183 results about "Rolling angle" patented technology

The invention provides a stabilization control method and system for a fuel filling ship at a zero navigational speed. The method provided by the invention comprises the following steps: acquiring rolling data and wind wave data of a refueling ship under a target sea condition in a fuel refueling process in real time; the rolling data and the wind wave data are processed based on a radial basis function neural network, and the precession angular velocity of the stabilization gyro device under the target sea condition is obtained through calculation; the stabilization gyro device is controlled to rotate at the precession angular speed, the stabilization gyro device generates a gyroscopic effect and generates rolling moment, the rolling moment is converted into stabilizing moment for resisting rolling through precession motion, and the rolling angle of the filling ship is reduced; and the stabilizing moment is opposite to the vibration direction of the rolling. According to the method and system, rolling control of the filling ship at the low speed or the zero navigational speed is achieved, the stabilization torque can be adjusted in real time according to sea wave changes, and the stability and safety of the filling ship in the fuel filling operation process are improved.

The invention relates to the technical field of aero-engines, in particular to a lubricating oil tank liquid level detection method, device and equipment, a medium and a program product. The method comprises the steps that the measured liquid level of oil in a lubricating oil tank of the aircraft is obtained, and current attitude information of the aircraft is obtained; the attitude information comprises a pitch angle, a yaw angle and a roll angle; based on a standard corresponding relation, a target standard liquid level corresponding to the current attitude information is determined, and the standard corresponding relation is the corresponding relation between the attitude information and the liquid level of the aircraft when oil in a lubricating oil tank is in a normal state; and determining a detection result based on the size relationship between the target standard liquid level and the measured liquid level. By means of the method, the precision of liquid level measurement of the oil in the lubricating oil tank can be improved.

The invention discloses a fixed-wing aircraft large-angle-of-attack landing control method based on a dynamic reference trajectory, and belongs to the technical field of fixed-wing aircraft autonomous landing, and the method comprises the following steps: establishing a landing control strategy based on a kinematic model and an attitude angle conversion model in the fixed-wing aircraft autonomous landing process, designing two-stage annular auxiliary line radiuses of an air flight stage of the fixed-wing aircraft, designing an expected track yaw angle, an expected track inclination angle and an expected roll angle of the fixed-wing aircraft, solving the expected yaw angle and the expected pitch angle, and constructing a control law of an attack angle when the fixed-wing aircraft is about to land; aiming at a flight stage and a taxiing stage of fixed-wing aircraft landing, a proper angular velocity controller and an air speed controller are designed based on a dynamic reference trajectory, so that the pose of the fixed-wing aircraft and a large attack angle when the fixed-wing aircraft is about to land are controlled, the dangerous conditions of collision and stall are avoided, and the safety of the fixed-wing aircraft is improved. And the accuracy and safety of autonomous landing control are improved.

The invention discloses a roll angle measuring method for a continuous variable roll test of a transient wind tunnel, and belongs to the technical field of wind tunnel test. Comprising the following steps: step S1, early-stage preparation: selecting two angle measurement sensors to carry out measurement work, and installing the two angle measurement sensors and an installation platform according to a preset angle; s2, carrying out field calibration operation; carrying out field calibration on the coefficient of the angle measurement sensor by adopting a fixed attack angle and variable roll angle mode; and S3, data interception, discretization and application during a step variable attack angle continuous variable roll angle test: in a wind tunnel test process, continuously collecting output voltage values of the two angle measurement sensors, calculating roll angles corresponding to the output voltage values, intercepting roll angle range values corresponding to the two angle measurement sensors according to needs, and calculating the roll angle range values of the two angle measurement sensors according to the intercepted roll angle range values. Splicing into a target measurement range; compared with a traditional step variable rolling and fixed rolling variable attack angle test mode, the scheme has the advantage that the test efficiency is greatly improved.

The invention relates to an aircraft test flight multi-scene gravity center adjusting method and system. The method comprises the following steps: receiving target gravity center data required by test flight and a rolling angle and a pitching angle of an aircraft; judging whether the rolling angle is greater than a rolling angle threshold value or not; when the rolling angle is larger than the rolling angle threshold value, the transverse pipe valve between the water tanks is closed, and gravity center adjustment is suspended; when the rolling angle is smaller than or equal to the rolling angle threshold value, a transverse pipe valve between the water tanks is opened, and whether the pitching angle is smaller than a first pitching angle threshold value or not is judged; when the pitching angle is smaller than a first pitching angle threshold value, the gravity center is adjusted in a water tank external circulation mode; when the pitching angle is larger than or equal to a first pitching angle threshold value and smaller than a second pitching angle threshold value, the second pitching angle threshold value is larger than the first pitching angle threshold value, and gravity center adjustment is conducted in a water tank internal circulation mode; when the pitching angle is larger than or equal to the second pitching angle threshold value, the gravity center adjustment is suspended.

