Aircraft power system with composite heat dissipation function and aircraft

Abstract

The invention relates to the technical field of aircrafts, and particularly discloses an aircraft power system with a composite heat dissipation function and an aircraft, the system comprises a power motor and a fairing, the power motor comprises a stator and a rotor rotationally connected with the stator through a rotating shaft, and the stator comprises a support component and an iron core-winding assembly; the support component is provided with a first containing cavity for containing the iron core-winding assembly and a second containing cavity located at the bottom of the rotating shaft, and the first containing cavity is communicated with the second containing cavity through a first flow channel arranged in the support component; the fairing is arranged on the end face of the rotor, a liquid cooling heat exchange cavity is formed in the exposed wall body of the fairing, and the liquid cooling heat exchange cavity communicates with the first containing cavity through a second flow channel penetrating through the rotor. The output end of the liquid pump is rotationally connected with a liquid guide pipe coaxial with the rotating shaft, one end of the liquid guide pipe penetrates through the rotating shaft to communicate with the liquid cooling heat exchange cavity, the power system integrates an internal circulation liquid cooling structure in the power motor and the fairing, and on the basis that efficient heat dissipation is achieved, additional flight wind resistance is not increased.

Description

Technical Field

[0001] This application relates to the field of aircraft technology, and in particular to an aircraft power system with composite heat dissipation and an aircraft. Background Technology

[0002] As electric aircraft rapidly develop towards higher payload and longer endurance, the power density of their core drive motors continues to increase significantly. When the drive motors are running, they generate a lot of heat under long-term, high-intensity working conditions. If the heat cannot be dissipated in time, it will not only lead to a decline in the performance of the power system, but also seriously restrict the flight safety and mission endurance of the aircraft. Against this background, single air cooling is no longer sufficient to meet the heat dissipation requirements of high-power motors, and liquid cooling technology with higher heat dissipation efficiency is gradually becoming the mainstream choice.

[0003] In related technologies, due to the small size of the aircraft's drive motor, cooling structures such as independent radiators and liquid supply lines are mainly added outside the motor. However, if the cooling structure is built-in, it will occupy a lot of internal space in the aircraft fuselage. If the cooling structure is exposed, it will disrupt the aerodynamic streamline, resulting in a significant increase in the flight drag coefficient and thus a significant reduction in flight time. Although related technologies have attempted to integrate the coolant flow channel with the motor housing, it is still connected to the fuselage liquid inlet or temperature control module through exposed pipes, adding multiple sealing nodes and greatly increasing the probability of coolant leakage. The integration of the flow channel is achieved only by thickening the motor housing, resulting in a significant increase in the overall weight of the motor compared to the original structure, which increases the wind resistance of the aircraft's power system. Summary of the Invention

[0004] This application aims to propose an aircraft power system and aircraft with composite heat dissipation, so as to at least solve the technical problems in the prior art where the integration of the cooling radiator structure increases the overall space ratio of the power system and the flight drag, and the exposed arrangement of coolant pipes makes coolant leakage easy.

[0005] In a first aspect, this application provides an aircraft propulsion system with composite heat dissipation, comprising: A power motor includes a stator and a rotor rotatably connected to the stator via a rotating shaft. The stator includes a support member and a core-winding assembly. The support member has a first cavity for accommodating the core-winding assembly and a second cavity located at the bottom of the rotating shaft. The first cavity communicates with the second cavity through a first flow channel provided in the support member. A shroud is provided on the end face of the rotor. A liquid-cooled heat exchange chamber is provided in the exposed wall of the shroud. The liquid-cooled heat exchange chamber is connected to the first cavity through a second flow channel that penetrates the rotor. The second cavity is equipped with a liquid pump, and the output end of the liquid pump is rotatably connected to a liquid guide pipe that is coaxial with the rotating shaft. One end of the liquid guide pipe passes through the rotating shaft and communicates with the liquid-cooled heat exchange cavity.

[0006] In some embodiments, the exposed wall of the fairing is a hollow sandwich structure, which forms the liquid-cooled heat exchange cavity; the liquid-cooled heat exchange cavity is continuously distributed along the circumference and axial direction of the exposed wall of the fairing and is located within the thickness range of the exposed wall.

[0007] In some embodiments, the first cavity is an annular cavity and is arranged around the axis of the support member. The support member is provided with a bearing chamber with an open end along the central axis. The rotating shaft extends coaxially from the open end into the bearing chamber through a bearing sleeve to form a sealed second cavity at the bottom end of the bearing chamber.

