Aircraft electric heat dissipation system and control method

Abstract

The invention discloses an aircraft electric heat dissipation system and a control method, the electric heat dissipation system comprises a liquid cooling mechanism in a power arm, and the liquid cooling mechanism comprises a radiator main body; an upper cover plate is arranged at the top of the heat dissipation shell, a lower cover plate is arranged at the bottom of the power arm, an air inlet corresponding to the upper cover plate is formed in the top of the power arm, an air outlet is formed in the bottom of the power arm, and the lower cover plate is located in the air outlet; the air inlet is located in the orthographic projection range of the rotation track of the aircraft blades, when the aircraft blades rotate, the upper cover plate and the lower cover plate are opened, and the heat dissipation air channel of the heat dissipation shell communicates with the external environment based on the air inlet and the air outlet. The liquid cooling mechanism and the air cooling mechanism are arranged in the power arm, heat dissipation is conducted based on combination of liquid cooling and air cooling, the heat dissipation efficiency is improved, the opening degree of the air inlet cover and the opening degree of the air outlet cover are dynamically adjusted, heat dissipation adjustment of a lift system is achieved, and the heat dissipation efficiency of the aircraft is improved.

Description

Technical Field

[0001] This invention relates to the field of aircraft technology, and more specifically to an electric-driven heat dissipation system and control method for aircraft. Background Technology

[0002] eVTOL (Electric Vertical Takeoff and Landing) is a new type of electric aircraft capable of vertical takeoff and landing. In compound wing aircraft, lift is generally provided by a lift system during the vertical takeoff and landing and transition phases. The lift system provides lift to the aircraft by driving the rotor blades of the power arm through the motor. Because the drive motor of the lift system has a high operating power when the aircraft is performing vertical takeoff, landing and hovering, it is easy to generate a lot of heat, and it is necessary to cool the drive motor of the lift system in a timely manner.

[0003] Patent CN202210128285.1 discloses a heat dissipation structure and an electric vertical takeoff and landing (eVTOL) aircraft. It employs air cooling to cool the motor controller, with vents on the aircraft fuselage and a ventilation assembly including a vent cover connected to the fuselage. When the vent cover is open, airflow generated by the lift propeller is allowed to flow through the motor controller and out of the fuselage to cool it. However, the ventilation assembly disrupts the aircraft's aerodynamic shape, increasing drag. Furthermore, air cooling has lower heat dissipation capacity than liquid cooling, making it unsuitable for meeting the high-power cooling requirements.

[0004] The patent text with patent number CN202310124038.9 discloses an eVTOL aircraft drive motor cooling system that uses liquid cooling to cool the electric drive. The outer surface of the heat sink has a curved structure, which reduces the drag of the aircraft during cruise flight. However, the contact area between the heat sink and the air is smaller than that of a conventional radiator, and the heat dissipation capacity of the heat sink is weaker than that of a conventional radiator. Summary of the Invention

[0005] The purpose of this invention is to overcome the shortcomings of the prior art. This invention provides an electric drive heat dissipation system and control method for an aircraft. Liquid cooling and air cooling mechanisms are set inside the power arm. With the adjustment of the opening of the air inlet and air outlet on the power arm, an air cooling heat dissipation channel is formed, which improves the heat dissipation efficiency of the drive motor of the lift system, while avoiding interference of the heat dissipation system with the flight of the aircraft.

[0006] This invention provides an electric drive cooling system for an aircraft, the electric drive cooling system comprising: a liquid cooling mechanism disposed inside the power arm of the aircraft, the power arm having a heat dissipation shell disposed inside the heat dissipation shell, the liquid cooling mechanism comprising a heat sink body disposed inside the heat dissipation shell;

[0007] The heat dissipation housing is provided with an upper cover plate at the top, the power arm is provided with a lower cover plate at the bottom, the power arm is provided with an air inlet corresponding to the upper cover plate at the top, the power arm is provided with an air outlet at the bottom, and the lower cover plate is located inside the air outlet.

[0008] The air inlet is located within the orthographic projection range of the aircraft propeller rotation trajectory. When the aircraft propeller rotates, the upper cover and the lower cover are opened, and the heat dissipation duct of the heat dissipation shell is connected to the external environment based on the air inlet and the air outlet.

[0009] Furthermore, the liquid cooling mechanism includes a drive pump disposed at the lower end of the drive motor, a cooling tank disposed at one end of the radiator body, and a first cooling pipe and a second cooling pipe disposed between the drive pump and the cooling tank.

[0010] The first cooling pipe and the second cooling pipe are filled with coolant.

[0011] Furthermore, the radiator body includes a plurality of heat dissipation fins arranged in an array, and one end of any one of the heat dissipation fins is located in the cooling groove.

