid to circulate between the heat dissipation structural member and the heat storage pool; the heat storage pool internally stores the cooling liquid; the cooling liquid can circulate between the heat storage pool and the aircraft skin and conduct heat to the external environment. The application can transfer heat to the external environment by driving the cooling liquid to circulate between the heat dissipation structural member, the heat storage pool and the aircraft skin, so that effective and rapid heat dissipation of the electronic equipment is realized.

Description

[0001] This application claims priority to application number "202411741202.1" filed on November 29, 2024, entitled "A Liquid Cooling Circulation Heat Dissipation System for Electronic Devices". Technical Field

[0002] This invention relates to the field of aircraft equipment heat dissipation technology, and in particular to a cooling system for long-endurance, high-power airborne electronic equipment. Background Technology

[0003] In recent years, as electronic devices have become increasingly integrated, their heat generation has increased exponentially. At the same time, the size and weight of electronic devices have been continuously reduced, and miniaturization and lightweighting have become fundamental requirements for electronic devices.

[0004] Under the combined effect of these two factors, the heat flux density of electronic devices has increased dramatically, and the operating time has continued to increase. Thermal control technology has become a key technology that restricts the further development of detection and guidance technology.

[0005] Active liquid circulation systems have high heat exchange capacity and are increasingly widely used in electronic equipment cooling systems. However, in practical applications, airborne equipment has limited space and cannot be equipped with water chillers.

[0006] Therefore, there is a need to provide a cooling system for high-power airborne electronic equipment with long flight time. Summary of the Invention

[0007] Based on the above analysis, the present invention aims to provide a cooling system for long-endurance, high-power airborne electronic equipment to solve the problems of high heat generation power and low heat dissipation efficiency of electronic equipment with long operating time in existing subsonic aircraft.

[0008] The objective of this invention is mainly achieved through the following technical solutions:

[0009] A cooling system for long-endurance, high-power airborne electronic equipment, used to dissipate heat from electronic equipment inside an aircraft; comprising: heat dissipation structural components, a drive pump, a heat storage tank, and an aircraft skin;

[0010] The electronic device comes into contact with the heat dissipation structure, and heats the heat dissipation structure while cooling the electronic device through heat conduction.

[0011] The heat storage tank contains coolant; the heat dissipation structure has an internal coolant flow channel for the flow of coolant.

[0012] The drive pump is used to drive the coolant to flow between the heat dissipation structure and the heat storage tank.

[0013] The coolant can flow between the heat storage tank and the aircraft skin and conduct heat to the external environment.

[0014] Furthermore, the heat dissipation structure is a ring-shaped structure that surrounds the outside of the electronic device; the heat dissipation structure is also provided with a coolant inlet and a coolant outlet that connect to the internal flow channel of the coolant.

[0015] Furthermore, the aircraft skin is provided with parallel flow channels for the circulation of coolant.

[0016] Furthermore, the heat dissipation structure includes: an upper retaining ring, a middle retaining ring, and a lower retaining ring; the upper retaining ring, the middle retaining ring, and the lower retaining ring are fixed by welding.

[0017] Furthermore, the lower surface of the upper ring is provided with a first L-shaped flow channel, a first linear flow channel, and a second L-shaped flow channel.

[0018] Furthermore, the middle layer retaining ring is provided with a first U-shaped flow channel and a second U-shaped flow channel penetrating its upper and lower surfaces.

[0019] Furthermore, the first L-shaped flow channel, the first U-shaped flow channel, the first linear flow channel, the second U-shaped flow channel, and the second L-shaped flow channel are connected in sequence, and the second linear flow channel and the third linear flow channel can respectively close the bottom slots of the first U-shaped flow channel and the second U-shaped flow channel to form the internal flow channel of the coolant.

[0020] Furthermore, the upper surface of the lower retaining ring is provided with a second linear flow channel and a third linear flow channel.

[0021] Furthermore, the internal cavity of the heat storage tank is filled with a phase change medium, and when the coolant flows through the heat storage tank, the phase change medium can exchange heat with the coolant.

[0022] Furthermore, the heat storage tank includes: an upper cover plate, a lower cover plate, heat-conducting pipes, and an outer cylinder; the outer cylinder is fixedly connected between the upper cover plate and the lower cover plate to form the internal cavity of the heat storage tank.

[0023] Furthermore, an upper liquid cavity is provided inside the upper cover plate, and a lower liquid cavity is provided inside the lower cover plate. A heat-conducting pipe is vertically arranged between the upper cover plate and the lower cover plate, and the upper and lower ends of the heat-conducting pipe are connected to the upper liquid cavity and the lower liquid cavity.

[0024] The technical solution of this invention can achieve at least one of the following effects:

[0025] 1. The cooling system for long-endurance, high-power airborne electronic equipment of the present invention uses two heat dissipation methods, namely heat storage tank and skin heat exchange, to dissipate heat from the electronic equipment, thereby improving the heat dissipation capacity.

