An aircraft electrically driven heat dissipation system and control method

Abstract

The application discloses an aircraft electric drive heat dissipation system and a control method. The electric drive 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 a 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 arranged at the top of the power arm, an air outlet is arranged at the bottom of the power arm, and the lower cover plate is located in the air outlet. The air inlet is located in the range of the orthographic projection of the rotating track of the blades of the aircraft. When the blades of the aircraft rotate, the upper cover plate and the lower cover plate are opened, and the heat dissipation air duct of the heat dissipation shell is in communication with the external environment based on the air inlet and the air outlet. By arranging the liquid cooling and air cooling mechanisms in the power arm, the heat dissipation efficiency is improved based on the liquid cooling combined with the air cooling. The opening degree of the air inlet cover and the air outlet cover is dynamically adjusted, the heat dissipation adjustment of the lift system is realized, and the heat dissipation efficiency of the aircraft is improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aircraft, in particular to an aircraft electric drive heat dissipation system and a control method. BACKGROUND

[0002] eVTOL (Electric Vertical Takeoff and Landing) is a new type of electric aircraft capable of vertical takeoff and landing. In a compound wing aircraft, the lift system generally provides lift for vertical takeoff and landing and conversion phase. The lift system provides lift for the aircraft by driving the blades of the power arm to rotate by the motor. Since the driving motor of the lift system has high power during vertical takeoff and landing and hovering of the aircraft, a large amount of heat is generated, and the driving motor of the lift system needs to be cooled in time.

[0003] The patent with the patent number CN202210128285.1 discloses a heat dissipation structure and an electric vertical takeoff and landing aircraft. The heat dissipation structure cools the motor controller by air cooling. A ventilation opening is provided on the aircraft body, and a ventilation assembly including a ventilation cover is connected to the body. When the ventilation cover is opened, the airflow generated by the lift propeller flows through the motor controller and out of the body to cool the motor controller. However, the ventilation assembly destroys the aerodynamic shape of the aircraft, resulting in an increase in flight resistance. In addition, the air cooling has lower heat dissipation capacity than liquid cooling, which is not conducive to meeting the heat dissipation demand of high power.

[0004] The patent with the patent number CN202310124038.9 discloses an eVTOL aircraft driving motor cooling system. The heat dissipation system cools the electric drive by liquid cooling. The outer surface of the heat dissipation water tank is a curved surface structure, which reduces the resistance of the aircraft during cruising flight. However, the contact area of the heat dissipation water tank with air is smaller than that of a conventional heat sink, and the heat dissipation capacity of the heat dissipation water tank is weaker than that of a conventional heat sink. SUMMARY

[0005] The present application aims to overcome the shortcomings of the prior art. The present application provides an aircraft electric drive heat dissipation system and a control method. The liquid cooling and air cooling mechanisms are arranged inside the power arm. The opening of the air inlet and air outlet on the power arm is adjusted to form an air cooling heat dissipation air duct, thereby improving the heat dissipation efficiency of the driving motor of the lift system and avoiding interference of the heat dissipation system with the flight of the aircraft.

[0006] The present application provides an aircraft electric drive heat dissipation system, which comprises a liquid cooling mechanism arranged inside a power arm of an aircraft. The power arm is provided with a heat dissipation shell. The liquid cooling mechanism comprises a heat sink body arranged in the heat dissipation shell.

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

[0008] The air inlet is located in the range of the orthogonal projection of the flight vehicle blade rotation track, the upper cover plate and the lower cover plate are opened when the blade of the flight vehicle rotates, and the heat dissipation air duct of the heat dissipation shell is in communication with the external environment based on the air inlet and the air outlet.

[0009] Further, the liquid cooling mechanism comprises a driving pump arranged at the lower end of the driving motor, a cooling tank arranged at one end of the radiator main body, and a first cooling pipe and a second cooling pipe arranged between the driving pump and the cooling tank.

[0010] The first cooling pipe and the second cooling pipe are provided with cooling liquid.

