A method for detecting dual mode smooth switching of an aircraft

Abstract

The application discloses a method for detecting smooth switching of dual modes of an aircraft, which comprises the following steps: determining a stage of the aircraft based on a motion posture of the aircraft and position information of the aircraft relative to a cable; determining a current control state of the aircraft by combining the stage of the aircraft and a normal contact pressure of a roller inside the aircraft; and switching a dominant right of a free flight control mode or a cable-holding control mode of the aircraft by using the current control state. The method can provide a switching reference by monitoring the motion posture of the aircraft and the position information of the aircraft relative to the cable in real time, determine the stage of the aircraft, determine the current control state of the aircraft by combining the obtained stage information and the normal contact pressure of the roller inside the aircraft, switch the dominant right of the free flight control mode and the cable-holding control mode based on the current control state, eliminate instruction grabbing and conflict between the two sets of control systems, and realize smooth transition between the dual modes.

Description

Technical Field

[0001] This application relates to the fields of bridge cable inspection, flight robot control, and UAV near-structure operation technology, specifically to a method for smooth switching between dual modes of an inspection aircraft. Background Technology

[0002] Cable components such as bridge stay cables, suspension bridge main cables and suspenders are exposed to wind, rain, corrosion, vibration and alternating loads for a long time, which can easily lead to problems such as rust, sheath damage, fatigue cracks and broken wires. Therefore, regular close-range inspections are necessary.

[0003] In related technologies, to improve the convenience of cable inspection operations, inspection aircraft equipped with a central hollow channel and an openable / closable cable gripping mechanism are often used to inspect cables, relying on rotor aerodynamics to achieve cable gripping and release actions. However, such inspection aircraft often treat free flight control and cable gripping and control as independent modules, making it difficult to establish a unified and continuous control method covering the entire process of flight approach, cable gripping entry, stable cable gripping, and safe release. This leads to control gaps that are prone to occur during mode switching. Summary of the Invention

[0004] This application provides a method for smooth dual-mode switching of a detection aircraft, which can solve the technical problem that detection aircraft in related technologies often regard free flight control and cable gripping and cable-following control as independent modules, making it difficult to establish a unified and continuous control method covering the entire process of flight approach, cable gripping entry, stable cable-following and safe detachment, resulting in control discontinuity during mode switching.

[0005] In a first aspect, embodiments of this application provide a method for detecting smooth switching between two modes of an aircraft. The method includes: determining the current stage of the aircraft based on its motion attitude and its position relative to the cable; determining the current control state of the aircraft by combining the current stage of the aircraft with the normal contact pressure of the rollers inside the aircraft; and using the current control state to control the switching of dominance between the free flight control mode and the cable-clutch control mode of the aircraft.

[0006] The beneficial effects of the technical solutions provided in this application include: By monitoring the aircraft's motion attitude and position relative to the cable in real time, the current stage of the aircraft is determined, providing a preliminary benchmark for mode switching. Then, the obtained stage information is combined with the normal contact pressure of the rollers inside the aircraft to dynamically evaluate and determine the current control state of the aircraft. Based on the current control state, the dominance of the free flight control mode and the cable-attached control mode is orderly allocated and switched, thereby eliminating command contention and conflict between the two control systems. This achieves a smooth and safe transition between the two modes and solves the technical problem in related technologies where the detection aircraft often treats free flight control and cable-attached control as independent modules, which easily leads to control gaps during mode switching. Attached Figure Description

[0007] Figure 1 This is a flowchart illustrating the first embodiment of the method for smooth switching between dual modes of an aircraft according to this application; Figure 2 For this application Figure 1 A detailed flowchart of step S3; Figure 3 For this application Figure 1 A detailed flowchart of step S1; Figure 4 This is a schematic diagram of the flight approach segment of this application; Figure 5 This is a schematic diagram of the opening alignment section of this application; Figure 6 This is a schematic diagram of the touch-to-enter segment of this application; Figure 7 This is a schematic diagram of the closed cable-holding and stable cable-stayed sections in this application; Figure 8 This is a block diagram of the control structure for the six-stage continuous mode smooth switching algorithm of this application; Figure 9 This is a block diagram of the attitude generation and thrust allocation logic driven by force field consistency in this application.

[0008] In the picture: 1. Aircraft; 2. Cable. Detailed Implementation

[0009] To enable those skilled in the art to better understand the present application, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present application, and not all embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present application.

[0010] To make the objectives, technical solutions, and advantages of this application clearer, the embodiments of this application will be described in further detail below with reference to the accompanying drawings.

[0011] In a first aspect, embodiments of this application provide a method for detecting smooth switching between dual modes of an aircraft.

[0012] In one embodiment, reference is made to Figure 1 , Figure 1 This is a flowchart illustrating the first embodiment of the method for smooth switching between dual modes of an aircraft according to this application. Figure 1 As shown, the method for detecting smooth switching between two modes of an aircraft includes: S1: Based on the motion attitude of aircraft 1 and its pose relative to cable 2, determine the current stage of aircraft 1. It should be understood that the motion attitude of aircraft 1 and its pose relative to cable 2 can be continuously collected during the flight phase of aircraft 1. The motion attitude includes at least the position and velocity of aircraft 1 based on the world coordinate system (NED), and the fuselage attitude angle, angular velocity, acceleration, rotor speed, open mechanism status, and normal contact pressure of cable 2 based on the body coordinate system. The pose relative to cable 2 includes at least the normal distance deviation, subnormal distance deviation, and spatial orientation of cable 2 of aircraft 1 relative to cable 2.

[0013] S2: Determine the current control state of aircraft 1 by combining the current stage of aircraft 1 with the normal contact pressure of the internal rollers of aircraft 1. That is, at different stages of aircraft 1, the controller determines the required current control state of aircraft 1 based on the different normal contact pressure values ​​obtained.

[0014] S3: Utilizing the current control state, control the switching of dominance between the free flight control mode and the cable-stayed control mode of aircraft 1. In other words, based on the current control state required by aircraft 1, the control weights of the flight propulsion system and the roller clamping mechanism can be dynamically allocated to achieve a smooth handover of control input commands between the two control modes, effectively avoiding command conflicts and mechanical shocks during mode switching.

[0015] In this embodiment, by monitoring the motion attitude of the aircraft 1 and its position relative to the cable 2, the current stage of the aircraft 1 is determined, providing a preliminary benchmark for mode switching. Then, the obtained stage information is combined with the normal contact pressure of the rollers inside the aircraft 1 to dynamically evaluate and determine the current control state of the aircraft 1. Based on the current control state, the dominance of the free flight control mode and the cable-clutch control mode is orderly allocated and switched, thereby eliminating command contention and conflict between the two control systems, realizing a smooth and safe transition between the two modes, and solving the technical problem in related technologies that the aircraft often treats the free flight control and cable-clutch control as independent modules, which easily leads to control gaps during mode switching.

