Light electric aircraft flight control system and method

Abstract

The invention discloses a light electric aircraft flight control system and method, and belongs to the technical field of electric aircraft control. The number of the control assemblies is multiple, and the multiple control assemblies are arranged on the electric aircraft rack at intervals in the circumferential direction; the plurality of control panels are arranged in one-to-one correspondence with the execution ends of the plurality of control assemblies, and the wind speed sensor is used for acquiring current wind speed data; the wind direction sensor is used for acquiring current wind direction data; and the airframe attitude measurement unit is used for acquiring the current attitude data of the light electric aircraft. According to the method, the wind speed data is monitored in real time, the wind direction data is collected periodically, the wind field interference level is judged in combination with the preset threshold value, an attitude adjustment mechanism can be rapidly triggered when the wind speed exceeds the standard or the wind direction suddenly changes, and the light electric aircraft can complete operation under more complex meteorological conditions.

Description

TECHNICAL FIELD

[0001] The application discloses a light electric aircraft flight control system and method, and belongs to the technical field of electric aircraft control. BACKGROUND

[0002] With the development of low-altitude economy and unmanned aerial vehicle technology, light electric aircrafts are widely used in aerial photography, logistics transportation, emergency rescue and other scenes due to their low energy consumption, low noise and flexible operation. However, light electric aircrafts are more sensitive to external environmental disturbances (especially wind speed and direction changes) due to their light weight and limited power reserves, and are prone to problems such as unstable attitude and excessive center of gravity shift.

[0003] Traditional flight control systems mostly adopt centralized control architecture, and the response speed to wind field changes is slow. In complex wind conditions (such as gusts and sudden changes in wind direction), it is difficult to quickly adjust the output of each power component to maintain the balance of the fuselage, which may lead to a decrease in flight accuracy or even a risk of loss of control. In addition, existing flight control methods do not adequately consider the coupling relationship between center of gravity shift and wind field disturbance, and the adaptability and reliability of the control strategy need to be improved. SUMMARY

[0004] The purpose of the present application is to solve the problem that the existing traditional flight control system mostly adopts centralized control architecture and is difficult to quickly adjust the output of each power component to maintain the balance of the fuselage, and to propose a light electric aircraft flight control system and method.

[0005] The technical solution of the present application is as follows: According to the first aspect of the embodiment of the present application, a light electric aircraft flight control system is provided, comprising: an electric aircraft frame; a control component, the control component comprising a plurality of control components, the plurality of control components being circumferentially spaced apart on the electric aircraft frame; a control board, the control board comprising a plurality of control boards, the plurality of control boards being one-to-one corresponding to the execution end of the plurality of control components; a wind speed sensor, the wind speed sensor being used to obtain current wind speed data; a wind direction sensor, the wind speed sensor being used to obtain current wind direction data; and a body attitude measurement unit, the body attitude measurement unit being used to obtain current attitude data of the light electric aircraft.

[0006] Further, the light electric aircraft flight control system further comprises: a first support leg, the first support leg being spaced apart from the control component and the control board, one end of the first support leg being connected to the electric aircraft frame, the first support leg being arranged at an angle with the electric aircraft frame; and a second support leg, the second support leg being spaced apart from the first support leg, the control component and the control board, one end of the second support leg being connected to the electric aircraft frame, the second support leg being arranged at an angle with the electric aircraft frame, and the second support leg being arranged at an angle with the first support leg.

[0007] Further, the electric aircraft fuselage comprises: an electric aircraft fuselage body, the electric aircraft fuselage body having a plurality of connection ends, the plurality of connection ends being arranged in a circumferential direction along the electric aircraft fuselage body; and a plurality of connecting plates, the plurality of connecting plates being arranged in one-to-one correspondence with the plurality of connection ends, one end of each connecting plate being connected to a connection end.

[0008] Further, the control assembly comprises: a first sliding rail, the first sliding rail being connected to the connecting plate and extending in a first preset direction; a second sliding rail, the second sliding rail being connected to the connecting plate and being arranged on the same side of the first sliding rail, the second sliding rail being arranged in a spaced manner with the first sliding rail and extending in the first preset direction; a support frame, the support frame being connected to the connecting plate and being arranged on the same side of the second sliding rail and the first sliding rail, respectively, the support frame being arranged in a spaced manner with the second sliding rail and the first sliding rail, respectively; a first sliding block, the first sliding block being movably connected to the first sliding rail, the first sliding block being arranged in a spaced manner with the second sliding rail and the support frame, respectively; a second sliding block, the second sliding block being movably connected to the second sliding rail, the second sliding block being arranged in a spaced manner with the first sliding rail, the first sliding block, and the support frame, respectively; a motor support frame, the motor support frame being connected to the connecting plate, the motor support frame being arranged on the same side of the second sliding rail and the first sliding rail, respectively, the motor support frame being arranged close to one of the second sliding rail and the first sliding rail; a drive motor, the drive motor being connected to the motor support frame; a drive lead screw, one end of the drive lead screw being connected to a main shaft of the drive motor, the drive lead screw being movably connected to the first sliding block and the first sliding block, respectively; a control shaft, the control shaft being movably connected to the support frame; a drive gear, the drive gear being sleeved on the control shaft, the drive gear being engaged with the drive lead screw, one end of at least one control plate being connected to one end of the drive lead screw, the drive lead screw being controlled to rotate so as to drive the drive gear to rotate and drive the control plate to rotate by a corresponding angle.