[0011] Further, the radiator main body comprises a plurality of heat dissipation fins, the plurality of heat dissipation fins are arranged in an array, and one end of any heat dissipation fin is located in the cooling tank.

[0012] Further, the cross-sectional area of the heat dissipation shell is the same as the cross-sectional area of the radiator main body.

[0013] Further, the upper cover plate is arranged on the heat dissipation shell based on a hinged mechanism, and the lower cover plate is arranged on the power arm based on a hinged mechanism.

[0014] The upper cover plate is arranged as a single cover plate structure opened along one side, or the upper cover plate comprises a first sub-cover plate and a second sub-cover plate symmetrically distributed, and the first sub-cover plate and the second sub-cover plate form a double door structure.

[0015] The lower cover plate is arranged as a single cover plate structure opened along one side, or the lower cover plate comprises a third sub-cover plate and a fourth sub-cover plate symmetrically distributed, and the third sub-cover plate and the fourth sub-cover plate form a double door structure.

[0016] Further, the opening direction of the upper cover plate is towards the heat dissipation shell, the opening direction of the lower cover plate is towards the outside of the bottom of the power arm, and the upper cover plate and the lower cover plate are located in the orthogonal projection area of the flight vehicle blade rotation track, and the downward airflow generated based on the rotation of the flight vehicle blade realizes the flip opening of the doors.

[0017] Further, the flight vehicle is provided with a first steering engine and a second steering engine, the first steering engine is drivingly connected with the upper cover plate, and the second steering engine is drivingly connected with the lower cover plate.

[0018] The application provides an aircraft electric drive heat dissipation method, which is suitable for the aircraft electric drive heat dissipation system.

[0019] S11: according to the flight task of the aircraft, the upper cover plate and the lower cover plate are driven to be turned over, and the air inlet and the air outlet are adjusted to a fully open state;

[0020] S12: the opening and closing control mode of the upper cover plate and the lower cover plate is detected, if it is a passive control mode, step S13 is entered, if it is an active mode, step S14 is entered;

[0021] S13: the turning folding opening amplitude of the upper cover plate and the lower cover plate is adjusted in linkage by adjusting the output rotating speed of the driving motor;

[0022] S14: the real-time temperature of the driving motor is detected, and the turning folding opening amplitude of the upper cover plate and the lower cover plate is adjusted according to the real-time temperature.

[0023] Further, the adjustment of the turning folding opening amplitude of the upper cover plate and the lower cover plate according to the output rotating speed of the driving motor comprises:

[0024] In combination with the flight task of the aircraft, the output power of the driving motor of the aircraft is adjusted, and the turning folding opening amplitude of the upper cover plate and the lower cover plate is adjusted based on the rotating speed of the driving motor driving the propeller blade of the aircraft.

[0025] Further, the detection of the real-time temperature of the driving motor and the adjustment of the turning folding opening amplitude of the upper cover plate and the lower cover plate according to the real-time temperature comprise:

[0026] The real-time temperature of the driving motor is detected, and the real-time temperature of the driving motor is obtained;

[0027] The real-time temperature is compared with a preset safety temperature threshold, if the real-time temperature is higher than the safety temperature threshold, the turning folding opening amplitude of the upper cover plate and the lower cover plate is reduced in proportion;

[0028] If the real-time temperature is lower than the safety temperature threshold, the upper cover plate and the lower cover plate are completely closed.

[0029] The application provides an aircraft electric drive heat dissipation system and a control method, liquid cooling and air cooling mechanisms are arranged in the power arm, the driving motor of the lifting system and the heat dissipation mechanism are integrated in the power arm, so that the driving motor and the heat dissipation system do not interfere with the size of the aircraft. The driving motor of the lifting system is cooled by the liquid cooling mode, the heat dissipation efficiency of the driving motor is improved, and the opening degree of the air inlet cover and the air outlet cover is dynamically adjusted during the flight of the aircraft, the heat dissipation adjustment of the lifting system is realized, and the heat dissipation efficiency of the aircraft is improved. BRIEF DESCRIPTION OF DRAWINGS

[0030] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the accompanying drawings needed to be used in the description of the embodiments or the prior art will be briefly introduced. Obviously, the accompanying drawings in the following description only only some embodiments of the present application, and for those skilled in the art, other drawings can be obtained without creative labor on the basis of these drawings.