[0008] In some embodiments, an axially arranged air inlet duct is provided between the bearing chamber and the first cavity, and a plurality of reinforcing plates are arranged radially along the support member in the air inlet duct. The two ends of the reinforcing plates extend to the bearing chamber and the first cavity, respectively, and the first flow channel is provided in the reinforcing plates.

[0009] In some embodiments, the rotor includes a front end cover coaxially disposed on the end face of the support member. The bottom surface of the front end cover is provided with an air outlet duct arranged radially and communicating with the air inlet duct. The air outlet duct is arranged circumferentially around the front end cover, and a plurality of partitions arranged radially along the support member are provided in the air outlet duct. A portion of the second flow channel is opened in the partitions and extends to communicate with the first cavity.

[0010] In some embodiments, multiple heat dissipation fins are arranged circumferentially along the support member inside the air inlet duct, and each heat dissipation fin is arranged radially along the support member.

[0011] In some embodiments, the rotating shaft is provided with a first through hole and a second through hole in sequence along the axial direction, and a stepped section is formed at the connection between the first through hole and the second through hole. A sealing ring adapted to the stepped section is provided on the outer wall of the liquid guide tube.

[0012] In some embodiments, the liquid guide tube is provided with a positioning part near the outer wall of the sealing ring, the outer diameter of the positioning part is larger than the inner diameter of the first through hole, and one end of the positioning part is provided with a limiting ring extending to the first through hole.

[0013] In some embodiments, a guide plate is provided at the end of the liquid guide tube opposite to the positioning part, and the guide plate is provided with a plurality of radially arranged liquid distribution holes; and The top of the shroud is provided with an annular groove adapted to the guide plate. The outer periphery of the guide plate is at least partially embedded in the annular groove and is interference-fitted with the groove wall. The inner sidewall of the annular groove is provided with a plurality of liquid inlet holes spaced apart along the circumference. Each liquid inlet hole is connected to the liquid distribution hole and the liquid-cooled heat exchange chamber.

[0014] Compared with the prior art, the technical solution provided in the first aspect of this application has at least the following beneficial effects or advantages: The aircraft propulsion system with composite heat dissipation provided in this application integrates a liquid-cooled heat exchange chamber into the exposed wall of the fairing. A first cavity, constructed via a support structure, directly accommodates the core-winding assembly, the heat-generating core. A second cavity houses a liquid pump. A closed, fully integrated liquid-cooled circulation path is constructed using first and second flow channels and a liquid guide pipe penetrating the rotating shaft. On one hand, the first cavity fully accommodates the core-winding assembly, and the second cavity temporarily stores and guides the coolant, allowing the coolant to fully absorb the large amount of heat generated by the high-power-density motor. This, combined with the liquid-cooled heat exchange chamber... The efficient heat dissipation enables continuous cooling through internal coolant circulation, effectively preventing performance degradation of the power system and ensuring flight safety and endurance of the aircraft. On the other hand, it eliminates the need for additional independent radiators and exposed liquid delivery pipelines, avoiding the problems of excessive fuselage space occupied by internal cooling structures or disruption of aerodynamic streamlines by exposed structures, thus reducing flight drag and minimizing flight time loss. Furthermore, the integrated design of the first cavity, second cavity, and support components allows for channel integration without thickening the motor housing, avoiding weight redundancy and reducing additional sealing nodes caused by exposed pipelines, thereby reducing the risk of coolant leakage.

[0015] Secondly, embodiments of this application provide an aircraft, the aircraft including an aircraft power system with composite heat dissipation as described in any of the first aspects above.

[0016] It is understandable that the beneficial effects of the second aspect mentioned above can be found in the relevant descriptions in the first aspect mentioned above, and will not be repeated here.

[0017] Additional aspects and advantages of this application will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of this application. Attached Figure Description

[0018] To more clearly illustrate the technical solutions of the embodiments of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0019] Figure 1This is an exploded structural diagram of a power system provided according to an embodiment of this application; Figure 2 This is a schematic diagram of the power system provided according to an embodiment of this application; Figure 3 It is based on Figure 2 A cross-sectional view along the AA direction; Figure 4 It is based on Figure 3 A magnified view of part A in the middle; Figure 5 This is a structural schematic diagram of a support component provided according to an embodiment of this application; Figure 6 This is a structural schematic diagram of the front cover plate provided according to an embodiment of this application; Figure 7 This is a schematic diagram of the structure of a liquid pump provided according to an embodiment of this application; Figure 8 This is a cross-sectional view of the rotating shaft along the axis provided in the embodiments of this application; Figure 9 This is a schematic diagram of the liquid guide tube provided according to an embodiment of this application.