[0009] Further, the control assembly further comprises: a connecting plate, the first sliding block and the second sliding block being connected through the connecting plate.

[0010] According to a second aspect of the embodiments of the present application, a flight control method of a light electric aircraft is provided, applied to the flight control system of the light electric aircraft of the first aspect, and includes: in response to a flight start signal of the light electric aircraft, acquiring current wind speed data and periodically acquiring wind direction data; based on the current wind speed data and the wind direction data, determining a first determination result by using a preset wind speed value range and a wind direction change value range; in response to the first determination result being that the current wind speed data is greater than the preset wind speed value range, or a current wind direction change angle value obtained by using wind direction data at a current time and wind direction data at a previous time in the wind direction data is greater than the wind direction change value range, acquiring current attitude data; based on the current attitude data, obtaining a current center of gravity offset; based on the current center of gravity offset, determining a second determination result by using a preset center of gravity offset; and in response to the second determination result being that the current center of gravity offset is greater than the preset center of gravity offset, generating a control instruction set, which is used to control the driving motor to perform a corresponding target action.

[0011] Further, the flight control method of the light electric aircraft further includes: in response to the first determination result being that the current wind speed data is less than or equal to the preset wind speed value range, or the wind direction change angle value is less than or equal to the wind direction change value range, repeatedly acquiring the current wind speed data and the wind direction data; and / or in response to the second determination result being that the current center of gravity offset is less than or equal to the preset center of gravity offset, repeatedly acquiring the current wind speed data and the wind direction data.

[0012] Further, the control instruction set includes: based on the current wind speed data, the wind direction data and the current center of gravity offset, obtaining a movement displacement of each first slider and second slider in the plurality of control components and a rotation angle of each control plate in the plurality of control plates; and based on the movement displacement of each first slider and second slider in the plurality of control components and the rotation angle of each control plate in the plurality of control plates, determining a motor step number of a corresponding driving motor, and sending the motor step number to the corresponding driving motor.

[0013] Further, the control instruction set includes: based on the current wind speed data, the wind direction data and the current center of gravity offset, obtaining a movement displacement of each first slider and second slider in the plurality of control components and a rotation angle of each control plate in the plurality of control plates; and based on the movement displacement of each first slider and second slider in the plurality of control components and the rotation angle of each control plate in the plurality of control plates, determining a motor step number of a corresponding driving motor, and sending the motor step number to the corresponding driving motor.

[0014] The present application has the following advantages: The application provides a light electric aircraft flight control system and method, which can quickly trigger an attitude adjustment mechanism when the wind speed exceeds the standard or the wind direction suddenly changes by monitoring the wind speed data in real time and periodically collecting the wind direction data, combining the preset threshold to judge the wind field interference level, significantly enhancing the stability of the aircraft in complex wind conditions, calculating the gravity center offset by obtaining the current attitude data, and generating a targeted control instruction based on the preset threshold, which can effectively compensate the influence of the gravity center offset on the balance of the fuselage, improve the flight attitude control precision, adopt a distributed architecture of one-to-one correspondence between multiple control components and control boards, reduce the data transmission delay, realize the rapid control of the driving motor, reduce the single point failure risk through the redundant design of multiple components, and improve the overall reliability of the system, and through the optimization of the wind field adaptability and the attitude control performance, the light electric aircraft can complete the work under more complex weather conditions.

[0015] It should be understood that the above general description and the following detailed description are only exemplary and explanatory, and cannot limit the application. BRIEF DESCRIPTION OF DRAWINGS

[0016] Figure 1 is a structural diagram of a first embodiment of a light electric aircraft flight control system according to an exemplary embodiment.

[0017] Figure 2 is a partial structural diagram of a second embodiment of a light electric aircraft flight control system according to an exemplary embodiment.

[0018] Figure 3 is a partial structural diagram of a third embodiment of a light electric aircraft flight control system according to an exemplary embodiment.

[0019] Figure 4 is a structural diagram of a fourth embodiment of a light electric aircraft flight control system according to an exemplary embodiment.

[0020] Figure 5 is a flowchart of a light electric aircraft flight control method according to an exemplary embodiment. DETAILED DESCRIPTION

[0021] The technical solutions of the application will be described clearly and completely below with reference to the drawings. Obviously, the described embodiments are part of the embodiments of the application, rather than all the embodiments. Based on the embodiments in the application, all other embodiments obtained by those skilled in the art without creative labor fall within the protection scope of the application.

[0022] In the description of the present invention, it should be noted that the terms "center", "up", "down", "left", "right", "vertical", "horizontal", "inside", "outside", etc., indicating directions or positional relationships, are based on the directions or positional relationships shown in the accompanying drawings. They are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific direction, be constructed and operated in a specific direction. Therefore, they cannot be understood as limiting the present invention.