[0031] Figure 1 is a structural schematic diagram of an electric drive heat dissipation system of an aircraft in an embodiment of the present application;

[0032] Figure 2 is a schematic diagram of an open state of a heat dissipation air duct of an electric drive heat dissipation system of an aircraft in an embodiment of the present application;

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

[0034] Figure 4 is a structural schematic diagram of a passive control electric drive heat dissipation system of an aircraft in an embodiment of the present application;

[0035] Figure 5 is a structural schematic diagram of an active control electric drive heat dissipation system of an aircraft in an embodiment of the present application;

[0036] Figure 6 is a working circuit schematic diagram of an active control electric drive heat dissipation system in an embodiment of the present application;

[0037] Figure 7 is a flow chart of a control method of an electric drive heat dissipation system of an aircraft in an embodiment of the present application. DETAILED DESCRIPTION

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

[0039] Embodiment one:

[0040] Please refer to Figures 1 to 4The embodiment of the present application provides a kind of aircraft electric drive heat dissipation system, the electric drive heat dissipation system includes: the liquid cooling mechanism being arranged in the inside of aircraft power arm 1, the inside of the power arm 1 is provided with heat dissipation shell 6, the liquid cooling mechanism includes the radiator main body 7 being arranged in the heat dissipation shell 6, the drive motor 2 of the aircraft is arranged in the inside of the power arm 1, and the drive motor 2 is drivingly connected with the paddle of the aircraft, the paddle of the aircraft is rotated based on the drive of the drive motor 2, to satisfy the vertical lift and hovering of the aircraft and other flight actions.

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

[0042] Further, at least one drive motor 2 is arranged in the power arm 1, the output shaft of the drive motor 2 is connected to the propeller 3 of the aircraft, the propeller 3 of the aircraft is rotated based on the drive of the drive motor 2, and the airflow downward at the top end of the power arm 1 can be formed based on the rotation of the paddle of the propeller 3, so as to cooperate with the liquid cooling mechanism inside the power arm 1 to perform heat dissipation operation.

[0043] The air inlet 4 is located in the orthographic projection range of the aircraft paddle rotation track, when the paddle of the aircraft rotates, the upper cover plate 41 and the lower cover plate 51 are opened, and the heat dissipation air duct of the heat dissipation shell 6 is communicated with the external environment based on the air inlet 4 and the air outlet 5, the liquid cooling mechanism dissipates heat based on the radiator main body 7 to the drive motor 2 of the aircraft, by setting the upper cover plate 41 and the lower cover plate 51 to an open state, so that the heat dissipation shell 6 is in a communication state with the external environment, thereby forming a heat dissipation air duct, cooperating with the airflow generated by the rotation of the paddle of the aircraft, to drive the external environment airflow to flow from the heat dissipation air duct formed by the heat dissipation shell 6, so as to realize heat exchange between the radiator main body 7 and the external environment airflow.

[0044] Specifically, the liquid cooling mechanism includes a driving pump arranged inside the power arm 1, a cooling tank arranged at one end of the radiator main body 7, and a first cooling pipe 71 and a second cooling pipe 72 arranged between the driving pump and the cooling tank, the first cooling pipe 71 and the second cooling pipe 72 are arranged with cooling liquid, the driving pump is arranged at the lower end of the driving motor 2, the cooling liquid in the first cooling pipe 71 flows in the direction of the driving motor 2 along the radiator main body 7, and the cooling liquid in the second cooling pipe 72 flows in the direction of the driving motor 2 along the radiator main body 7, so that the radiator main body 7 and the driving motor 2 can form a liquid cooling circulation cooling based on the cooling liquid flowing in the first cooling pipe 71 and the second cooling pipe 72, thereby realizing efficient heat absorption of the driving motor 2, that is, realizing efficient cooling of the driving motor 2.