The invention belongs to the field of aviation structure strength tests, and particularly relates to a testing machine pitching and rolling angle adjusting method and system for a full-aircraft drop test, and the method comprises the steps: adjusting the lengths of a front lifting rope l1 and a rear lifting rope l2 through a first servo motor; the lengths of the adjusted front lifting rope l1 and the adjusted rear lifting rope l2 are calculated through the cosine law according to the geometrical relationship; the lengths of the left lifting rope 13 and the right lifting rope 14 are adjusted through a second servo motor; and the lengths of the adjusted left lifting rope 13 and the right lifting rope 14 can be calculated through the cosine law according to the geometrical relationship. The pitching and rolling postures can be automatically adjusted, and the problems that an existing method is low in efficiency and poor in precision are solved. According to test experience, according to an existing posture adjusting method, replacement needs about half a day to one day every time, the number of times of putting every day does not exceed ten at most, and manual operation is needed on wings or a fuselage every time. And after automatic adjustment is achieved, the target can be achieved in several minutes through adjustment of the motor, efficiency is high, and safety is higher.

The invention discloses a compound control method for wind interference resistance of a quad-rotor unmanned aerial vehicle, and relates to the technical field of engineering control. The method comprises the following steps: acquiring a speed variable, an attitude angular speed variable and a control parameter when the quad-rotor unmanned aerial vehicle is in a wind field; inputting the speed variable, the attitude angular speed variable and the control parameters into a nonlinear disturbance observer, and determining an external disturbance torque and an external disturbance force generated by the wind field on the quadrotor unmanned aerial vehicle; obtaining the position control rate of the four-rotor unmanned aerial vehicle according to the external disturbing force generated by the wind field on the four-rotor unmanned aerial vehicle; according to the speed component and the expected yaw angle of the quad-rotor unmanned aerial vehicle in the wind field, an expected pitch angle and an expected roll angle are obtained through calculation; according to the external interference torque, the expected pitch angle, the expected roll angle and the expected yaw angle, the attitude control rate of the four-rotor unmanned aerial vehicle is obtained; and controlling the position and the attitude of the quadrotor unmanned aerial vehicle through the position control rate and the attitude control rate of the quadrotor unmanned aerial vehicle.

The invention discloses a wind tunnel multi-body multi-degree-of-freedom separation simulation test system which comprises an attack angle mechanism, a radial radiation type support, a right-angle supporting rod, a yaw assembly, a linear electric cylinder, a pitching assembly, a rolling assembly, a sublevel external store and a main level model, and the attack angle mechanism installed on a wind tunnel test section achieves attitudes such as the attack angle, the rolling angle and the yaw angle of the main level model. The radial radiation type support, the right-angle supporting rod, the yaw assembly, the linear electric cylinder, the pitching assembly, the rolling assembly and the sublevel external store form a series connection type six-degree-of-freedom movement device, and each sublevel external store can manually or automatically achieve six-degree-of-freedom control relative to the main-level model. The plurality of sublevel external store and the driving system are arranged along the circumferential direction as required relative to the main-level model, each driving device is controlled to move, the multi-degree-of-freedom position and attitude change of the plurality of sublevel external store is realized in a wind tunnel test section, and the simulation of the motion trail of the sublevel external store and the load measurement of the pneumatic relative interference under the multi-body and multi-attitude condition are realized.

The invention relates to the technical field of aircrafts, in particular to a sensor parameter adjustment method and device, equipment and a medium, and the method comprises the steps: obtaining the acceleration data of an unmanned plane, the acceleration data of a holder, and the sensing data; and acquiring a detected pitch angle and a detected roll angle based on the sensor parameters and the sensing data. The pitching angle and the rolling angle of the holder are obtained, and the pitching angle and the rolling angle of the unmanned aerial vehicle are obtained. And obtaining an estimated pitch angle according to the pitch angle of the holder and the pitch angle of the unmanned aerial vehicle, and obtaining an estimated roll angle according to the roll angle of the holder and the roll angle of the unmanned aerial vehicle. If the angle difference between the estimated pitch angle and the detected pitch angle is larger than or equal to an angle difference threshold value, or the angle difference between the estimated roll angle and the detected roll angle is larger than or equal to the angle difference threshold value, sensor parameters are adjusted. According to the scheme, the parameters of the sensor can be adjusted in a wide range of scenes, normal use of the holder is ensured, and the user experience is improved.