[0023] In the description of the present invention, it should be noted that, unless otherwise expressly specified or limited, the terms "mounted," "connected," and "connected" should be understood in a broad sense. For example, they may refer to fixed, detachable, or integral connections; mechanical or electrical connections; direct or indirect connections through an intermediate medium; and internal communication between two components. Those skilled in the art will understand the specific meanings of the above terms in the present invention based on the specific circumstances.

[0024] Example 1: Figure 1 As shown, a light electric aircraft flight control system according to an exemplary embodiment is shown, comprising: an electric aircraft frame 1; a control component 2, the control component 2 including multiple control components 2, the multiple control components 2 being circumferentially spaced apart and arranged on the electric aircraft frame 1; a control board 4, the control board 4 including multiple control boards 4, the multiple control boards 4 being arranged in a one-to-one correspondence with the execution ends of the multiple control components 2; a wind speed sensor, the wind speed sensor being used to obtain current wind speed data; a wind direction sensor, the wind speed sensor being used to obtain current wind direction data; and an airframe attitude measurement unit, the airframe attitude measurement unit being used to obtain current attitude data of the light electric aircraft.

[0025] In this embodiment, multiple control boards 4 correspond one-to-one with the execution terminals of control components 2, forming a distributed control architecture. This avoids information congestion in a single processing unit in centralized control, enabling rapid response to control commands and reducing signal transmission and processing delays. This makes it particularly suitable for flight scenarios requiring high real-time performance. The configuration of multiple control components and control boards creates a natural redundancy structure. If a control component or control board fails, other functioning units can compensate to a certain extent, reducing the risk of system failure caused by a single point of failure and improving flight safety.

[0026] The control assembly is circumferentially spaced apart on the frame and can exert a control force on the aircraft from multiple directions. The wind speed sensor captures real-time wind speed data, and the wind direction sensor dynamically obtains wind direction change information. The combination of the two provides precise wind field environment parameters for the flight control system. This enables the aircraft to timely perceive complex wind conditions such as gusts and sudden changes in wind direction, providing reliable data support for subsequent wind-resistant control strategies and avoiding attitude instability due to insufficient environmental perception. The body attitude measurement unit uses an inertial sensor, which can obtain real-time attitude data of the aircraft, including pitch angle, roll angle, heading angle, and other key parameters. These data directly reflect the balance state and spatial position of the fuselage and are the core basis for the flight control system to judge the center of gravity deviation and develop attitude adjustment schemes, ensuring that the control command can accurately match the actual state of the fuselage. The wind speed and direction data obtained by the sensor and the body attitude data form a complete perception closed loop, and the flight control system can perform comprehensive analysis based on these real-time data.

[0027] Further, as shown in Figure 2 The light electric aircraft flight control system further includes: a first support leg 3, which is spaced apart from the control assembly 2 and the control board 4, one end of the first support leg 3 is connected to the electric aircraft frame 1, and the first support leg 3 is arranged at an angle with the electric aircraft frame 1; a second support leg 4, which is spaced apart from the first support leg 3, the control assembly 2, and the control board 4, one end of the second support leg 4 is connected to the electric aircraft frame 1, and the second support leg 4 is arranged at an angle with the electric aircraft frame 1, and the second support leg 4 is arranged at an angle with the first support leg 3. The first support leg 3 and the second support leg 4 are connected at an angle with the electric aircraft frame 1, and an angle is also formed between them, forming a stable triangular structure (or a triangular support system) of multi-directional support. This spatial layout can disperse the weight of the fuselage and external loads to multiple support points, significantly improving the overall structural stiffness and anti-deformation ability of the frame, especially effectively buffering the ground impact force during the take-off and landing stages.

[0028] In this example, the electric aircraft frame 1 includes: an electric aircraft frame body 102 and a connecting plate 101, the electric aircraft frame body 102 has a plurality of connecting ends, the plurality of connecting ends are circumferentially spaced apart along the electric aircraft frame body 102, and the connecting plate 101 includes a plurality of connecting plates 101, which are arranged one by one corresponding to the plurality of connecting ends, and one end of the connecting plate 101 is connected to the connecting end. The connecting ends are circumferentially spaced apart along the frame body 102, and cooperate with the corresponding connecting plates 101 to uniformly disperse the forces transmitted by each component (such as the control assembly 2 and the support leg) to the frame body 102, avoiding local stress concentration, enhancing the overall structural strength and fatigue resistance of the frame, and prolonging the service life.