[0026] 2. The cooling system for long-endurance, high-power airborne electronic equipment of the present invention improves the heat exchange capacity of the skin and reduces the flow resistance by optimizing the internal flow channels of the skin and designing them as parallel flow channels.

[0027] 3. The cooling system for long-endurance, high-power airborne electronic equipment of the present invention, by designing the heat dissipation structure as a multi-layered assembly structure, and setting multiple bends in the internal flow channels of the coolant inside the heat dissipation structure, achieves sufficient heat exchange of the coolant inside the heat dissipation structure, thereby achieving rapid heat dissipation of the electronic equipment inside the heat dissipation structure.

[0028] In this invention, the above-described technical solutions can be combined with each other to achieve more preferred combinations. Other features and advantages of this invention will be set forth in the following description, and some advantages may become apparent from the description or be learned by practicing the invention. The objects and other advantages of this invention can be realized and obtained from what is particularly pointed out in the description and drawings. Attached Figure Description

[0029] The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Throughout the drawings, the same reference numerals denote the same parts.

[0030] Figure 1 This is a schematic diagram of the structure of the cooling system for a long-endurance, high-power airborne electronic device according to Embodiment 1 of the present invention;

[0031] Figure 2 This is a schematic diagram of the heat dissipation structure of the cooling system for the long-endurance, high-power airborne electronic equipment according to Embodiment 1 of the present invention;

[0032] Figure 3 This is a schematic diagram of the internal parallel flow channel structure of the aircraft skin in Embodiment 1 of the present invention;

[0033] Figure 4 This is a schematic diagram of the three-layer ring structure of the heat dissipation structure component in Embodiment 1 of the present invention;

[0034] Figure 5 This is a schematic diagram of the upper ring structure of the heat dissipation structure;

[0035] Figure 6 A schematic diagram of the front structure of the middle layer retaining ring of the heat dissipation structure;

[0036] Figure 7 A schematic diagram of the back structure of the middle layer retaining ring of the heat dissipation structure;

[0037] Figure 8 This is a schematic diagram of the lower layer of the heat dissipation structure.

[0038] Figure 9 This is a schematic diagram of the structure of the heat storage tank in Embodiment 2 of the present invention;

[0039] Figure 10 for Figure 9 A schematic diagram of the heat pipe structure in the heat storage tank;

[0040] Figure 11 for Figure 9 A cross-sectional view of the upper cover plate of the thermal storage tank.

[0041] Figure 12 This is a schematic diagram of the nested annular parallel flow channel structure of the aircraft skin in Embodiment 3 of the present invention;

[0042] Figure 13 for Figure 12 A magnified view of a portion of the first confluence channel;

[0043] Figure 14 for Figure 12 A magnified view of the second confluence channel.

[0044] Figure label:

[0045] 1-Electronic equipment; 2-Heat dissipation structural components; 3-Drive pump; 4-Aircraft skin; 5-Heat storage tank; 201-Coolant inlet; 202-Coolant outlet; 203-Coolant internal flow channel; 401-Parallel flow channel;

[0046] 21-Upper layer retaining ring; 22-Middle layer retaining ring; 23-Lower layer retaining ring; 24-Positioning pin hole;

[0047] 204 - First L-shaped flow channel; 205 - First U-shaped flow channel; 206 - First linear flow channel; 207 - Second U-shaped flow channel; 208 - Second L-shaped flow channel; 209 - Second linear flow channel; 210 - Third linear flow channel;

[0048] 41-Liquid inlet port; 42-Liquid inlet channel; 43-Annular parallel channel; 44-Cross-shaped outlet channel; 45-Outlet port; 431-First annular channel; 432-First confluence channel; 433-Second annular channel; 434-Second confluence channel; 435-Third annular channel; 436-Third confluence channel; 437-Fourth annular channel; 4320-Concentric annular channel; 4321-Outer annular channel; 4322-Inner annular channel; 4341-First heat dissipation ring; 4342-Second heat dissipation ring;

[0049] 51-Upper cover plate; 52-Upper liquid chamber; 53-Lower cover plate; 54-Lower liquid chamber; 55-First coolant connector; 56-Heat pipe; 57-Second coolant connector; 58-Outer cylinder; 59-Heat pipe connection hole; 561-Heat-conducting flow pipe; 562-Heat-conducting sphere; 563-Through hole. Detailed Implementation

[0050] The preferred embodiments of the present invention will now be described in detail with reference to the accompanying drawings, which constitute a part of the present invention and are used together with the embodiments of the present invention to illustrate the principles of the present invention, but are not intended to limit the scope of the present invention.

[0051] Example 1

[0052] One specific embodiment of the present invention discloses a cooling system for long-endurance, high-power airborne electronic equipment, such as... Figure 1 As shown, it includes: a heat dissipation structure 2, a drive pump 3, a heat storage tank 5, and an aircraft skin 4; an electronic device 1 is in contact with the heat dissipation structure 2, heating the heat dissipation structure 2 and cooling the electronic device 1 through heat conduction; the heat storage tank 5 stores coolant; the heat dissipation structure 2 is provided with a coolant flow channel 203 for the flow of coolant; the drive pump 3 is used to drive the coolant to flow between the heat dissipation structure 2 and the heat storage tank 5; the coolant can flow between the heat storage tank 5 and the aircraft skin 4 to conduct heat to the external environment.