[0016] Furthermore, in one embodiment, see... Figure 2 and Figure 7 As shown, the switching of control over the free flight control mode or the cable-stayed control mode of the control aircraft 1 may include: S31: A mode fusion factor λ(t) is introduced in the transition interval of the switch of dominance; in the embodiments of this application, λ(t) can represent the weight of the free flight control input, and 1-λ(t) represents the weight of the cable gripping control input; during the process of the aircraft 1 entering the cable gripping state from the free flight state, λ(t) smoothly decreases from 1 to 0; during the process of the aircraft 1 exiting the cable gripping state to the free flight state, λ(t) smoothly increases from 0 to 1.

[0017] S32: Based on the model fusion factor, the free flight control input, the cable-stayed control input, and the stage transition compensation term are continuously weighted to obtain the unified control input u(t). Specifically, the model fusion factor λ(t) can satisfy: , ; in, To switch the smoothing coefficient, Let t be the current control time, and σ be the switching direction coefficient. When aircraft 1 transitions from the cable-clamping state to the free flight state, σ = 1, so that λ(t) increases smoothly from 0 to 1 over time. When aircraft 1 transitions from the free flight state to the cable-clamping state, σ = -1, so that λ(t) increases smoothly from 1 to 0 over time. λ(t) is used as the weight of the free flight control input, and 1-λ(t) is used as the weight of the cable-clamping control input. The free flight control input, the cable-clamping control input, and the stage transition compensation term are continuously weighted to obtain a unified control input. When λ(t) approaches 1, the free flight control input dominates. When λ(t) approaches 0, the cable-clamping control input dominates.

[0018] S33: Generate the target acceleration, target attitude, target total thrust, target torque, and rotor speed distribution commands for aircraft 1 using the unified control input u(t). It should be understood that the unified control input u(t) is the comprehensive control quantity actually used by aircraft 1 in the current control cycle, used to generate the target acceleration, target attitude, target total thrust, target torque, and rotor speed distribution commands. Specifically: ; in, For free flight control input, For cable gripping and cable control input, This is a phase transition compensation item.

[0019] Used to represent the contribution of free-flight control input to the current control cycle. Used to represent the contribution of the cable-holding control input to the current control cycle. Used to compensate for normal deviation, attitude deviation and velocity abrupt changes that occur during phase switching; When λ(t) increases, the aircraft 1 increases the proportion of free flight control input in the unified control input, and enhances attitude holding, hovering stability and disengagement recovery control according to u(t); when λ(t) decreases, the aircraft 1 increases the proportion of cable gripping control input in the unified control input, and enhances cable speed control, low normal pressure control and cable gripping constraint control according to u(t).

[0020] Furthermore, the stage transition compensation term can be constructed based on the safety boundary function and velocity damping as follows: ; in, The safety boundary function is used to characterize the safe operating area of ​​the aircraft 1 relative to the cable 2; the safety boundary function is constructed based at least on the normal deviation of the aircraft 1 relative to the cable 2, the normal contact pressure, and the allowable safety threshold. The gradient of the safety boundary function is used to represent the direction that increases the safety margin. The controller generates a compensation control quantity based on this gradient direction, so that the aircraft 1 adjusts away from the cable 2 or reduces the contact pressure when approaching the safety boundary, thereby avoiding excessive normal offset or excessive contact pressure. For boundary compensation gain, For velocity damping gain, This is the state velocity vector.

[0021] The safety boundary function is: ; in, To allow for a safe distance in the normal direction, The normal deviation of aircraft 1 relative to cable 2. The normal contact pressure safety threshold, For the current normal contact pressure, This refers to the weighting coefficient for the stress term; when When >0, it indicates that aircraft 1 is within the safety boundary; when When it approaches 0, it indicates that aircraft 1 is approaching the safety boundary; when When the value is less than 0, it indicates that there is a risk of excessive normal offset or excessive normal contact pressure for aircraft 1.

[0022] Furthermore, in one embodiment, see... Figure 3 As shown, the phases in which the aircraft 1 is currently located include the flight approach phase, the opening alignment phase, the light touch entry phase, the closing grip phase, the stable cable-along phase, and the opening disengagement phase. The determination of the current phase of the aircraft 1 based on its motion attitude and position relative to the cable 2 includes: S11: See also Figure 4 As shown, if the radial distance between the center of the hollow channel in the middle of the aircraft 1 and the centerline of the cable 2 is greater than a preset distance threshold, then the aircraft 1 is determined to be in the flight approach phase; that is, the flight approach phase can be the stage where the operator controls the aircraft 1 to gradually approach the target cable 2 after takeoff. In this stage, the aircraft 1 approaches the target cable 2 in free flight, and the aircraft 1 keeps the opening mechanism in the open or pre-open state, and the opening direction is towards the target cable 2.

[0023] S12: After the aircraft 1 enters the flight approach phase, if the radial distance between the center of the hollow channel in the middle of the aircraft 1 and the center line of the cable 2 is less than a preset distance threshold, then the aircraft 1 is determined to have entered the opening alignment phase. In this embodiment of the application, it is set that when the radial distance between the center of the hollow channel in the middle of the aircraft 1 and the center line of the cable 2 is less than 40 cm, the aircraft 1 is determined to have entered the opening alignment phase. After entering the opening alignment phase, the aircraft 1 can be controlled to hover for a period of time. During this time, the camera device on the aircraft 1 is used to take some pictures, and the captured picture data is used to establish the local coordinate system of the cable 2. Specifically, when the aircraft 1 is suspended in the air, the centerline of the target cable 2 is identified based on airborne vision, and the geometric center of the hollowed-out channel in the middle of the aircraft 1 is determined. The geometric center of the hollowed-out channel is projected onto the centerline of the cable 2 to obtain the nearest projection point on the centerline of the cable 2. Using the projection point as the origin of the local coordinate system of the cable 2, a local coordinate system (TNB) for the cable 2 is established: the T-axis is along the direction of the centerline of the cable 2, the N-axis points from the origin to the geometric center of the hollowed-out channel in the middle of the aircraft 1, and the B-axis is obtained by cross-product of the T-axis and the N-axis according to the right-hand rule. Thus, the normal distance deviation and the sub-normal distance deviation are both determined based on the local coordinate system of the cable 2.

[0024] S13: After aircraft 1 enters the opening alignment section, if the z-axis of aircraft 1 is parallel to the axis of cable 2, then aircraft 1 is determined to have entered the light-touch entry section. (See also...) Figure 5 As shown, the opening alignment stage can be a phase in which aircraft 1 precisely adjusts the opening position and body attitude. In the opening alignment stage, aircraft 1 adjusts the positional relationship between the center of the central hollow channel and the centerline of cable 2, and adjusts the angle between the z-axis of the body and the direction of cable 2. When the opening is aligned with the centerline of cable 2, the z-axis of the body is aligned with the direction of cable 2, and the x-axis of the body is perpendicular to the direction of cable 2, it enters the touch-entry stage. Preferably, during the transition from the opening alignment stage to the touch-entry stage, aircraft 1 can have a relatively low speed.