[0029] In one exemplary embodiment, as shown in Figure 3 and Figure 4As shown, the control assembly 2 comprises: a first sliding rail 201, a second sliding rail 211, a support frame 207, a first sliding block 202, a second sliding block 204, a motor support frame 206, a drive motor 205, a drive lead screw 208, a control shaft 210 and a drive gear 209, the first sliding rail 201 is connected with the connecting plate 101, the first sliding rail 201 is arranged in extension along a first preset direction, the second sliding rail 211 is connected with the connecting plate 101, the second sliding rail 211 is arranged on the same side of the first sliding rail 201, the second sliding rail 211 is arranged in interval with the first sliding rail 201, the second sliding rail 211 is arranged in extension along the first preset direction, the support frame 207 is connected with the connecting plate 101, the support frame 207 is arranged on the same side of the second sliding rail 211 and the first sliding rail 201 respectively, the support frame 207 is arranged in interval with the second sliding rail 211 and the first sliding rail 201 respectively, the first sliding block 202 is movably connected with the first sliding rail 201, the first sliding block 202 is arranged in interval with the second sliding rail 211 and the support frame 207 respectively, the second sliding block 204 is movably connected with the second sliding rail 211, the second sliding block 204 is arranged in interval with the first sliding rail 201, the first sliding block 202 and the support frame 207 respectively, the motor support frame 206 is connected with the connecting plate 101, the motor support frame 206 is arranged on the same side of the second sliding rail 211 and the first sliding rail 201 respectively, the motor support frame 206 is arranged close to one of the second sliding rail 211 and the first sliding rail 201, the drive motor 205 is connected with the motor support frame 206, one end of the drive lead screw 208 is connected with the spindle of the drive motor 205, the drive lead screw 208 is movably connected with the first sliding block 202 and the first sliding block 202 respectively, the control shaft 210 is movably connected with the support frame 207, the drive gear 209 is sleeved on the control shaft 210, the drive gear 209 is engaged with the drive lead screw 208, one end of at least one control board 4 is connected with one end of the drive lead screw 208, the drive lead screw 208 is controlled to rotate, so that the drive gear 209 rotates to drive the control board 4 to rotate by a corresponding angle.

[0030] By applying the embodiment, the first sliding rail 201 and the second sliding rail 211 are arranged in interval and parallel along the same preset direction, and are fixedly connected with the connecting plate 101, cooperating with the auxiliary support of the support frame 207, a rigid track foundation is provided for the sliding of the first sliding block 202 and the second sliding block 204, effectively reducing the shaking or deviation in the movement process. This double-sliding rail guiding design greatly improves the straightness and stability of the assembly movement, providing a basic guarantee for the subsequent transmission precision.

[0031] The drive motor 205 directly drives the lead screw 208, which then meshes with the drive gear 209, ultimately driving the control shaft 210 and the control board 4 to rotate, forming a rigid "motor-lead screw-gear-shaft" transmission chain. This mechanical transmission method eliminates losses in the flexible links and offers high transmission efficiency. Furthermore, by controlling the speed and direction of the drive motor, the rotation angle of the control board 4 can be precisely controlled (e.g., for small angle adjustments), meeting the high-precision attitude fine-tuning requirements of the flight control system.

[0032] The first slider 202 and the second slider 204 move on independent slide rails respectively, and the components are spaced apart to avoid motion interference from a structural perspective. At the same time, the movable connection design of the drive screw 208 and the double sliders can evenly disperse the forces during the transmission process, reduce jamming or wear caused by unilateral force, and ensure smooth movement of the entire component with strong anti-interference ability.

[0033] Furthermore, the control assembly 2 further includes a connecting plate 203, through which the first slider 202 and the second slider 204 are connected. The connecting plate 203 rigidly connects the two independent sliders, ensuring that the movement of the first slider 202 and the second slider 204 on their respective slide rails (the first slide rail 201 and the second slide rail 211) is strictly synchronized, avoiding movement misalignment caused by gaps between the two rails, differential wear of the sliders, etc., ensuring balanced force on the drive screw 208 and subsequent transmission components, and reducing additional torque or vibration caused by asynchronous movement.

[0034] Example 2: Figure 5 A light electric aircraft flight control method is shown according to an exemplary embodiment, which is implemented by a terminal, which includes at least a CPU. The method includes: Step S101, in response to a light electric aircraft flight start signal, obtaining current wind speed data and periodically obtaining wind direction data; In step S101, when the flight control system receives a flight start signal (such as a remote control takeoff command or an autonomous flight program start command), it immediately activates the environmental perception module. The wind speed sensor prioritizes real-time sampling, quickly acquiring the instantaneous wind speed value in the current airspace in meters per second (m / s), which serves as baseline data for the initial wind field conditions. Simultaneously, the wind direction sensor simultaneously activates and continuously collects wind direction information at a preset interval, such as every 0.5 seconds or 1 second. Wind direction information is typically expressed as an angle between 0° and 360°, relative to true north.

[0035] Step S102, based on the current wind speed data and multiple wind direction data, a judgment is performed using a preset wind speed value range and a wind direction change value range to obtain a first judgment result; In step S102, The instantaneous wind speed value obtained by the wind speed sensor at the start time is denoted as V, with the unit of m / s, representing the wind intensity of the current airspace.

[0036] A plurality of wind direction data: a continuous wind direction sequence collected at a preset period, denoted as , with the unit of °, wherein is the earliest collected value, is the current latest value, used to reflect the dynamic change trend of the wind direction.

