Aircraft

Abstract

The invention relates to an aircraft. The aircraft comprises a first driving system, a second driving system used for driving the aircraft to move horizontally and an operation system. The first driving system comprises two driving motors which are coaxially arranged and two groups of wind wings which are arranged in a conjugate manner; coaxial cavities are respectively formed in the two driving motors; the two sets of wind wings are arranged in the circumferential directions of rotors of the two driving motors correspondingly. The second driving system is arranged in the first driving system along the axial direction; the operation system defines a first loading bin arranged at the bottom end of the first driving system and a water spraying device; a first loading space for accommodating water is defined in the first loading bin; a nozzle of the water spraying device is arranged in the first loading bin and communicates with the first loading space. The stability of the aircraft during navigation can be improved.

Description

Technical Field

[0001] The present application relates to the technical field of aircraft, and more particularly, to an aircraft. Background Art

[0002] Helicopter-type aircraft can achieve navigation operations through operator control of the wind vanes. However, this type of aircraft in related technologies has a poor ability to maintain natural stability, which is not conducive to its application in special operating environments such as aerial firefighting and rescue operations that require stable operation. Summary of the Invention

[0003] The technical problem to be solved by this application is to provide an aircraft in view of the above-mentioned defects of the prior art.

[0004] The technical solution adopted by the present application to solve the technical problem is to construct an aircraft, including: The first drive system includes two coaxially arranged drive motors and two sets of conjugately arranged wind blades; the two drive motors each have a coaxial cavity; the two sets of wind blades are respectively arranged along the circumference of the rotors of the two drive motors; a second drive system for driving the aircraft to move horizontally, the second drive system being axially disposed within the first drive system; and The operating system is defined by a first loading bin and a water spraying device arranged at the bottom end of the first driving system; the first loading bin is defined by a first loading space for storing water; the nozzle of the water spraying device is arranged in the first loading bin and is connected to the first loading space.

[0005] In some embodiments, the operating system further includes a second loading bin coaxially disposed on the top end of the first driving system, wherein the second loading bin defines a second loading space.

[0006] In some embodiments, rescue equipment is arranged in the second loading space; And / or, the second loading space is provided with seats in a centrally symmetrical manner.

[0007] In some embodiments, at least a portion of the wall of the second loading bin is a transparent wall.

[0008] In some embodiments, an AI navigation system is disposed in the second loading compartment; and a display structure of the AI ​​navigation system is disposed on the transparent wall.

[0009] In some embodiments, at least two support mechanisms are provided at the bottom of the first loading bin, the support height of which can be adjusted according to the flatness of the ground.

[0010] In some embodiments, the support mechanism includes a cylinder and a servo motor that drives the cylinder to extend and retract; a plurality of exhaust valves are provided on the cylinder wall of the cylinder; the exhaust valves are used to open exhaust when the air pressure in the cylinder is greater than a threshold value.

[0011] In some embodiments, the first loading chamber is made of a material comprising a polyimide composite material.

[0012] In some embodiments, a lighting system is provided at the bottom end of the first loading bin.

[0013] In some embodiments, a control system is further included, at least a portion of the control system is disposed in the cavity.

[0014] In some embodiments, a volute is further included; the first drive system is coaxially disposed in the volute.

[0015] In some embodiments, a ratio of an outer diameter of the drive motor to an axial length of the drive motor is greater than or equal to 10 and less than or equal to 300.

[0016] The implementation of the technical solution constructed in this application has at least the following beneficial effects: By configuring the aircraft to be centrally symmetrical along its axis, the present invention ensures that the aircraft's center of gravity remains on its axis, making the aircraft's navigational attitude more stable and reducing the probability of misalignment such as tilting and capsizing. By providing a cavity, the present invention increases the moment of inertia of the drive motor and rotor, enabling the aircraft to generate a greater restoring torque and improve the aircraft's stability during navigation, facilitating its application in areas requiring stable operation, such as firefighting and rescue operations. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] The present application will be further described below with reference to the accompanying drawings and embodiments, in which: Figure 1 is a schematic structural diagram of an aircraft in the first embodiment of the present application; Figure 2 yes Figure 1 A schematic cross-sectional structure diagram of the first drive system, the second drive system, and the first loading bin; Figure 3 yes Figure 1 The schematic cross-sectional structure of the first and second drive systems of the aircraft is shown when the aircraft is turning; Figure 4 yes Figure 1 A top view of the first drive system in FIG. Figure 5 yes Figure 2 A schematic cross-sectional structure diagram of the first drive motor; Figure 6 yes Figure 1 Schematic diagram of the aerodynamic field that can be generated by the first drive system of the aircraft during navigation; Figure 7 yes Figure 1 The schematic diagram of the module relationship of the aircraft control system shown; Figure 8 is a schematic structural diagram of an aircraft in a second embodiment of the present application; Figure 9 It is a structural schematic diagram of an aircraft in the related art; Figure 10 It is a structural schematic diagram of another aircraft in the related art. DETAILED DESCRIPTION

[0018] In order to provide a clearer understanding of the technical features, objectives, and effects of the present application, the specific embodiments of the present application are now described in detail with reference to the accompanying drawings. The following description sets forth many specific details to facilitate a full understanding of the present application. However, the present application can be implemented in many other ways than those described herein, and those skilled in the art may make similar improvements without violating the scope of the present application. Therefore, the present application is not limited to the specific embodiments disclosed below.

[0019] In the description of this application, it should be understood that the terms "longitudinal", "transverse", "up", "down", "top", "bottom", "inside", "outside", etc., indicating the orientation or position relationship, are based on the orientation or position relationship shown in the accompanying drawings or the orientation or position relationship in which the product of this application is usually placed when in use. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation. Therefore, they should not be understood as limiting this application.

[0020] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be understood to indicate or imply relative importance or implicitly specify the number of technical features indicated. Thus, a feature specified as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of this application, "plurality" means at least two, such as two or three, unless otherwise specifically defined.

[0021] In this application, unless otherwise specified or limited, the terms "installed," "connected," "connect," "fixed," etc. should be understood in a broad sense. For example, they can refer to fixed connection, detachable connection, or integration; mechanical connection or electrical connection; direct connection or indirect connection through an intermediate medium; internal communication between two elements or interaction between two elements, unless otherwise specified. Those skilled in the art will understand the specific meanings of the above terms in this application based on specific circumstances.

[0022] In this application, unless otherwise expressly specified or limited, a first feature being "above" or "below" a second feature may mean that the first and second features are in direct contact, or that the first and second features are in indirect contact through an intermediate medium. Furthermore, a first feature being "above" a second feature may mean that the first feature is directly above or diagonally above the second feature, or simply means that the first feature is at a higher level than the second feature. A first feature being "below" a second feature may mean that the first feature is directly below or diagonally below the second feature, or simply means that the first feature is at a lower level than the second feature.

[0023] Figures 1 to 7 The aircraft 1 in the first embodiment of the present application is shown, which can be used in the field of firefighting and rescue to achieve stable aerial firefighting and rescue operations. The aircraft 1 includes a first drive system 10, a second drive system 20, an operating system 40 and a control system 30. Among them, the first drive system 10 is used to achieve the lifting and turning of the aircraft 1. The second drive system 20 is axially arranged in the first drive system 10, and is used to drive the aircraft 1 to achieve horizontal navigation. The control system 30 is electrically connected to the first drive system 10, the second drive system 20 and the operating system 40 respectively, and is used to control the normal operation of the aircraft 1. The operating system 40 is used to carry people, goods and water to facilitate firefighting and rescue operations.

[0024] The first drive system 10 is generally in the shape of a flat pancake, dish, or disc, with a larger outer diameter than its axial dimension. This structure is simple to set up and provides stable navigation, and can be applied to both medium-sized and large aircraft, as well as small aircraft.

[0025] Specifically, the first drive system 10 is arranged centrally and symmetrically along its axis and includes two drive motors and two sets of wind blades. The two sets of wind blades are arranged circumferentially along the rotors of the two drive motors, and the blades of the two sets are arranged conjugately. The two drive motors are arranged coaxially, and each drive motor is arranged in an annular shape, with a coaxial cavity 100 arranged within the ring to expand its outer diameter.

[0026] The control system 30 is electrically connected to the two drive motors respectively. The control system 30 controls the two drive motors to rotate in opposite directions, thereby achieving opposite rotation of the two conjugated groups of wind blades.

[0027] Among them, such as Figure 5 As shown, the ratio of the outer diameter D of the drive motor to the axial length L of the drive motor is greater than or equal to 10 and less than or equal to 300. Specifically, the ratio of the outer diameter D to the axial length L can be 10, 15, 20, 25, 30, 35, 40, 45, 50, 55, 60, 65, 70, 75, 80, 85, 90, 95, 100, 105, 110, 115, 120, 125, 130, 135, 140, 145, 150, 155, 160, 165, 170, 175, 180, 185, 190, 195, 200, 205, 210, 215, 220, 225, 230, 235, 240, 245, 250, 255, 260, 265, 270, 275, 280, 285, 290, 295, 300, or any other value within this range.