[0053] Furthermore, such as Figure 2 As shown, the heat dissipation structure 2 is a ring-shaped structure; in use, the heat dissipation structure 2 surrounds the outside of the electronic device; the heat dissipation structure 2 is provided with a coolant inlet 201, a coolant outlet 202 and a coolant inner flow channel 203, and coolant flows in the coolant inner flow channel 203.

[0054] In one specific embodiment of the present invention, such as Figure 4 As shown, the heat dissipation structure 2 is a split structure, including: an upper retaining ring 21, a middle retaining ring 22 and a lower retaining ring 23.

[0055] Specifically, such as Figure 5 , Figure 6 , Figure 7 , Figure 8As shown, the lower surface of the upper retaining ring 21 is provided with a first L-shaped flow channel 204, a first linear flow channel 206, and a second L-shaped flow channel 208; the middle retaining ring 22 is provided with a first U-shaped flow channel 205 and a second U-shaped flow channel 207 penetrating its upper and lower surfaces; the upper surface of the lower retaining ring 23 is provided with a second linear flow channel 209 and a third linear flow channel 210; after the upper retaining ring 21, the middle retaining ring 22, and the lower retaining ring 23 are welded together, the first L-shaped flow channel 204, the first U-shaped flow channel 205, the first linear flow channel 206, the second U-shaped flow channel 207, and the second L-shaped flow channel 208 are connected in sequence, and the second linear flow channel 209 and the third linear flow channel 210 can respectively close the bottom slots of the first U-shaped flow channel 205 and the second U-shaped flow channel 207, forming the coolant internal flow channel 203.

[0056] Furthermore, such as Figure 2 As shown, the heat dissipation structure 2 is further provided with a coolant inlet 201 and a coolant outlet 202 communicating with the coolant internal flow channel 203. Both the coolant inlet 201 and the coolant outlet 202 are connected to the outer surface of the heat dissipation structure 2. The coolant inlet 201 is used to connect to an inlet connector, which is used to introduce coolant into the coolant internal flow channel 203 of the heat dissipation structure 2. The coolant outlet 202 is used to connect to an outlet connector, which is used to discharge coolant from the coolant internal flow channel 203 of the heat dissipation structure 2.

[0057] Preferably, such as Figure 4 As shown, two coolant flow channels 203 are symmetrically arranged inside the heat dissipation structure 2. Figure 4 As shown, two internal coolant channels 203 are arranged within the heat dissipation structure 2, flowing from the front and rear sides respectively. This fully utilizes the usable solid structure within the heat dissipation structure 2 to enhance the efficiency of heat exchange and cooling. The working principle of the heat dissipation structure 2 of this invention is as follows: After the electronic device 1 starts working, it generates heat, which heats the heat dissipation structure 2. The pump 3 drives low-temperature coolant to flow into the internal coolant channels 203. The liquid coolant exchanges heat with the heat dissipation structure 2, carrying away the heat and thus achieving temperature control for the electronic device 1.

[0058] To ensure good heat exchange between the coolant and the heat dissipation structure, the liquid working fluid must have sufficient heat exchange area within the heat dissipation structure 2. In this invention, the design of the coolant flow channel 203 within the heat dissipation structure is as follows: One end of the first L-shaped flow channel 204 is connected to the coolant inlet 201, and the other end is connected to one end of the first U-shaped flow channel 205. One end of the first U-shaped flow channel 205 is connected to the first L-shaped flow channel 204, and the other end is connected to one end of the first linear flow channel 206. One end of the second U-shaped flow channel 207 is connected to the other end of the first linear flow channel 206, and the other end is connected to one end of the second L-shaped flow channel 208. The other end of the second L-shaped flow channel 208 is connected to the coolant outlet 202.

[0059] Furthermore, such as Figure 4 , Figure 8 As shown, the second linear flow channel 209 and the third linear flow channel 210 on the upper surface of the lower ring 23 are respectively aligned with the bottom cavities of the first U-shaped flow channel 205 and the second U-shaped flow channel 207 on the middle ring 22, which can seal the bottom cavities of the first U-shaped flow channel 205 and the second U-shaped flow channel 207.

[0060] Specifically, such as Figure 6 , Figure 7 As shown, the first U-shaped flow channel 205 and the second U-shaped flow channel 207 are arranged perpendicular to the end face of the middle layer ring 22. The bottom of the first U-shaped flow channel 205 and the second U-shaped flow channel 207 are connected to the lower surface of the middle layer ring 22, and both ends of the first U-shaped flow channel 205 and the second U-shaped flow channel 207 are connected to the upper surface of the middle layer ring 22. That is, the starting ends of the first U-shaped flow channel 205 and the second U-shaped flow channel 207 are located on the upper surface of the middle layer ring 22, pass through the middle layer ring 22 and connect to the lower surface of the middle layer ring 22, and the first U-shaped flow channel 205 and the second U-shaped flow channel 207 are connected to the side of the starting end by a slot at the bottom end of the middle layer ring 22, and then pass through the middle layer ring 22 again to connect to its upper surface as the end openings of the first U-shaped flow channel 205 and the second U-shaped flow channel 207.