[0045] Further, when the driving motor 2 is working, the driving pump can work synchronously, based on the driving pump driving the cooling liquid in the first cooling pipe 71 and the second cooling pipe 72 to flow, so that the cooling liquid in the first cooling pipe 71 flows to the driving pump position along the first cooling pipe 71, that is, the cooling liquid in the first cooling pipe 71 can flow through the lower end position of the driving motor 2, so that the cooling liquid can exchange heat with the driving motor 2, and based on the cooling liquid can satisfy the heat exchange of the working heat of the driving motor 2, thereby satisfying the driving working demand of the driving motor 2.

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

[0047] Further, when the driving motor 2 is working, the driving pump can work synchronously, based on the driving pump driving the cooling liquid in the first cooling pipe 71 and the second cooling pipe 72 to circulate and flow, so that the reciprocating motion of the cooling liquid between the driving motor 2 and the radiator main body 7 can realize liquid cooling, that is, when the cooling liquid flows through the driving motor 2, it can absorb part of the heat generated by the driving motor 2 through heat exchange, and the heat dissipation efficiency of the cooling liquid can be enhanced through the heat dissipation fins, so that the liquid cooling mechanism is located inside the power arm 1, and has sufficient heat dissipation efficiency to meet the working demand of the driving motor 2 of the aircraft propeller 3.

[0048] The heat dissipation fins are arranged as thin sheet structures with regular geometric shapes, the length direction of the heat dissipation fins is perpendicular to the extension direction of the cooling groove, for example, the heat dissipation fins have rectangular or trapezoidal cross sections, and the thickness of the heat dissipation fins is controlled in the range of 0.5-1.2 mm. The array arrangement is implemented as equidistant staggered arrangement, and the interval between adjacent heat dissipation fins is configured as 2-5 mm to form parallel flow channels. The connecting end of the heat dissipation fin and the cooling groove is processed as a wedge-shaped structure, and is embedded into the side wall of the cooling groove through interference fit or welding, and the embedding depth reaches 15%-30% of the total length of the heat dissipation fin. Further, the extension part of the heat dissipation fin outside the cooling groove is arranged to be in contact with the outer shell of the driving motor 2, and the contact area accounts for 40%-60% of the total surface area of the heat dissipation fin.

[0049] Specifically, when the cooling liquid flows through the cooling groove, heat is conducted to the heat dissipation fins embedded in the side wall of the cooling groove. Since the heat dissipation fins are arranged in an array, the cooling liquid flow path is divided into multiple parallel branches, the flow rate is increased to 0.8-1.5 m / s, and the turbulent flow effect enhances the heat transfer efficiency. The area of the heat dissipation fin surface washed by the cooling liquid is increased to 1.8-2.5 times that of the traditional structure, wherein the end portion embedded in the cooling groove directly absorbs the heat of the liquid cooling circuit, and the portion extending outside the cooling groove is subjected to secondary heat dissipation through air convection. Through the matching design of the interval and thickness of the heat dissipation fins, the overall thermal resistance of the radiator main body 7 is reduced to 0.05-0.08 ℃ / W, and the heat conduction efficiency is increased by 35%-50% compared with the radiator without array fins. Thus, the heat generated by the driving motor 2 is simultaneously conducted through the liquid cooling circuit forced convection and air natural convection double paths, solving the heat accumulation problem in a single heat dissipation mode.

[0050] Specifically, the cross-sectional area of the heat dissipation shell 6 is the same as the cross-sectional area of the radiator main body 7, the heat dissipation shell 6 is arranged as a rectangular structure, the heat dissipation channel of the heat dissipation shell 6 is limited based on the rectangular structure, and the downdraft generated by the rotation of the propeller 3 of the aircraft can flow in the vertical direction, so that the airflow generated by the rotation of the propeller 3 has sufficient flow rate and airflow volume to meet the heat exchange demand of the heat dissipation fins of the radiator main body 7.