[0025] S14: After aircraft 1 enters the light-touch entry phase, if the normal contact pressure of cable 2 is detected to be within a preset light-touch pressure range and the pressure change rate is lower than a preset threshold, then aircraft 1 is determined to have entered the closed-grip phase. (See also...) Figure 6 As shown, the light touch entry stage is the stage in which the aircraft 1 enters the cable 2 into the middle hollow channel at a limited normal velocity; when the pressure sensor installed on the opposite side of the opening or at the internal roller detects that the normal contact pressure of the cable 2 reaches the preset light touch pressure range, such as 5-10N, and the pressure change rate is lower than the preset threshold, it is determined that the aircraft 1 has entered the closed grip section.

[0026] S15: After aircraft 1 enters the closed cable gripping section, if the opening mechanism of aircraft 1 reaches the closed position, and the normal contact pressure is lower than the preset stable pressure upper limit, and the attitude deviation and cable velocity fluctuation are both less than the stable threshold, then aircraft 1 is determined to have entered the stable cable gripping section; see [link to relevant documentation]. Figure 7 As shown, the closed cable section is the stage in which the opening mechanism gradually closes from the open state; when the opening mechanism reaches the closed position, the cable 2 remains in the middle hollow channel or roller constraint area, and the normal contact pressure is lower than the preset stable pressure upper limit, it is determined that the aircraft 1 has entered the stable cable section.

[0027] S16: After aircraft 1 enters the stable cable-stayed section, if the detection task is detected as completed, or the normal contact pressure exceeds the limit, the attitude deviation is too large, or the opening mechanism is abnormal, then aircraft 1 is determined to enter the opening detachment section. The stable cable-stayed section is the stage in which aircraft 1 performs stable detection operations along the direction of cable 2; when the detection task is completed, it is necessary to cross an obstacle, actively detach from cable 2, or when there is an excessive attitude deviation, excessive normal contact pressure, excessive pressure change rate, abnormal opening mechanism, or abnormal rotor status, it enters the opening detachment section.

[0028] The opening detachment stage is the phase in which aircraft 1 opens the opening mechanism and recovers from the cable-clamping state to the free flight state. In this stage, the opening mechanism gradually opens, the control weight of the cable-clamping state decreases, and the control weight of the free flight state increases until the cable 2 completely detaches from the middle hollow channel. That is, when the normal mission ends, obstacle crossing requirement is triggered, or emergency detachment is triggered, a safe detachment trajectory can be generated according to the pre-set minimum impact principle, and the rotor thrust change rate and the opening mechanism change rate are jointly limited, so that aircraft 1 can smoothly recover from the cable-clamping state to the free flight state.

[0029] Furthermore, in one embodiment, after determining the current stage of the aircraft 1, the process may include: constructing a thrust field including the thrust generated by the rotor system. Aerodynamic disturbance field in the near cable 2 region Cable 2 constraint reaction field Mode switching additional force field The total target force field, including the gravitational field mg, is denoted as... Based on the overall target force field, the target thrust direction, target attitude matrix, and rotor speed distribution commands are generated. It should be understood that the control objective of the overall target force field is to track the detection trajectory along the cable, suppress the normal contact pressure of cable 2, compensate for aerodynamic disturbances near cable 2 and constraint reaction forces of cable 2, suppress normal overshoot and attitude abrupt changes through stage transition compensation terms, and limit the rate of change of rotor thrust, target attitude, and opening mechanism to ensure continuous changes in thrust, attitude, and opening action during stage switching, thereby reducing the impact of mode switching.

[0030] See Figure 8 and Figure 9 As shown, in one embodiment, the thrust field generated by the rotor system... scalar of total thrust generated by the rotor system and rotor thrust vector in the airframe coordinate system It is concluded that the thrust field generated by the rotor system is... Rotor thrust vector in the airframe coordinate system and the total thrust scalar generated by the rotor system satisfy: , , ; in, Let i be the rotational speed of the i-th rotor. R is the rotor thrust coefficient, and R∈SO(3) is the rotation matrix from the body coordinate system to the world inertial coordinate system; The cable 2 constrains the reaction field Represented in the world coordinate system as ,and ,in, , ; Where m is the number of internal rollers, Let the normal constraint reaction force of cable 2 on the j-th roller be represented by an equivalent spring-damped nonlinear contact model. Let j be the normal direction of the j-th roller contact point, which is the unit vector of the normal constraint reaction force in the world coordinate system. Let j be the frictional or rolling resistance of the j-th roller. For cable direction 2 in the world coordinate system, This refers to the compression amount of the roller. The roller compression speed, The equivalent normal contact stiffness coefficient, This is the equivalent normal contact damping coefficient. For nonlinear contact stiffness coefficient; when When ≤0, it indicates that the roller has not made effective contact with cable 2. =0, The equivalent friction coefficient or rolling resistance coefficient between the j-th roller and cable 2. Used to indicate the direction of motion along the cable. The velocity damping coefficient is along cable direction 2. The speed of movement along cable 2; The aerodynamic disturbance field Recorded as: ; in: Relative airflow velocity The linear damping coefficient is... This is the second-order drag coefficient. The additional disturbance force generated by the interaction between the rotor wake and cable 2 The disturbance force is caused by local eddies, and there is an additional disturbance force. satisfy , ; in, The relative distance between aircraft 1 and cable 2. and For the disturbance amplitude parameter, The wake attenuation coefficient is... The angular frequency of the eddy current disturbance. This represents the phase of the eddy current disturbance. and The direction of the disturbance is denoted by t, and t represents the current control time. The mode switching additional force field for: ; in, Here, x represents the transition weight matrix for mode fusion, and x is the current state vector of the system. This is the reference state vector for the current stage. For safety boundary compensation weights, The gradient of the safety boundary function is used to represent the direction that increases the safety margin; Used to generate a compensation force away from cable 2 or to reduce contact pressure when the difference between aircraft 1 and the safety boundary is less than or equal to a preset value; λ(1-λ) is used to enhance transition compensation during the intermediate stage of mode switching and weaken it at the endpoint of stable mode. For velocity damping weight, Let the state velocity vector be... Used to suppress speed abrupt changes and oscillations during the switching process.

[0031] It should be understood that when aircraft 1 is operating in the area near cable 2, the controller does not directly use a single attitude angle as the starting point, but first constructs the overall target force field. The overall target force field unifies the description of rotor thrust, aerodynamic disturbance, cable 2 constraint reaction force, mode switching additional force, and gravity. When obtained... Then, the controller calculates the target thrust direction. And combined with cable 2 directions Construct the target attitude matrix Subsequently, by constraining the force field error, attitude error, and angular velocity error through the joint error energy function V, the target force field is made consistent with the actual body response. This is achieved after obtaining the target's total thrust T and target torque. , , Then, the rotational speed commands for the four rotors are obtained by controlling the allocation matrix A. If the obtained rotor speed exceeds the upper limit of the motor, is lower than the lower limit of the motor, or is unattainable, the controller will saturate the target torque and target thrust, and prioritize the safety of normal pressure and attitude stability.