[0028] like Figure 2 As shown, two drive motors are defined as a first drive motor 11 and a second drive motor 12, and two sets of wind blades are defined as a first wind blade 13 and a second wind blade 14. The first drive motor 11 and the second drive motor 12 are both annular and coaxially arranged. The first wind blade 13 is arranged circumferentially of the first drive motor 11, and the second wind blade 14 is arranged circumferentially of the second drive motor 12.

[0029] During the specific operation of aircraft 1, by controlling the two drive motors to rotate at the same speed and in opposite directions, the two sets of conjugate airfoils rotate at the same speed and in opposite directions, generating a tubular aerodynamic force. This tubular aerodynamic force exerts a vertical downward pressure on the ground and the atmosphere, thereby generating the lift required for takeoff and landing of aircraft 1. By controlling the two drive motors to rotate at different speeds and in opposite directions, the two sets of conjugate airfoils rotate at different speeds and in opposite directions, generating a controllable reaction torque, thereby achieving steering of aircraft 1. By controlling the operation of the second drive system 20, horizontal propulsion of aircraft 1 can be achieved.

[0030] It's important to understand that aircraft blades require relatively high torque but relatively low speed requirements. Aircraft drive motors in related art, limited by their size and weight, struggle to directly output high torque. Therefore, the drive motors are typically configured to operate in a high-speed range to increase their power density. A speed reducer is then used to reduce the speed to the required range for the blades, thereby achieving amplified torque.

[0031] Therefore, in the related art, the power of the driving motor can be obtained by formula 1.

[0032] Formula 1 in, is the power of the driving motor; is the gear transmission efficiency between the drive motor and the reducer; is the torque of the driving motor; is the reduction ratio of the reducer; is the angular velocity of the drive motor.

[0033] For the present application, by setting the cavity 100 to expand the outer diameter of the first drive system 10 and controlling the size ratio of the outer diameter D and the axial length L of the drive motor, the torque that the drive motor can output can be effectively improved, and there is no need to arrange a reducer, so that the power of the drive motor can be obtained through Formula 2.

[0034] Formula 2 in, is the power of the first drive motor 11 and the second drive motor 12; is the torque of the first drive motor 11 and the second drive motor 12; is the angular velocity of the first drive motor 11 and the second drive motor 12 .

[0035] That is, the aircraft 1 constructed in the present application can provide the required torque to the first wind wing 13 and the second wind wing 14, and the efficiency of the drive motor can be improved. 1 / u , and at the same time the overall weight of the aircraft 1 can be greatly reduced.

[0036] It is important to understand that for each aircraft to operate stably, two natural stability conditions must be met, which are defined as the first natural stability condition and the second natural stability condition.

[0037] Among them, the first natural stability condition requires that the aircraft has the ability to maintain its original equilibrium state when it is in a calm atmospheric environment and in an unpowered control state (such as free fall).

[0038] It should be understood that an aircraft is in a calm atmospheric environment, which can be understood as the navigation environment in which it is located has no significant airflow disturbance, gentle temperature gradient, uniform air pressure distribution, uniform air humidity distribution, no rain, snow, thunderstorms, cumulus clouds, etc., that is, there is no environment that will produce obvious external forces on the aircraft.

[0039] like Figure 9As shown, there is a compound wing aircraft in the related art, which has a horizontally rotating wind wing and a vertically rotating wind wing, and the vertically rotating wind wing is arranged on one side of the axis of the horizontally rotating wind wing. Its structural setting makes its center of gravity located on one side of the axis of the horizontal wind wing. When the aircraft is sailing stably, its horizontal wind wing extends horizontally. However, when it changes from the stable sailing state to the unpowered control state, since its center of gravity is located on one side of the axis of the horizontal wind wing, it will tilt, so that the plane where the horizontal wind wing is located is tilted. It is difficult for the aircraft to land in its original stable posture, and it may even lose its posture and fall in rotation, which is a high risk.

[0040] The present application arranges the first drive system 10 in a centrally symmetrical arrangement along its axis, and arranges the second drive system 20 along the axis to the first drive system 10, so that the center of gravity of the aircraft 1 is approximately located on its axis. When the aircraft 1 is in a stable navigation state, the direction in which its axis extends is parallel to the direction of gravity. When it changes from the stable navigation state to an unpowered control state, since the direction in which its axis extends is parallel to the direction of gravity and its center of gravity is located on the axis, the aircraft 1 will be more likely to maintain its posture in a stable navigation state in a calm atmospheric environment, and perform free fall in a relatively stable posture.

[0041] By providing a coaxial cavity 100 within the drive motor, the present invention allows the aircraft 1 to generate significant air resistance while also maintaining a certain degree of buoyancy during its descent. Compared to other aircraft of the same mass, the acceleration during its descent is less than the acceleration due to gravity, resulting in a slower descent. Due to its slow and smooth landing, the impact force per unit area of ​​the aircraft 1 is significantly reduced, thereby improving safety.

[0042] The present application can further increase the air resistance encountered by the aircraft 1 during the descent process by setting the ratio of the outer diameter of the driving motor to its axial length to be greater than or equal to 10, thereby reducing the descent speed.

[0043] The second natural stability condition requires that the aircraft, in the flight control state, be able to respond to changes in the environment and the aircraft itself and have the ability to maintain balance.

[0044] It is important to understand that environmental changes can be understood as changes in the aircraft's flight environment, such as airflow disturbances, changes in temperature gradients, changes in air pressure distribution, changes in air humidity distribution, rain, snow, thunderstorms, cumulus clouds, and other conditions. In other words, these conditions can cause changes in the forces acting on the aircraft. Changes in the aircraft itself can be understood as changes in the power output of the aircraft's propulsion system, such as external impacts on the aircraft, the sudden drop of loaded cargo, changes in the aircraft's flight attitude, or operational commands issued by the operator.

[0045] Furthermore, the second natural stability condition can be expressed by formula three.

[0046] Formula 3 in, is the component of the restoring moment in the horizontal direction; is the component of the overturning moment in the horizontal direction.

[0047] It should be understood that the restoring moment is the moment automatically generated by the aircraft 1 to return to its original equilibrium position when it is subjected to an external force and deviates from its equilibrium state. The overturning moment is the moment that may cause the aircraft to deviate from its equilibrium state or even lose control due to changes in the environment or the aircraft itself.

[0048] The component of the restoring moment that the aircraft 1 can generate in the horizontal direction is It can be calculated by formula 4.

[0049] Formula 4 in, is the equivalent angular momentum of aircraft 1; is the angular velocity of the driving motor; is the equivalent moment of inertia; The angle at which the axis of the drive motor deviates from the direction of gravity.

[0050] It is important to understand that the rotation of the wings of aircraft in related art all exhibits a gyroscopic effect. This allows the restoring torque generated by this rotation to be used to align the aircraft's center of mass with the Earth's center, maintaining a stable attitude and improving flight stability. However, due to the limitations of their rotational inertia, the angular momentum provided by the rotor of the drive motor and the rotation of the wings is also relatively limited, hindering the realization of the second natural stability condition and, consequently, hindering the aircraft's ability to maintain a stable attitude through the gyroscopic effect.

[0051] By providing cavity 100, the present application increases the outer diameter of first drive system 10. Compared to motors of equivalent mass in related art, at the same speed, the drive motor of the present application can have a greater moment of inertia. As the outer diameter of the drive motor increases, the inner and outer diameters of the blade of the present application also increase, resulting in a greater moment of inertia. Furthermore, the present application can generate a greater restoring torque, thereby ensuring smoother flight of aircraft 1 compared to related art.

[0052] For example, there are Figure 10The related art shown includes eight motors. The eight motors are divided into four groups, each coaxially arranged. Each motor is equipped with two blades. The maximum wheelbase of the four motor groups (i.e., the distance between two motor groups on a diagonal) is 2330 mm. The outer diameter of each motor is 155 mm. Each blade is 1570 mm long and weighs 1.6 kg.

[0053] On this basis, based on its structural setting, the inner diameter of the rotor of each motor is less than 155mm. If the motor is simplified and the density of steel is regarded as the density of the motor rotor, the moment of inertia of each motor can be roughly calculated to be less than 0.1 , or even negligible.

[0054] Based on this data, assuming that the center of mass of each blade is at 40% of its length, the equivalent radius of gyration of each blade is 0.4*1.57m, or 0.628m. Further, it can be concluded that the moment of inertia of a single blade is approximately .

[0055] Furthermore, it can be calculated that Figure 10 The total moment of inertia of the related art shown is approximately 16*0.631+0.1*8 , that is, 10.896 .

[0056] On the other hand, if the dimensions of this application are the same as Figure 10 Since the first drive system 10 of the present application only includes two coaxially arranged drive motors, the maximum wheelbase of the four motors in the related art is taken as the outer diameter of the two drive motors of the present application, that is, 2330 mm.

[0057] Based on the structure of the first drive system 10 of the present application, due to the existence of the cavity 100, it is assumed that the inner diameters of the first drive motor 11 and the second drive motor 12 of the present application are both 2290 mm, and the axial thicknesses are both 40 mm. If the rotors of the first drive motor 11 and the second drive motor 12 are simplified to hollow flat round tubes, and the density of steel is also regarded as the density of the rotor. Based on the structures of the first drive motor 11 and the second drive motor 12, it can be calculated that the masses of the first drive motor 11 and the second drive motor 12 are respectively approximately 45.575 kg. Furthermore, it can be calculated that the equivalent moment of inertia of a single motor is approximately 60.8 .