[0037] The safe wind speed range according to the wind resistance performance of the light electric aircraft is denoted as [V min , V max ]. For example, in combination with the parameters such as the body weight and rotor power, the upper limit of the safe wind speed V max =10 m / s is set, and exceeding this value may lead to attitude loss of control, and the lower limit V min =0 m / s is set, which is the state of no wind or slight wind.

[0038] The allowed wind direction change angle range set based on the flight stability requirement is denoted as [Δθ min , Δθ max ]. For example, Δθ max =15° is set, and if the wind direction changes more than 15° within a short time, it is determined as “sudden change of wind direction”, and Δθ min =0°, which means no change of wind direction.

[0039] By converting the abstract “wind field state” into calculable wind speed value and wind direction change angle, in combination with the preset threshold, the interference degree is accurately quantified, the subjectivity of relying on experience judgment is avoided, and the evaluation of the environment by the flight control system is more objective. The first judgment result directly determines whether the flight control system needs to start the strong interference response mechanism: if it is “strong interference state of wind field”, the subsequent attitude monitoring and gravity center adjustment process is triggered; if it is “steady state of wind field”, the normal flight control logic is maintained, and unnecessary energy consumption and adjustment actions are reduced.

[0040] In step S103, in response to the first judgment result being that the current wind speed data is greater than the preset wind speed value range, or, the current wind direction change angle value is obtained by the wind direction data at the current time and the wind direction data at the last time from the plurality of wind direction data, and the current wind direction change angle value is greater than the wind direction change value range, the current attitude data is obtained. In step S103, the current wind speed data V is directly compared with the preset wind speed value range [V min , V max ]: If V>V max : it is determined as “excessive wind speed”, which belongs to the strong interference state of wind field.

[0041] If V≤V max: The wind speed is within the safe range, and the next step is to determine the wind direction change.

[0042] In step S103, a "wind direction change angle value" is calculated based on multiple wind direction data and compared with the wind direction change value range, specifically including: Calculate the wind direction difference between adjacent moments: take the latest two wind direction data (current moment) and (previous moment), calculate the absolute difference If Δθ>180°, for example, if the angle changes suddenly from 350° to 10°, the actual change is 20°, then 360°-Δθ is taken as the effective change angle to avoid misjudgment due to the circular angle characteristics.

[0043] Comparison with the wind direction change range: If Δθ>Δθ max If the calculated Δθ=20°>preset 15°, it is determined as "wind direction mutation", which belongs to the state of strong wind field interference; if Δθ≤Δθ max , it is judged that the wind direction changes steadily.

[0044] Based on the results of wind speed judgment and wind direction change judgment, the first judgment results are divided into two categories: Strong wind interference status: This result is output when any of the following conditions are met, triggering the subsequent attitude monitoring and center of gravity adjustment process to obtain the current attitude data: Current wind speed V> preset wind speed range upper limit V max ; Current wind speed V≤V max , but the wind direction change angle Δθ> the preset wind direction change value range upper limit Δθ max .

[0045] Wind farm steady state: This result is output when the following conditions are met at the same time: Current wind speed V≤preset wind speed range upper limit V max ; Wind direction change angle Δθ≤preset wind direction change value upper limit Δθ max .

[0046] Step S104, obtaining the current center of gravity offset based on the current posture data; In S104, the current attitude data is collected in real time by an aircraft attitude measurement unit (such as an IMU inertial measurement unit, which includes an accelerometer, gyroscope, magnetometer, etc.). The center of gravity offset is usually expressed as "the offset distance (Δx, Δy) of the actual center of gravity position relative to the design reference point" or "the offset angle (γ)". The calculation process needs to combine the structural parameters of the aircraft (such as wheelbase, rotor layout) and the mechanical model to obtain the current center of gravity offset. This process is common knowledge for those skilled in the art and will not be described in detail.

[0047] When the aircraft is in acceleration or deceleration flight state, the motion acceleration will interfere with the attitude data. At this time, through the angular velocity data and the preset flight state model (such as hovering, flat flying, climbing), the "attitude deviation caused by the center of gravity deviation" and the "attitude adjustment caused by the active flight control" are distinguished, the influence of the motion acceleration on the center of gravity calculation is eliminated, and finally the accurate current center of gravity deviation is obtained.

[0048] In step S105, based on the current center of gravity deviation, the preset center of gravity deviation is judged to obtain a second judgment result.

[0049] In step S105, the current center of gravity deviation is usually expressed as a three-dimensional coordinate deviation (Δx, Δy, Δz) or an angle deviation (such as a roll deviation angle γ, a pitch deviation angle β), reflecting the imbalance degree of the fuselage in space. For example, Δx = 0.15m indicates that the center of gravity deviates from the reference point by 15cm in the X-axis direction, and γ = 3° indicates that the fuselage deviates by 3° in the roll direction.

[0050] Based on the acceleration, angular velocity and other physical quantities in the attitude data, the center of gravity deviation is solved through a dynamic model such as Newton-Euler equation, the interference of flight attitude adjustment is eliminated, and only the passive deviation caused by wind field interference or load change is reserved.