[0032] In this embodiment, the target total force field is used to uniformly describe the main forces acting on the aircraft 1 during its operation near cable 2, cable engagement, stabilization along cable, and disengagement from the opening. It is also used to generate the target thrust direction, target attitude matrix, target total thrust, and rotor speed distribution commands. This ensures that the aircraft 1 maintains low contact pressure, low attitude abrupt changes, and low mode-switching shocks during these processes. In other words, through the target total force field, the controller can simultaneously achieve cable trajectory tracking, cable 2 normal contact pressure suppression, near-cable 2 aerodynamic disturbance compensation, cable 2 constraint reaction force compensation, and mode-switching shock suppression within the same control framework. Specifically, the aircraft 1's controller can decompose the total target force field into a tangential component along the cable 2 direction, a normal component perpendicular to the cable 2 direction, and an attitude correction component for attitude stabilization. The tangential component is mainly used to adjust the velocity and tracking acceleration of aircraft 1 along the cable; the normal component is mainly used to adjust the relative distance, contact degree, and roller compression between aircraft 1 and cable 2; the attitude correction component is mainly used to adjust the target attitude matrix, aircraft pitch angle, roll angle, yaw angle, and target angular velocity. The calculation results of the total target force field are transformed into comprehensive adjustment commands for the motion, attitude, thrust, and contact state of aircraft 1.

[0033] During the approach phase or the phase near cable 2, the controller adjusts the target trajectory, target thrust direction, and approach speed of aircraft 1 based on the aerodynamic disturbance compensation term in the overall target force field and the relative position error of cable 2. When the overall target force field indicates that aircraft 1 is subject to wake disturbance, lateral drift, or insufficient normal distance, the controller reduces the normal velocity approaching cable 2, increases the stabilizing component parallel to cable 2, and corrects the roll and pitch attitude through rotor speed distribution to keep aircraft 1 within the safe approach path.

[0034] During the alignment phase, the controller adjusts the target attitude matrix, nose direction, lateral position of the aircraft 1, and alignment status of the opening mechanism relative to cable 2 based on the calculation results of the total target force field. If the normal component or attitude correction component in the total target force field indicates that there is an angular, offset, or alignment error between aircraft 1 and cable 2, the controller issues attitude fine-tuning, lateral displacement correction, and thrust differential distribution commands to gradually satisfy the entry conditions for the central passage of aircraft 1, the opening mechanism, and the direction of cable 2.

[0035] During the initial approach phase, the controller adjusts the normal contact speed, total target thrust, and contact pressure of aircraft 1 based on the calculation results of the constraint reaction field of cable 2 and the roller contact model. When the roller compression or normal contact force is below the target range, the controller appropriately increases the normal target force towards cable 2; when the roller compression or contact pressure approaches the upper limit, the controller reduces the normal contact force, lowers the approach speed, and suppresses secondary collisions through attitude and rotor speed corrections.

[0036] During the tethered engagement phase, the controller adjusts the hovering thrust, attitude damping, closing speed, and rotor speed distribution of aircraft 1 based on the cable 2 constraint reaction force compensation term and the mode switching additional force field in the overall target force field. If a sudden change in attitude, a sudden increase in normal pressure, or an increase in the closing resistance of the opening mechanism occurs during the tethering process, the controller reduces the closing action speed, limits the target attitude change rate, and increases the damping control weight to ensure that aircraft 1 maintains low impact and low contact pressure during the tethering process.

[0037] In the stable cable-line section, the controller adjusts the cable-line traction force and velocity based on the tangential component of the total target force field, adjusts the roller contact pressure and cable 2 constraint reaction force compensation based on the normal component, and adjusts the rotor thrust distribution and attitude angle based on the aerodynamic disturbance component. If the cable-line velocity is too low, the target thrust component along cable 2 is increased; if the cable-line velocity is too high or the contact pressure increases, the cable-line acceleration is reduced and the normal contact force is decreased; if disturbed by wind, wake, or vortex, the airframe stability is maintained through rotor differential speed and target attitude matrix correction.

[0038] During the detachment phase, the controller adjusts the unloading direction, detachment speed, detachment mechanism action speed, and target attitude of aircraft 1 based on the total target force field calculation results. The controller first reduces the normal contact force and roller compression, then generates a smooth yielding force field away from cable 2, and limits the rate of change of total thrust and rate of change of attitude, so that aircraft 1 avoids re-collision with the cable, sudden attitude jumps, or sudden changes in rotor thrust during the detachment process.

[0039] When the actual position, attitude, velocity along the cable, contact pressure, or rotor speed does not meet the target requirements for the current stage, the controller constructs an error correction term based on the total target force field and recalculates the target thrust direction, target attitude matrix, total target thrust, target torque, and rotor speed command. This correction process is continuously executed in a closed loop until the trajectory error, attitude error, normal distance error, contact pressure error, and velocity error of aircraft 1 converge to the allowable range.

[0040] When any monitored quantity exceeds the warning threshold, the controller prioritizes adjusting the objects directly related to the risk based on the type of exceedance: when contact pressure or roller compression exceeds the limit, it prioritizes adjusting the normal contact force and cable velocity; when attitude error or angular velocity exceeds the limit, it prioritizes adjusting the target attitude matrix, target torque, and attitude change rate; when the rotor speed is close to saturation, it prioritizes adjusting the target total thrust, target torque, and rotor speed distribution; when the relative distance between cables 2 is less than the safe distance, it prioritizes generating a yielding force field away from cable 2. If the monitored quantity continues to exceed the safety threshold, the controller stops the current stage of action, enters the safety protection mode, and executes deceleration, unloading, hovering, yielding, or opening disengagement commands.

[0041] Furthermore, in one embodiment, the target attitude matrix is ​​denoted as... , , , and ; in, The direction of the target thrust axis is determined by the overall target force field. Let z be the target airframe's z-axis. When the rotor thrust is output along the z-axis of the airframe, the z-axis of the target airframe is aligned with the direction of the target thrust axis. The target heading reference vector, In the direction of the target aircraft nose, Let y be the target body's y-axis as determined by the right-hand rule. It should be understood that the target attitude matrix can also be denoted as the target rotation matrix.

[0042] In this embodiment of the application, the total target force field is obtained. Afterwards, based on Calculate the target thrust direction, and then introduce the target heading reference vector in the world coordinate system. The target heading reference vector can be based on the cable's 2-direction unit vector. Confirmed, will Projected to perpendicular to Construct the target attitude axis on the plane. .

[0043] Furthermore, in one embodiment, the rotor speed allocation command is solved by a control allocation matrix: Recorded as , ; Where T is the target total thrust. , , These represent the target roll, pitch, and yaw moments, respectively, and A is the control allocation matrix. That is, after obtaining the total target thrust T and the target moment... , , Subsequently, the rotor speed is solved using the control allocation matrix. In this embodiment, the control allocation matrix is ​​a universal rotor execution allocation matrix, which can be used in different control stages. The controller determines the source of the target total thrust and target torque input to the control allocation matrix based on the current stage identifier and mode fusion factor. When in the dominance switching transition interval, the target total thrust and target torque are generated by the unified control input and mode switching additional terms; when in the stable cable-following section, the target total thrust and target torque are generated by the cable-following tracking term, contact pressure suppression term, cable 2 constraint reaction force compensation term, and aerodynamic disturbance compensation term in the total target force field. The rotor speed allocation command generated thereby serves the smooth transition of mode switching and the stable cable-following travel control, respectively.