[0020] Figure label: 100. Power system; 101. First flow channel; 102. Second flow channel; 103. Fastening nut; 10. Power motor; 11. Stator; 111. Support component; 1111. First cavity; 1112. Bearing chamber; 11121. Second cavity; 1113. Reinforcing plate; 1114. Air inlet duct; 1115. Heat dissipation fins; 112. Core-winding assembly; 12. Rotor; 121. Bearing sleeve; 1211. Bearing; 122. Shaft; 1221. First through hole; 1222. Second through hole; 1223. Trapezoidal section; 123. Front cover; 1231. Air outlet duct; 1232. Partition plate; 20. Fairing; 21. Liquid-cooled heat exchange chamber; 22. Annular groove; 221. Liquid inlet; 23. Reinforcing rib; 231. Flow guide hole; 30. Liquid pump; 301. Input end; 302. Output end; 31. Liquid guide tube; 311. Positioning part; 3111. Limiting ring; 312. Sealing ring; 313. Flow guide plate; 3131. Liquid distribution hole; 40. Propeller blades. Detailed Implementation

[0021] The embodiments of this application are described in detail below. The embodiments described with reference to the accompanying drawings are exemplary. It should be understood that the specific embodiments described herein are merely for explaining this application and are not intended to limit this application.

[0022] It should be noted that when a component is said to be "fixed to" another component, it can be directly attached to the other component or there may be an intervening component. When a component is said to be "connected to" another component, it can be directly connected to the other component or there may be an intervening component.

[0023] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this application belongs. The terminology used herein in the specification of this application is for the purpose of describing particular embodiments only and is not intended to be limiting of the application. The term "and / or" as used herein includes any and all combinations of one or more of the associated listed items.

[0024] Please see Figures 1 to 4 This embodiment provides an aircraft propulsion system with composite heat dissipation. The propulsion system 100 may include a power motor 10, a fairing 20, a liquid pump 30, and a propeller blade 40. The power motor 10 includes a stator 11 and a rotor 12 rotatably connected to the stator 11 via a rotating shaft 122. The stator 11 includes a support member 111 and a core-winding assembly 112. The support member 111 has a first cavity 1111 for accommodating the core-winding assembly 112 and a second cavity 11121 located at the bottom of the rotating shaft 122. The first cavity 1111 communicates with the second cavity 11121 through a first flow channel 101 provided in the support member 111. Specifically, the first cavity... The first cavity 1111 is formed by two coaxially arranged rings of different diameters on the support member 111, and enclosed by the end cover and bottom cover (not shown in the figure) of the rotor 12. The first cavity 1111 is an annular cavity adapted to the shape of the core-winding assembly 112, coaxially arranged around the axis of the support member 111, and can completely enclose the core-winding assembly 112. In this way, the coolant flowing into the first cavity 1111 can completely immerse the core-winding assembly 112. The second cavity 11121 is located directly below the connection between the rotating shaft 122 and the support member 111. At the same time, the second cavity 11121 is a sealed cavity and communicates with the first cavity 1111.

[0025] Furthermore, the fairing 20 is located on the end face of the rotor 12. The exposed wall of the fairing 20 contains a liquid-cooled heat exchange chamber 21. It should be understood that the exposed wall refers to the wall structure of the fairing 20 that is in contact with the outside air. For example, if the fairing 20 in this application is bullet-shaped, its entire arc-shaped wall can be regarded as an exposed wall. The liquid-cooled heat exchange chamber 21 is connected to the first cavity 1111 through the second flow channel 102 that passes through the rotor 12. The second cavity 11121 contains a liquid pump 30. The output end 302 of the liquid pump 30 is rotatably connected to a liquid guide pipe 31 that is coaxially arranged with the rotating shaft 122. One end of the liquid guide pipe 31 passes through the rotating shaft 122 and is connected to the liquid-cooled heat exchange chamber 21.