[0058] Assuming that the number of blades of the first wind wing 13 and the second wind wing 14 of the present application is eight, that is, the aircraft 1 of the present application also includes a total of 16 blades. Assuming that the size of each blade is Figure 10The relevant technical parameters shown remain consistent, that is, the blade length is 1.57m, the mass is 1.6kg, and the center of mass is at 40% of its length.

[0059] Then, due to the structural setting of the first drive system 10 of the present application, the blades are set on the rotor of the motor, and it can be seen that the equivalent turning radius of each blade is m, which is 1.793 m. Further, it can be calculated that the moment of inertia of a single blade of the aircraft 1 of the present application is approximately equal to 5.144 .

[0060] Furthermore, it can be calculated that the total moment of inertia of the aircraft 1 under the specific parameters of this application is approximately 16*5.144+60.8*2 , that is, 203.904 .

[0061] It is obvious that, while pursuing the same external parameters as much as possible, the moment of inertia of the aircraft 1 constructed in this application is dozens of times that of the related art. Even if the moment of inertia of the motor is not considered in this application, the moment of inertia of only 16 blades is Figure 10 Furthermore, the component of the restoring torque that can be generated by the present application in the horizontal direction is It is also several times that of the related art.

[0062] It should be understood that during the flight, if the component of the overturning moment on the aircraft is Reached 203.904 , which may be affected by extreme environments such as extreme crosswinds, sudden strong gusts, typhoons, hurricanes, severe extratropical cyclones, thunderstorm downbursts, orographic turbulence, and the passage of cold fronts. The probability of such environments occurring is low, and the probability of using Aircraft 1 is even lower if such extreme conditions exist (or if meteorological monitoring indicates they will occur in the future).

[0063] Therefore Figure 1 The horizontal component of the restoring moment generated by the aircraft 1 during normal flight is The large restoring torque it can generate is sufficient to cope with the environmental changes that aircraft 1 may encounter on a daily basis. Based on the gyroscopic effect, the center of mass of this application can be stably pointed at the center of the earth. Even if an external force is applied, causing the axis of aircraft 1 to tilt with the direction of gravity, the large restoring torque will eliminate the tilt based on the gyroscopic effect. Therefore, the body shaking of aircraft 1 when subjected to external forces will be significantly reduced, significantly improving the stability during navigation. Therefore, compared with related technologies, it is more effective in coping with environmental changes.

[0064] Even if the aircraft 1 suddenly fails during normal navigation, the instantaneous power changes or loses power, resulting in loss of balance and the axis deviating from the direction of gravity. Gradually changes from 0 to non-zero. At this time, since the rotational kinetic energy provided by the rotational inertia of the aircraft 1 is still large, it can still provide a large restoring torque based on its own structural characteristics. In other words, the gyroscopic stabilizing torque generated by its rotational kinetic energy still exists, and it can still naturally cause its axis to deviate from the direction of gravity by an angle of Gradually decrease to the minimum. Therefore, compared with related technologies, it is better at responding to its own changes.

[0065] It is important to understand that, for example Figure 9 and Figure 10 As shown, in related art, the structural arrangement between the aircraft's drive motor and blades places the blade roots roughly at the axis of rotation. This results in a low linear velocity at the blade roots, which in turn reduces the lift contribution of at least a portion of each blade. Furthermore, because the blade roots are roughly at the axis of rotation, at least a portion of the airflow generated by each blade acts on the aircraft, generating harmful turbulence. This turbulence then interacts with the surrounding sudden airflow, hindering the aircraft's stable flight.

[0066] By providing the cavity 100 and controlling the ratio of the outer diameter D of the drive motor to its axial length to be between 10 and 300, the present application enables the drive motor to have a relatively large diameter. By arranging the two sets of wind blades on the outer peripheries of the two drive motors, the two sets of wind blades also have a larger equivalent diameter. By arranging the two sets of wind blades on the outer peripheries of the rotors of the two drive motors having the cavity 100, the roots of the blades of the two sets of wind blades also have a higher linear velocity, thereby increasing their contribution to lift.

[0067] At the same time, by providing only two coaxial drive units, the present application can be considered both the drive structure of aircraft 1 and the fuselage of the aircraft. By arranging the drive units in a centrally symmetrical arrangement along the axis, the aircraft 1 constructed in the present application can be roughly centrally symmetrical along the axis, ensuring that its center of gravity is approximately located on its axis. By coaxially arranging two sets of wind blades on the periphery of two sets of drive motors, the present application avoids the interaction between the airflow generated by the rotation of the wind blades and the fuselage of aircraft 1, thereby reducing the generation of turbulence, further improving the stability of the aircraft's navigation, and facilitating the safety and comfort of passengers.

[0068] It should be understood that the lift efficiency generated by the wings of aircraft in related technologies is relatively low. Figure 9 and Figure 10In the related art shown, although the wind vanes generate downward airflow, most of the airflow is dissipated into the environment in the form of wake, and the kinetic energy is not completely converted into lift. The efficiency of converting kinetic energy into lift is only about 14%. For example, a winged aircraft has a curved upper surface and a flat lower surface. When air flows over the wing, the upper surface has a fast flow rate and low pressure, while the lower surface has a slow flow rate and high pressure. The pressure difference creates an upward lift. This arrangement makes the airflow generated during navigation roughly parallel to the ground, and the lift efficiency is about 20%. In other words, the energy loss of the aircraft in the related art is relatively high, accounting for about 75% to 86% of the total power of its drive motor, and the efficiency of converting kinetic energy into lift is only about 14% to 25%.

[0069] By providing a cavity 100 and controlling the ratio of the outer diameter D of the drive motor to its axial length to be between 10 and 300, the drive motor can be provided with a relatively large diameter. By arranging two sets of blades on the outer peripheries of the two drive motors, the two sets of blades also have a large equivalent diameter. By coaxially arranging the two drive motors and two sets of conjugate blades with opposite torsional directions, and controlling the two sets of blades to rotate in opposite directions, a stable tubular downwash airflow is generated. This creates a larger hollow area within the downwash airflow, resulting in a higher air density in the middle of the airflow. This results in a significant density difference between the upper and lower sides of the aircraft 1 during flight, creating the illusion that the aircraft 1 is positioned above a denser air layer, significantly enhancing lift. Because this lift is perpendicular and tangential to the aircraft's horizontal flight direction, there is no mutual constraint between lift and windage. This arrangement results in a high lift efficiency for the aircraft 1, improving the lift-to-power ratio (i.e., the ratio of lift to drive efficiency) by at least approximately 2 to 4 times. Furthermore, the larger the diameter of the first driving system 10 is, the greater the lift improvement is.

[0070] The increase in lift is conducive to providing the aircraft 1 with better load-bearing performance, making the aircraft 1 more conducive to stably carrying out firefighting and rescue operations.

[0071] It should be understood that, based on the above, due to the volume of the cavity 100 of the aircraft 1, the component of the restoring moment in the horizontal direction Both need to be as large as possible, and the volume of the cavity 100 is proportional to the square of the radius of the aircraft 1. The component of the restoring moment in the horizontal direction is It is also proportional to the square of the radius of the aircraft 1. Therefore, the present application sets a range of the ratio of the outer diameter D of the drive motor to its axial length L to control the shape of the aircraft 1 to be flat, which can facilitate the satisfaction of two natural stability conditions.

[0072] In other words, the aircraft 1 constructed in the present application satisfies both natural stability conditions based on its own structure, rather than on the operator's operational control of the aircraft 1 during flight. In other words, the present application maintains the natural stability conditions automatically, rather than relying on manual control by the operator. In contrast, in related art, stable flight of an aircraft relies on real-time manual control by the operator to maintain stable flight.

[0073] Therefore, compared with the related art, the aircraft 1 constructed in the present application is easier to maintain the first natural stability condition and the second natural stability condition under various working conditions.

[0074] Further, based on Figure 1 In the aircraft 1 shown, if the speed of the first drive motor 11 or the second drive motor 12 is controlled to change by 1 revolution per unit time, so that the differential speed of the two motors rotates in opposite directions, the aircraft 1 can be turned. Then, when the speed of one of the drive motors is controlled to change, the reaction torque that the drive motor can generate is .

[0075] For a flywheel torque motor with an outer diameter of 2330 mm, it generates The current required for the torque is only at the milliampere level, which is very small. Therefore, the aircraft 1 can achieve agile steering by controlling the speed change between the first drive motor 11 and the second drive motor 12.

[0076] The present application can realize the navigation operation of the aircraft 1 by setting up the first drive system 10, the second drive system 20 and the control system. Its structure is simpler and has higher safety and stability.

[0077] In some embodiments, the radius of the cavity 100 is larger than the radial dimension of the driving motor. Figure 5 The axial cross-section diagram of the drive motor is shown. Now define the radius of the cavity 100 as R and the radial thickness of the drive motor as M. Figure 5 As shown, R is significantly larger than M.

[0078] By limiting the relationship between the radius R of cavity 100 and the radial dimension M of the drive motor, cavity 100 can be made larger. Furthermore, in the unpowered control state of the first natural stability condition, when aircraft 1 is in free fall, cavity 100 can provide aircraft 1 with a significant air buoyancy force, thereby causing the acceleration of aircraft 1 to be significantly less than the acceleration due to gravity, allowing it to descend to the ground at a slower rate.