[0012] Furthermore, the cross-sectional area of ​​the heat dissipation housing is the same as the cross-sectional area of ​​the heat sink body.

[0013] Furthermore, the upper cover plate is mounted on the heat dissipation housing based on a hinge mechanism, and the lower cover plate is mounted on the power arm based on a hinge mechanism;

[0014] The upper cover is configured as a single cover structure that opens along one side, or the upper cover includes a symmetrically distributed first sub-cover and second sub-cover, the first sub-cover and the second sub-cover forming a double door structure.

[0015] The lower cover is configured as a single cover structure that opens along one side, or the lower cover includes a symmetrically distributed third sub-cover and a fourth sub-cover, the third sub-cover and the fourth sub-cover forming a double-door structure.

[0016] Furthermore, the opening direction of the upper cover is towards the heat dissipation housing, and the opening direction of the lower cover is towards the bottom outer side of the power arm. The upper cover and the lower cover are located within the orthogonal projection area of ​​the propeller rotation trajectory of the aircraft, and the flip-opening is achieved based on the downward airflow generated by the rotation of the propeller of the aircraft.

[0017] Furthermore, the aircraft is equipped with a first servo and a second servo, the first servo being driven and connected to the upper cover plate, and the second servo being driven and connected to the lower cover plate.

[0018] This invention provides a method for cooling an aircraft's electric drive system. The method is applicable to the aforementioned aircraft electric drive cooling system and includes:

[0019] S11: According to the flight mission of the aircraft, drive the upper and lower cover plates to flip and adjust the air inlet and air outlet to the fully open state;

[0020] S12: Detect the opening and closing control mode of the upper cover and the lower cover. If it is a passive control mode, proceed to step S13; if it is an active mode, proceed to step S14.

[0021] S13: By adjusting the output speed of the drive motor, the folding and opening range of the upper and lower cover plates can be adjusted in a coordinated manner.

[0022] S14: Detect the real-time temperature of the drive motor, and adjust the folding opening range of the upper and lower cover plates according to the real-time temperature.

[0023] Furthermore, adjusting the folding opening range of the upper and lower cover plates according to the output speed of the drive motor includes:

[0024] Based on the flight mission of the aircraft, the output power of the aircraft's drive motor is adjusted, and the opening range of the upper and lower covers is adjusted according to the rotation speed of the propeller blades driven by the drive motor.

[0025] Furthermore, the method of detecting the real-time temperature of the drive motor and adjusting the folding opening range of the upper and lower cover plates based on the real-time temperature includes:

[0026] The system detects when the drive motor is in a non-working state and acquires the real-time temperature of the drive motor.

[0027] The real-time temperature is compared with a preset safe temperature threshold. If the real-time temperature is higher than the safe temperature threshold, the opening range of the upper and lower cover plates is reduced proportionally.

[0028] If the real-time temperature is lower than the safe temperature threshold, then the upper cover and the lower cover should be completely closed.

[0029] This invention provides an electric drive cooling system and control method for an aircraft. By incorporating liquid cooling and air cooling mechanisms within the power arm, the drive motor and cooling system of the lift system are integrated inside the power arm, avoiding interference between the drive motor and cooling system and the aircraft's external dimensions. Liquid cooling is used to dissipate heat from the drive motor of the lift system, improving its cooling efficiency. Furthermore, during flight, the opening sizes of the air inlet and outlet covers are dynamically adjusted to regulate the heat dissipation of the lift system, further enhancing the aircraft's cooling efficiency. Attached Figure Description

[0030] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0031] Figure 1 This is a schematic diagram of the structure of the aircraft electric drive cooling system in an embodiment of the present invention;

[0032] Figure 2 This is a schematic diagram of the cooling air duct of the aircraft electric drive cooling system in the embodiment of the present invention in the open state;

[0033] Figure 3 This is a schematic diagram of the cooling air duct of the aircraft electric drive cooling system in the closed state in an embodiment of the present invention;

[0034] Figure 4 This is a schematic diagram of the passive control electric drive cooling system of the aircraft in this embodiment of the invention;

[0035] Figure 5 This is a schematic diagram of the active control electric drive cooling system of the aircraft in this embodiment of the invention;

[0036] Figure 6 This is a schematic diagram of the working circuit of the active control electric drive cooling system in an embodiment of the present invention;

[0037] Figure 7 This is a flowchart of the control method for the aircraft electric drive cooling system in an embodiment of the present invention. Detailed Implementation

[0038] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0039] Example 1:

[0040] Please refer to Figures 1 to 4This invention provides an electric drive cooling system for an aircraft. The electric drive cooling system includes: a liquid cooling mechanism disposed inside a power arm 1 of the aircraft; a heat dissipation shell 6 disposed inside the power arm 1; a heat dissipation body 7 disposed inside the heat dissipation shell 6; and a drive motor 2 of the aircraft disposed inside the power arm 1. The drive motor 2 drives and connects to the propeller blades of the aircraft. The propeller blades of the aircraft rotate based on the drive motor 2 to meet the vertical take-off and landing and hovering flight actions of the aircraft.