[0061] Furthermore, positioning pin holes 24 are provided on the upper retaining ring 21, the middle retaining ring 22 and the lower retaining ring 23; pins are installed in the positioning pin holes 24 to position the relative positions of the upper retaining ring 21, the middle retaining ring 22 and the lower retaining ring 23.

[0062] In this invention, since the heat dissipation structure 2 is in direct contact with the electronic device 1, it is required to have good sealing performance. Furthermore, to ensure that the flow resistance of the complex flow channel is less than the pressure head of the driving pump, the heat dissipation structure cannot be formed using 3D molding. Due to the complexity of the internal flow channel, conventional brazing splicing molding is also not feasible. Therefore, the heat dissipation structure 2 was ultimately designed as a three-section structure fixed by welding to achieve proper molding. The specific processing flow is as follows: First, process the upper retaining ring 21, the middle retaining ring 22, and the lower retaining ring 23 according to requirements. To ensure the reliability of welding, the welding width should not be less than 3mm at any welding point. Leave machining allowance in positions prone to deformation for subsequent machining after welding. Machining the solder according to the dimensions of the welding surface, and installing positioning pins in the directional pin holes 24 of the upper retaining ring 21, the middle retaining ring 22, and the lower retaining ring 23 for positioning and assembly. Press the upper retaining ring 21, the middle retaining ring 22, and the lower retaining ring 23 together, and then perform vacuum brazing. After welding, pressure testing and leak detection are performed. The pressure test pressure should be greater than the pump head to ensure no leakage occurs. After passing the pressure test, final processing is performed, and leak testing is performed again after processing to ensure its reliability.

[0063] Furthermore, such as Figure 3 As shown, the aircraft skin 4 has parallel flow channels 401 inside for the flow of coolant.

[0064] In one specific embodiment of the present invention, in the electronic cooling active circulation system, the drive pump 3 is required to have a simple structure, high reliability, convenient maintenance, long service life, low cost, small size, large head, and stable performance.

[0065] In one specific embodiment of the present invention, the heat storage tank 5 is structured similarly to a water tank, and the inlet and outlet positions of the heat storage tank 5 are designed to be bottom inlet and top outlet, thereby ensuring that the low-temperature coolant flows out continuously from the inside of the heat storage tank 5.

[0066] In one specific embodiment of the present invention, the aircraft skin 4 is the external load-bearing structure of the entire aircraft. In addition to its structural load-bearing function, the aircraft skin 4 also undertakes the function of convective heat exchange with the outside world.

[0067] Therefore, the aircraft skin 4 must have a sufficiently large internal coolant flow channel area to ensure effective convective heat transfer while maintaining structural strength. Simultaneously, its internal flow resistance must be as low as possible to accommodate the driving capability of the drive pump 3. In this embodiment, the aircraft skin 4 is internally equipped with multiple parallel flow channels 401 connected in parallel, such as... Figure 3 As shown; simultaneously, process fillets are added to the parallel flow channels 401 to increase the heat exchange area and reduce flow resistance. Specifically, in this embodiment, the parallel flow channels 401 are configured as multi-layered parallel U-shaped flow channels.

[0068] The aircraft skin 4 of the present invention is made of corrosion-resistant and wear-resistant materials, such as stainless steel or special alloys, according to the properties and flow rate of the coolant.

[0069] In this embodiment, the coolant internal flow channel 203 of the heat dissipation structure 2, the heat storage tank 5, and the parallel flow channel 401 in the aircraft skin 4 are sequentially connected to form a circulation path, and the coolant is driven to circulate in the circulation path by the drive pump 3. Specifically, the heat dissipation structure 2 is connected to the internal flow channel of the aircraft skin 4 through the liquid inlet connector, so that the low-temperature coolant flows into the coolant internal flow channel 203 to cool the heat dissipation structure 2. The high-temperature coolant after absorbing heat and heating up flows into the heat storage tank 5 to exchange heat with the heat-absorbing material in the heat storage tank 5 for a first-stage cooling. The cooled coolant flows into the parallel flow channel 401 inside the aircraft skin 4 for a second cooling to become a low-temperature coolant. The low-temperature coolant flows into the heat dissipation structure 2 again for the next cycle of cooling.