[0051] Specifically, the embodiment of the present application proposes to integrate the liquid cooling mechanism in the power arm 1, and through the linkage design of the upper and lower cover plates 51 and the propeller airflow, the heat dissipation air duct is passively opened by the airflow generated by the rotation of the propeller in the vertical take-off and landing stage, and the liquid cooling and air cooling double heat dissipation mechanisms are combined, so that the aircraft aerodynamic shape is maintained, and high-efficiency heat dissipation under high-power working condition is realized.

[0052] The working process and principle of the present application is that the aircraft electric drive heat dissipation system includes a liquid cooling mechanism arranged inside the power arm 1 of the aircraft. The power arm 1 is internally provided with a heat dissipation shell 6, and the liquid cooling mechanism includes a heat sink body 7 arranged in the heat dissipation shell 6. The top of the heat dissipation shell 6 is provided with an upper cover plate 41, and the bottom is provided with a lower cover plate 51. The top of the power arm 1 is provided with an air inlet 4 corresponding to the upper cover plate 41, and the bottom is provided with an air outlet 5, and the lower cover plate 51 is located in the air outlet 5.

[0053] The air inlet 4 is located in the range of the orthogonal projection of the rotating track of the propeller of the aircraft. When the propeller of the aircraft rotates, the upper cover plate 41 and the lower cover plate 51 are opened, and the heat dissipation air duct of the heat dissipation shell 6 is in communication with the external environment through the air inlet 4 and the air outlet 5. The downward airflow generated by the rotation of the propeller drives the opening of the upper cover plate 41 and the lower cover plate 51, forming a natural heat dissipation channel. The airflow enters the heat dissipation shell 6 from the air inlet 4, and after taking away the heat through the heat sink body 7, it is discharged from the air outlet 5.

[0054] Specifically, the upper cover plate 41 is arranged on the heat dissipation shell 6 based on a hinge mechanism 8, and the upper cover plate 41 can be configured as a single cover plate structure opened on one side, and the airflow generated by the rotation of the propeller 3 drives the opening.

[0055] Further, the upper cover plate 41 can also be provided as a double cover plate structure, including a first sub-cover plate and a second sub-cover plate symmetrically distributed, the first sub-cover plate and the second sub-cover plate form a double door structure, the opening direction of the first sub-cover plate and the second sub-cover plate is towards the heat dissipation shell 6, and the first sub-cover plate and the second sub-cover plate are located in the orthogonal projection area of the rotating track of the propeller of the aircraft, and the downward airflow generated by the rotation of the propeller of the aircraft realizes the opening of the double doors. This design integrates the liquid cooling mechanism inside the power arm 1, uses the aircraft body structure to form the heat dissipation shell 6, and maintains the integrity of the aerodynamic shape. The upper and lower through air duct design utilizes the airflow generated by the rotation of the propeller to realize passive heat dissipation. The air inlet 4 is arranged in the orthogonal projection area of the rotating track of the propeller, maximizing the utilization of the negative pressure effect generated by the rotation of the propeller to guide the airflow into the heat dissipation shell 6. The opening and closing state of the upper cover plate 41 and the lower cover plate 51 is dynamically related to the propeller rotating speed, realizing the automatic matching of the heat dissipation efficiency and the motor working state.

[0056] Specifically, the lower cover plate 51 can be provided as a single cover plate structure opened on one side, or the lower cover plate 51 includes a third sub-cover plate and a fourth sub-cover plate symmetrically distributed, the third sub-cover plate and the fourth sub-cover plate form a double door structure, and the action principle of the lower cover plate 51 is the same as that of the upper cover plate 41, which will not be repeated here.

[0057] Further, the opening direction of the upper cover plate 41 is towards the heat dissipation shell, the opening direction of the lower cover plate 51 is towards the bottom outside of the power arm 1, and the upper cover plate 41 and the lower cover plate 51 are located in the orthographic projection area of the blade rotation track of the aircraft, and the downward airflow generated by the rotation of the blade of the aircraft realizes the turnover opening.