[0041] The heat dissipation housing 6 is provided with an upper cover plate 41 at the top, and the power arm 1 is provided with a lower cover plate 51 at the bottom. The power arm 1 is provided with an air inlet 4 at the top corresponding to the upper cover plate 41, and an air outlet 5 at the bottom. The lower cover plate 51 is located inside the air outlet 5, that is, the bottom end of the heat dissipation housing 6 is fixed to the bottom of the power arm 1. When the lower cover plate 51 is in the closed state, the outer contour of the power arm 1 can maintain a smooth curved transition, so that the aircraft can maintain a smooth flight state.

[0042] Furthermore, at least one drive motor 2 is provided inside the power arm 1. The output shaft of the drive motor 2 is connected to the propeller 3 of the aircraft. The drive motor 2 drives the propeller 3 of the aircraft to rotate. The rotation of the propeller blades of the propeller 3 can form a downward airflow at the top of the power arm 1, so as to cooperate with the liquid cooling mechanism inside the power arm 1 for heat dissipation.

[0043] The air inlet 4 is located within the orthographic projection range of the aircraft propeller rotation trajectory. When the aircraft propeller rotates, the upper cover plate 41 and the lower cover plate 51 are opened, and the heat dissipation duct of the heat dissipation shell 6 is connected to the external environment based on the air inlet 4 and the air outlet 5. The liquid cooling mechanism dissipates heat from the aircraft's drive motor 2 based on the heat dissipator body 7. By setting the upper cover plate 41 and the lower cover plate 51 to the open state, the heat dissipation shell 6 is connected to the external environment, thereby forming a heat dissipation duct. In conjunction with the airflow generated by the rotation of the aircraft propeller, the airflow of the external environment flows through the heat dissipation duct formed by the heat dissipation shell 6, so as to realize the heat exchange between the heat dissipator body 7 and the external environment airflow.

[0044] Specifically, the liquid cooling mechanism includes a drive pump disposed inside the power arm 1, a cooling tank disposed at one end of the radiator body 7, and a first cooling pipe 71 and a second cooling pipe 72 disposed between the drive pump and the cooling tank. Coolant is disposed in the first cooling pipe 71 and the second cooling pipe 72. The drive pump is disposed at the lower end of the drive motor 2. The coolant in the first cooling pipe 71 flows along the radiator body 7 towards the drive motor 2, and the coolant in the second cooling pipe 72 flows along the drive motor 2 towards the radiator body 7. This allows the radiator body 7 and the drive motor 2 to form a liquid cooling circulation for heat dissipation based on the coolant flow in the first cooling pipe 71 and the second cooling pipe 72, thereby achieving efficient heat absorption of the drive motor 2, i.e., efficient heat dissipation and cooling of the drive motor 2.

[0045] Furthermore, when the drive motor 2 is working, the drive pump can work synchronously. Based on the flow of coolant in the first cooling pipe 71 and the second cooling pipe 72 driven by the drive pump, the coolant in the first cooling pipe 71 flows along the first cooling pipe 71 towards the drive pump position. That is, the coolant in the first cooling pipe 71 can flow through the lower end of the drive motor 2, so that the coolant can exchange heat with the drive motor 2. Based on the fact that the coolant can meet the heat exchange requirements of the drive motor 2, the drive motor 2 can be driven.

[0046] Specifically, the radiator body 7 includes a plurality of heat dissipation fins arranged in an array, and one end of any heat dissipation fin is located in the cooling tank. When the coolant in the first cooling pipe 71 and the second cooling pipe 72 flows through the cooling tank, the coolant can exchange heat with the heat dissipation fins. By exchanging heat with the coolant through the heat dissipation fins, the efficiency of heat exchange by the coolant can be improved.

[0047] Furthermore, when the drive motor 2 is working, the drive pump can work synchronously. Based on the drive pump driving the coolant in the first cooling pipe 71 and the second cooling pipe 72 to circulate, the reciprocating motion of the coolant between the drive motor 2 and the radiator body 7 can achieve liquid cooling. That is, when the coolant flows through the drive motor 2, it absorbs part of the heat generated by the drive motor 2 through heat exchange, and enhances the heat dissipation efficiency of the coolant through the heat dissipation fins. This ensures that the liquid cooling mechanism, located inside the power arm 1, has sufficient heat dissipation efficiency to meet the working requirements of the drive motor 2 of the propeller 3 of the aircraft.