[0079] At the same time, by limiting the size of the cavity 100 , based on the principle of minimum energy, the aircraft 1 can further meet the first natural stability condition during free fall and fall steadily and slowly at a horizontal angle.

[0080] In some embodiments, the air buoyancy generated by the cavity 100 , and the ratio of the weight of the aircraft 1 to its own weight is greater than or equal to 0.2 and less than or equal to 75. Specifically, the ratio may be 0.22, 0.25, 0.28, 0.3, 0.32, 0.35, 0.38, 0.4, 0.42, 0.45, 0.48, 0.5, 0.52, 0.55, 0.58, 0.6, 0.7, 0.8, 0.9, 1, 5, 10, 15, 20, 25, 30, 35, 40, 45, 50, 55, 60, 65, 70, 75, etc., or any other value within this range.

[0081] The two cavities 100 are now defined as a first cavity 101 and a second cavity 102 , respectively. The first cavity 101 is located in the first driving motor 11 , and the second cavity 102 is located in the second driving motor 12 .

[0082] In some embodiments, the first cavity 101 and / or the second cavity 102 can also be used to carry passengers or cargo. Of course, the first cavity 101 and / or the second cavity 102 can also be used to carry batteries (not shown), control system 30, communications equipment (not shown), navigation system (not shown), various accessories of the aircraft 1 (not shown), and at least part of the electric servo pump of the water spraying system. Therefore, the provision of the cavity 100 can also provide carrying space for the aircraft 1, improving the space utilization of the aircraft 1.

[0083] It should be understood that the passenger seats and / or equipment items arranged in the first cavity 101 and / or the second cavity 102 should be arranged as symmetrically as possible along the axis to maintain the stability of the aircraft 1 during navigation.

[0084] In some embodiments, the first drive system 10 may further include an electric balancing device (not shown in the figure), which is electrically connected to the control system 30 and is used to achieve dynamic balance of the aircraft 1 when loaded.

[0085] Specifically, the electric balancing device may be disposed in the first cavity 101 and / or the second cavity 102 .

[0086] It should be understood that the electric balancing device can be implemented using existing technologies such as an active mass balance system (AMBS), which is not specifically limited here.

[0087] For example Figure 2 and Figure 3As shown, in some embodiments, the first drive motor 11 includes a first rotor and a first stator 111. The first rotor includes a first iron core 112 and two first magnetic steels 113. The first iron core 112 is annular, and its end face facing away from the second drive motor 12 is recessed to form an annular stator slot with a U-shaped cross-section. The two annular first magnetic steels 113 are respectively fixed to the opposite slot walls of the stator slot of the first iron core 112. The first stator 111 extends into the stator slot and is radially located at the midpoint between the two first magnetic steels 113.

[0088] The second drive motor 12 includes a second rotor and a second stator 121. The second rotor includes a second core 122 and two second magnets 123. The second core 122 is annular, with a U-shaped stator slot formed in a recessed pattern on its end surface facing away from the first drive motor 11. The two annular second magnets 123 are fixed to opposing walls of the stator slots of the second core 122. The second stator 121 extends into the stator slots and is radially located midway between the two second magnets 123.

[0089] It should be understood that the first drive motor 11 and the second drive motor 12 may specifically be existing flywheel torque motors, so they will not be described in further detail here.

[0090] By configuring the two drive motors as flywheel torque motors, the flywheel torque motors can be utilized to generate a larger moment of inertia, thereby further improving the ability of the aircraft 1 constructed in the present application to maintain the second natural stability condition.

[0091] In some other embodiments, the first drive motor 11 and the second drive motor 12 may also be other existing motors such as a brushless DC motor, a permanent magnet synchronous motor, etc.

[0092] In some embodiments, the first stator 111 and the second stator 121 can be formed by curing a polymer material into a ring-shaped stator winding to form an iron-coreless stator structure, thereby reducing the weight of the aircraft 1 and facilitating a lightweight design. Of course, the first stator 111 and the second stator 121 can also be a stator structure with an iron core.

[0093] In some embodiments, the first driving system 10 further includes a frame 15. The first driving motor 11 and the second driving motor 12 are symmetrically disposed on opposite sides of the frame 15 along a direction perpendicular to the axis.

[0094] Specifically, the frame 15 includes two mounting frame assemblies. These two mounting frame assemblies are arranged perpendicular to the axis of the first drive system 10 in a mirror-symmetrical manner, each supporting two drive motors. The two mounting frame assemblies are defined as a first mounting frame assembly 151 and a second mounting frame assembly 152. The first drive motor 11 is mounted on the first mounting frame assembly 151, and the second drive motor 12 is mounted on the second mounting frame assembly 152.

[0095] The first mounting frame assembly 151 comprises a first base 1511, a first mounting frame, and a first mounting bearing 1512. The first base 1511 is plate-shaped and perpendicular to the axis. The first mounting frame is located on the side of the first base 1511 facing away from the second mounting frame assembly 152 and is generally hollow and barrel-shaped. The first stator 111 of the first drive motor 11 is fixed to the first mounting frame, and the first rotor is rotatably mounted relative to the first mounting frame assembly 151 via the first mounting bearing 1512.

[0096] Specifically, the first mounting bracket includes a plate-shaped first portion 1513 and a tubular second portion 1514. The second portion 1514 can be in the form of a flat circular tube, with an axial length matching the axial length of the first drive motor 11. One end of the second portion 1514 is fixed to the first base 1511 and extends perpendicularly to the first base 1511. The first portion 1513 is generally annular and plate-shaped, with its inner periphery fixed to the end of the second portion 1514 away from the first base 1511.

[0097] The second portion 1514 and the first base 1511 together define the first cavity 101. The first mounting bearing 1512 is disposed on a side of the second portion 1514 facing away from the first cavity 101 and radially located between the first rotor of the first drive motor 11 and the second portion 1514. One end of the first stator 111 of the first drive motor 11 is fixed to a side of the first portion 1513 axially facing the first base 1511, and the other end extends in a direction parallel to the axis so as to extend into the stator slot of the first rotor.

[0098] By securing the first stator 111 to the first portion 1513 and positioning the first rotor on the first mounting bearing 1512, the first stator 111 and the first rotor are axially aligned, ensuring that the first stator 111 and the first rotor are axially aligned and that the air gap between them is uniform. This arrangement also reduces the axial dimensions of the aircraft 1 while maintaining rotational stiffness, further contributing to a lightweight and compact design for the aircraft 1.

[0099] The second mounting frame assembly 152 includes a second base 1521, a second mounting frame, and a second mounting bearing 1522. The second base 1521 is plate-shaped and perpendicular to the axis. The second mounting frame is located on the side of the second base 1521 facing away from the first mounting frame assembly 151 and is generally hollow and barrel-shaped. The second stator 121 of the second drive motor 12 is fixed to the second mounting frame, and the second rotor is rotatably mounted relative to the second mounting frame assembly 152 via the second mounting bearing 1522.

[0100] Specifically, the second mounting bracket includes a plate-shaped third portion 1523 and a tubular fourth portion 1524. The fourth portion 1524 can be a flat, circular tube with an axial length that matches the axial length of the second drive motor 12. One end is fixed to the second base 1521 and extends perpendicularly to the second base 1521. The third portion 1523 is generally annular, plate-shaped, with its inner periphery fixed to the end of the fourth portion 1524 away from the second base 1521.

[0101] The fourth portion 1524 and the second base 1521 together define the second cavity 102. The second mounting bearing 1522 is disposed on a side of the fourth portion 1524 facing away from the second cavity 102 and radially located between the second rotor of the second drive motor 12 and the fourth portion 1524. One end of the second stator 121 of the second drive motor 12 is fixed to a side of the third portion 1523 axially facing the second base 1521, and the other end extends in a direction parallel to the axis so as to extend into the stator slot of the second rotor.

[0102] By securing the second stator 121 to the third portion 1523 and positioning the second rotor on the second mounting bearing 1522, the second stator 121 and the second rotor can be axially positioned, ensuring that the second stator 121 and the second rotor are axially aligned and that the air gap between them is uniform. This arrangement also reduces the axial dimensions of the aircraft 1 while maintaining rotational stiffness, further contributing to a lightweight and compact design for the aircraft 1.

[0103] It should be understood that the first portion 1513 and the second portion 1514 can be fixed to each other detachably or non-detachably by welding, bolting, integral molding, etc. The third portion 1523 and the fourth portion 1524 are similar and are not specifically limited herein.

[0104] It should be understood that the first stator 111 can be fixed on the first part 1513 by first embedding the first stator 111 in the first part 1513 and then achieving it through existing technologies such as bonding and curing of the molecular material. The second stator 121 is fixed on the third part 1523 in the same way, and will not be elaborated on here.

[0105] It should be understood that the first mounting bearing 1512 and the second mounting bearing 1522 may specifically adopt existing cross ball bearings, double row bearings, or other types of common bearings, and no specific limitation is made here.

[0106] It should be understood that the specific shapes of the first substrate 1511 and the second substrate 1521 can be various shapes such as a prototype, polygon, ellipse, irregular, etc., and are not specifically limited here.