[0070] The cooling system of this invention operates as follows: After the aircraft enters its flight path, the drive pump 3 is powered on first, cooling the coolant in the heat storage tank 5 through convection heat exchange. After the electronic equipment 1 starts operating, it generates heat, which heats the heat dissipation structure 2 through contact heat dissipation. The heat dissipation structure 2 has internal coolant flow channels 203. Under the action of the drive pump 3, the coolant circulates between the heat storage tank 5, the aircraft skin 4, and the heat dissipation structure 2. Simultaneously, the coolant in the heat storage tank 5, driven by the drive pump 3, continuously flows to the subsonic aircraft skin 4. Through convection heat exchange between the low-temperature coolant in the heat storage tank 5 and the aircraft skin 4 with the external environment, the temperature of the electronic equipment 1 is controlled. After the electronic equipment finishes operating, the drive pump 3 continues to operate, further cooling the coolant in the heat storage tank 5, awaiting the electronic equipment 1 to operate again.

[0071] Example 2

[0072] In one specific embodiment of the present invention, the heat storage tank 5 in Embodiment 1 is further improved by design:

[0073] like Figure 9 , Figure 10 , Figure 11 As shown, the heat storage tank 5 includes: an upper cover plate 51, a lower cover plate 53, a heat-conducting pipe 56, an outer cylinder 58, a first coolant connector 55, and a second coolant connector 57. The upper and lower ends of the outer cylinder 58 are fixedly connected to the upper cover plate 51 and the lower cover plate 53, and the three together form the internal cavity of the heat storage tank 5.

[0074] Specifically, the upper cover plate 51 is provided with an upper liquid cavity 52, the lower cover plate 53 is provided with a lower liquid cavity 54, and the heat pipe 56 is vertically arranged between the upper cover plate 51 and the lower cover plate 53, and the upper and lower ends of the heat pipe 56 are connected to the upper liquid cavity 52 and the lower liquid cavity 54.

[0075] Furthermore, a first coolant connector 55 is provided on the outer side of the upper cover plate 51, and the first coolant connector 55 communicates with the upper liquid chamber 52 inside the upper cover plate 51 for introducing coolant into the upper liquid chamber 52. A second coolant connector 57 is provided on the outer side of the lower cover plate 53, and the second coolant connector 57 communicates with the lower liquid chamber 54 inside the lower cover plate 53 for discharging coolant from the lower liquid chamber 54.

[0076] Specifically, the first coolant connector 55 is connected to the coolant outlet 202 of the heat dissipation structure 2 via the heat storage tank inlet pipe, thereby enabling the coolant flowing through the coolant channel 203 of the heat dissipation structure 2 to be introduced into the heat storage tank 5. In other words, the first coolant connector 55 can introduce the high-temperature coolant that has exchanged heat with the heat dissipation structure 2 into the heat storage tank 5 for cooling.

[0077] Specifically, the second coolant connector 57 can introduce coolant into the internal flow channel of the aircraft skin 4 through the heat storage tank outlet pipe.

[0078] Furthermore, the upper cover plate 51 and the lower cover plate 53 of the heat storage tank 5 are connected as one unit by the outer cylinder 58, and the cavity formed by the three is filled with a phase change medium; in this embodiment, the phase change medium is paraffin wax in a liquid state.

[0079] In this embodiment, as Figure 10 As shown, the heat pipe 56 includes a heat-conducting flow pipe 561 and a plurality of heat-conducting balloons 562 arranged in parallel outside the heat-conducting flow pipe 561.

[0080] Specifically, multiple heat-conducting balloons 562 are arranged side by side along the axial direction of the heat-conducting tube 561; and the multiple heat-conducting balloons 562 are distributed at equal intervals on the heat-conducting tube 561.

[0081] Preferably, the distance between two adjacent heat-conducting balloons 562 is smaller than the diameter of the heat-conducting balloon 562.

[0082] Specifically, such as Figure 10 As shown, the heat-conducting balloon 562 has a spherical shell structure, and the heat-conducting balloon 562 is entirely sleeved on the outside of the heat-conducting and diverting tube 561; and multiple through holes 563 are opened on the heat-conducting balloon 562.

[0083] Specifically, two sets of through holes 563 are symmetrically arranged on the upper and lower sides of the heat-conducting balloon 562; and multiple through holes 563 in each set are equally spaced along the circumferential direction of the heat-conducting tube 561. Preferably, the number of through holes 563 in each set is four or five.

[0084] Preferably, the through hole 563 is a circular, elliptical, or U-shaped hole.

[0085] In this embodiment, by providing multiple through holes 563, the inner and outer spaces of the heat-conducting balloon 562 can be connected. When the phase change medium is injected into the internal cavity of the heat storage tank 5, the phase change medium can flow into the interior of the heat-conducting balloon 562 through the through holes 563, thereby filling the space between the heat-conducting balloon 562 and the heat-conducting flow pipe 561.

[0086] Preferably, the upper cover plate 51 and the lower cover plate 53 are both split structures with identical structural components. Specifically, the upper cover plate 51 includes a top plate, a bottom plate, and an upper liquid cavity 52 disposed between the top plate and the bottom plate; the upper liquid cavity 52 is a closed rectangular cavity used for the flow of coolant, thereby enabling the coolant to be distributed to multiple heat pipes 56.