[0048] The heat dissipation fins are configured as thin sheet structures with regular geometric shapes, their length direction perpendicular to the extension direction of the cooling groove, for example, using rectangular or trapezoidal cross-sections, with a thickness controlled within the range of 0.5-1.2 mm. The array arrangement is implemented as an equally spaced staggered arrangement, with the spacing between adjacent heat dissipation fins configured at 2-5 mm to form parallel flow channels. The connection end between the heat dissipation fins and the cooling groove is machined into a wedge-shaped structure, embedded into the sidewall of the cooling groove by interference fit or welding, with an embedding depth reaching 15%-30% of the total length of the heat dissipation fins. Furthermore, the extension portion of the heat dissipation fins outside the cooling groove is arranged to contact the housing of the drive motor 2, with the contact area accounting for 40%-60% of the total surface area of ​​the heat dissipation fins.

[0049] Specifically, when the coolant flows through the cooling tank, heat is conducted to the embedded heat dissipation fins through the sidewalls of the cooling tank. Because the heat dissipation fins are arranged in an array, the coolant flow path is divided into multiple parallel branches, increasing the flow velocity to 0.8-1.5 m / s, and the turbulence effect enhances heat transfer efficiency. The area of ​​the heat dissipation fins washed by the coolant increases to 1.8-2.5 times that of traditional structures. The ends embedded in the cooling tank directly absorb heat from the liquid cooling circuit, while the portions extending outside the cooling tank undergo secondary heat dissipation through air convection. Through the matching design of the heat dissipation fin spacing and thickness, the overall thermal resistance of the radiator body 7 is reduced to 0.05-0.08℃ / W, and the heat conduction efficiency is improved by 35%-50% compared to radiators without arrayed fins. Thus, the heat generated by the drive motor 2 is simultaneously dissipated through both forced convection in the liquid cooling circuit and natural air convection, solving the problem of heat accumulation in a single heat dissipation mode.

[0050] Specifically, the cross-sectional area of ​​the heat dissipation shell 6 is the same as that of the heat sink body 7. The heat dissipation shell 6 is set as a rectangular structure. Based on the rectangular structure, the heat dissipation channel of the heat dissipation shell 6 is restricted. Based on the downward airflow generated by the rotation of the propeller 3 of the aircraft, it can be ensured that the airflow generated by the rotation of the propeller 3 has sufficient flow rate and airflow volume to meet the heat exchange requirements of the heat dissipation fins of the heat sink body 7.

[0051] Specifically, this embodiment of the invention proposes to integrate the liquid cooling mechanism inside the power arm 1. Through the linkage design of the upper and lower cover plates 51 and the airflow of the propeller blades, the airflow generated by the rotation of the propeller blades is used to passively open the heat dissipation duct during the vertical take-off and landing phase. Combining the dual heat dissipation mechanism of liquid cooling and air cooling, the aerodynamic shape of the aircraft is maintained, and efficient heat dissipation under high power conditions is achieved.

[0052] The working process and principle of this application are as follows: the aircraft's electric drive cooling system includes a liquid cooling mechanism installed inside the aircraft's power arm 1. A heat dissipation shell 6 is installed inside the power arm 1, and the liquid cooling mechanism includes a radiator body 7 installed within the heat dissipation shell 6. An upper cover plate 41 is installed at the top of the heat dissipation shell 6, and a lower cover plate 51 is installed at the bottom. An air inlet 4 corresponding to the upper cover plate 41 is installed at the top of the power arm 1, and an air outlet 5 is installed at the bottom. The lower cover plate 51 is located within the air outlet 5.

[0053] The air inlet 4 is located within the orthographic projection range of the aircraft's propeller rotation trajectory. When the propeller rotates, the upper cover 41 and lower cover 51 open, and the heat dissipation duct of the heat sink 6 connects to the external environment through the air inlet 4 and the air outlet 5. The downward airflow generated by the propeller rotation drives the upper cover 41 and lower cover 51 to open, forming a natural heat dissipation channel. The airflow enters the heat sink 6 from the air inlet 4, carries away heat through the radiator body 7, and is then discharged from the air outlet 5.

[0054] Specifically, the upper cover plate 41 is mounted on the heat dissipation housing 6 based on the hinge mechanism 8. The upper cover plate 41 can be configured as a single cover plate structure that opens on one side, and is opened by the airflow generated by the rotation of the propeller blades of the propeller 3.