[0051] It refers to the maximum allowed center of gravity deviation range preset according to the structural strength, flight stability requirements and safety threshold of the light electric aircraft, which is the basis for judging whether the fuselage is in an "acceptable balanced state". Combined with the parameters such as the weight of the aircraft, the payload, the rotor lift characteristics and the like, the simulation and flight test verification are determined. For example: in static hovering, the maximum coordinate deviation is ±0.1m (Δx, Δy), and the angle deviation is ±2° (γ, β); in dynamic flight (such as flat flying, climbing), because the air flow disturbance is larger, the preset threshold can be relaxed to ±0.15m, ±3°.

[0052] The judgment process is realized through "quantitative comparison of actual deviation and preset threshold", and the specific steps are as follows: The system obtains the current center of gravity deviation after filtering processing, such as Δx = 0.12m, γ = 2.5°, and simultaneously calls the preset center of gravity deviation threshold in the corresponding flight mode, such as Δx max = 0.1m, γ max = 2°.

[0053] The multiple dimensions of the center of gravity deviation are respectively subjected to threshold verification: If it is a coordinate deviation: judge whether Δx exceeds the preset Δx max , whether Δy exceeds the preset Δy max , and whether Δz exceeds the preset Δz max ; If it is an angle deviation: judge whether the roll angle γ exceeds the preset γmax whether the pitch angle β exceeds a preset β max .

[0054] According to the comparison result, the second judgment result is divided into two categories: when the current center of gravity offset is greater than the preset center of gravity offset: that is, the offset in at least one dimension exceeds the safe range (for example, Δx = 0.12m > Δx max = 0.1m, or γ = 2.5° > γ max = 2°), indicating that the balance of the fuselage has been significantly disturbed, and the adjustment mechanism needs to be started immediately.

[0055] when the current center of gravity offset is less than or equal to the preset center of gravity offset: the offset in all dimensions is within the safe threshold, and the fuselage is in an acceptable balance state, and no strong adjustment action needs to be started, only regular monitoring needs to be maintained.

[0056] The second judgment result is the key basis for the flight control system to generate control instructions: when the result is that the current center of gravity offset is greater than the preset center of gravity offset, step S106 is triggered, and the system calculates the moving displacement of the sliders of each control component and the rotation angle of the control board based on the current wind speed, wind direction and center of gravity offset, generates the step number or rotation angle instruction of the driving motor, and compensates the center of gravity offset by adjusting the attitude of the rotor to restore the balance of the fuselage.

[0057] when the result is that the current center of gravity offset is less than or equal to the preset center of gravity offset, the system returns to the wind field monitoring link (repeatedly acquires wind speed and wind direction data) to avoid unnecessary adjustment actions and reduce energy consumption and mechanical wear and tear.

[0058] Step S106, in response to the second judgment result that the current center of gravity offset is greater than the preset center of gravity offset, a control instruction set is generated, and the control instruction set is used to control the driving motor to perform a corresponding target action.

[0059] In S106, based on the current wind speed data, wind direction data and current center of gravity offset, the moving displacement of each first slider and second slider in multiple control components is obtained, and the rotation angle of each control board of multiple control boards is obtained. In S106, the direction and size of the interference of the wind on the fuselage are determined by the wind speed and wind direction, and the degree and direction of the imbalance of the fuselage are determined in combination with the center of gravity offset, and finally the “opposite correction force / torque” needed to be generated by each control component is calculated, and then the correction force / torque is converted into the moving displacement of the slider and the rotation angle of the control board through the mechanical structure parameters of the control component, so as to offset the interference of the wind field and correct the center of gravity offset.

[0060] Firstly, the total interference that the fuselage needs to offset is determined in combination with the wind speed, wind direction and the center of gravity offset, including: calculating the wind force in the fuselage coordinate system according to the wind speed V and the wind direction θ. The correction moment required to restore the center of gravity to the reference position is calculated according to the center of gravity offset ΔG. For example, the right deviation of the center of gravity requires the left control assembly to generate additional upward lift to form a left moment to pull back the center of gravity. Since multiple control assemblies are arranged circumferentially, the total correction force / moment needs to be proportionally distributed to the control assembly in the corresponding direction according to the interference direction and the center of gravity offset direction. The first slider 202 and the second slider 204 move synchronously on the first slide rail 201 and the second slide rail 211 through the connecting plate 203, and the displacement directly affects the “force arm length” or “force point position” of the control assembly, thereby changing the effect of the correction force. The movement of the slider is driven by the drive screw 208, and the pitch (P) of the screw determines the relationship between the displacement and the rotation of the motor. If the displacement of the slider needs to be d, the number of turns n of the screw needs to be d / P, and then the displacement of the slider is converted.