[0107] like Figure 2 and Figure 3 As shown, the first base 1511 and the second base 1521 are arranged to be movable relative to each other in a direction perpendicular to the axis. Along the axial direction of the aircraft 1, a drive space 150 is further defined between the first base 1511 and the second base 1521. The second drive system 20 can be accommodated in the drive space 150.

[0108] It should be understood that the driving space 150 can be defined by partially recessing one of the first base 1511 and the second base 1521 to form a groove, with the other of the two forming a cover to define the groove. Alternatively, the first base 1511 and the second base 1521 can be recessed to form a groove, and the two grooves can be butted together to form the groove. This is not specifically limited here.

[0109] It should be understood that when the second drive system 20 itself has a guiding function, the relative movement between the first base 1511 and the second base 1521 along a predetermined direction can be achieved by the second drive system 20. In this embodiment, the first base 1511 and the second base 1521 can contact and rub against each other. Alternatively, the first base 1511 and the second base 1521 can be arranged in parallel and spaced apart to reduce the coefficient of friction. Alternatively, a lubricating coating, balls, rollers, or other friction-reducing structures can be provided between the contact surfaces. This is not specifically limited here.

[0110] Of course, a guide mechanism may be further provided between the first base 1511 and the second base 1521 to improve the guiding effect and facilitate relative movement between the first base 1511 and the second base 1521 along a preset direction.

[0111] Specifically, the guide mechanism can be a guide rail structure formed by convex or concave portions of the first base 1511 and the second base 1521. A rolling mechanism with a low friction coefficient, such as a ball bearing or roller bearing, can be disposed between the friction surfaces within the guide mechanism to convert sliding friction into rolling friction, thereby reducing the friction coefficient. Alternatively, the sliding contact surfaces of the guide mechanism can be coated with a lubricating oil layer to reduce the friction coefficient. Alternatively, the guide mechanism can be implemented using existing guide structures, such as a lead screw. This is not specifically limited here.

[0112] In some embodiments, a battery (not shown in the figure), a control system 30, communication equipment (not shown in the figure), a navigation system (not shown in the figure), various aircraft 1 accessories (not shown in the figure), and at least one of the electric servo pumps of the water spraying device can also be arranged in the drive space 150.

[0113] Continue to read Figure 2 and Figure 3 In some embodiments, the first wind wing 13 and the second wind wing 14 each include a plurality of blades, and the plurality of blades are evenly spaced along the circumference of the rotor to ensure that the aircraft 1 is centrally and symmetrically arranged along the axial direction.

[0114] Specifically, the first wind blade 13 is radially disposed at one end of the first core 112 away from the second portion 1514 , and the second wind blade 14 is radially disposed at one end of the second core 122 away from the fourth portion 1524 .

[0115] By directly attaching the two sets of wind blades to the iron cores of the two drive motor rotors (i.e., first iron core 112 and second iron core 122), the wind blades can be assembled more easily, simplifying the structure of aircraft 1, reducing costs, and achieving lightweighting. By arranging the two sets of wind blades on the outer periphery of the two drive motor rotors, the rotation of the wind blades provides additional rotational inertia, further improving aircraft 1's ability to maintain two natural stability conditions.

[0116] It should be understood that the assembly of the wind blade and the iron core of the rotor of the drive motor can be achieved through existing assembly methods such as bolt connection, welding, and glue connection, and no specific limitation is made here.

[0117] In some embodiments, the circumference of the driving motor can be divided into a plurality of angular range periods, and an even number of fan blades are arranged at intervals in each angular range period.

[0118] Specifically, the angular range period may be greater than or equal to 10° and less than or equal to 20°. For example, the angular range period may be 12°, 15°, 18°, or any other value within this range. The even number of blades may be 2, 4, etc., and is not specifically limited here.

[0119] like Figure 4 As shown, in some embodiments, the ratio of the length H of the fan blade to the outer diameter D of the drive motor is greater than or equal to 0.4 and less than or equal to 0.8. Specifically, the ratio can be 0.45, 0.48, 0.5, 0.53, 0.55, 0.58, 0.6, 0.63, 0.65, 0.68, 0.7, 0.73, 0.75, 0.78, or any other value within this range.

[0120] For example Figure 2 and Figure 3As shown, in some embodiments, the second drive system 20 is axially disposed between the first drive motor 11 and the second drive motor 12 and may include at least one horizontal drive assembly. The horizontal drive assembly is disposed within the drive space 150 and drives the first base 1511 and the second base 1521 to move relative to each other in a direction perpendicular to the axis, thereby driving the first drive motor 11 and the second drive motor 12 to move relative to each other perpendicular to the axis.

[0121] It is important to understand that the flight control of an aircraft is based on the intensity and changes of the wake vortex it generates. Figure 9 and Figure 10 As shown, in related art aircraft, the wake vortices generated by the various wings interact with each other. Therefore, achieving flight control, such as lift, horizontal movement, and steering, requires the operator to simultaneously control multiple control variables while also maintaining the stability of the aircraft's attitude to constantly meet two natural stability conditions. The interplay of multiple control variables places high demands on the operator's operational skills and increases the difficulty of maintaining stable flight for aircraft 1.

[0122] For example, based on Figure 9 In the related technology shown, during navigation, operators are required to control the synchronous rotation of the horizontally rotating blades and the vertically rotating blades. Because the vertically rotating blades are located to one side of the horizontally rotating blades, the wakes generated by the two blades interact with each other, causing the control variables of the two blades to couple with each other.

[0123] At this time, how to ensure the rapid vertical take-off and landing of the aircraft during the take-off and landing phases under the premise of mutual coupling of various control variables, how to accurately output the output power of the motors corresponding to each wind wing when hovering, and how to output stable power to the motors corresponding to each wind wing during horizontal navigation to ensure the stable speed and direction of the aircraft, all require the operator to achieve through his operations.

[0124] Therefore, aircraft in related art require operators to possess highly sophisticated operating techniques, synchronously adjusting multiple, coupled control variables to ensure stable flight. This also requires aircraft in related art to be subject to special constraints, such as prohibiting changes in flight direction when descending close to the ground, to ensure safe and stable flight.

[0125] like Figure 2 and Figure 3 As shown, the present application sets up a first drive system 10, two drive motors set in coaxial mirrors, and two sets of conjugately set wind blades, so that the first drive system 10 can generate a tubular aerodynamic field along the circumference of the first drive system 10 during a single operation.

[0126] When the first drive motor 11 and the second drive motor 12 are displaced from each other and their axes are arranged parallel and spaced apart, the original tubular air dynamic field will change accordingly and the axis of the dynamic field will be tilted. Figure 3 As shown, the aerodynamic field, originally oriented along the axis of gravity, tilts after the first drive motor 11 and the second drive motor 12 are misaligned. This axis is then decomposed into a vertical component P extending parallel to the direction of gravity and a horizontal component Q extending perpendicular to the direction of gravity. The vertical component P provides lift for aircraft 1, enabling aircraft 1 to take off, land, or hover based on the magnitude of the vertical component P. The horizontal component Q provides propulsion for aircraft 1, enabling it to move horizontally.

[0127] In other words, the present application achieves horizontal movement of aircraft 1 by altering the aerodynamic field generated by first drive system 10, fundamentally avoiding the problem of coupled wakes between different drive systems. This reduces the generation of harmful airflow, such as turbulence, and further enhances aircraft 1's ability to maintain two naturally stable conditions.

[0128] This configuration also simplifies the operator's operation of aircraft 1. To take off or land aircraft 1, the two drive motors of first drive system 10 need only be controlled to rotate at the same speed and in opposite directions. To turn aircraft 1, the two drive motors of first drive system 10 need only be controlled to rotate at a differential speed and in opposite directions. To control horizontal movement of aircraft 1, while maintaining the operation of first drive system 10, the second drive system 20 needs only be controlled to offset the first drive motor 11 from the second drive motor 12.

[0129] It is important to understand that if Figure 6 As shown, the tubular aerodynamic field generates a vertical downward pressure effect relative to the ground and the atmosphere, thereby producing a unique wind field similar to that of a straw hat. Based on Bernoulli's principle, this unique wind field can provide aircraft 1 with approximately 20% to 30% additional lift, giving the aircraft 1 constructed in this application the ability to glide similar to fixed-wing aircraft.

[0130] That is, the configuration of the first drive system 10 and the second drive system 20 of the present application eliminates the restrictions on the special landing constraints of helicopters in the related art. While safely and stably navigating, the second drive system 20 can also be driven to move horizontally when landing close to the ground, thereby changing the flight direction.

[0131] The first drive system 10 and the second drive system 20 constructed in the present application generate wakes during operation that are more ideal than those in related technologies, and the wakes are more stable, safe, controllable, and efficient.

[0132] It is important to understand that something like Figure 9 and Figure 10 The presence of wake vortices in the related technologies shown above not only affects the aircraft's own flight stability but also significantly occupies flight space, thereby indirectly reducing the number of aircraft that can be accommodated within a given airspace during flight. This application significantly reduces the airspace occupied by aircraft 1 during flight by confining the wind flow generated by the aircraft 1 to a tubular aerodynamic field, thereby increasing aircraft flight density and alleviating the problem of airspace congestion.