[0026] It should be noted that the liquid pump 30 can be a micro pump, which can be obtained from existing technology. The specific working principle will not be elaborated here. The second cavity 11121 is designed to provide a closed space for the internal installation of the liquid pump 30. Simultaneously, the input end 301 of the liquid pump 30 needs to be connected to the second cavity 11121, and the output end 302 of the liquid pump 30 can be arranged coaxially with the rotating shaft 122 and the liquid guide tube 31. This facilitates the installation of the liquid pump 30 and the liquid guide tube 31, and ensures the original functionality of the power motor 10 as much as possible. With minimal structural changes, a circulating liquid cooling structure can be arranged within the power system 100. Furthermore, since the corresponding shroud 20 in the industry typically has a conical tip, and this tip is often located on the central axis of the rotor 12, the outlet end of the liquid guide pipe 31 can be precisely connected to the tip of the shroud 20. This allows the liquid-cooled heat exchange chamber 21 to fill the entire exposed wall, facilitating the full flow of coolant through the liquid-cooled heat exchange chamber 21 and achieving higher heat exchange efficiency, thereby improving the cooling efficiency of the power motor 10.

[0027] It should also be noted that the fairing 20 can be fixed to the end face of the rotor 12 by bolts or welding, which can be selected according to actual needs. The liquid-cooled heat exchange chamber 21 undergoes forced convection heat exchange with the external airflow through the exposed wall of the fairing 20 to complete heat dissipation. In order to realize the connection between the liquid-cooled heat exchange chamber 21 and the first cavity 1111, a second flow channel 102 is provided through the rotor 12, so that the coolant after being cooled by the natural environment can flow from the liquid-cooled heat exchange chamber 21 into the first cavity 1111. When the coolant in the first cavity 1111 absorbs the heat dissipated by the iron core-winding assembly 112, it flows to the second cavity 11121 through the first flow channel 101.

[0028] The power system 100 provided in this embodiment integrates the liquid-cooled heat exchange chamber 21 into the exposed wall of the shroud 20. The first cavity 1111, provided by the support member 111, completely accommodates the core-winding assembly 112. The second cavity 11121 houses the liquid pump 30. A closed liquid-cooled circulation path is constructed within the power system 100 body through the first flow channel 101, the second flow channel 102, and the liquid guide pipe 31 penetrating the rotating shaft 122. On the one hand, the first cavity 1111 completely accommodates the core-winding assembly 112, and the second cavity 11121 temporarily stores and guides the coolant, allowing the coolant to fully absorb the large amount of heat generated by the high-power-density motor. The high-efficiency heat dissipation of the liquid-cooled heat exchange chamber 21 enables continuous cooling through internal circulation of the coolant, effectively preventing performance degradation of the power system 100 and ensuring flight safety and endurance of the aircraft. On the other hand, there is no need to add an independent radiator and exposed liquid delivery pipelines, which avoids the problem of excessive fuselage space occupied by the internal cooling structure or the damage to aerodynamic streamline caused by exposure, reducing flight drag and thus reducing flight time loss. Furthermore, through the integrated design of the first cavity 1111, the second cavity 11121 and the support component 111, the flow channel integration can be achieved without thickening the motor housing, avoiding weight redundancy, and reducing the additional sealing nodes caused by exposed pipelines, thus reducing the risk of coolant leakage.

[0029] Continue reading Figure 1 Optionally, the power system 100 also includes a propeller blade 40, which can be fixedly mounted on the end face of the rotor 12 by bolts. The propeller blade 40 rotates synchronously with the rotor 12 and the fairing 20. Meanwhile, the liquid guide pipe 31 passes through the connecting hole opened in the center of the propeller blade 40. The propeller blade 40 can be obtained from the prior art, and its specific structure will not be described in detail here.

[0030] Continue reading Figure 3 In some embodiments, the exposed wall of the shroud 20 is a hollow sandwich structure, which forms a liquid-cooled heat exchange cavity 21. The liquid-cooled heat exchange cavity 21 is continuously distributed along the circumference and axial direction of the exposed wall of the shroud 20 and is located within the thickness range of the exposed wall. This arrangement ensures the maximization of the heat exchange area without destroying the streamlined structure of the shroud 20. At the same time, in order to improve the stability of the connection between the shroud 20 and the rotor 12, a plurality of reinforcing ribs 23 are provided at intervals along the circumference of the inner wall of the shroud 20. The reinforcing ribs 23 are in close contact with the end face of the rotor 12. A flow guide hole 231 is provided inside each reinforcing rib 23. The flow guide hole 231 communicates with the through hole opened on the rotor 12 to form a second flow channel 102.

[0031] With this configuration, forced convection heat exchange is achieved between the exposed wall of the fairing 20 and the external airflow, allowing the coolant in the liquid-cooled heat exchange chamber 21 to dissipate heat. Multiple second flow channels 102 are formed around the fairing 20 and the rotor 12, and these channels are spaced apart along the circumference. At the same time, both ends are connected to the liquid-cooled heat exchange chamber 21 and the first cavity, respectively, so that the multiple second flow channels 102 form a flow divider. As a result, coolant continuously flows vertically downwards along the circumference in the first cavity, avoiding uneven cooling in the first cavity.