[0061] According to the correction moment allocated to the control assembly, in combination with the lift that the current control plate can provide, and according to the current position of the slider, the displacement that the slider needs to move is obtained. The rotation angle of the control plate 4 directly affects the lift or resistance it generates, and is the core execution parameter for offsetting the wind force and correcting the center of gravity. The lift of the control plate is positively correlated with the angle between the rotation angle and the reference plane. According to the correction force allocated to the control assembly, the angle at which the control plate needs to be rotated can be directly obtained.

[0062] In S106, based on the movement displacement of each first slider and second slider in the plurality of control assemblies, and the rotation angle of each control plate of the plurality of control plates, the motor step number of the corresponding drive motor is determined, and the motor step number is sent to the corresponding drive motor.

[0063] Further, the movement displacement of each first slider and second slider in the plurality of control assemblies, and the rotation angle of each control plate of the plurality of control plates, obtain the motor step number of the corresponding drive motor through formula (1): (1) Wherein, is the movement displacement of each first slider and second slider, N is the motor step number, is the total amount of pulses received and executed by the motor in one movement cycle, P is the pitch of the screw, is the motor step subdivision coefficient, r is the reduction ratio coefficient, the reduction ratio coefficient of the drive gear and the drive screw, is the transmission error coefficient, which reflects the comprehensive error caused by the reverse clearance, elastic deformation or backlash of the screw.

[0064] In one example embodiment, in S106, based on the current wind speed data, wind direction data and the current center of gravity offset, the movement displacement of each first slider and second slider in the plurality of control components is obtained, and the rotation angle of each control plate of the plurality of control plates is obtained. Based on the movement displacement of each first slider and second slider in the plurality of control components and the rotation angle of each control plate of the plurality of control plates, the motor rotation angle of the corresponding driving motor is determined, and the motor rotation angle is sent to the corresponding driving motor.

[0065] Further, based on the movement displacement of each first slider and second slider in the plurality of control components and the rotation angle of each control plate of the plurality of control plates, the motor rotation angle is obtained by formula (2): (2) Wherein, is the movement displacement of each first slider and second slider, P is the pitch of the lead screw, is the motor step subdivision coefficient, r is the reduction ratio coefficient, the reduction ratio coefficient of the driving gear and the driving lead screw, is the transmission error coefficient, used to reflect the comprehensive error caused by the reverse clearance, elastic deformation or backlash of the lead screw, is the motor rotation angle, is the corresponding pulse number of the motor per revolution.

[0066] By applying the embodiment, through real-time monitoring of wind speed data and periodic acquisition of wind direction data, the wind field interference level is judged in combination with the preset threshold value, the attitude adjustment mechanism can be quickly triggered when the wind speed exceeds the standard or the wind direction suddenly changes, the stability of the aircraft in complex wind conditions is significantly enhanced, the influence of the center of gravity offset on the balance of the fuselage is effectively compensated by obtaining the current attitude data to calculate the center of gravity offset and generating targeted control instructions based on the preset threshold value, the flight attitude control precision is improved, the distributed architecture of one-to-one correspondence between the plurality of control components and the control plates is adopted, the data transmission delay is reduced, the quick control of the driving motor is realized, the single point failure risk is reduced by the multi-component redundant design, and the overall reliability of the system is improved. By optimizing the wind field adaptability and attitude control performance, the light electric aircraft can complete the work under more complex meteorological conditions.

[0067] Although the embodiments of the application have been disclosed as above, it is not limited to the application listed in the specification and the embodiments. It can be fully applied to various fields suitable for the application. Additional modifications can be easily realized by those skilled in the art. Therefore, the application is not limited to specific details and the figures shown and described herein without departing from the general concept defined by the claims and the equivalent scope.

Claims

1. A light electric aircraft flight control system, characterized in that: include: Electric aircraft frame (1); A control component (2), the control component (2) comprising a plurality of control components (2), the plurality of control components (2) being arranged on the electric aircraft frame (1) at circumferential intervals; A control panel (4), the control panel (4) comprising a plurality of panels, the plurality of control panels (4) being arranged in a one-to-one correspondence with the execution ends of the plurality of control components (2); A wind speed sensor is used to obtain current wind speed data; Wind direction sensor, the wind speed sensor is used to obtain current wind direction data; The body attitude measurement unit is used to obtain current attitude data of the light electric aircraft.

2. The light electric aircraft flight control system according to claim 1, characterized in that: The light electric aircraft flight control system further includes: a first supporting foot (3), the first supporting foot (3) being spaced apart from the control assembly (2) and the control panel (4), one end of the first supporting foot (3) being connected to the electric aircraft frame (1), and the first supporting foot (3) being arranged at an angle to the electric aircraft frame (1); A second supporting foot (4), the second supporting foot (4) is spaced apart from the first supporting foot (3), the control assembly (2) and the control panel (4), one end of the second supporting foot (4) is connected to the electric aircraft frame (1), the second supporting foot (4) and the electric aircraft frame (1) are arranged at an angle, and the second supporting foot (4) and the first supporting foot (3) are arranged at an angle.