[0055] Furthermore, the upper cover plate 41 can also be configured as a double-cover plate structure, including a symmetrically distributed first sub-cover plate and a second sub-cover plate. The first sub-cover plate and the second sub-cover plate form a double-door structure, with the opening direction of the first sub-cover plate and the second sub-cover plate facing the heat dissipation shell 6. The first sub-cover plate and the second sub-cover plate are located within the orthogonal projection area of ​​the aircraft's propeller rotation trajectory, and the door opening is achieved based on the downward airflow generated by the aircraft's propeller rotation. This design integrates the liquid cooling mechanism inside the power arm 1 and utilizes the aircraft body structure to form the heat dissipation shell 6, maintaining the integrity of the aerodynamic shape. The vertically connected air duct design utilizes the airflow generated by the propeller rotation to achieve passive heat dissipation. The air inlet 4 is arranged in the orthogonal projection area of ​​the propeller rotation trajectory, maximizing the use of the negative pressure effect generated by the propeller rotation to guide the airflow into the heat dissipation shell 6. The opening and closing states of the upper cover plate 41 and the lower cover plate 51 are dynamically correlated with the propeller rotation speed, achieving automatic matching between heat dissipation efficiency and motor operating state.

[0056] Specifically, the lower cover plate 51 can be configured as a single cover plate structure that opens along one side, or the lower cover plate 51 includes a symmetrically distributed third sub-cover plate and a fourth sub-cover plate, the third sub-cover plate and the fourth sub-cover plate forming a double door structure. The operating principle of the lower cover plate 51 is the same as that of the upper cover plate 41, and will not be described in detail here.

[0057] Furthermore, the opening direction of the upper cover plate 41 faces the heat dissipation housing, and the opening direction of the lower cover plate 51 faces the bottom outer side of the power arm 1. The upper cover plate 41 and the lower cover plate 51 are located within the orthogonal projection area of ​​the propeller rotation trajectory of the aircraft, and the flip-opening is achieved based on the downward airflow generated by the rotation of the propeller of the aircraft.

[0058] Furthermore, the hinge mechanism 8 can be configured as a torsion spring and a hinge. Based on the cooperation of the torsion spring and the hinge, the hinge mechanism can maintain a horizontally extended state, allowing the upper cover plate 41 or the lower cover plate 51 to remain closed. The lower cover plate 51 can rotate around the axis connected to the power arm 1, thereby switching between closed and open states. When closed, the lower cover plate 51 completely covers the air outlet 5, preventing gas from flowing into or out of the air outlet 5. At the same time, together with the power arm 1, it forms a smooth aerodynamic shape, reducing flow resistance and improving the flight stability of the aircraft.

[0059] When drive motor 2 is working, it drives propeller 3 to rotate via its output shaft. Simultaneously, a pump within drive motor 2 drives coolant to circulate within drive motor 2, the first cooling pipe 71, the radiator body 7, and the second cooling pipe 72. The rotation of the lift propeller 3 generates a downward airflow. The upper cover plate 41 and lower cover plate 51 can be passively opened by the airflow generated by the lift propeller 3, allowing the airflow to pass through the air duct and radiator. The heat generated by drive motor 2 is transferred to the coolant, which then transfers heat to the air in the radiator, thus cooling drive motor 2.

[0060] Specifically, the top of the power arm 1 is provided with a plurality of air inlets 4, which are configured as circular openings. By forming the air inlets 4 with a plurality of circular openings, the overall structural mass distribution of the power arm 1 is made uniform, the size of the opening area of ​​the power arm 1 is reduced, and the stability of the power arm 1 structure is maintained.

[0061] Specifically, in this embodiment, when the propeller blades of the propeller 3 rotate at high speed, the downward airflow forms a local high-pressure zone within the projected area, and the airflow pressure acts on the outer surface of the cover plate. Since the opening direction of the cover plate faces the inside of the housing, the airflow pressure difference pushes the cover plate to flip inward into the housing, avoiding the increase in resistance caused by flipping outward. The correspondence between the rotational speed of the propeller blades of the propeller 3 and the upper cover plate 41 and the lower cover plate 51 is shown in the following table:

[0062] blade speed 1500rpm 2000rpm 2500rpm Top cover opening angle ＞65° ＞75° ＞85° Lower cover opening angle ＞60° ＞70° ＞80°

[0063] In summary, when the drive motor 2 drives the propeller 3 to rotate, and the propeller 3 rotates at 1500 rpm, the opening angle of the upper cover 41 is not less than 65 degrees, and the opening angle of the lower cover 51 is not less than 60 degrees; when the propeller 3 rotates at 2000 rpm, the opening angle of the upper cover 41 is not less than 75 degrees, and the opening angle of the lower cover 51 is not less than 70 degrees; when the propeller 3 rotates at 2500 rpm, the opening angle of the upper cover 41 is not less than 85 degrees, and the opening angle of the lower cover 51 is not less than 80 degrees. Furthermore, since the upper cover 41 is located below the blades of the propeller 3, the downward airflow generated when the blades of the propeller 3 rotate can directly act on the upper cover 41. The lower cover 51 opens based on the downward airflow flowing through the radiator body 7, resulting in an opening angle of the lower cover 51 that is smaller than that of the upper cover 41.