[0032] Continue reading Figure 4 In some embodiments, the first cavity 1111 is an annular cavity and is arranged around the axis of the support member 111. The support member 111 has a bearing chamber 1112 with one end open along its central axis. The rotating shaft 122 extends coaxially from the open end into the bearing chamber 1112 through the bearing sleeve 121 to form a sealed second cavity 11121 at the bottom end of the bearing chamber 1112. Specifically, the bearing chamber 1112 is a cylindrical cavity, and one end of the bearing chamber 1112 is open, while the other end is a closed bottom structure. The rotating shaft 122 passes coaxially through the bearing sleeve 121, which is interference-fitted into the inner wall of the open end of the bearing chamber 1112. After extending from the open end, 122 extends along the axis of the bearing chamber 1112 to the bottom end. It is enclosed by the bearing sleeve 121 and the bottom end face of the bearing chamber 1112 to form a sealed second cavity 11121. At the same time, the bearing sleeve 121 and the bearing chamber 1112 are fixedly connected by bolts. A bearing 1211 is provided in the bearing sleeve 121. In this embodiment, two bearings 1211 are provided and cooperate with the rotating shaft 122, thereby realizing the rotational connection between the rotating shaft 122 and the support member 111. With this arrangement, the second cavity 11121 is located at the bottom end of the bearing chamber 1112, which is convenient for processing and for the installation of the liquid pump 30, the liquid guide pipe 31 and the rotating shaft 122.

[0033] Please see Figure 3 and Figure 5In some embodiments, an axially arranged air inlet duct 1114 is provided between the bearing chamber 1112 and the first cavity 1111. The air inlet duct 1114 contains several reinforcing plates 1113 arranged radially along the support member 111. The two ends of the reinforcing plates 1113 extend to the bearing chamber 1112 and the first cavity 1111, respectively. A first flow channel 101 is provided within the reinforcing plates 1113. Specifically, to further optimize structural integration and air-cooling effect, an air inlet duct 1114 arranged axially along the support member 111 is also provided between the bearing chamber 1112 and the first cavity 1111. The structural components 111 are evenly spaced around the periphery and are equipped with heat dissipation features, forming multiple parallel air-cooling channels. Several reinforcing plates 1113 are arranged radially along the support components 111 inside the air inlet duct 1114. The two ends of the reinforcing plates 1113 are fixedly connected to the outer wall of the bearing chamber 1112 and the inner wall of the first cavity 1111, respectively. This not only enhances the overall structural rigidity of the support components 111 and resists the centrifugal force and vibration generated when the motor rotates at high speed, but also reserves a channel inside the reinforcing plates 1113 to form the first flow channel 101, thus achieving a combination of structural reinforcement and liquid cooling flow guidance without the need for additional flow channels to occupy space.

[0034] Continue reading Figure 3 and Figure 5 Optionally, multiple heat dissipation fins 1115 are also arranged around the support member 111 within the air inlet duct 1114. Each heat dissipation fin 1115 is arranged radially along the support member 111. When the power motor 10 is running, part of the heat conducted from the iron core-winding assembly 112 to the support member 111 is carried away by liquid cooling, while the other part is transferred to the heat dissipation fins 1115. The external airflow enters the air inlet duct 1114 under the negative pressure of the rotating propeller blade 40, and fully contacts the heat dissipation fins 1115 to complete heat exchange. Then, it is discharged from the air outlet duct 1231, forming a highly efficient air-cooled heat dissipation.

[0035] Please see Figure 3 and Figure 6 In some embodiments, the rotor 12 includes a front end cover 123 coaxially disposed on the end face of the support member 111. The bottom surface of the front end cover 123 is provided with an air outlet duct 1231 arranged radially and communicating with the air inlet duct 1114. The air outlet duct 1231 is arranged circumferentially around the front end cover 123, and a plurality of partitions 1232 arranged radially along the support member 111 are provided inside the air outlet duct 1231. A portion of the second flow channel 102 is opened in the partitions 1232 and extends to communicate with the first cavity 1111. Specifically, the air outlet ducts 1231 are evenly spaced along the circumference of the front end cover 123, and the inner end of each air outlet duct 1231 is connected to the air outlet of the air inlet duct 1114, while the outer end extends to the outer periphery of the front end cover 123, forming a complete air-cooling circulation path from the air inlet duct 1114 to the air outlet duct 1231. Meanwhile, the air outlet duct 1231 is provided with several baffles 1232 arranged radially along the support member 111. The baffles 1232 can not only enhance the structural strength of the front cover 123, but also open a part of the second flow channel 102 in the baffles 1232, so that the second flow channel 102 can extend from the first cavity 1111 to the outer periphery of the front cover 123, and finally connect with the liquid-cooled heat exchange cavity 21 of the rectifier 20, realizing the integrated integration of the liquid-cooled flow channel and the rotor 12 structure, avoiding the setting of additional pipelines.