The invention provides a hovering control method, device and equipment for a tilt-configuration aircraft and a medium. The method comprises the steps that an acceleration instruction of the aircraft is obtained in real time; the method comprises the following steps: acquiring a tilt angle of a tilt mechanism of the aircraft, acquiring a control distribution matrix when the tilt angle is greater than a preset angle, and converting an acceleration instruction into a target control instruction by using the control distribution matrix, thereby performing hovering control on the aircraft based on the target control instruction. The target control instruction comprises a roll angle control instruction, a tension control instruction and a pitch angle control instruction. Compared with the prior art, hovering control does not need to depend on a hovering datum point, the current flight state of the aircraft is considered, control is conducted by controlling the distribution matrix, hovering control can be matched with the actual state of the aircraft, and the risk of the aircraft in the hovering process is reduced.

The application discloses a launching projectile support for irregular projectile body self-defined rolling angle and an assembling method thereof, which comprises a support body, the support body comprising an upper shell, a lower shell, positioning columns and a lever rod, the upper shell and the lower shell being coaxially and threadedly connected at the ends, a cavity with a closed end and an open end being formed between the upper shell and the lower shell, a projectile body being arranged in the middle of the cavity, the positioning columns filling the remaining space of the cavity and supporting the projectile body, the lever rod comprising a lifting rod, a flat plate lever and a supporting column, the flat plate lever being horizontally arranged on the end face of the lower shell, the lifting rod being arranged in the upper shell along the axial direction, one end of the lifting rod being in contact with one end of the flat plate lever, and the supporting column being arranged in the lower shell along the axial direction, one end of the supporting column being in contact with the other end of the flat plate lever. The application can well position various irregular projectile bodies, thereby meeting the demand for various specifications of projectile bodies in experiments, and the application is convenient for experiments, saves manpower and material resources, and guarantees the accuracy of experimental results and the effectiveness of data.

The present application belongs to the technical field of aircraft landing control on a deck, and specifically relates to a method and system for power compensation of an aircraft landing on a deck at a constant angle of attack. The method comprises: using the angle of attack #imgabs0#, normal overload #imgabs1#, pitch angle rate #imgabs2#, longitudinal instruction #imgabs3#, and roll angle #imgabs4# as feedback signals to calculate the throttle stick angle compensation instruction #imgabs5#: #i mgabs6#; among them, #imgabs7# is the reference angle of attack, #imgabs8# is the angle of attack integral gain coefficient, #imgabs9# is the angle of attack proportional gain coefficient, #imgabs10# is the normal overload gain coefficient, #imgabs11# is the pitch angle rate gain coefficient, #imgabs12# is the longitudinal command feedforward gain coefficient, and #imgabs13# is the roll angle gain coefficient. The throttle stick is controlled by calculating the throttle stick angle compensation command #imgabs14#.

The present invention provides a dynamic flight simulation control system for a manned centrifuge, belonging to the field of flight control, including: The roll module obtains the first load that the centrifuge needs to simulate through the flight attitude based on the offset of the cockpit joystick, and calculates the cockpit roll angle when the load direction is consistent with the human head-foot axis based on the first load; The offset module obtains the cockpit offset angle at which the human head-foot axis deviates from the load direction based on the first load, the maximum offset angle, and the maximum load; the maximum load is the maximum value among the first load, the second load, and the third load, the second load is obtained through the centrifuge model based on the angle of the cockpit joystick shake, and the third load is obtained based on the maximum offset angle; The control module is used to control the cockpit rotation angle as the sum of the cockpit roll angle and the cockpit offset angle. The present invention reduces the roll angle by setting the maximum offset angle φ and making the cockpit roll from (45 + φ)°; uses a conditional function to flexibly set the relationship between the load and the offset angle, reducing the Coriolis effect.

The application designs a high-speed aircraft cluster cooperative guidance method based on space-time synchronous attraction domain. An artificial potential field is designed to keep a proper distance between aircrafts, avoid collision while gathering; a speed matching item is designed based on consistency theory to realize dynamic consistency of aircraft speed; a guidance law based on resistance acceleration-speed profile is designed to meet the cooperative requirements of terminal range, height, speed and other aircrafts, and the combination of the three realizes efficient autonomous clustering and guidance of the aircraft. The application can make the aircraft cluster reach the target point with high precision under the conditions of meeting complex constraints such as heat flow, overload, dynamic pressure, collision avoidance, clustering, dynamic consistency of speed and the like through the design of attraction domain, angle of attack and roll angle control.