[0087] Specifically, such as Figure 11 As shown, an upper liquid chamber 52 is provided inside the upper cover plate 51. Multiple heat pipe connection holes 59 are arrayed on the bottom plate of the upper liquid chamber 52, and these connection holes 59 connect the upper liquid chamber 52 to the internal cavity of the heat storage tank 5. The upper end of the heat-conducting flow pipe 561 is fixedly installed in the heat pipe connection hole 59. Correspondingly, multiple heat pipe connection holes 59 are provided on the top plate of the lower cover plate 53 for connecting the lower end port of the heat-conducting flow pipe 561.

[0088] In practice, the high-temperature coolant, after exchanging heat with the heat dissipation structure 2, flows from the first coolant connector 55 into the upper liquid chamber 52 of the upper cover plate 51, and then flows through multiple heat pipe inlets 59 into the heat-conducting guide pipes 561 of multiple heat pipes 56, transferring heat to the heat pipes 56. The heat is then transferred to the phase change medium inside the heat storage tank 5 through the heat-conducting guide pipes 561 and multiple heat-conducting balloons 562. The phase change medium absorbs heat and undergoes a phase change, cooling the coolant in the heat-conducting guide pipes 561. Finally, the cooled coolant flows into the lower liquid chamber 54 and exits the heat storage tank 5 through the second coolant connector 57. Furthermore, after exiting the heat storage tank 5, the coolant can flow into the internal flow channels of the aircraft skin 4 for further cooling.

[0089] In this embodiment, multiple arrayed heat-conducting pipes 56 are arranged inside the heat storage tank 5, and multiple heat-conducting spheres 562 are arranged on the heat-conducting pipes 561. The multiple heat-conducting spheres 562 on the multiple heat-conducting pipes 56 form a rectangular cubic lattice of heat exchange nodes, realizing multi-point heat dissipation. When the high-temperature coolant transfers heat to the heat-conducting pipes 56, the inner and outer walls of the heat-conducting spheres 562 on the heat-conducting pipes 56 can quickly conduct the heat to the phase change medium, realizing rapid cooling of the high-temperature coolant by the heat storage tank 5.

[0090] In this embodiment, by increasing the heat storage capacity of the heat storage tank 5, the temperature of the coolant flowing into the aircraft skin 4 can be reduced as much as possible, thereby weakening the heating effect of the high-temperature coolant on the aircraft skin 4 and avoiding the phenomenon of reduced structural strength of the aircraft skin 4 due to excessive heating.

[0091] Example 3

[0092] In a specific embodiment of the present invention, the parallel flow channel 401 of the aircraft skin 4 in Embodiment 1 is further improved by design:

[0093] In this embodiment, the parallel flow channel 401 of the aircraft skin 4 is designed as a multi-layer nested annular parallel flow channel.

[0094] like Figure 12 As shown, the multi-layer nested annular parallel flow channel of this embodiment includes: liquid inlet 41, liquid inlet flow channel 42, annular parallel flow channel 43, cross-shaped liquid outlet flow channel 44 and liquid outlet 45.

[0095] Specifically, one end of the liquid inlet channel 42 is connected to the liquid inlet port 41, and the other end is connected to the outermost channel of the annular parallel channel 43; the liquid inlet port 41 is connected to the heat storage tank 5 through the coolant pipeline, and is used to introduce coolant into the inner channel of the aircraft skin 4.

[0096] Preferably, such as Figure 12 As shown, four liquid inlet channels 42 are equally spaced along the circumference of the annular parallel channels 43, and each of the four liquid inlet channels 42 is connected to a liquid inlet port 41 at its starting end, allowing coolant to flow into the annular parallel channels 43 through the four liquid inlet channels 42 for heat exchange with the aircraft skin 4. Furthermore, the liquid inlet port 41 is connected to the outlet pipe of the heat storage tank 5 via four skin liquid inlet pipes and a five-way pipe structure, allowing coolant flowing out of the heat storage tank 5 to flow into the five-way pipe structure via the outlet pipe and be distributed to the four skin liquid inlet pipes, thus enabling it to flow into the annular parallel channels 43 through the four liquid inlet channels 42 for heat exchange and cooling with the aircraft skin 4.

[0097] The multi-layer nested annular parallel flow channel of this embodiment has multiple liquid inlet ports 41 designed in the inlet section to introduce the coolant into the annular parallel flow channel 43 in multiple ways for heat exchange and cooling. In addition, by setting multiple first confluence flow channels 432 as branching branches, the multiple branching branches are connected in parallel with each other, which can improve the uniformity and stability of the coolant flow in the aircraft skin 4.

[0098] Specifically, the annular parallel flow channel 43 has a multi-layered nested annular structure; the cross-shaped liquid outlet flow channel 44 is located in the middle of the annular parallel flow channel 43, and all four ports of the cross-shaped liquid outlet flow channel 44 are connected to the innermost flow channel of the annular parallel flow channel 43; the liquid outlet interface 45 is located in the center of the cross-shaped liquid outlet flow channel 44, and is connected to the annular parallel flow channel 43 through the four branches of the cross-shaped liquid outlet flow channel 44, so that the coolant after being cooled by the aircraft skin 4 can be discharged through the liquid outlet interface 45.