[0133] For example Figure 2 and Figure 3 As shown, in some embodiments, the horizontal drive assembly is connected to the first base 1511 and the second base 1521, respectively, and includes a mounting base 21 and at least one driving mechanism 22. The mounting base 21 is fixed to the wall of the second base 1521 forming the driving space 150. The driving mechanism 22 is fixed to the wall of the first base 1511 forming the driving space 150 and is movably arranged relative to the mounting base 21 in a direction perpendicular to the axis.

[0134] Specifically, the drive mechanism 22 can be a drive motor, whose output shaft can be coaxially provided with a lead screw. The mounting base 21 can be a nut seat, with the lead screw movably inserted into the nut seat. Based on the principle of relative movement of the screw and nut, the drive motor drives the lead screw to rotate, thereby achieving relative movement between the lead screw and the nut seat, and thus relative movement between the drive mechanism 22 and the mounting base 21.

[0135] Furthermore, the horizontal drive assembly may also include two drive mechanisms 22, which are coaxially mounted on the mounting base 21 to improve drive efficiency. When the first drive motor 11 and the second drive motor 12 are coaxially arranged, the mounting base 21 can be located on the axis of the first drive system 10, and the two drive mechanisms 22 are symmetrically distributed to ensure that the aircraft 1 is centrally symmetrical along the axis.

[0136] In some other embodiments, the driving mechanism 22 can also adopt existing driving devices with retractable output shafts, such as servo motors and linear push rod motors, by fixing its output shaft to the mounting seat 21 and realizing relative movement between the driving mechanism 22 body and the mounting seat 21 through the extension and retraction of the output shaft.

[0137] In some other embodiments, the driving mechanism 22 may also be implemented by using existing technologies such as lifts, push rod motors, screw motors, cylinders, hydraulic cylinders, etc. It should be understood that the driving mechanism 22 may be implemented by using lifts, push rod motors, screw motors, cylinders, hydraulic cylinders, etc., which will not be elaborated one by one here.

[0138] In some other embodiments, the number of the horizontal driving components may also be set to a plurality of two, three, etc., which are symmetrically arranged in the driving space 150 .

[0139] In some other embodiments, the mounting base 21 may also be disposed on the first base 1511 , and the driving mechanism 22 may be disposed on the second base 1521 .

[0140] It should be understood that the tubular aerodynamic field is mostly formed at the bottom of the wind blade.

[0141] Therefore, in some other embodiments, the second drive system 20 may also be at least one fan coaxially disposed at the top of the aircraft 1, with the blowing direction perpendicular to the axial direction of the first drive system 10. By arranging the fan at the top of the aircraft 1, it can be located at a location where the tubular aerodynamic field strength is relatively weak, avoiding the relatively strong aerodynamic field formed at the bottom of the first drive system 10, thereby significantly reducing the mutual influence between the lateral wake generated by the fan and the tubular aerodynamic field generated by the first drive system 10.

[0142] Specifically, the second drive system 20 may include one fan, or may include two, three, four or more fans, which may be centrally symmetrically arranged at the ends of the first drive system 10 with the axis of the first drive system 10 as the symmetry line.

[0143] It should be understood that the fan may specifically adopt existing fan types such as a turbine fan, a magnetic levitation fan, etc., and no specific limitation is made here.

[0144] like Figure 7 As shown, in some embodiments, the control system 30 may include a controller 31, a lift control unit 32, a steering control unit 33, a horizontal motion control unit 34, a vertical position detection unit 35, a horizontal position detection unit 36, a posture detection unit 37 and a dual-axis motor driver 38.

[0145] The lift control unit 32, steering control unit 33, horizontal motion control unit 34, vertical position detection unit 35, horizontal position detection unit 36, and attitude detection unit 37 are electrically connected to the controller 31. The vertical position detection unit 35 is also electrically connected to the lift control unit 32. The horizontal position detection unit 36 ​​is also electrically connected to the horizontal motion control unit 34. The attitude detection unit 37 is also electrically connected to the steering control unit 33. The dual-axis motor driver 38 is electrically connected to the lift control unit 32, the steering control unit 33, the first drive motor 11, and the second drive motor 12.

[0146] The vertical position detection unit 35 is used to detect the vertical position of the aircraft 1 along the direction extending from gravity, and is used to output the detected vertical position information to the lift control unit 32. The horizontal position detection unit 36 ​​is used to detect the horizontal position of the aircraft 1 along the direction perpendicular to gravity, and is used to output the detected horizontal position information to the horizontal motion control unit 34. The attitude detection unit 37 is used to detect the vertical attitude position of the aircraft 1, and is used to output the detected vertical attitude position information to the rotation control unit 33.

[0147] The controller 31 is used to output control signals to the lift control unit 32 , the steering control unit 33 and the horizontal motion control unit 34 .

[0148] The lift control unit 32 receives control signals from the controller 31, generates a lift control signal based on the vertical position information output by the vertical position detection unit 35, and outputs the lift control signal to the dual-axis motor driver 38. Based on the lift control signal, the dual-axis motor driver 38 drives the first drive motor 11 and the second drive motor 12 to adjust their output power. The configuration of the lift control unit 32 enables feedback control, minimizing vertical position deviation of the aircraft 1.

[0149] The horizontal motion control unit 34 is configured to receive control signals from the controller 31, generate a horizontal position control signal based on the horizontal position information output by the horizontal position detection unit 36, and output the horizontal position control signal to the second drive system 20, allowing the second drive system 20 to make adjustments based on the horizontal position control signal. The configuration of the horizontal motion control unit 34 enables feedback control to minimize horizontal position deviations of the aircraft 1.

[0150] The steering control unit 33 is configured to receive control signals from the controller 31, generate a steering control signal based on the vertical attitude position information output by the attitude detection unit 37, and output the steering control signal to the dual-axis motor driver 38. Based on the steering control signal, the dual-axis motor driver 38 drives the first drive motor 11 and the second drive motor 12 to adjust their output power. The configuration of the steering control unit 33 enables feedback control to minimize vertical attitude deviation of the aircraft 1.

[0151] Compared with the related art, the control system 30 constructed in the present application further realizes the decoupling between the lifting, steering and horizontal movement of the aircraft 1 by setting a lift control unit 32, a steering control unit 33 and a horizontal motion control unit 34, in conjunction with the setting of the first drive system 10 and the second drive system 20, thereby simplifying the control of the aircraft 1 during navigation.

[0152] It should be understood that the dual-axis motor driver 38 has differential speed control capabilities. By providing the dual-axis motor driver 38, the first drive motor 11 and the second drive motor 12 can be driven synchronously, while also facilitating high-precision drive of both motors and enabling differential speed parameter settings for both motors to achieve opposite differential speed rotation, thereby enabling sensitive steering.

[0153] Furthermore, the configuration of the first drive system 10 and the second drive system 20 of the present application also simplifies the structure of the control system 30, eliminating the need to add a large amount of software and hardware around the safety performance of the aircraft 1. While ensuring the safety, stability and reliability of the navigation process of the aircraft 1, it further improves the lightweight design of the aircraft 1 and reduces costs.

[0154] It should be understood that the vertical position detection unit 35, the horizontal position detection unit 36, and the attitude detection unit 37 can be flexibly arranged on the aircraft 1. For example, they can be arranged on the inner and outer walls of the outer shell 60 of the aircraft 1, or on the first drive system 10 or the second drive system 20, or in the cavity 100, etc., without specific limitation herein.

[0155] The vertical position detection unit 35 can be implemented by using existing technologies such as a radio altimeter, an ultrasonic rangefinder, and a barometric altimeter, which is not specifically limited here.

[0156] The horizontal position detection unit 36 ​​can be implemented by using existing technologies such as global navigation satellite system (GNSS), inertial navigation system (INS), Doppler radar, vision / laser SLAM, etc., which are not specifically limited here.

[0157] The attitude detection unit 37 can be implemented by using existing technologies such as an inertial measurement unit (IMU), an attitude and heading reference system (AHRS), and a visual attitude detection unit, and is not specifically limited here.

[0158] The lift control unit 32 , steering control unit 33 , and horizontal motion control unit 34 can specifically utilize existing hardware platforms such as high-performance MCUs, DSPs (digital signal processors), and dedicated flight control modules, equipped with existing control algorithms such as PID control, fuzzy logic control, and adaptive control, as well as existing communication protocols, software architectures, power drive, and signal output modules to achieve their respective functions. This is not specifically limited here.

[0159] The dual-axis motor driver 38 can be implemented by using existing driver types such as an open-loop driver, a closed-loop driver, or based on the type of two drive motors, by using existing driver types such as a dual-axis torque motor driver, a dual-axis servo motor driver, or a dual-axis brushless DC motor (BLDC) driver, without specific limitation here.

[0160] like Figure 1 and Figure 2 As shown, in some embodiments, the operation system 40 includes a first loading compartment 41, a second loading compartment 42, and a water spraying device. The first loading compartment 41 is coaxially disposed at the bottom end of the first drive system 10, defining a first loading space 411 for storing firefighting water. The water spraying device is at least partially disposed above the first loading compartment 41 and communicates with the first loading space 411. It is configured to spray the firefighting water within the first loading space 411 at a predetermined pressure and in a predetermined direction for firefighting operations. The second loading compartment 42 is coaxially disposed at the top end of the first drive system 10 and is configured to carry goods or personnel.