A missile rolling attitude control method with integrator dynamic limiting, belonging to the field of flight control technology. The present invention includes: a guidance command calculation unit calculates the rolling command of the missile and sends it to the rolling attitude controller, and a sensitive element unit measures the rolling angular velocity of the missile and sends it to the rolling attitude controller; the rolling attitude controller calculates the upper and lower bounds of the integrator output value according to the maximum available rudder deflection; limits the integrator output according to the upper and lower bounds; calculates the variable structure controller output according to the information after the limiting process; calculates the rolling rudder command according to the variable structure controller output and the rolling angular velocity; the rolling rudder command is sent to the rudder system to control the movement of the missile. The present invention solves the problem that when the missile rolls at a large angular velocity, due to the too large integrator output value, the rolling rudder deflection is in the limited state for a long time, and the rudder deflection needs a long time to get out of the saturation area, resulting in a large amplitude oscillation of the missile rolling angular velocity and unstable attitude.

The invention provides a high-precision roll angle measurement method and system, and relates to the technical field of optical measurement, and the method comprises the following steps: respectively measuring corresponding first roll angle data, second roll angle data and environment data; performing preprocessing and error estimation on the first roll angle data and the second roll angle data in combination with the environmental parameters, and generating a corresponding to-be-processed sequence according to a data processing result; constructing a roll angle measurement model, and inputting the to-be-processed sequence into the roll angle measurement model for processing to obtain a corresponding roll angle measurement result; according to the invention, through error compensation and data fusion, part of high-precision and high-cost hardware sensor requirements are replaced, and the measurement precision can be effectively improved.

The application discloses a roll aircraft control method based on a two-loop roll attitude pilot, in which a roll angle error and a roll angle rate are obtained in real time, a roll rudder deflection control instruction corresponding to a unit roll angle error and a roll attitude pilot damping coefficient are further weighted, so that a rudder deflection control instruction is obtained and is transmitted to a rudder, the rudder works based on the rudder deflection control instruction, and the roll aircraft is controlled to fly to a target.

The invention provides an aircraft roll angle reconstruction method and system, and belongs to the technical field of aircraft control, the method comprises the steps that aircraft state signals are acquired, and the state signals comprise a ground speed signal Vg, a vacuum speed signal Vt, a yaw angle rate signal R, a pitch angle rate signal Q and a pitch angle signal Pitch; and under the condition that the roll angle signal fails, according to a lift force and gravity balance principle in a rolling state of the airplane, an approximate roll angle reconstruction signal is constructed based on a ground speed signal or a vacuum speed signal, a yaw angle rate signal, a pitch angle rate signal and a pitch angle signal.

This invention provides a high-dynamic space-based cooperative guidance method and system for multi-level virtual target points, relating to the field of aircraft control technology. It involves acquiring initial flight state data of each aircraft and position data of the target point; generating at least one virtual target point based on the initial flight state data and position data, using a zero heading angle constraint; generating roll angle control command data for each aircraft based on the initial flight state data and the virtual target point, using a two-sided azimuth error corridor; and adjusting the attitude of each aircraft according to the roll angle control command data to ensure that each aircraft meets the terminal heading angle constraint and reaches the target point. This method enables precise control of the terminal heading angle of high-dynamic aircraft in complex dynamic environments, thereby enabling multiple aircraft to perform cooperative strike missions from different spatial orientations.

The invention belongs to the technical field of control over landing of an aircraft on a deck, and particularly relates to an attack angle constant approach power compensation method and system for landing of the aircraft on the deck. The attack angle constant approach power compensation method for landing of the aircraft on the deck comprises the steps that an attack angle # imgabs0 #, a normal overload # imgabs1 #, a pitch angle rate # imgabs2 #, a longitudinal instruction # imgabs3 # and a roll angle # imgabs4 # are used for compensating the attack angle constant approach power of the aircraft on the deck according to the attack angle # imgabs0 #, the normal overload # imgabs1 #, the pitch angle rate # imgabs2 #, the longitudinal instruction # imgabs3 # and the roll angle # imgabs4 #; a throttle lever angle compensation instruction # imgabs5 #: # imgabs6 # is calculated as a feedback signal; wherein the # imgabs7 is a reference angle of attack, and the # imgabs8 is an angle of attack integral gain coefficient; # imgabs9 # is an incidence angle proportional gain coefficient; # imgabs10 # is a normal overload gain coefficient; the # imgabs11 # is a pitch angle rate gain coefficient; # imgabs12 # is a longitudinal instruction feed-forward gain coefficient; the # imgabs13 # is a roll angle gain coefficient, and the throttle lever is controlled according to the calculated throttle lever angle compensation instruction # imgabs14 #.