[0036] It should be explained that, in this embodiment, through the arrangement of the air inlet duct 1114 and the air outlet duct 1231, when the aircraft is running, the high-speed rotation of the propeller blades 40 will create a negative pressure field in the root region. This negative pressure field can actively guide the airflow from the external environment into the air inlet duct 1114 along the axial direction of the support member 111. The air inlet duct 1114 is evenly distributed around the circumference of the support member 111. The airflow entering the duct will fully contact the multiple heat dissipation fins 1115 that are pre-set in the air inlet duct 1114 and extend radially along the support member 111. The heat dissipation fins 1115 and the outer wall of the first cavity 1111 are in full contact. The close fit allows for rapid heat transfer from the core-winding assembly 112 to the support member 111 during operation, enabling the airflow to efficiently carry away heat as it flows through the fin gaps. Subsequently, the airflow that has completed heat exchange continues to flow along the inlet air duct 1114 and is finally discharged through the outlet air duct 1231, which is opened on the bottom surface of the rotor front cover 123 and extends radially. The outlet air duct 1231 is arranged circumferentially around the front cover 123, with the outlet facing the outer periphery of the front cover 123, to avoid interference between the discharged airflow and the rotating airflow of the propeller blades 40, thus forming a passive air-cooled heat dissipation system with stable heat exchange efficiency without the need for additional power drive. Meanwhile, the air-cooling system works synergistically with the liquid-cooling structure integrated inside the power system 100. The liquid-cooling structure is mainly used to precisely absorb and efficiently dissipate heat from core heat sources such as the iron core-winding assembly 112, while the air-cooling structure helps to remove heat from the support component 111. Together, they form a composite heat dissipation system that enhances heat dissipation from the core heat source through liquid cooling and supplements heat dissipation through air cooling in the auxiliary area. This ensures the heat dissipation requirements of the high-power-density motor without requiring additional components such as air-cooled fans or external radiators outside the power system 100, thus avoiding any aerodynamic changes or additional wind resistance to the power system 100. This achieves a design that enables efficient heat dissipation and low-drag lightweight design for the aircraft power system 100. Please see Figure 7Optionally, for the liquid pump 30, its input end 301 port can be located on its top surface and communicate with the second cavity 11121, and the output end 302 port is arranged along the central axis of the rotating shaft 122. One end of the output end 302 port is provided with a connecting nozzle that can extend coaxially to one end of the liquid guide tube 31. At the same time, the output end 302 port and the liquid guide tube 31 form a rotatable connection. Since the liquid guide tube 31 rotates synchronously with the rotating shaft 122, and the liquid pump 30 is fixedly connected to the second cavity 11121, it is ensured that the liquid guide tube 31 can still deliver coolant during rotation.

[0037] Please see Figure 8 and Figure 9 In some embodiments, the rotating shaft 122 is provided with a first through hole 1221 and a second through hole 1222 sequentially along the axial direction. The connection between the first through hole 1221 and the second through hole 1222 forms a stepped portion 1223. The outer wall of the liquid guide tube 31 is provided with a sealing ring 312 adapted to the stepped portion 1223. It should be noted that the liquid guide tube 31 can be a steel shaft tube. To ensure the sealing performance and coaxiality between the liquid guide tube 31 and the rotating shaft 122, the rotating shaft 122 is provided with a first through hole 1221 and a second through hole 1222 sequentially along the axial direction. The inner diameter of the first through hole 1221 is larger than that of the second through hole 1222, and the connection between the two forms an annular trapezoidal portion 1223. A sealing ring 312 is fitted on the outer wall of the liquid guide tube 31 at the position corresponding to the trapezoidal portion 1223. The sealing ring 312 can be made of rubber. One side of it is tightly fitted with the trapezoidal portion 1223, and the other side is interference-fitted with the outer wall of the liquid guide tube 31. It can effectively prevent coolant from leaking from the gap between the liquid guide tube 31 and the rotating shaft 122, and at the same time play a radial positioning role.