[0058] Further, the hinge mechanism 8 can be provided as a torsion spring and a hinge, based on the cooperation of the torsion spring and the hinge, the hinge mechanism can maintain a horizontal unfolded state, so that the upper cover plate 41 or the lower cover plate 51 maintains a closed state, and the lower cover plate 51 can rotate around the shaft connected with the power arm 1, so as to switch between the closed and opened states. When the lower cover plate 51 is closed, it completely covers the air outlet 5, and does not allow gas to flow into or out of the air outlet 5, while together with the power arm 1, it forms a smooth aerodynamic shape, reduces the flow resistance, and improves the flight stability of the aircraft.

[0059] When the driving motor 2 works, the propeller 3 is driven to rotate through the output shaft, and at the same time, the cooling liquid is driven to circulate in the driving motor 2, the first cooling pipe 71, the heat dissipation main body 7 and the second cooling pipe 72 through the pump in the driving motor 2. The upward airflow generated by the rotation of the lift propeller 3 passively opens the upper cover plate 41 and the lower cover plate 51 under the action of the airflow generated by the lift propeller 3, so that the airflow generated by the lift propeller 3 flows through the air duct and the radiator. The heat generated by the driving motor 2 is transferred to the cooling liquid, and the cooling liquid transfers the heat to the air in the radiator, thereby realizing the cooling of the driving motor 2.

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

[0061] Specifically, in this embodiment, when the blades of the propeller 3 rotate at a high speed, the downward airflow forms a local high pressure area in the orthographic projection area, and the airflow pressure acts on the outer surface of the cover plate. Since the opening direction of the cover plate is towards the inside of the shell, the pressure difference of the airflow pushes the cover plate to turn over towards the inside of the shell, avoiding the increase of resistance caused by outward turning, and the corresponding relationship between the rotation speed of the blades of the propeller 3 and the upper cover plate 41 and the lower cover plate 51 is as follows:

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

[0063] In summary, when the driving motor 2 drives the propeller 3 to rotate, and the rotating speed of the propeller 3 is 1500 rpm, the opening angle of the upper cover plate 41 is not less than 65 degrees, and the opening angle of the lower cover plate 51 is not less than 60 degrees; when the rotating speed of the propeller 3 is 2000 rpm, the opening angle of the upper cover plate 41 is not less than 75 degrees, and the opening angle of the lower cover plate 51 is not less than 70 degrees; when the rotating speed of the propeller 3 is 2500 rpm, the opening angle of the upper cover plate 41 is not less than 85 degrees, and the opening angle of the lower cover plate 51 is not less than 80 degrees. And since the position of the upper cover plate 41 is 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 plate 41, and the lower cover plate 51 is driven to open based on the downward airflow flowing through the radiator main body 7, so that the opening angle of the lower cover plate 51 is smaller than the opening angle of the upper cover plate 41.

[0064] It is ensured that the downward airflow generated by the blades can directly act on the surface of the cover plate, and the passive opening of the cover plate is realized by using the pressure difference of the airflow as the driving force, without the need for additional driving mechanism, thereby simplifying the system structure. Moreover, based on the linkage design of the downward airflow turning the door, the opening amplitude of the cover plate and the rotating speed of the blades form a positive correlation, so that the cooling air duct is adaptively adjusted according to the cooling demand, the influence on the flight performance is minimized while the cooling efficiency is ensured. Finally, the design scheme solves the contradiction between the opening efficiency of the cover plate and the aerodynamic performance, realizes the dynamic balance between the cooling efficiency and the flight resistance, and improves the overall performance of the aircraft.