3. The light electric aircraft flight control system according to claim 1 or 2, characterized in that: The electric aircraft frame (1) comprises: An electric aircraft frame body (102), the electric aircraft frame body (102) having a plurality of connection ends, the plurality of connection ends being arranged at intervals along the circumference of the electric aircraft frame body (102); A connecting plate (101), comprising a plurality of connecting plates (101), the plurality of connecting plates (101) being arranged in a one-to-one correspondence with the plurality of connecting ends, and one end of the connecting plate (101) being connected to the connecting end.

4. The light electric aircraft flight control system according to claim 3, characterized in that: The control component (2) comprises: a first slide rail (201), the first slide rail (201) being connected to the connecting plate (101), and the first slide rail (201) extending along a first preset direction; a second slide rail (211), the second slide rail (211) being connected to the connecting plate (101), the second slide rail (211) being arranged on the same side as the first slide rail (201), the second slide rail (211) being spaced apart from the first slide rail (201), and the second slide rail (211) extending along the first preset direction; a support frame (207), the support frame (207) being connected to the connecting plate (101), the support frame (207) being arranged on the same side as the second slide rail (211) and the first slide rail (201), and the support frame (207) being spaced apart from the second slide rail (211) and the first slide rail (201); a first slider (202), the first slider (202) being movably connected to the first slide rail (201), and the first slider (202) being spaced apart from the second slide rail (211) and the support frame (207); a second slider (204), the second slider (204) being movably connected to the second slide rail (211), and the second slider (204) being spaced apart from the first slide rail (201), the first slider (202) and the support frame (207); a motor support frame (206), the motor support frame (206) being connected to the connecting plate (101), the motor support frame (206) being arranged on the same side as the second slide rail (211) and the first slide rail (201), and the motor support frame (206) being arranged close to one of the second slide rail (211) and the first slide rail (201); A driving motor (205), the driving motor (205) being connected to the motor support frame (206); A driving screw (208), one end of which is connected to the main shaft of the driving motor (205), and the driving screw (208) is movably connected to the first slider (202) and the first slider (202); a control shaft (210), the control shaft (210) being movably connected to the support frame (207); A driving gear (209) is sleeved on the control shaft (210), the driving gear (209) is engaged with the driving screw (208), and one end of at least one of the control plates (4) is connected to one end of the driving screw (208), and the driving screw (208) is controlled to rotate so that the driving gear (209) rotates and drives the control plate (4) to rotate by a corresponding angle.

5. The light electric aircraft flight control system according to claim 4, characterized in that: The control component (2) further comprises: A connecting plate (203), wherein the first slider (202) and the second slider (204) are connected via the connecting plate (203).

6. A light electric aircraft flight control method, applied to the light electric aircraft flight control system according to any one of claims 1 to 5, characterized in that: include: In response to a light electric aircraft flight start signal, obtaining current wind speed data and periodically obtaining wind direction data; Based on the current wind speed data and the plurality of wind direction data, a first judgment result is obtained by performing a judgment through a preset wind speed value range and a wind direction change value range; In response to the first judgment result being that the current wind speed data is greater than the preset wind speed value range, or obtaining a current wind direction change angle value by comparing the wind direction data at the current moment and the wind direction data at the previous moment among the plurality of wind direction data, and the current wind direction change angle value is greater than the wind direction change value range, obtaining current posture data; Based on the current posture data, obtaining a current center of gravity offset; Based on the current center of gravity offset, a judgment is made by using a preset center of gravity offset to obtain a second judgment result; In response to the second judgment result that the current center of gravity offset is greater than the preset center of gravity offset, a control instruction set is generated, where the control instruction set is used to control the drive motor to perform a corresponding target action.

7. The light electric aircraft flight control method according to claim 6, characterized in that: The light electric aircraft flight control method further includes: In response to the first judgment result being that the current wind speed data is less than or equal to the preset wind speed value range, or the wind direction change angle value is less than or equal to the wind direction change value range, repeatedly acquiring the current wind speed data and the wind direction data; and / or; In response to the second judgment result being that the current center-of-gravity offset is less than or equal to the preset center-of-gravity offset, the current wind speed data and the wind direction data are repeatedly acquired.

8. The light electric aircraft flight control method according to claim 6, characterized in that: The control instruction set includes: Based on the current wind speed data, the wind direction data, and the current center of gravity offset, obtaining the movement displacement of each first slider and second slider in the plurality of control assemblies, and the rotation angle of each control panel of the plurality of control panels; Based on the movement displacement of each first slider and second slider in the multiple control components and the rotation angle of each control board of the multiple control boards, the motor step number corresponding to the drive motor is determined, and the motor step number is sent to the corresponding drive motor.

9. The light electric aircraft flight control method according to claim 6, characterized in that: The control instruction set includes: Based on the current wind speed data, the wind direction data, and the current center of gravity offset, obtaining the movement displacement of each first slider and second slider in the plurality of control assemblies, and the rotation angle of each control panel of the plurality of control panels; Based on the movement displacement of each first slider and second slider in the multiple control components and the rotation angle of each control board of the multiple control boards, the motor rotation angle corresponding to the drive motor is determined and sent to the corresponding drive motor.

Images