The application discloses a kind of double hulls lift-float integrated aircraft roll adjusting method, it is related to aircraft design technical field, using the form of auxiliary air bag control, main air bag control and full-motion horizontal tail control combination to carry out double hulls lift-float integrated aircraft roll adjusting, using the mode that auxiliary air bag is connected with atmosphere environment and air is pumped out, main air bag is connected and helium is pumped out, full-motion horizontal tail deflection, changes the lift distribution of aircraft, buoyancy distribution, so that aircraft generates roll torque, adjusts aircraft roll angle.The double hulls lift-float integrated aircraft roll adjusting method that the application adopts has significantly improved the aircraft roll adjusting efficiency, is conducive to improving the maneuverability and safety of aircraft, compared with the existing aircraft attitude adjusting mode, combines the characteristics of double hulls lift-float integrated aircraft, aircraft roll adjusting mechanism work can generate greater moment change, aircraft has very high adjusting capacity to attitude.

The application discloses a kind of high-efficiency omni-directional decoupling annular vector tilt-rotor aircraft and control method thereof, aircraft includes fuselage, fuselage driving gear set, four tilting mechanisms, four annular vector tilt-rotor power groups and electronic equipment etc.Parts.The aircraft described in the application can realize six degrees of freedom omni-directional pose decoupling flight, the attitude angle is omni-directional arbitrary angle in it, including roll angle, pitch angle and yaw angle, can realize arbitrary attitude hovering and flight of aircraft.The aircraft described in the application solves the controllable attitude angle limit of conventional variable tilt-rotor mechanism from structure, improves the ability of complex space activity of aircraft, and fundamentally solves the problem of power offset and power redundancy, reduces energy loss, improves power efficiency, prolongs endurance time, and is suitable for application scenarios requiring strong anti-interference and air physical interaction capability, and has important significance and value in the future.

The invention provides a carrier rocket rolling attitude angle control method, which comprises the following steps of: adding an angle deviation amplitude limiter and an integrator with amplitude limiting in a basic rolling attitude angle control loop to obtain an optimized rolling attitude angle control loop; the rolling attitude angle and the rolling angle speed are input into an optimized rolling attitude angle control loop, and a rolling channel control swing angle without an integration link and an output result of a rolling angle deviation band amplitude limiting integrator are obtained; and adding the roll channel control swing angle without the integration link and the output result of the roll angle deviation band amplitude limiting integrator to serve as a roll channel attitude control swing angle instruction, and controlling the roll attitude angle of the carrier rocket according to the roll channel attitude control swing angle instruction. On one hand, attitude control errors caused by constant value interference can be eliminated, the rolling attitude control precision is improved, and on the other hand, the rolling angular velocity during attitude maneuver can be limited, and the situation that the rolling angular velocity is too large is avoided.

Proposed are a system and a method for monitoring a safety status of a ship using roll motion data. The system and the method may be configured as follows. Rolling period data, average list angle data, and significant rolling angle data may be received from a rolling period computation part, an average list angle computation part, and a significant rolling angle computation part. It may be determined whether the rolling period data is equal to or less than a maximum unique rolling period. On the basis of the maximum unique rolling period, a controller may determine whether a rolling period is in a normal, caution, or danger state.

Systems and methods are provided for detecting and warning of aircraft rollover conditions. The systems include one or more sensors onboard the aircraft and a controller onboard the aircraft that is in operable communication with the one or more sensors. The controller is configured to, by one or more processors: receive input data from the one or more sensors and, optionally, one or more other sources, process the input data to dynamically determine a roll angle of the aircraft, an angular roll rate of the aircraft, and a critical rollover angle of the aircraft, determine whether current or potential static or dynamic rollover conditions of the aircraft exist based, at least in part, on the roll angle, the angular roll rate, and the critical rollover angle, and generate an alert in response to determining that the current or potential static or dynamic rollover conditions exist.