[0044] Furthermore, in one embodiment, after the aircraft 1 enters the opening alignment section, the method further includes: determining whether the opening width of the opening mechanism of the aircraft 1 is greater than or equal to the sum of the outer diameter of the cable 2 and the preset safety gap; calculating the opening alignment error based on the normal deviation, sub-normal deviation, and attitude direction deviation between the aircraft 1 and the cable 2 relative to the cable 2; if the opening width of the opening mechanism of the aircraft 1 is greater than or equal to the sum of the outer diameter of the cable 2 and the preset safety gap, and the opening alignment error is less than a preset value, then controlling the aircraft 1 to enter the light-touch entry section. Specifically, the cable gripping entry conditions include the opening geometry conditions as follows:

[0045] in, The width of the opening. The outer diameter of the cable is 2. A safety gap is reserved for cable 2 when it enters the central hollow channel.

[0046] The opening alignment error is: , in, The distance deviation along the normal N-axis from the center of the hollowed-out channel in the middle of aircraft 1 relative to the origin of the local coordinate system of cable 2. The distance deviation of the center of the hollow channel in the middle of the device from the origin of the local coordinate system of cable 2 on the secondary normal B-axis. Let z be the unit vector of the body along the z-axis in the machine system. Let x be the unit vector of the organism's x-axis. Let be the unit vector representation of the cable in direction 2 within the machine system. For coaxial error weighting, The weight of the nose vertical error; and The cable's 2-direction unit vector in the world coordinate system is obtained by transforming it through the body attitude rotation matrix, specifically: ,in, This is the rotation matrix from the world coordinate system to the machine coordinate system.

[0047] when , If the rate of change of normal contact pressure is lower than the set threshold, the system allows aircraft 1 to enter the touch entry section from the opening alignment section.

[0048] like If the aircraft does not enter the light-touch entry phase, but instead proceeds according to the normal deviation... Subnormal deviation Adjust the center position of the central hollow channel and the attitude of the aircraft by adjusting the deviation of the attitude and direction, so that the center of the central hollow channel continues to align with the center line of cable 2, and the z-axis of the aircraft gradually becomes coaxial with the direction of cable 2, and the x-axis of the aircraft gradually becomes perpendicular to the direction of cable 2.

[0049] After entering the light-touch entry section, the aircraft 1 slowly approaches the cable 2 along the normal direction of the cable 2 at a speed lower than the preset upper limit, so that the cable 2 gradually enters the middle hollow channel; when the cable 2 has entered the middle hollow channel, the normal contact pressure is within the preset light-touch pressure range and the normal contact pressure change rate is lower than the set threshold, the system allows the aircraft 1 to enter the closed grip section from the light-touch entry section.

[0050] If the rate of change of normal contact pressure exceeds the set threshold or the normal contact pressure exceeds the preset upper limit during the light touch entry process, the aircraft 1 will reduce its normal approach speed and generate a yield control amount in the normal direction away from the cable 2, so that the normal contact pressure drops below the preset warning value; if the pressure cannot drop within the preset time, the cable entry will be stopped and the aircraft will enter the opening release section.

[0051] Furthermore, in one embodiment, the step of determining that the aircraft 1 has entered the opening disengagement phase if the detection task is completed, or the normal contact pressure exceeds the limit, the attitude deviation is too large, or the opening mechanism is abnormal, includes: constructing an emergency judgment vector and a comprehensive risk function based on the normal contact pressure, attitude deviation, rotor state, wind disturbance estimate, normal offset, and opening mechanism state; and executing early warning level, suppression level, disengagement level, or return-to-home level emergency control according to the comprehensive risk function. It should be understood that during the six stages described above, the controller of aircraft 1 can construct an emergency judgment vector and a comprehensive risk function based on normal contact pressure, attitude deviation, rotor state, wind disturbance estimate, normal offset, and opening mechanism state throughout the entire process. Based on the comprehensive risk function, it can execute early warning, suppression, separation, or return-to-home emergency control, and can be applied in key stable cable-stayed sections. The normal contact pressure is obtained by the pressure detection module, the attitude deviation is obtained by the inertial measurement unit and attitude estimator, the rotor state is obtained by the electronic speed controller or speed feedback, the wind disturbance estimate is obtained by the disturbance observer, the normal offset is obtained by the vision and ranging modules, and the opening mechanism state is obtained by the encoder, limit switch, or opening state detection module.

[0052] Preferably, the emergency decision vector is: ; in, , , , , , ; For normal contact pressure, This is the upper limit of the normal contact pressure. This represents the attitude deviation of the current attitude relative to the target attitude. This represents the upper limit of attitude deviation. Let i be the rotational speed of the i-th rotor. The rotor failure threshold, This is an estimate of the wind disturbance. The wind disturbance threshold, This represents the normal offset of aircraft 1 relative to cable 2. The normal safety offset threshold, For the state variables of the opening mechanism. This is a binary risk assessment value, set to 1 when the corresponding risk condition is met, and 0 otherwise; the emergency assessment vector Used to characterize abnormal conditions of aircraft 1 in terms of normal pressure, attitude, rotor, wind disturbance, normal offset, and open mechanism status.

[0053] The preferred comprehensive risk function is: ; in, The emergency decision vector The i-th binary risk assessment quantity, The normal contact pressure change rate, The threshold for the rate of change of normal contact pressure. For attitude error, This is the attitude error threshold. , and Risk weights; comprehensive risk function Based on the aforementioned emergency judgment vector, continuous risk quantities of normal contact pressure change rate and attitude error are further introduced to evaluate the overall risk level of the current working condition. When the first preset threshold is reached, an early warning level emergency control is triggered; when When the second preset threshold is reached, the suppression-level emergency control is triggered; when When the third preset threshold is reached, the escape-level emergency control is triggered; when the aircraft 1 completes the escape and its flight capability meets the return-to-home conditions, the return-to-home emergency control is triggered.

[0054] In this embodiment of the application, the first preset threshold, the second preset threshold, and the third preset threshold are respectively... , and And satisfy: ,when At that time, aircraft 1 maintained normal operation along the cable; when When this is triggered, an early warning-level emergency control is activated. Vehicle 1 reduces its velocity along the cable, increases its attitude control weight, and enhances normal pressure suppression. When this occurs, the suppression-level emergency control is triggered. The aircraft 1 is restricted from advancing along the cable, and a yield control amount is generated in the normal direction away from cable 2, causing the normal contact pressure to drop below the preset warning value; when When the time is right, the escape-level emergency control is triggered, the aircraft 1 opens the opening mechanism and executes the safe escape trajectory, so that the cable 2 is detached from the hollow channel in the middle of the aircraft 1; when the aircraft 1 has detached from the cable 2 and the flight capability meets the return conditions, the return-level emergency control is triggered, and the aircraft 1 executes the return, hovering and waiting for recovery or forced landing strategy.