[0099] like Figure 12 As shown, the annular parallel flow channel 43 of the present invention includes multiple nested annular flow channels, and adjacent annular flow channels are connected by multiple parallel converging flow channels, and the multiple converging flow channels are equally spaced along the circumferential direction of the annular flow channels.

[0100] Preferably, such as Figure 12 As shown, the annular parallel flow channel 43 includes four nested annular flow channels: a first annular flow channel 431, a second annular flow channel 433, a third annular flow channel 435, and a fourth annular flow channel 437.

[0101] Specifically, multiple first converging channels 432 are provided between the first annular flow channel 431 and the second annular flow channel 433 for connection; the multiple first converging channels 432 are equally spaced along the circumferential direction of the second annular flow channel 433; preferably, the first converging channel 432 is an arc-shaped flow channel.

[0102] Specifically, multiple second confluence channels 434 are provided between the second annular flow channel 433 and the third annular flow channel 435 for connection; the multiple second confluence channels 434 are equally spaced along the circumference of the third annular flow channel 435; preferably, the second confluence channel 434 is an arc-shaped flow channel.

[0103] Specifically, multiple third confluence channels 436 are provided between the third annular flow channel 435 and the fourth annular flow channel 437 for connection; the multiple third confluence channels 436 are equally spaced along the circumference of the fourth annular flow channel 437; preferably, the third confluence channels 436 are arc-shaped channels.

[0104] Furthermore, the end of the liquid inlet channel 42 is connected to the first annular channel 431, and the four liquid inlet channels 42 are arranged at 90° intervals along the outer side of the first annular channel 431. The four liquid inlet ports of the cross-shaped liquid outlet channel 44 are connected to the inner side of the fourth annular channel 437, and the four branches of the cross-shaped liquid outlet channel 44 are interconnected in the middle.

[0105] In this embodiment, a cross-shaped liquid outlet channel 44 is designed, with an outlet port 45 located at the center of the cross-shaped liquid outlet channel 44. The coolant, after being cooled by the annular parallel channel 43, is gathered at the outlet port 45, and then flows back into the heat dissipation structure 2 for heat absorption. In this embodiment, by setting an inlet channel 42 on the outside of the annular parallel channel 43 and a cross-shaped liquid outlet channel 44 on its inside, the parallel connection and multi-layer nested conduction of multiple converging channel branches are realized, improving the heat exchange efficiency of the coolant. At the same time, due to the rapid spread of the coolant inside the aircraft skin 4, the temperature rise of the skin structure is effectively reduced, avoiding the decrease in structural performance such as stiffness and strength caused by the skin temperature rise.

[0106] In practice, coolant flows into the first annular channel 431 through four inlet channels 42, and is then diverted from the first annular channel 431 to multiple first confluence channels 432. The coolant then flows from the multiple first confluence channels 432 into the second annular channel 433. Further, the coolant flows from the second annular channel 433 to multiple second confluence channels 434, and then from the multiple second confluence channels 434 into the third annular channel 435. The coolant in the third annular channel 435 is diverted into multiple third confluence channels 436, and then flows through the multiple third confluence channels 436 into the fourth annular channel 437. Finally, the coolant in the fourth annular channel 437 flows through the cross-shaped outlet channel 44 into the outlet port 45, and can also flow through the skin outlet pipe into the coolant inner channel 203 of the heat dissipation structure 2, thus dissipating heat for the heat dissipation structure 2 and the electronic equipment 1.

[0107] In this embodiment, when the coolant inside the aircraft skin 4 flows in the annular flow channel and the arc-shaped confluence flow channel, the overall flow channel structure is mostly arc-shaped and can be divided by multiple confluence flow channels, which can effectively reduce flow resistance and realize the smooth diversion and confluence of coolant. The flow channel design conforms to the principles of fluid dynamics, which is conducive to the smooth flow of coolant inside the annular parallel flow channel 43, thereby ensuring the heat exchange effect.

[0108] Furthermore, in order to enhance the heat exchange effect between the coolant and the aircraft skin structure, the aforementioned annular parallel flow channel 43 is further improved in design:

[0109] In this embodiment, a concentric annular channel 4320 is provided on the first confluence channel 432. For example... Figure 13 As shown, the concentric annular flow channel 4320 includes an outer annular flow channel 4321 and an inner annular flow channel 4322 arranged concentrically. Specifically, the outer annular flow channel 4321 is connected to the first confluence flow channel 432, and the first confluence flow channel 432 passes through the outer annular flow channel 4321 and is connected to the inner annular flow channel 4322. The coolant flows from the first confluence flow channel 432 into the outer annular flow channel 4321 and the inner annular flow channel 4322, and then flows back into the first confluence flow channel 432, and flows into the second annular flow channel 433 via the first confluence flow channel 432.