[0038] Continue reading Figure 8 Optionally, one end of the rotating shaft 122 extends out of the bearing sleeve 121 and is connected to a fastening nut 103. The fastening nut 103 completely fixes the rotating shaft 122 and the bearing sleeve 121 in the vertical direction. That is, the rotating shaft 122 and the bearing sleeve 121 can only rotate in the axial direction. A sealing ring can be provided at the bottom of the bearing sleeve 121 and at the connection between the bearing sleeve 121 and the rotating shaft 122 to ensure the sealing of the second cavity 11121.

[0039] Furthermore, the outer wall of the liquid guide tube 31 near the sealing ring 312 protrudes outward to form a positioning part 311. The outer diameter of the positioning part 311 is larger than the inner diameter of the first through hole 1221, and an annular limiting ring 3111 extends from the end of the positioning part 311 toward the sealing ring 312. The limiting ring 3111 is embedded in the first through hole 1221 and fits against the outer wall of the liquid guide tube 31. This not only restricts the axial displacement of the liquid guide tube 31 and prevents it from moving under the pressure of the liquid pump 30 and the centrifugal force of rotation, but also further enhances the installation stability of the sealing ring 312 and reduces the risk of seal failure.

[0040] Please see Figure 5 and Figure 9 In some embodiments, a guide plate 313 is provided at one end of the liquid guide tube 31 away from the positioning part 311, and the guide plate 313 is provided with a plurality of radially arranged liquid distribution holes 3131; and an annular groove 22 adapted to the guide plate 313 is provided at the top of the rectifier 20, wherein the outer periphery of the guide plate 313 is at least partially embedded in the annular groove 22 and is interference-fitted with the groove wall of the annular groove 22, and a plurality of circumferentially spaced liquid inlet holes 221 are provided on the inner side wall of the annular groove 22, and each liquid inlet hole 221 is connected to the liquid distribution hole 3131 and the liquid cooling heat exchange chamber 21 respectively.

[0041] Specifically, a guide plate 313 is provided at one end of the liquid guide tube 31 away from the positioning part 311. The guide plate 313 is disc-shaped, and its diameter is adapted to the space of the annular groove 22 provided inside the top of the shroud 20. Multiple liquid distribution holes 3131 are evenly distributed radially on the guide plate 313. The inner end of each liquid distribution hole 3131 is connected to the inner cavity of the liquid guide tube 31, and the outer end extends to the outer periphery of the guide plate 313. Correspondingly, an annular groove 22 adapted to the guide plate 313 is provided on the inner wall surface of the top of the shroud 20. The outer periphery of the guide plate 313 is at least partially embedded in the annular groove 22, and is interference-fitted with the groove wall of the annular groove 22, thereby facilitating the liquid guide tube 313. 1. It is manufactured as an independent component. At the same time, the synchronous rotation of the guide plate 313 and the rectifier 20 ensures the sealing of the connection part. Multiple liquid inlet holes 221 are provided on the inner side wall of the annular groove 22. One end of each liquid inlet hole 221 is connected to the outer end of at least one liquid distribution hole 3131, and the other end is connected to the liquid cooling heat exchange chamber 21. This allows the coolant transported by the liquid guide pipe 31 to be evenly distributed to each liquid inlet hole 221 through the liquid distribution hole 3131 of the guide plate 313, and then smoothly enter the liquid cooling heat exchange chamber 21. This avoids the accumulation of coolant in local areas, ensures that the coolant flow rate inside the liquid cooling heat exchange chamber 21 is uniform, and improves the overall heat exchange efficiency.

[0042] In some embodiments, an aircraft is also provided, the aircraft including the aircraft propulsion system with composite heat dissipation as described in any of the above embodiments.

[0043] In the description of this application, it should be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", "counterclockwise", "axial", "radial", "circumferential", etc., indicating the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation on the invention.

[0044] In the description of this specification, the references to terms such as "one embodiment," "some embodiments," "illustrative embodiment," "example," "specific example," or "some examples," etc., refer to specific features, structures, materials, or characteristics described in connection with that embodiment or example, which are included in at least one embodiment or example of this application. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example.

[0045] Obviously, the described embodiments are only a part of the embodiments of this application, and not all of the embodiments. The reference to "embodiment" herein means that a specific feature, structure, or characteristic described in connection with an embodiment can be included in at least one embodiment of this application. The appearance of this phrase in various places in the specification does not necessarily indicate the same embodiment, nor is it an independent or alternative embodiment mutually exclusive with other embodiments. It will be explicitly and implicitly understood by those skilled in the art that the embodiments described herein can be combined with other embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of this application without inventive effort are within the scope of protection of this application.