[0055] The preferred safe escape trajectory is: ; in, To trigger the initial position of spacecraft 1 upon separation, To trigger the initial velocity of aircraft 1 upon separation, Let be the safe exit acceleration; t be the current moment during the exit process. The safe exit acceleration is obtained through the following optimization objective: , Where 'a' represents the candidate escape acceleration. To allow for deceleration beyond the upper limit, This represents the predicted normal contact pressure under candidate separation acceleration. This is the upper limit of the normal contact pressure. The predicted rate of change of normal contact pressure, This represents the upper limit of the rate of change of normal contact pressure. The predicted attitude error, This represents the upper limit of attitude error. , , and To optimize weights, The rate of change of the state of the opening mechanism. This represents the upper limit of the rate of change of the state of the opening mechanism.

[0056] Preferably, when performing emergency control of the opening disengagement section or disengagement level, the following constraints are simultaneously met: , ; Where d / dt represents the derivative with respect to time, This is the upper limit of the rotor thrust variation rate. This is the upper limit of the rate of change of the opening mechanism. This limits the rate of change of rotor thrust and the rate of change of the opening mechanism's state, ensuring a smooth decrease in normal contact pressure during disengagement and preventing secondary cable impact.

[0057] Furthermore, in one embodiment, before generating the target acceleration, target attitude, target total thrust, target torque, and rotor speed distribution commands of the aircraft 1 using the unified control input u(t), the method further includes: A joint error energy function is constructed, which characterizes the combined deviation of the current force state, attitude state, and angular velocity state of the aircraft 1 relative to the target state, and is denoted as: ; ; ; in, For force field error, For the current actual or estimated total force field of aircraft 1, For the overall target force field; Let R be the attitude error, and R be the current attitude matrix. The target attitude matrix; For angular velocity error, , and It is a positive definite weight matrix used to adjust the weights of force field error, attitude error and angular velocity error in the joint error energy function; This represents the mapping from an antisymmetric matrix to a vector.

[0058] In other words, to ensure that the generated target attitude is trackable, force field error and attitude error can be introduced: and Then, the joint error energy function is further constructed. The joint error energy function is used to reduce the force field error, attitude error, and angular velocity error together.

[0059] This application provides a control process for an aircraft 1 to transition from free flight to cable-bound operation and then detach again: First, the aircraft 1 uses airborne vision to identify the centerline and direction of the cable 2, obtains the relative distance through a ranging device, obtains the body attitude and angular velocity through an inertial measurement unit, obtains the normal contact pressure of the cable 2 through a pressure detection module, and obtains the status of the opening mechanism through an opening status detection module.

[0060] During the approach phase, aircraft 1, in free flight mode, approaches the cable 2 reference point and sets an upper limit on its normal velocity. The opening mechanism remains fully open to prevent accidental contact with cable 2 during the approach. When aircraft 1 approaches to a preset safe distance and its normal velocity, tangential error, and attitude error meet the entry conditions, it enters the opening alignment phase.

[0061] During the opening alignment phase, aircraft 1 adjusts its position and attitude to align the center of the central hollow channel with the centerline of cable 2, ensure the geometric direction of the opening surface is consistent with the local normal of cable 2, make the z-axis of the aircraft approximately coaxial with the direction of cable 2, and make the x-axis of the aircraft approximately perpendicular to the direction of cable 2. The system continuously reduces the opening alignment error. ;when At that time, enter the tap entry segment.

[0062] During the light-touch entry phase, aircraft 1 employs a compliant contact control law: ; in, Here, k is the normal acceleration, c is the normal deviation feedback gain, and ρ is the normal velocity damping gain. This control law suppresses peak contact force, ensuring the normal pressure increases slowly within a preset safety range. If the normal pressure growth rate exceeds a threshold, or if the attitude error rapidly amplifies within a short period, the system aborts the entry action and reverts to the opening alignment section. It should be understood that... It can serve as the basis for generating the normal component in the overall target force field, and can be used to further calculate the target thrust direction, target attitude matrix, target total thrust and rotor speed distribution command. It is the stage control component after the overall target force field is decomposed.

[0063] In the closed grip section, once cable 2 is fully inside the central hollow channel, the opening mechanism gradually closes from the open state. At this time, the mode fusion factor λ(t) continuously changes from 1 to 0, the free flight control weight gradually decreases, and the grip along the cable control weight gradually increases. During the closing process, normal pressure and attitude error are continuously constrained to prevent cable 2 impact, aircraft deflection, or abnormal friction at the opening edge due to the opening closure.

[0064] In the stable cable-line section, aircraft 1 is in a cable-line propulsion state constrained by cable 2. The velocity along cable 2 is defined as: ; The system provides the target velocity along the cable based on the requirements of the detection task. Acceleration along the cable is generated through closed-loop velocity control: ; Preferred, As the basis for generating the tangential component in the overall target force field, it can be further used to calculate the target thrust direction, target attitude matrix, target total thrust and rotor speed distribution command. It is also the stage control component after the overall target force field is decomposed.

[0065] Meanwhile, lateral disturbances and attitude deviations are corrected in real time by an extended state observer and a nonlinear damping controller. When a parameter approaches a threshold, the system prioritizes parameter suppression and compensation rather than immediately exiting the system.

[0066] During the opening disengagement phase, when the task is completed, the obstacle crossing requirement is triggered, or the normal exit conditions are met, the system reverse-activates the mode fusion factor, gradually transferring control from cable-grip control back to free-flight control. The opening mechanism gradually opens, and the opening state variable... The opening gradually increases; the weight of free flight control increases, while the weight of cable-holding control decreases. During the opening process, the system maintains a continuous decrease in normal pressure and causes the fuselage attitude to converge towards a stable free flight attitude. When the opening is fully open and cable 2 is completely detached from the central hollow channel, the system switches to free flight mode and enters the attitude recovery and path replanning phase.

[0067] This embodiment uses a prototype quadcopter cable 2 testing aircraft with a wheelbase of approximately 450 mm and a cable-holding roller mechanism as an example for explanation. Among the parameters described below, the overall mass, opening width, roller compression, cable 2 outer diameter, and cable 2 tilt angle were obtained from prototype weighing, mechanism measurement, and cable 2 identification results; the rotor thrust coefficient was obtained from a static thrust test fitting on a rotor bench; and the remaining control parameters were calculated from the desired transition time, access margin, and escape displacement constraint.

[0068] The total mass is m = 1.60 kg, and the standard gravitational acceleration is taken as... The total lift required for hovering is: =1.62 × 9.80665 = 15.89 N. When the thrust is distributed evenly among the four rotors, the hovering thrust of a single rotor is... =3.97 N. The rotor thrust model is: When the hovering speed is approximately When = 705 rad / s, we can obtain = Therefore, the total thrust of the four rotors at this speed is approximately 15.88 N, which is basically consistent with the hovering requirement of 15.89 N, indicating that the rotor thrust parameters are well matched with the overall mass parameters.

[0069] The pattern fusion factor is adopted as follows: Assume the desired control weight is approximately =Within 2.0 s, it transitions from 0.98 to 0.02. Taking the endpoint error ε = 0.02, then: .