The application relates to the technical field of aero-engines, in particular to a lubricating oil tank liquid level detection method, device, equipment, medium and program product. The method comprises the following steps: acquiring a measured liquid level of oil liquid in a lubricating oil tank of an aircraft, and acquiring current attitude information of the aircraft; the attitude information comprises a pitch angle, a yaw angle and a roll angle; based on a standard corresponding relationship, a target standard liquid level corresponding to the current attitude information is determined; the standard corresponding relationship is a corresponding relationship between the attitude information of the aircraft and the liquid level when the oil liquid in the lubricating oil tank is in a normal state; and based on the size relationship between the target standard liquid level and the measured liquid level, a detection result is determined. The method can improve the accuracy of the liquid level measurement of the oil liquid in the lubricating oil tank.

The application discloses a roll angle measurement method for continuous variable roll test of a temporary flush wind tunnel, and belongs to the technical field of wind tunnel testing. The method comprises the following steps: step S1, preliminary preparation, two angle measurement sensors are selected for measurement, and the two angle measurement sensors are installed at a predetermined angle with a mounting platform; step S2, on-site calibration operation; on-site calibration of the angle measurement sensor coefficient is performed in a mode of constant attack angle and variable roll angle; step S3, data interception, discretization and application during step-by-step variable attack angle and continuous variable roll angle test, output voltage values of the two angle measurement sensors are continuously collected during the wind tunnel test, and corresponding roll angles are calculated, and the corresponding roll angle range values of the two angle measurement sensors are intercepted and spliced into a target measurement range as required; compared with a traditional step variable roll and constant roll variable attack angle test mode, the scheme greatly improves test efficiency.

The invention belongs to the field of aircraft drop simulation and test, and particularly relates to an aircraft sinking speed determination method and system under the condition of no lift force and no mass reduction, and the method comprises the steps: determining parameters needed by the dynamics simulation of an aircraft landing gear according to a test working condition; determining the mechanical relationship of the undercarriage according to the parameters required by simulation, and building a full-aircraft landing simulation model; performing lift simulation on the full-aircraft landing simulation model, and determining a vertical ground load of the aircraft landing gear; extracting a vertical ground load of an aircraft landing gear, performing lift-free drop simulation on the full-aircraft landing simulation model, and iteratively determining a putting height according to the ground load; and calculating the sinking speed of the full-aircraft drop test according to the putting height. Compared with the prior art, the method has the advantages that lift force simulating loading equipment is omitted, initial launching attitudes such as the pitch angle, the roll angle and the yaw angle of the aircraft are fully considered, the method can be used for guiding initial launching of a drop test, the large overload risk of free falling impact is reduced, and the number of times of adjusting the launching height can be reduced.

The application is a kind of lateral asymmetry online trimming control method of hypersonic aircraft, which comprises the following steps: firstly, constructing a lateral asymmetry hypersonic aircraft model; then, setting a sideslip angle and a roll angle judgment threshold of the hypersonic aircraft in the lateral asymmetry hypersonic aircraft model, detecting the lateral asymmetry of the hypersonic aircraft, judging whether the lateral asymmetry exists, if the lateral asymmetry exists, judging whether the hypersonic aircraft is controllable according to the detected abnormal information, if the hypersonic aircraft is uncontrollable, dynamically controlling the rudder of the hypersonic aircraft through control law reconstruction, and giving a systematic solution of the lateral asymmetry of the hypersonic aircraft from the aspects of asymmetry model construction, detection method, controllability analysis and automatic trimming control law design, so that the roll and sideslip caused by the asymmetry can be corrected without increasing additional trimming surfaces, and the flight safety of the hypersonic aircraft is ensured.

The application belongs to the field of aircraft flight control system, and particularly relates to an anti-collision control method based on artificial potential field, collision conditions of two aircrafts are set according to relative motion conditions and current flight data of the two aircrafts; the collision conditions of the two aircrafts are obtained, a potential field function is set, and a distance between the two aircrafts is designed as a virtual repulsive potential field; whether there is a collision risk is judged through the potential field function, if yes, a negative gradient in the potential field function is calculated, is converted into repulsive force between the two aircrafts, the repulsive force between the two aircrafts is solved into an aircraft roll angle instruction and a speed increment instruction, an anti-collision instruction is formed to control the two aircrafts; a relative position vector and a closest time of the two aircrafts are obtained, whether the anti-collision condition is met is judged, and the anti-collision is exited after the condition is met; by adopting the potential field function and performing instruction tracking, it is ensured that the two aircrafts can keep a certain relative distance and do not collide.