[0046] Although embodiments of this application have been shown and described, those skilled in the art will understand that various changes, modifications, substitutions and alterations can be made to these embodiments without departing from the principles and spirit of this application, the scope of which is defined by the claims and their equivalents.

Claims

1. An aircraft propulsion system with composite heat dissipation, characterized in that, include: A power motor includes a stator and a rotor rotatably connected to the stator via a rotating shaft. The stator includes a support member and a core-winding assembly. The support member has a first cavity for accommodating the core-winding assembly and a second cavity located at the bottom of the rotating shaft. The first cavity communicates with the second cavity through a first flow channel provided in the support member. A shroud is provided on the end face of the rotor. A liquid-cooled heat exchange chamber is provided in the exposed wall of the shroud. The liquid-cooled heat exchange chamber is connected to the first cavity through a second flow channel that penetrates the rotor. The second cavity is equipped with a liquid pump, and the output end of the liquid pump is rotatably connected to a liquid guide pipe that is coaxial with the rotating shaft. One end of the liquid guide pipe passes through the rotating shaft and communicates with the liquid-cooled heat exchange cavity.

2. The aircraft propulsion system with composite heat dissipation according to claim 1, characterized in that, The exposed wall of the fairing is a hollow sandwich structure, which forms the liquid-cooled heat exchange cavity. The liquid-cooled heat exchange cavity is continuously distributed along the circumference and axial direction of the exposed wall of the fairing and is located within the thickness range of the exposed wall.

3. The aircraft propulsion system with composite heat dissipation according to claim 1, characterized in that, The first cavity is an annular cavity and is arranged around the axis of the support member. The support member is provided with a bearing chamber with one end open along the central axis. The rotating shaft extends coaxially from the open end into the bearing chamber through a bearing sleeve to form a sealed second cavity at the bottom end of the bearing chamber.

4. The aircraft propulsion system with composite heat dissipation according to claim 3, characterized in that, An axially arranged air inlet duct is provided between the bearing chamber and the first cavity. Several reinforcing plates are arranged radially along the support member in the air inlet duct. The two ends of the reinforcing plates extend to the bearing chamber and the first cavity, respectively. The first flow channel is provided in the reinforcing plates.

5. The aircraft propulsion system with composite heat dissipation according to claim 4, characterized in that, The rotor includes a front end cover coaxially disposed on the end face of the support member. The bottom surface of the front end cover is provided with an air outlet duct arranged radially and communicating with the air inlet duct. The air outlet duct is arranged circumferentially around the front end cover, and a plurality of partitions are arranged radially along the support member inside the air outlet duct. A portion of the second flow channel is opened in the partitions and extends to communicate with the first cavity.

6. The aircraft propulsion system with composite heat dissipation according to claim 4, characterized in that, Multiple heat dissipation fins are arranged circumferentially along the support member inside the air inlet duct, and each heat dissipation fin is arranged radially along the support member.

7. The aircraft propulsion system with composite heat dissipation according to claim 1, characterized in that, The rotating shaft is provided with a first through hole and a second through hole in sequence along the axial direction. The connection between the first through hole and the second through hole forms a stepped section. The outer wall of the liquid guide tube is provided with a sealing ring adapted to the stepped section.

8. The aircraft propulsion system with composite heat dissipation according to claim 7, characterized in that, The liquid guide tube is provided with a positioning part on the outer wall near the sealing ring. The outer diameter of the positioning part is larger than the inner diameter of the first through hole, and a limiting ring extending to the first through hole is provided at one end of the positioning part.

9. The aircraft propulsion system with composite heat dissipation according to claim 8, characterized in that, A flow guide plate is provided at the end of the liquid guide tube opposite to the positioning part, and the flow guide plate is provided with a plurality of radially arranged liquid distribution holes; and The top of the shroud is provided with an annular groove adapted to the guide plate. The outer periphery of the guide plate is at least partially embedded in the annular groove and is interference-fitted with the groove wall. The inner sidewall of the annular groove is provided with a plurality of liquid inlet holes spaced apart along the circumference. Each liquid inlet hole is connected to the liquid distribution hole and the liquid-cooled heat exchange chamber.

10. An aircraft, characterized in that, The aircraft includes an aircraft propulsion system with composite heat dissipation as described in any one of claims 1 to 9.

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