[0070] When aircraft 1 transitions from free flight to cable-following mode, σ = -1 is set to smoothly decrease λ(t) from close to 1 to close to 0. When aircraft 1 returns from cable-following mode to free flight, σ = 1 is set to smoothly increase λ(t) from close to 0 to close to 1. Thus, the proportions of free flight control input and cable-following control input in the unified control input u(t) change continuously, avoiding abrupt changes in control mode.

[0071] Among the access conditions for the sling, the sampling machine opening width =82 mm, outer diameter of target cable 2 =60 mm. Considering ranging error, assembly error, residual attitude error, and cable 2 sway, the safe clearance is taken as: =12 mm, then: Substituting the values, we get: 82 mm ≥ 60 mm + 12 mm = 72 mm, indicating that the opening width meets the entry condition of cable 2.

[0072] In the alignment error of the opening, the normal deviation is taken. =4 mm, subnormal deviation =5 mm; the angle between the z-axis of the aircraft and the direction of cable 2 is 3°; the perpendicularity error between the x-axis of the aircraft and the direction of cable 2 is taken as =0.04. The attitude error calculation weight is taken as... =0.12m / rad, =0.08 m, then: Substituting the values, we get: ≈9.5mm, if the opening is aligned with the threshold. =15 mm, then It satisfies the alignment conditions for entering the light-touch entry section from the opening alignment section.

[0073] Lightly touch the entry point, with a light touch pressure range of 1.0 N ≤ ≤3.0 N, the pressure change rate threshold is: ≤4.0 N / s. When the pressure sensor detects that the normal contact pressure has entered the above range and the pressure change rate has not exceeded the threshold, it indicates that cable 2 has smoothly entered the middle hollow channel and can enter the closed cable clamping section.

[0074] In the stable cable section, the normal contact pressure warning value is taken. =8 N, upper limit of stable pressure =10N, exceeding the trigger pressure threshold. =12 N. When the normal contact pressure exceeds the warning value, the flight controller reduces the speed along the cable and generates a yield control quantity in the normal direction away from cable 2; when the normal contact pressure reaches the disengagement trigger threshold or fails to descend, it enters the opening disengagement phase.

[0075] The roller constraint reaction force adopts an equivalent spring-damped contact model. The equivalent contact stiffness of the four rollers is taken as: =2500 N / m, contact damping is: =6 N·s / m, in the stable section along the cable, the compression of the four rollers is approximately: δ=[0.52,0.47, 0.55, 0.50] mm, and the calculated total normal contact pressure is approximately: ≈5.3 N, which is below the upper limit of stable pressure of 10 N, indicating that the aircraft 1 can maintain low pressure while running along the cable.

[0076] In the opening separation section, the safe separation displacement is taken as follows: =0.15 m, expected detachment time: =1.5 s, based on a smooth acceleration and deceleration trajectory, the peak separation acceleration is estimated to be approximately: The corresponding additional release force is approximately: The additional force is much smaller than the total thrust of 15.89N when the aircraft is hovering, so a low-impact escape can be achieved through slight attitude and thrust adjustments.

[0077] The parameters in this embodiment are derived from prototype weighing, rotor rig calibration, mechanism dimension measurement, pressure sensor calibration, and roller compression testing, and are used to illustrate the engineering scope and calculation process of the control method of this invention. In practical applications, each parameter can be recalibrated according to the specific aircraft platform 1, the structure of the cable-holding mechanism, the outer diameter of the cable 2, and the on-site working conditions.

[0078] It should be noted that the sequence numbers of the embodiments in this application are for descriptive purposes only and do not represent the superiority or inferiority of the embodiments.

[0079] The terms "comprising" and "having," and any variations thereof, in the specification, claims, and accompanying drawings of this application are intended to cover non-exclusive inclusion. For example, a process, method, system, product, or apparatus that includes a series of steps or units is not limited to the listed steps or units, but may optionally include steps or units not listed, or may optionally include other steps or units inherent to such process, method, product, or apparatus. The terms "first," "second," and "third," etc., are used to distinguish different objects, etc., and do not indicate a sequence, nor do they limit "first," "second," and "third" to different types.

[0080] In the description of the embodiments of this application, terms such as "exemplary," "for example," or "for instance" are used to indicate examples, illustrations, or explanations. Any embodiment or design described as "exemplary," "for example," or "for instance" in the embodiments of this application should not be construed as being more preferred or advantageous than other embodiments or designs. Specifically, the use of terms such as "exemplary," "for example," or "for instance" is intended to present the relevant concepts in a concrete manner.

[0081] In the description of the embodiments of this application, unless otherwise stated, " / " means "or". For example, A / B can mean A or B. The "and / or" in the text is merely a description of the relationship between related objects, indicating that there can be three relationships. For example, A and / or B can mean: A exists alone, A and B exist simultaneously, and B exists alone. In addition, in the description of the embodiments of this application, "multiple" means two or more.

[0082] In some processes described in the embodiments of this application, multiple operations or steps are included in a specific order. However, it should be understood that these operations or steps may not be executed in the order they appear in the embodiments of this application, or they may be executed in parallel. The sequence number of the operation is only used to distinguish different operations, and the sequence number itself does not represent any execution order. In addition, these processes may include more or fewer operations, and these operations or steps may be executed sequentially or in parallel, and these operations or steps may be combined.

[0083] Through the above description of the embodiments, those skilled in the art can clearly understand that the methods of the above embodiments can be implemented by means of software plus necessary general-purpose hardware platforms. Of course, they can also be implemented by hardware, but in many cases the former is a better implementation method. Based on this understanding, the technical solution of this application, in essence, or the part that contributes to the prior art, can be embodied in the form of a software product.

[0084] The above are merely preferred embodiments of this application and do not limit the patent scope of this application. Any equivalent structural or procedural transformations made using the content of this application's specification and drawings, or direct or indirect applications in other related technical fields, are similarly included within the patent protection scope of this application.

Claims

1. A method for detecting smooth switching between dual modes of an aircraft, characterized in that, The method for smooth switching between dual modes of the detection aircraft includes: Based on the aircraft's motion attitude and its position relative to the cable, the current stage of the aircraft is determined. By combining the current stage of the aircraft with the normal contact pressure of the rollers inside the aircraft, the current control state of the aircraft is determined; Using the current control state, control the switching of the aircraft's free flight control mode or cable-stayed control mode.

2. The method for detecting smooth switching between dual modes of an aircraft as described in claim 1, characterized in that, The switching of control over the free flight control mode or the cable-stayed control mode of the controlled aircraft includes: A mode fusion factor λ(t) is introduced during the transition period of the shift in dominance; Based on the mode fusion factor, the free flight control input, the cable grip and cable-stayed control input, and the stage transition compensation term are continuously weighted to obtain the unified control input u(t); The target acceleration, target attitude, target total thrust, target torque, and rotor speed distribution commands of the aircraft are generated using the unified control input u(t).