[0064] This design ensures that the downward airflow generated by the propeller blades directly impacts the cover surface, using the airflow pressure difference as the driving force to passively open the cover, eliminating the need for an additional drive mechanism and simplifying the system structure. Furthermore, the linkage design based on the downward airflow-driven door opening creates a positive correlation between the cover opening amplitude and the propeller rotation speed, enabling the cooling duct to adaptively adjust to cooling demands, minimizing the impact on flight performance while ensuring cooling efficiency. Finally, this design resolves the conflict between cover opening efficiency and aerodynamic performance, achieving a dynamic balance between cooling efficiency and flight drag, thus improving the overall performance of the aircraft.

[0065] Example 2:

[0066] Please refer to Figure 5 and Figure 6 This invention provides an electric-driven heat dissipation system for an aircraft. Specifically, in this embodiment, the aircraft is equipped with a first servo motor 9 and a second servo motor. The first servo motor 9 is driven and connected to the upper cover plate 41, and the second servo motor is driven and connected to the lower cover plate 51. The first servo motor 9 and the second servo motor are respectively installed on both sides inside the heat dissipation housing 6. The output shaft of the servo motor is fixedly connected to one end of a linkage mechanism, and the other end of the linkage mechanism is hinged to the cover plate. When it is necessary to open the cover plate, the servo motor drives the linkage mechanism to move, pushing the cover plate to rotate around the hinge point, thereby opening the cover plate. When it is necessary to close the cover plate, the servo motor moves in the opposite direction, pulling the linkage mechanism to return the cover plate to the closed position.

[0067] The control signals for the servo motors are sent by the aircraft's main controller. Based on parameters such as the aircraft's operating status, ambient temperature, and motor temperature, the main controller calculates the required opening angle of the cover and sends the corresponding control signals to the servo motors. The servo motors then precisely control the rotation angle of their output shafts according to the received control signals, thereby achieving precise adjustment of the cover opening angle.

[0068] Furthermore, to achieve more precise control, an angle sensor is installed on the cover to detect the actual opening angle of the cover in real time. The data from the angle sensor is fed back to the main controller, which compares the actual angle with the target angle and corrects for any deviation, thus forming a closed-loop control system.

[0069] In addition, to improve system reliability, an elastic buffer device is installed between the servo and the cover plate. When an external force obstructs the movement of the cover plate, the elastic buffer device can absorb part of the impact force, preventing damage to the servo or linkage mechanism.

[0070] Through the above technical solution, this application achieves active and precise control of the heat dissipation system cover. Therefore, heat dissipation efficiency can be dynamically matched with flight conditions, overcoming the high dependence on airflow inherent in passive control methods. Even at low speeds or when no propellers are rotating, the cover can be actively opened as needed to ensure effective heat dissipation. Simultaneously, by precisely controlling the opening and closing state of the cover, the impact on the aircraft's aerodynamic performance can be minimized while ensuring effective heat dissipation. Furthermore, the servo-driven method offers fast response speed, allowing for real-time adjustment of the cover's opening and closing angle according to flight conditions and heat dissipation requirements, thus improving the flexibility and adaptability of the heat dissipation system.

[0071] Example 3:

[0072] Please refer to Figure 7 This invention provides a control method for an electric drive cooling system, applicable to the electric drive cooling system of an aircraft, and the cooling method includes:

[0073] S11: According to the flight mission of the aircraft, drive the upper cover plate 41 and the lower cover plate 51 to flip, and adjust the air inlet 4 and the air outlet 5 to the fully open state.

[0074] Specifically, in the passive control mode, during the vertical take-off and landing phase of the aircraft, the propeller blades of the propeller 3 are driven to rotate by the drive motor 2, so that the downward airflow generated by the rotation of the propeller blades can drive the upper cover plate 41 and the lower cover plate 51 to open, thereby forming a heat dissipation duct in the power arm 1.

[0075] Furthermore, under the main control mode, before the aircraft takes flight, the upper cover 41 can be actively opened by the first servo motor 9, and the lower cover 51 can be actively opened by the second servo motor.

[0076] Furthermore, the aircraft can be configured to adjust the opening angles of the upper cover 41 and the lower cover 51 according to different flight missions, so that the heat dissipation duct can meet the flight requirements of different flight missions.

