Research On Control Method For The Transition Mode Of The Small Tilt-Rotor UAV

Small tilt-rotor UAVs facilitate the use in applications and have great potential applications with the advantages of low cost,high speed and long-endurance flight in fixed-wing mode,and vertical takeoff and landing,hovering,and low speed flight in rotor mode.The small tilt-rotor UAV has these advantages because it has the transition mode to realize the switching between the fixed-wing mode and the rotor mode through rotor tilting.Meanwhile,there are great challenges to the modeling and flight control in the transition modes.In terms of modeling,it is necessary to consider both the vector effect of thrust of the rotor mode and the aerodynamic characteristics of the fuselage of the fixed-wing mode.With the tilting of the rotor,the time-variant mutual airflow interference between the rotor and the wings lead to the change of the position of the center of gravity of the fuselage.In terms of control,the tilt-rotor UAV in the transition mode presents nonlinear and multi-channel coupling characteristics and input non-affinity due to the vector effect of thrust,and the differences or even contradictions between the control logic of the rotor mode and the fixed-wing mode and the different number of types of actuators even exacerbate the difficulty in smooth transition of state and control logic of the control system.In addition,different from only motion control of the system state in ordinary rotor UAVs and fixed-wing UAVs,the transition mode also requires strategies to plan the tilting movement of the rotor to ensure flight safety.The control system needs to be designed according to the strategies,so as to ensure a smooth flight in the transition mode.In this thesis,based on the research on the key technologies of tilt-rotor UAVs and the development of cluster combat platform,the research on the design,nonlinear dynamics modeling and flight control method of the small tilt-rotor UAVs is carried out.The main research work and innovative points are as follows:1,With the optimal design of the small tilt-rotor UAV system,the mutual airflow interference between the rotor and the fixed wings is reduced,and the changes of the center of gravity and moment of inertia of the craft caused by the tilting of rotor motor assembly are eliminated;and compared with other tilt-rotor UAVs in the prior art,the small tilt-rotor UAV in this research has a higher endurance-weight ratio.Firstly,the fixed-wing configuration,rotor configuration,tilting mechanism and battery of the tilt-rotor UAV system are optimized,which reduces the mutual interference of airflow between the rotor and the fixed wings,eliminates the changes of the center of gravity and moment of inertia of the craft caused by the tilting of rotor motor assembly,and satisfies the index of total endurance load.Therefore,the difficulty in modeling and controller design is reduced from the stage of system design and construction.Then,the flight control system composed of open source flight control and commercial high performance sensors,and the fusion processing of multi-sensor data by the Kalman filter algorithm provide reliable data guarantee for system identification and control system design.According to the characteristics of the system,the control logic for each mode and the mission profile for the full-mode flight are designed.Compared with other tilt-rotor UAVs,the constructed system has a higher endurance-weight ratio and a satisfactory payload-weight ratio,which lay foundation for subsequent modeling,control method design and strategy design in this thesis.2,The method of combining dynamics modeling and system identification is used for modeling of the small tilt-rotor UAV;and the consistency between the dynamics model and the actual system response curve verifies the accuracy of the model.Based on the fixed-wing UAV modeling method,the modeling method combines dynamics modeling and system identification.First of all,the Newton-Euler method is employed for modeling of the forces and moments of the fuselage and the rotor;and combined with the rigid-body dynamics modeling method,the dynamics modeling of different modes is represented by the same model.Then,the unknown non-aerodynamic and aerodynamic coefficients in the model are respectively determined by the system identification method.In the process of determining the non-aerodynamic coefficients,the improved compound pendulum method is used to improve the problem of sensitivity of the moment of inertia measurement to the cycloid length.The servo calibration method is used to solve the problem that the sensor can not be installed.The rotor tension and torque coefficient are obtained by means of measurement fitting.Then the aerodynamic coefficients are identified based on the non-aerodynamic coefficient measurement results.In the identification of the unknown aerodynamic coefficients of the model,the input signal and the experimental procedure are designed;the preprocessing unifies the data sampling frequency,removes the high frequency noise,and acquires the angular acceleration data through the combination of global smooth and polynomial fitting derivation;and the equation error method is adopted for pneumatic coefficient identification.The identification results show that the partial derivatives of the aerodynamic coefficients are similar and correspond to the actual fuselage design.The verification results show that the fitting degree is high and the root mean square error is small,which proves the accuracy of the aerodynamic coefficient identification.Finally,the consistency between the dynamic model and the actual system response curve further proves that the established model can represent the dynamic characteristics of the system and lays foundation for the control method research.3,The tilting angle jump strategy is optimized and perfected,the control logic weighting coefficient of rotor and fixed wings based on this strategy is designed,and the transition mode iterative control allocation method is put forward,so as to realize a fast transition mode.Used in this system,the tilting angle jump strategy widely used in the open-source flight control causes attitude and altitude vibration,and can not guarantee the flight safety of this system;and due to the tilting effect of the tilting angle,the constant allocation matrix widely used in the open-source flight control is not suitable for the transition mode.In view of these two problems,this thesis preserves the controller design and control system framework of the open-source flight control.First,the tilting angle jump strategy is optimized and perfected,and then the control logic weighting coefficient of rotor and fixed wings based on this strategy is designed,and iterative allocation method,which is based on transition mode dynamics and separately calculates rotor speed and tilting angle,is put forward.The attitude control simulation of the tilting process and the actual experimental results show that the iterative allocation can ensure the system stability and realize more accurate attitude tracking than the constant allocation matrix.The transition mode simulation and the actual flight experiment show that the small tilt-rotor UAV system achieves the transition mode flight and solves the problems of coexistence of two actuators,redundancy of the control mechanisms and input non-affinity.The transition mode flight takes 6s and has an altitude variance of 6m,thereby demonstrating the effectiveness of iterative allocation and of optimized strategy.4,A cascade control system based on tilting angle gradient strategy is proposed.The position loop,speed loop,angle loop and control allocation of the system are designed respectively according to different airspeeds,a high-speed section and a low-speed section,thereby enhancing the anti-interference ability and lateral position control ability of the transition mode,and ensuring less time consumption while significantly improving the stability of the transition mode.Aiming at the problems of poor anti-disturbance ability,great fluctuation of the transition mode attitude and lateral position diverges in the experiments adopting the tilting angle jump strategy,the cascade control system consisting of position control,speed control,angle control,angular velocity control and distributors is first designed by using the tilting angle gradient strategy.Dynamics design of high-speed section and low-speed section controllers are conducted with the position,speed and angle loop in the lateral direction and radial direction respectively.The angular velocity and speed loop are controlled by the sliding mode control method based on disturbance observer.Then,the control allocation employs the weight coefficient adjustment with airspeed as a variable to realize the transition of two different kinds of control logic and actuators of the rotor mode and the fixed-wing mode,and the non-affine input is allocated by way of variable substitution.Finally,the termination conditions and pitch angle expectation in the tilting process of the tilting angle gradient strategy are designed.Non-interference and interference simulation experiments show that the transition mode takes less than 10 s,and the altitude variance does not exceed 0.6m;the attitude is stable,and the actuators are basically non-saturated;the anti-interference ability and lateral position control ability of the transition mode are enhanced;and the rapidity of the transition mode is ensured and the stability of the transition mode flight is significantly improved.Therefore,the effectiveness of the control system and control strategy is validated.