[0110] In this embodiment, the coolant is further divided into two annular outer ring channel 4321 and inner ring channel 4322 in the first confluence channel 432 as heat dissipation sub-channels. The number and size of the sub-channels can be adjusted as needed to adjust the heat dissipation capacity of the aircraft skin 4, thereby improving the adaptability of the cooling system of the present invention.

[0111] In this embodiment, as Figure 14 As shown, at least two heat dissipation annular channels are arranged side by side in the second confluence channel 434; preferably, a first heat dissipation ring 4341 and a second heat dissipation ring 4342 are arranged side by side on the second confluence channel 434.

[0112] Specifically, the first heat dissipation ring 4341 and the second heat dissipation ring 4342 are connected in series in the second confluence channel 434, such as Figure 14 As shown, the coolant is diverted from the second annular flow channel 433 to the multi-channel second confluence flow channel 434, and then flows through the second confluence flow channel 434 in sequence through the first heat dissipation ring 4341 and the second heat dissipation ring 4342 before flowing into the third annular flow channel 435.

[0113] In this embodiment, the aircraft skin 4 is internally configured with multi-layered nested annular parallel flow channels as internal liquid flow channels. These channels include multiple parallel confluence channels and multi-layered nested annular channels, which are interconnected through multiple confluence channels. The rotating and converging flow channel structure in the annular parallel flow channels 43 enables the rapid distribution of coolant to multiple branch confluence channels, and allows the coolant to flow sequentially between the multi-layered nested annular channels. On one hand, this achieves tiered heat exchange and cooling of the coolant, allowing it to quickly spread on the outside of the annular parallel flow channels 43, preventing high-temperature coolant from stagnating in localized areas of the aircraft skin 4, thus improving heat exchange efficiency. On the other hand, the flow velocity of the coolant slows down when it flows into the outer annular channels and confluence channels with larger areas, extending the heat exchange time with the aircraft skin 4 and improving the heat exchange effect.

[0114] The above description is only a preferred embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention.

Claims

1. A cooling system for long-endurance, high-power airborne electronic equipment, used to dissipate heat from electronic equipment (1) inside an aircraft; characterized in that, include: Heat dissipation structure (2), drive pump (3), heat storage tank (5) and aircraft skin (4); The electronic device (1) comes into contact with the heat dissipation structure (2), and the heat dissipation structure (2) is heated by heat conduction while the electronic device (1) is cooled down. The heat storage tank (5) contains coolant; the heat dissipation structure (2) has a coolant flow channel for the flow of coolant. The drive pump (3) is used to drive the coolant to flow between the heat dissipation structure (2) and the heat storage tank (5); The coolant can flow between the heat storage tank (5) and the aircraft skin (4) and conduct heat to the external environment.

2. The cooling system for long-endurance, high-power airborne electronic equipment according to claim 1, characterized in that, The heat dissipation structure (2) is a ring-shaped structure that surrounds the outside of the electronic device (1); the heat dissipation structure (2) is also provided with a coolant inlet (201) and a coolant outlet (202) that connect to the coolant internal flow channel (203).

3. The cooling system for long-endurance, high-power airborne electronic equipment according to claim 1, characterized in that, The aircraft skin (4) is provided with parallel flow channels (401) for the flow of coolant.

4. The cooling system for long-endurance, high-power airborne electronic equipment according to claim 1, characterized in that, The heat dissipation structure (2) includes an upper retaining ring (21), a middle retaining ring (22), and a lower retaining ring (23); the upper retaining ring (21), the middle retaining ring (22), and the lower retaining ring (23) are fixed by welding.

5. The cooling system for long-endurance, high-power airborne electronic equipment according to claim 4, characterized in that, The lower surface of the upper ring (21) is provided with a first L-shaped flow channel (204), a first linear flow channel (206), and a second L-shaped flow channel (208).

6. The cooling system for long-endurance, high-power airborne electronic equipment according to claim 5, characterized in that, The middle layer ring (22) is provided with a first U-shaped flow channel (205) and a second U-shaped flow channel (207) that penetrate its upper and lower surfaces.

7. The cooling system for long-endurance, high-power airborne electronic equipment according to claim 6, characterized in that, The upper surface of the lower ring (23) is provided with a second linear flow channel (209) and a third linear flow channel (210).

8. The cooling system for long-endurance, high-power airborne electronic equipment according to any one of claims 1-7, characterized in that, The internal cavity of the heat storage tank (5) is filled with a phase change medium. When the coolant flows through the heat storage tank (5), the phase change medium can exchange heat with the coolant.

9. The cooling system for long-endurance, high-power airborne electronic equipment according to claim 8, characterized in that, The heat storage tank (5) includes: an upper cover plate (51), a lower cover plate (53), a heat-conducting pipe (56), and an outer cylinder (58).

10. The cooling system for long-endurance, high-power airborne electronic equipment according to claim 9, characterized in that, The outer cylinder (58) is fixedly connected between the upper cover plate (51) and the lower cover plate (53) to form the internal cavity of the heat storage tank (5).

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