[0077] S12: Detect the opening and closing control mode of the upper cover plate 41 and the lower cover plate 51. If it is a passive control mode, proceed to step S13; if it is an active mode, proceed to step S14.

[0078] S13: By adjusting the output speed of the drive motor 2, the folding and opening range of the upper cover plate 41 and the lower cover plate 51 can be adjusted in a coordinated manner.

[0079] Specifically, in conjunction with the flight mission of the aircraft, the output power of the aircraft's drive motor 2 is adjusted, and based on the rotation speed of the propeller 3 driven by the drive motor 2, the folding opening range of the upper cover 41 and the lower cover 51 is adjusted.

[0080] Specifically, when the aircraft performs high-load flight missions, the actual power of the drive motor 2 and propeller 3 is close to 100% of the rated power. The wind pressure generated by the propeller 3 can drive the opening angle of the upper cover 41 and the lower cover 51 to be no less than 85 degrees, thereby enhancing the heat dissipation capacity of the radiator body 7. When the aircraft performs low-load flight missions, the actual power of the drive motor 2 and propeller 3 can be less than 80% of the rated power. The wind pressure generated by the propeller 3 is smaller, driving the opening angle of the upper cover 41 to be less than 75 degrees and the opening angle of the lower cover 51 to be less than 70 degrees, thus meeting the heat dissipation requirements of the drive motor 2 under lower loads.

[0081] S14: Detect the real-time temperature of the drive motor 2, and adjust the folding opening range of the upper cover plate 41 and the lower cover plate 51 according to the real-time temperature.

[0082] Specifically, the system detects that the drive motor 2 is in a non-working state and obtains the real-time temperature of the drive motor 2;

[0083] The real-time temperature is compared with a preset safe temperature threshold. If the real-time temperature is higher than the safe temperature threshold, the folding opening range of the upper cover plate 41 and the lower cover plate 51 is reduced proportionally.

[0084] If the real-time temperature is lower than the safe temperature threshold, then the upper cover plate 41 and the lower cover plate 51 are completely closed.

[0085] Specifically, before the aircraft performs a flight mission, the work tasks that the drive motor 2 needs to perform are obtained according to the flight mission of the aircraft. Before the drive motor starts working, the upper cover plate 41 and the lower cover plate 51 are rotated 90 degrees from the initial angle by the first servo motor 9 and the second servo motor to fully open the upper cover plate 41 and the lower cover plate 51, so as to avoid the airflow generated when the drive motor 2 starts causing the servo motor torque to be too large.

[0086] After the electric drive stops working, based on the real-time temperature and temperature change rate of the electric drive, gradually reduce the opening angle of the air inlet 4 cover and the air outlet 5 cover:

[0087] Furthermore, when the drive motor 2 is working, the first servo motor 9 keeps the upper cover plate 41 in a fully open state, and the second servo motor controls the lower cover plate 51 to be in a fully open state. After the drive motor 2 stops working, if the electric drive temperature is still higher than the set safe range, and the electric drive temperature change rate is negative, that is, the real-time temperature of the drive motor 2 is gradually decreasing, then the angle of the upper cover plate 41 and the lower cover plate 51 is reduced. The angle is reduced by 5 degrees to 30 degrees each time, until the electric drive temperature is within the safe range and no longer rises. At this point, the upper cover plate 41 and the lower cover plate 51 can be in a fully closed state.

[0088] When the temperature of the drive motor 2 is within the set safe range, and the real-time temperature of the drive motor 2 does not rise again within 10 seconds, the first servo motor 9 and the second servo motor are controlled to return to the initial angle, and the upper cover plate 41 and the lower cover plate 51 are completely closed, so as to wait for the heat dissipation control adjustment to perform the next flight mission.

[0089] Intelligent management of residual heat after drive motor 2 stops is achieved. By dynamically matching the opening degree of the cover with the real-time temperature, it prevents accelerated aging of motor components caused by insufficient passive heat dissipation under high temperature conditions, and avoids additional aerodynamic drag caused by the continuous opening of cover 51 under low temperature conditions. This solution effectively balances heat dissipation efficiency and aircraft cruise performance through a closed-loop control mechanism, and solves the control blind spot problem of traditional heat dissipation systems when the motor is not working.

[0090] Furthermore, the above description provides a detailed overview of the electric drive cooling system and control method for an aircraft provided by the embodiments of the present invention. Specific examples have been used to illustrate the principles and implementation methods of the present invention. The description of the above embodiments is only for the purpose of helping to understand the method and core ideas of the present invention. At the same time, for those skilled in the art, there will be changes in the specific implementation methods and application scope based on the ideas of the present invention. Therefore, the content of this specification should not be construed as a limitation of the present invention.