[0065] Embodiment two:

[0066] Please refer to Figure 5 and Figure 6 , the embodiment of the present application provides a kind of aircraft electric drive cooling system, specifically, in the embodiment, the first steering gear 9 and second steering gear are provided in the aircraft, the first steering gear 9 is driven and is connected with the upper cover plate 41, and the second steering gear is driven and is connected with the lower cover plate 51.The first steering gear 9 and the second steering gear are respectively installed on the two sides inside the cooling shell 6.The output shaft of steering gear is fixedly connected with one end of connecting rod mechanism, and the other end of connecting rod mechanism is hinged with cover plate.When needing to open cover plate, steering gear drives connecting rod mechanism to move, and push cover plate rotates around hinge point, realize the opening of cover plate.When needing to close cover plate, steering gear reversely moves, and pulls connecting rod mechanism, so that cover plate returns to closed position.

[0067] The control signal of steering gear is sent by the main controller of aircraft, and the main controller calculates the required cover plate opening angle according to the working state of aircraft, environmental temperature, motor temperature and other parameters, and sends corresponding control signal to steering gear.Steering gear accurately controls the rotating angle of output shaft according to the received control signal, to accurately adjust the opening angle of cover plate.

[0068] Further, in order to achieve more precise control, an angle sensor is installed on the cover plate to detect the actual opening angle of the cover plate in real time. The data of the angle sensor is fed back to the main controller, which compares the actual angle with the target angle and makes corrections if there is a deviation, forming a closed-loop control system.

[0069] In addition, in order to improve the reliability of the system, a elastic buffer device is provided between the rudder and the cover plate. When encountering external forces that hinder the movement of the cover plate, the elastic buffer device can absorb part of the impact force to prevent the rudder or the linkage mechanism from being damaged.

[0070] Through the above technical solutions, the present application realizes active and precise control of the cover plate of the heat dissipation system. Therefore, the heat dissipation efficiency and the flight condition can be dynamically matched, overcoming the problem of high dependence on airflow in passive control mode. Even at low speed or without blade rotation, the cover plate can be actively opened as needed to ensure heat dissipation effect. At the same time, by precisely controlling the opening and closing state of the cover plate, the influence on the aerodynamic performance of the aircraft can be minimized while ensuring heat dissipation effect. In addition, the rudder driving mode has fast response speed, which can adjust the opening angle of the cover plate in real time according to the flight state and heat dissipation demand, improving the flexibility and adaptability of the heat dissipation system.

[0071] Embodiment three:

[0072] Please refer to Figure 7 The present application improves a control method for an electrically driven heat dissipation system, which is suitable for the electrically driven heat dissipation system of the aircraft, and the heat dissipation method comprises:

[0073] S11: According to the flight task of the aircraft, the upper cover plate 41 and the lower cover plate 51 are driven to flip, and the air inlet 4 and the air outlet 5 are adjusted to a fully open state.

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

[0075] Further, in the main control mode, before the aircraft flies, the upper cover plate 41 can be actively driven to open by the first rudder 9, and the lower cover plate 51 can be actively driven to open by the second rudder.

[0076] Further, the aircraft can set the opening angle of the upper cover plate 41 and the lower cover plate 51 according to different flight tasks, so that the heat dissipation air duct can meet the flight requirements of different flight tasks of the aircraft.

[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, then enter step S13, if it is an active mode, then enter step S14;

[0078] S13: Adjust the output speed of the driving motor 2 to realize the linkage adjustment of the folding opening amplitude of the upper cover plate 41 and the lower cover plate 51;

[0079] Specifically, in combination with the flight task of the aircraft, the output power of the driving motor 2 of the aircraft is adjusted, and based on the rotation rate of the driving motor 2 driving the propeller 3 blade of the aircraft, the folding opening amplitude of the upper cover plate 41 and the lower cover plate 51 is adjusted.