3. The method for detecting smooth switching between dual modes of an aircraft as described in claim 1, characterized in that, The phases in which the aircraft is located include the flight approach phase, the opening alignment phase, the touch-in phase, the closing grip phase, the stable cable-along phase, and the opening disengagement phase. Determining the current phase of the aircraft based on its motion attitude and position relative to the cable includes: If the radial distance between the center of the hollow channel in the middle of the aircraft and the center line of the cable is greater than the preset distance threshold, the aircraft is determined to be in the flight approach segment. After the aircraft enters the flight approach phase, if the radial distance between the center of the hollow channel in the middle of the aircraft and the center line of the cable is less than a preset distance threshold, it is determined that the aircraft has entered the opening alignment phase. After the aircraft enters the opening alignment section, if the z-axis of the aircraft body is parallel to the cable axis, it is determined that the aircraft has entered the light touch entry section. After the aircraft enters the light-touch entry phase, if the cable normal contact pressure is detected to be within the preset light-touch pressure range and the pressure change rate is lower than the preset threshold, it is determined that the aircraft has entered the closed-grip phase. After the aircraft enters the closed grip section, if the aircraft's opening mechanism reaches the closed position, and the normal contact pressure is lower than the preset stable pressure upper limit, and the attitude deviation and along-the-cable velocity fluctuation are both less than the stable threshold, then the aircraft is determined to have entered the stable along-the-cable section. After the aircraft enters the stable cable section, if the detection task is completed, or the normal contact pressure exceeds the limit, the attitude deviation is too large, or the opening mechanism is abnormal, the aircraft is determined to have entered the opening separation section.

4. The method for detecting smooth switching between dual modes of an aircraft as described in claim 1, characterized in that, After determining the current stage of the aircraft, the following is included: Constructing a thrust field including the rotor system Aerodynamic disturbance field in the near-cable region Cable-constrained reaction field Mode switching additional force field The total target force field, including the gravitational field mg, is denoted as... ; The target thrust direction, target attitude matrix, and rotor speed distribution command are generated based on the total target force field.

5. The method for detecting smooth switching between dual modes of an aircraft as described in claim 4, characterized in that: The thrust field generated by the rotor system scalar of total thrust generated by the rotor system and rotor thrust vector in the airframe coordinate system It is concluded that the thrust field generated by the rotor system is... Rotor thrust vector in the airframe coordinate system and the total thrust scalar generated by the rotor system satisfy: ， ， ； in, Let i be the rotational speed of the i-th rotor. R is the rotor thrust coefficient, and R∈SO(3) is the rotation matrix from the body coordinate system to the world inertial coordinate system; The cable constraint reaction field Represented in the world coordinate system as ,and , , ; Where m is the number of internal rollers, The magnitude of the normal constraint reaction force of the cable on the j-th roller is represented by an equivalent spring-damped nonlinear contact model. Let j be the normal direction of the j-th roller contact point, which is the unit vector of the normal constraint reaction force in the world coordinate system. Let j be the frictional or rolling resistance of the j-th roller. The cable direction in the world coordinate system. This refers to the compression amount of the roller. The roller compression speed, The equivalent normal contact stiffness coefficient, This is the equivalent normal contact damping coefficient. For nonlinear contact stiffness coefficient; when When the value is ≤0, it indicates that the roller has not made effective contact with the cable. =0, Let j be the equivalent coefficient of friction or rolling resistance between the j-th roller and the cable. Used to indicate the direction of motion along the cable. The velocity damping coefficient along the cable direction is... The speed of movement along the cable; The aerodynamic disturbance field Recorded as: ; in: Relative airflow velocity The linear damping coefficient is... This is the second-order drag coefficient. The additional disturbance force generated by the interaction between the rotor wake and the cable. The disturbance force is caused by local eddies, and there is an additional disturbance force. satisfy , ; in, The relative distance between the aircraft and the cable. and For the disturbance amplitude parameter, The wake attenuation coefficient is... The angular frequency of the eddy current disturbance. This represents the phase of the eddy current disturbance. and The direction of the disturbance is denoted by t, and t represents the current control time. The mode switching additional force field for: ; in, Here, x represents the transition weight matrix for mode fusion, and x is the current state vector of the system. This is the reference state vector for the current stage. For safety boundary compensation weights, The gradient of the safety boundary function is used to represent the direction that increases the safety margin; Used to generate a compensating force away from the cable or to reduce contact pressure when the difference between the aircraft and the safety boundary is less than or equal to a preset value; λ(1-λ) is used to enhance the transition compensation during the intermediate stage of mode switching and weaken it at the endpoint of stable mode. For velocity damping weight, Let the state velocity vector be... Used to suppress speed abrupt changes and oscillations during the switching process.

6. The method for detecting smooth switching between dual modes of an aircraft as described in claim 5, characterized in that: The target attitude matrix is ​​denoted as , , , and ; in, The direction of the target thrust axis is determined by the overall target force field. Let z be the target airframe's z-axis. When the rotor thrust is output along the z-axis of the airframe, the z-axis of the target airframe is aligned with the direction of the target thrust axis. The target heading reference vector, In the direction of the target aircraft nose, Let y be the target body determined according to the right-hand rule.

7. The method for detecting smooth switching between dual modes of an aircraft as described in claim 6, characterized in that, The rotor speed distribution command is solved by the control distribution matrix: Recorded as , ; Where T is the target total thrust, , , These represent the target roll, pitch, and yaw moments, respectively, and A is the control allocation matrix.

8. The method for detecting smooth switching between dual modes of an aircraft as described in claim 3, characterized in that, After the aircraft enters the opening alignment section, the following is also included: Determine whether the opening width of the aircraft's opening mechanism is greater than or equal to the sum of the cable's outer diameter and the preset safety clearance; The opening alignment error is calculated based on the aircraft's normal deviation relative to the cable, subnormal deviation, and attitude direction deviation between the aircraft axis and the cable direction. If the opening width of the aircraft's opening mechanism is greater than or equal to the sum of the cable's outer diameter and the preset safety gap, and the opening alignment error is less than the preset value, then the aircraft is controlled to enter the touch-entry phase.

9. The method for detecting smooth switching between dual modes of an aircraft as described in claim 3, characterized in that, If the detection task is completed, or the normal contact pressure exceeds the limit, the attitude deviation is too large, or the opening mechanism is abnormal, then the aircraft is determined to have entered the opening disengagement phase, including: An emergency judgment vector and a comprehensive risk function are constructed based on normal contact pressure, attitude deviation, rotor state, wind disturbance estimate, normal offset, and open mechanism state. Emergency control measures are implemented at the warning, suppression, departure, or return-to-base levels based on the comprehensive risk function.

10. The method for detecting smooth switching between dual modes of an aircraft as described in claim 2, characterized in that, Before generating the target acceleration, target attitude, target total thrust, target torque, and rotor speed distribution commands for the aircraft using the unified control input u(t), the following is also included: A joint error energy function is constructed, which characterizes the combined deviation of the aircraft's current force state, attitude state, and angular velocity state relative to the target state, and is denoted as: ； ； ； in, For force field error, For the current actual or estimated total force field of the aircraft, For the overall target force field; Let R be the attitude error, and R be the current attitude matrix. The target attitude matrix; For angular velocity error, , and It is a positive definite weight matrix used to adjust the weights of force field error, attitude error and angular velocity error in the joint error energy function; This represents the mapping from an antisymmetric matrix to a vector.

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