Claims

1. An electric-driven heat dissipation system for aircraft, characterized in that, The electric drive cooling system includes: a liquid cooling mechanism disposed inside the aircraft power arm, wherein a heat dissipation shell is disposed inside the power arm, and the liquid cooling mechanism includes a heat sink body disposed inside the heat dissipation shell; The heat dissipation housing is provided with an upper cover plate at the top, the power arm is provided with a lower cover plate at the bottom, the power arm is provided with an air inlet corresponding to the upper cover plate at the top, the power arm is provided with an air outlet at the bottom, and the lower cover plate is located inside the air outlet. The air inlet is located within the orthographic projection range of the aircraft propeller rotation trajectory. When the aircraft propeller rotates, the upper cover and the lower cover are opened, and the heat dissipation duct of the heat dissipation shell is connected to the external environment based on the air inlet and the air outlet.

2. The aircraft electric drive cooling system as described in claim 1, characterized in that, The liquid cooling mechanism includes a drive pump disposed at the lower end of the drive motor, a cooling tank disposed at one end of the radiator body, and a first cooling pipe and a second cooling pipe disposed between the drive pump and the cooling tank. The first cooling pipe and the second cooling pipe are filled with coolant.

3. The aircraft electric drive cooling system as described in claim 2, characterized in that, The radiator body includes a plurality of heat dissipation fins arranged in an array, and one end of any one of the heat dissipation fins is located in the cooling groove.

4. The aircraft electric drive cooling system as described in claim 1, characterized in that, The cross-sectional area of ​​the heat dissipation housing is the same as the cross-sectional area of ​​the heat sink body.

5. The aircraft electric drive cooling system as described in claim 1, characterized in that, The upper cover plate is mounted on the heat dissipation housing based on a hinge mechanism, and the lower cover plate is mounted on the power arm based on a hinge mechanism; The upper cover is configured as a single cover structure that opens along one side, or the upper cover includes a symmetrically distributed first sub-cover and second sub-cover, the first sub-cover and the second sub-cover forming a double door structure. The lower cover is configured as a single cover structure that opens along one side, or the lower cover includes a symmetrically distributed third sub-cover and a fourth sub-cover, the third sub-cover and the fourth sub-cover forming a double-door structure.

6. The aircraft electric drive cooling system as described in claim 5, characterized in that, The opening direction of the upper cover is towards the heat dissipation housing, and the opening direction of the lower cover is towards the bottom outer side of the power arm. The upper cover and the lower cover are located within the orthogonal projection area of ​​the aircraft's propeller rotation trajectory. The flip-opening is achieved based on the downward airflow generated by the aircraft's propeller rotation.

7. The aircraft electric drive cooling system as described in claim 1, characterized in that, The aircraft is equipped with a first servo and a second servo. The first servo is driven and connected to the upper cover plate, and the second servo is driven and connected to the lower cover plate.

8. A method for cooling an aircraft via electric drive, characterized in that, The electric drive cooling method is applicable to the aircraft electric drive cooling system as described in any one of claims 1 to 7, and the cooling method includes: S11: According to the flight mission of the aircraft, drive the upper and lower cover plates to flip and adjust the air inlet and air outlet to the fully open state; S12: Detect the opening and closing control mode of the upper cover and the lower cover. If it is a passive control mode, proceed to step S13; if it is an active mode, proceed to step S14. S13: By adjusting the output speed of the drive motor, the folding and opening range of the upper and lower cover plates can be adjusted in a coordinated manner. S14: Detect the real-time temperature of the drive motor, and adjust the folding opening range of the upper and lower cover plates according to the real-time temperature.

9. The aircraft electric drive heat dissipation method as described in claim 8, characterized in that, The adjustment of the folding opening range of the upper and lower cover plates according to the output speed of the drive motor includes: Based on the flight mission of the aircraft, the output power of the aircraft's drive motor is adjusted, and the opening range of the upper and lower covers is adjusted according to the rotation speed of the propeller blades driven by the drive motor.

10. The aircraft electric drive heat dissipation method as described in claim 8, characterized in that, The method of detecting the real-time temperature of the drive motor and adjusting the folding opening range of the upper and lower cover plates based on the real-time temperature includes: The system detects when the drive motor is in a non-working state and acquires the real-time temperature of the drive motor. The real-time temperature is compared with a preset safe temperature threshold. If the real-time temperature is higher than the safe temperature threshold, the opening range of the upper and lower cover plates is reduced proportionally. If the real-time temperature is lower than the safe temperature threshold, then the upper cover and the lower cover should be completely closed.

Images