[0161] Specifically, the first loading chamber 41 is generally hollow and circular and is fixed to the bottom end of the first drive system 10. The outer diameter of the first loading chamber 41 is generally compatible with the outer diameters of the first mounting frame assembly 151 and the second mounting frame assembly 152. The first loading chamber 41 is mounted on the side of the second mounting frame assembly 152 facing away from the first mounting frame assembly 151. The first loading space 411 can be a relatively enclosed space to facilitate the function of a water tank for storing water.

[0162] In other embodiments, the first loading space 411 may be a non-enclosed space, and at least one water tank may be disposed therein. The first loading space 411 may be defined by the water tank. The nozzle 43 of the water spraying device may be connected to the water tank and configured to spray the firefighting water in the water tank at a predetermined pressure in a predetermined direction.

[0163] In some other embodiments, the first loading bin 41 may also be in the shape of a hollow polygonal pie or polygonal column, polygon, hemisphere, hemi-ellipsoid, irregular shape or other shapes that are centrally symmetrical along the axis.

[0164] like Figure 2 As shown, in some embodiments, the first loading compartment 41 is axially connected to one end of the second mounting bracket assembly 152, so that the first loading space 411 and the second cavity 102 are connected to form a single space. This arrangement can increase the volume of the space for storing firefighting water, thereby facilitating longer firefighting operations.

[0165] In some embodiments, the material of the first loading compartment 41 may include a polyimide composite material, which has good corrosion resistance, sealing properties, and strong mechanical strength. The material of the first loading compartment 41 can facilitate the aircraft 1 to take off and land on the water surface, facilitating efficient water acquisition for firefighting and rescue operations.

[0166] In some embodiments, the water spraying device may be an existing electric private service water spraying device, which may include a servo motor, a water pump, a pressure sensor, at least one nozzle 43, and other components. The servo motor, pressure sensor, and other components may be electrically connected to the control system 30 so that, under the control of the control system 30, water is sprayed through the nozzle 43 at a preset water pressure in a preset direction, thereby achieving functions such as firefighting. This will not be elaborated on in detail here.

[0167] Specifically, the nozzle 43 can be installed through the first loading chamber 41, at least partially extending outside the first loading chamber 41 to spray water outward. The remaining components of the water spraying device can be installed inside the first loading chamber 41. Alternatively, the remaining components of the water spraying device can be partially installed inside the first loading chamber 41 to draw firefighting water from the first loading chamber 41 and spray it through the nozzle 43, with the remaining components installed inside the first cavity 101, the second cavity 102, or the drive space 150.

[0168] Furthermore, the number of the nozzles 43 may be at least two. At least two nozzles 43 may be evenly spaced along the circumference of the first loading bin 41 to increase the coverage area of ​​the water spraying direction and the flexibility of use.

[0169] In some other embodiments, the water spraying device may also adopt other existing water spraying devices such as existing pneumatic proportional control water spraying device, piezoelectric ceramic water spraying device, solenoid valve array water spraying device, hydraulic servo mechanical arm water spraying device, etc.

[0170] Of course, when the first loading space 411 is loaded with agricultural products such as pesticides, the aircraft 1 can also be used in the agricultural field, etc., and can be used in conjunction with a water spraying device to carry out pesticide spraying operations.

[0171] In some embodiments, the aircraft 1 may also include a lighting system (not shown in the figure), which can be electrically connected to the control system 30 and used for lighting under the control of the control system 30, which is beneficial for the aircraft 1 to perform firefighting and rescue in an environment with insufficient light.

[0172] Specifically, the lighting system may include at least one lighting lamp, which may be disposed at the bottom end of the first loading bin 41 .

[0173] It should be understood that the lighting system may adopt existing technologies, and the lighting lamps may adopt lamps such as HID (High Intensity Discharge Lamp) xenon lamps and LED lamps commonly used in existing aircraft, which will not be elaborated here.

[0174] like Figure 1As shown, in some embodiments, the second loading bin 42 is generally hollow, circular, or cylindrical and is fixed to the top of the first drive system 10. The outer diameter of the second loading bin 42 is generally compatible with the outer diameters of the first mounting frame assembly 151 and the second mounting frame assembly 152, and the second loading bin 42 is mounted on a side of the first mounting frame assembly 151 facing away from the second mounting frame assembly 152.

[0175] In some other embodiments, the second loading bin 42 may also be in the shape of a hollow polygonal pie or polygonal column, polygon, hemisphere, hemi-ellipsoid, irregular shape, or other shapes that are centrally symmetrical along the axis.

[0176] In some embodiments, the second loading compartment 42 may be a rescue compartment, and rescue equipment is arranged in the second loading space 421 .

[0177] It should be understood that the rescue equipment can be understood as medical devices, equipment, and related systems used to provide emergency medical intervention, injury assessment, vital sign maintenance, and transport support to injured personnel during firefighting and rescue operations on aircraft 1. For example, the rescue equipment may include seats, emergency beds 44, stretchers, first aid kits, automatic monitoring systems, air purification systems, and short-distance living supplies, etc., without specific limitations here.

[0178] When heavy equipment such as an ambulance bed 44 is placed in the second loading space 421, it can be placed on the axis of the aircraft 1 or arranged symmetrically along the axis to maintain the center of gravity of the aircraft 1 on its axis.

[0179] In some embodiments, the second loading compartment 42 can also be a manned compartment. A plurality of seats are arranged symmetrically along the axis in the second loading space 421 for carrying people, so as to facilitate the transportation of firefighting and rescue personnel, or to facilitate the command of firefighting commanders, thereby realizing the diversity of uses of the aircraft 1.

[0180] The multiple seats are arranged centrally and symmetrically along the axis. Specifically, the multiple seats can be arranged in a circle with the axis as the center. Alternatively, the multiple seats can be arranged in a matrix with the axis as the center line. This is not specifically limited here.

[0181] In some embodiments, at least a portion of the walls of the second loading bin 42 may be made of a transparent material to form a window structure, so that people in the second loading bin 42 can observe the external environment.

[0182] For example, if multiple seats are arranged in a circular pattern within second loading compartment 42, the seats can be positioned facing away from the center of the circle, allowing seated passengers to face the transparent wall of second loading compartment 42 and observe the surrounding landscape. This arrangement can also expand the application areas of aircraft 1. Aircraft 1 can be used for firefighting and rescue operations, facilitating firefighters' observation of disaster situations, as well as for sightseeing tours.

[0183] For example, Figure 1 In the illustrated embodiment, the second loading chamber 42 is entirely made of transparent material.

[0184] Furthermore, when part of the wall of the second loading bin 42 is a transparent wall, the transparent wall can be specifically arranged in a ring shape and arranged along the circumference of the second loading bin 42, so that people in the second loading space 421 can observe outward at any circumferential angle position.

[0185] It should be understood that when part of the warehouse wall of the second loading warehouse 42 is a transparent wall, its transparent wall and the warehouse wall of other parts can be assembled with each other detachably or non-detachably through various connection methods such as welding, gluing, one-piece molding, snap connection, bolt connection, etc., and no specific limitation is made here.

[0186] In some embodiments, the second loading space 421 may also be equipped with at least one of the existing system facilities such as an AI navigation display system, an electric door and window system, an air conditioning system, a communication system, a video system, a food supply system, a medical insurance system, an AI management service system, etc., which will not be elaborated here.

[0187] Among them, when the second loading warehouse 42 is equipped with a system with a display device such as an AI navigation display system and a video system, its display device can be arranged on the warehouse wall of the second loading warehouse 42.

[0188] Furthermore, it can be arranged on the wall of the second loading bin 42 made of transparent material.

[0189] In some other embodiments, the second loading space 421 may also be at least partially empty space for loading objects.

[0190] In some embodiments, the second loading compartment 42 is axially arranged to penetrate one end close to the first mounting frame assembly 151 so that the second loading space 421 is connected to the first cavity 101 to form an integrated space.

[0191] In some embodiments, the electric balancing device may also be disposed in the first loading space 411 or the second loading space 421 or the driving space 150 .

[0192] In other embodiments, the operating system 40 may include only the first loading compartment 41 and the water spray device. This configuration can further lower the center of gravity of the aircraft 1 and improve the aircraft 1's ability to withstand the first natural stability condition. In this embodiment, the second drive system 20 may include a fan mounted on top of the first drive system 10. The first cavity 101 and / or the second cavity 102 may house emergency equipment, seats, and other service equipment.

[0193] In some other embodiments, the operation system 40 may also include only the second loading compartment 42 and the water spraying device. The second loading space 421 can be used to load fire water, and the water spraying device is disposed on the second loading compartment 42 to spray the fire water in the second loading space 421.

[0194] It should be understood that the assembly of the first loading bin 41 and the second mounting frame assembly 152, and the assembly of the second loading bin 42 and the first mounting frame assembly 151 can be assembled to each other detachably or non-detachably through welding, bolt connection, threaded connection, one-piece molding and other processes, and no specific limitation is made here.

[0195] like Figure 1 As shown, in some embodiments, the aircraft 1 further includes a support system. The support system may include at least two support mechanisms 50 for providing balanced support during takeoff and landing of the aircraft 1 to ensure smooth takeoff and landing of the aircraft 1.