[0080] Specifically, when the aircraft performs a high-load flight task, the actual power of the driving motor 2 and the propeller 3 is close to 100% of the rated power, and the wind pressure generated by the propeller 3 can drive the opening angle of the upper cover plate 41 and the lower cover plate 51 to be not less than 85 degrees, thereby enhancing the heat dissipation capacity of the radiator main body 7; when the aircraft performs a low-load flight task, the actual power of the driving motor 2 and the propeller 3 can be less than 80% of the rated power, the wind pressure generated by the propeller 3 is smaller, the opening angle of the upper cover plate 41 is less than 75 degrees, and the opening angle of the lower cover plate 51 is less than 70 degrees, thereby meeting the heat dissipation demand of the driving motor 2 under lower load.

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

[0082] Specifically, the driving motor 2 is detected to be in a non-working state, and the real-time temperature of the driving motor 2 is obtained;

[0083] The real-time temperature is compared with a preset safety temperature threshold, if the real-time temperature is higher than the safety temperature threshold, then the folding opening amplitude of the upper cover plate 41 and the lower cover plate 51 is reduced in proportion;

[0084] If the real-time temperature is lower than the safety 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 task, the working task to be performed by the driving motor 2 is obtained according to the flight task of the aircraft, and before the driving motor starts to work, the upper cover plate 41 and the lower cover plate 51 are rotated 90 degrees from the initial angle by the first steering engine 9 and the second steering engine, and the upper cover plate 41 and the lower cover plate 51 are completely opened, thereby avoiding that the airflow generated when the driving motor 2 starts to work causes the torque of the steering engine to be too large.

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

[0087] Further, when the driving motor 2 is working, the upper cover plate 41 is kept in the fully open state by the first steering wheel 9, and the lower cover plate 51 is kept in the fully open state by the second steering wheel, after the driving motor 2 stops working, if the temperature of the electric drive is still higher than the set safety range, and the temperature change rate of the electric drive is negative, that is, the real-time temperature of the driving motor 2 is in the state of gradually cooling down, then the angle of the upper cover plate 41 and the lower cover plate 51 is reduced, and the angle is reduced by 5 to 30 degrees each time, until the temperature of the electric drive is in the safety range and no longer rises, and the upper cover plate 41 and the lower cover plate 51 can be in the fully closed state:

[0088] When the temperature of the driving motor 2 is in the set safety range, and the real-time temperature of the driving motor 2 does not rise within 10 seconds, the first steering wheel 9 and the second steering wheel are controlled to return to the initial angle, and the upper cover plate 41 and the lower cover plate 51 are fully closed, so as to wait for the next flight task to perform the heat dissipation control adjustment.

[0089] The residual heat intelligent management after the driving motor 2 stops working is realized, by dynamically matching the cover plate opening degree and the real-time temperature, the aging acceleration of the motor parts caused by insufficient passive heat dissipation under high temperature state is prevented, and the additional aerodynamic resistance caused by the continuous opening of the cover plate 51 under low temperature state is avoided. The scheme effectively balances the heat dissipation efficiency and the aircraft cruising performance through the closed-loop control mechanism, and solves the control blind area problem of the traditional heat dissipation system in the non-working state of the motor.

[0090] In addition, the above-mentioned aircraft electric drive heat dissipation system and control method provided by the embodiment of the application are described in detail, and the principles and implementation modes of the application are described by using specific examples in this paper. The above-mentioned embodiment is only used to help understand the method and its core idea of the application; at the same time, for the general technical personnel in the art, according to the idea of the application, the specific implementation mode and application range will be changed, and the above-mentioned description should not be understood as the limitation of the application.

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 air duct of the heat dissipation shell is connected to the external environment based on the air inlet and the air outlet. The liquid cooling mechanism includes a drive pump located at the lower end of the drive motor, a cooling tank located at one end of the radiator body, and a first cooling pipe and a second cooling pipe located between the drive pump and the cooling tank. The first cooling pipe and the second cooling pipe are filled with coolant; 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.

2. 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.

3. 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.

4. The aircraft electric drive cooling system as described in claim 3, 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.

5. 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.

6. 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 5, 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.

7. The aircraft electric drive heat dissipation method as described in claim 6, 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.

8. The aircraft electric drive heat dissipation method as described in claim 6, 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