[0196] Specifically, the support mechanism 50 can be implemented using existing mechanisms with automatic leveling capabilities, such as existing servo electric cylinders and hydraulic automatic leveling mechanisms. This provides support while also maintaining a stable attitude during takeoff and landing for the aircraft 1. Regardless of whether the ground at the takeoff and landing location of the aircraft 1 is uneven, such as convex, inclined, or concave, the takeoff and landing support system can flexibly adjust the length of each support mechanism 50 accordingly, thereby ensuring that the first drive system 10, operating system 40, and other components of the aircraft 1 are in a level attitude.

[0197] Alternatively, the support mechanism 50 may also adopt various existing cylinder structures such as air cylinders, hydraulic cylinders, etc., so as to play a supporting role while utilizing the compressible property of the fluid, so that the support mechanism 50 can provide elastic buffering at the moment of landing, thereby further improving the landing stability of the aircraft 1.

[0198] Of course, the support mechanism 50 may also adopt various existing support structures such as support columns that do not have a telescopic function, which will not be elaborated in detail here.

[0199] Specifically, the support mechanism 50 may include a servo motor, a cylinder, a sensor, and other structures (not shown). The servo motor is electrically connected to the control system 30 and is used to directly or indirectly drive the cylinder. The cylinder can adjust its length by extending and retracting its piston, thereby adjusting the length of the support mechanism 50. The sensor is electrically connected to the control system 30 and is used to output parameters such as sensed pressure or horizontal inclination to the control system 30, so that the control system 30 can control the servo motor to drive the cylinder piston to extend and retract based on the parameters output by the sensor.

[0200] This arrangement allows the support mechanisms 50 to not only provide support but also maintain a stable attitude for the aircraft 1. Regardless of whether the ground at the takeoff and landing location of the aircraft 1 is uneven, such as convex, inclined, or concave, the support system can flexibly adjust the length of each support mechanism 50 accordingly, thereby ensuring that the first drive system 10, operating system 40, etc. of the aircraft 1 are in a horizontal attitude.

[0201] By providing a cylinder, the compressibility of gas can be utilized, so that the support mechanism 50 can provide elastic buffering at the moment of landing, further improving the landing stability of the aircraft 1 and improving the comfort of the passengers inside.

[0202] In some embodiments, the support mechanism 50 may also include multiple exhaust valves (not shown). These valves are spaced apart along the cylinder cavity wall and automatically open to release the air pressure when the air pressure within the cylinder cavity exceeds a threshold. The exhaust valves can rapidly release axial kinetic energy, further reducing vibration and protecting the aircraft 1.

[0203] It should be understood that the support mechanism 50 can be implemented using an existing mechanism, which will not be elaborated on in detail here.

[0204] It should be understood that the number of support mechanisms 50 provided can be flexibly adjusted based on the configuration of the support mechanism 50. For example, if the horizontal cross-section of the support mechanism 50 has a certain area, there can be one support mechanism 50, coaxially disposed at the bottom of the aircraft 1. Alternatively, there can be two support mechanisms 50, symmetrically spaced along the axis of the aircraft 1 at the bottom. Of course, the number of support mechanisms 50 can also be three, four, or more, evenly distributed at the bottom of the aircraft 1 to enhance support and leveling stability.

[0205] Figure 8 FIG. 1 shows an aircraft 1A according to another embodiment of the present application. The aircraft 1A differs from the aircraft 1 according to the aforementioned embodiments primarily in that the aircraft 1A further includes a volute. The volute defines a duct, and the first drive system 10A is coaxially disposed within the duct, thereby achieving a wingless appearance for the aircraft 1A.

[0206] The volute also serves to gather wind, further enhancing the ideal aerodynamic field generated by aircraft 1. This creates a downwash with orderly streamline distribution, a large, thin tubular profile, a straight downwash pattern, slow downwash velocity changes, a complete, and strong, upright shape. This tubular aerodynamic field exhibits higher airflow volume and intensity, greater resistance to ambient airflow, and improved resilience to various wind shear conditions.

[0207] The volute can also prolong the duration of the wake and improve the efficiency of lift generation. At the same time, since the two sets of wind wings are accommodated in the volute, the volute can also protect the wind wings and improve the safety and reliability of the aircraft 1.

[0208] Specifically, the volute comprises a housing 60A and at least one connecting structure (not shown). The housing 60A has a generally annular outer contour and a hollow interior, thereby reducing the weight of the aircraft 1A while also allowing for the placement of equipment, passenger transport, and other equipment. The housing 60A defines a duct, within which the first drive system 10A is disposed, coaxially with the housing 60A, and secured to the housing 60A via a connecting structure 62A.

[0209] In some embodiments, the housing 60A may be annular.

[0210] It should be understood that the axial dimension of the housing 60A can be compatible with the axial dimension of the first drive system 10A, or slightly larger than the axial dimension of the first drive system 10A, and both sets of wind blades are inside the housing 60A.

[0211] In some other embodiments, the shell 60A may also be configured as a ring of other shapes such as a regular polygon.

[0212] In some embodiments, the connection structure 62A includes at least one first connector, at least one second connector, and multiple connecting rods (not shown). The first and second connectors are both annular. The first connector is fixedly mounted around the frame (not shown) and / or the operating system 40A. The second connector is fixed to the inner circumference of the housing 60A. Multiple connecting rods are spaced circumferentially between the first and second connectors, with their ends fixed to the first and second connectors, respectively, to connect the first drive system 10A to the housing 60A.

[0213] Specifically, the first connecting member is optionally arranged on one of the first mounting frame assembly (not shown in the figure), the second mounting frame assembly (not shown in the figure), the first loading bin (not shown in the figure) and the second loading bin 42A, and the second connecting member is arranged axially corresponding to the first connecting member, which is not specifically limited here.

[0214] Such an arrangement can minimize the influence of the connection structure on the tubular aerodynamic field generated by the first drive system 10A, thereby ensuring the smooth navigation of the aircraft 1A.

[0215] Furthermore, the number of the connection structures may be at least two, and they may be arranged at intervals along the axial direction of the aircraft 1 to improve the stability of the volute assembly.

[0216] In some embodiments, the connecting rods may be evenly spaced along the circumference between the first connecting member and the second connecting member to further reduce the impact on the aerodynamic field.

[0217] In some other embodiments, the connection structure may also include only a plurality of connecting rods, which are evenly or unevenly spaced along the circumference between the housing 60A and the frame and / or the operating system 40A, with their ends fixed to the housing 60A and the frame and / or the operating system 40A, respectively.

[0218] It can be understood that the above technical features can be used in any combination without limitation.

[0219] The above embodiments only express the specific implementation methods of the present application. The descriptions thereof are relatively specific and detailed, but they cannot be understood as limiting the scope of the patent of the present application. It should be pointed out that, for ordinary technicians in this field, the above technical features can be freely combined without departing from the concept of the present application, and several deformations and improvements can be made, which all fall within the scope of protection of the present application. Therefore, all equivalent transformations and modifications made to the scope of the claims of the present application should fall within the scope of coverage of the claims of the present application.

Claims

1. An aircraft, characterized in that: include: The first drive system includes two coaxially arranged drive motors and two sets of conjugately arranged wind blades; The two drive motors each have a coaxial cavity therein; the two groups of wind blades are respectively arranged along the circumference of the rotors of the two drive motors; a second drive system for driving the aircraft to move horizontally, the second drive system being axially arranged within the first drive system; as well as The operating system is defined by a first loading bin and a water spraying device arranged at the bottom end of the first driving system; the first loading bin is defined by a first loading space for storing water; the nozzle of the water spraying device is arranged in the first loading bin and is connected to the first loading space.

2. The aircraft according to claim 1, characterized in that The operating system further includes a second loading chamber coaxially arranged at the top end of the first driving system, wherein the second loading chamber defines a second loading space.

3. The aircraft according to claim 2, characterized in that Rescue equipment is arranged in the second loading space; And / or, the second loading space is provided with seats in a centrally symmetrical manner.

4. The aircraft according to claim 1, characterized in that At least a portion of the wall of the second loading bin is a transparent wall.

5. The aircraft according to claim 4, characterized in that An AI navigation system is arranged in the second loading compartment; and a display structure of the AI ​​navigation system is arranged on the transparent wall.

6. The aircraft according to claim 1, characterized in that At least two support mechanisms are provided at the bottom end of the first loading bin, the support height of which can be adjusted according to the flatness of the ground.

7. The aircraft according to claim 6, characterized in that The supporting mechanism includes a cylinder and a servo motor that drives the cylinder to extend and retract; a plurality of exhaust valves are provided on the cylinder wall of the cylinder; the exhaust valves are used to open and exhaust when the air pressure in the cylinder is greater than a threshold value.

8. The aircraft according to claim 1, characterized in that The first loading bin is made of a material comprising a polyimide composite material.

9. The aircraft according to claim 1, characterized in that A lighting system is provided at the bottom end of the first loading bin.

10. The aircraft according to any one of claims 1 to 9, characterized in that Also included is a control system, at least a portion of which is disposed within the cavity.

11. The aircraft according to any one of claims 1 to 9, characterized in that It also includes a volute; the first drive system is coaxially arranged in the volute.

12. The aircraft according to any one of claims 1 to 9, characterized in that The ratio of the outer diameter of the drive motor to the axial length of the drive motor is greater than or equal to 10 and less than or equal to 300.

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