Research On Some Key Technologies Of Flight Control System Of UAV

UAV plays an increasingly important role throughout the whole aviation ministry.Flight control system is one of the core components of the UAV, whose performancedirectly affects the flight performances, flying qualities and flight safety. On the basisof summarizing the research results and development status of related fields, thisthesis has a central research on a fixed-wing UAV, including dynamics modeling offlying and taxiing, model simplifying, control law design and implementation,disturbance observer design, clearance of control law, hardware in loop simulationand flight test.Firstly, considering the engine torque, thrust eccentric and ground angle, the sixdegree of freedom nonlinear equations of flying and taxiing is established.Taking theUAV moving along the centerline of the runway constantly without sliding as thereference movement, the actual movement is small perturbation motion based on thereference movement. The linear state space expression of flying, taxiing with threewheels and two wheels is derived using small perturbation principle, which lays afoundation for design and clearance of the control system.The robust servo LQR control method is studied, the longitudinal and lateralflight control law is designed combining the robust servo LQR and the classicalcontrol. The simulation results shows that the response of the method is more gentlethan conventional PID controller, with the overshoot of pitch angle responsedecreasing by50%, and the overshoot of roll angle response decreasing by50%. Thephenomenon of a larger angle rate because of servo command mutation is suppressedin initial response, the requirements of available overload is greatly reduced. The control structure of flight path tracking loop uses the feedback of yaw rate, yaw angle,yaw speed and yaw distance. The simulation results show that this control method cantrack the default flight path accurately.Disturbance observer is applied to the flight control law design to eliminate theeffects of hinge moment to control accuracy during flight, the independence ofdisturbance observer design and controller design is proved, the robust stability ofdisturbance observer is analyzed. Taking the circle flight mode for example, thesimulation results show that the control law with a disturbance observer can eliminatethe3.2°roll angle steady-state error generated by the control law without disturbanceobserver, the control accuracy of the roll angle is improved. The nonlinear control lawdiscretion issues in engineering implementation is solved using the block discretemethod and the servos transition issues when the control law switching is solvedusing single-mode transient suppression method.For the situation of extensive pole-zero distribution of the taxiing model and theactual parameters are hard to obtained accurately, longitudinal control of three wheelstaxiing is designed to ensure that the front and rear wheels pressure ratio is constantthroughout the deflection of elevator. The model reference adaptive control law, givenparameter control law and gain scheduling control law are designed respectively forthe yaw rate, yaw angle and yaw distance loop of lateral control. The PD gainadaptive control law and given parameter control law are designed respectively for theyaw angle and yaw distance loop of two wheels taxiing lateral control. The simulationresults show that the control law can eliminate the deviation quickly when there existsinitial yaw angle2.5°and initial yaw distance0.2m, so as to the UAV taxis alongcenterline of the runway.To increase the success rate of flight test, the control law is cleared and verifiedbased on unstable eigenvalues criterion using polynomial approach when there existsthe inertia, pneumatic and other uncertainties during flight. The clearance results showthat the available angle of attack throughout the whole flight envelope in nominalstate is [-2°,18°], the available angle of attack range with single uncertain parameterperturbation is [0°,15°]，the available angle of attack range with multiple uncertainparameters simultaneously perturbation is reduced to [0°,11°].Finally, the research works mentioned above are verified through hardware inloop simulation for taxiing test and flight test. The simulation results agree well withthe digital simulation, the actual pitch angle control steady-state error is0.2°in outfield flight test, the actual roll angle control steady-state error is0.4°, the actualheight control steady-state error is1m. All the above meet the military standardrequirements. The outfield autonomous taxiing test is completed to verify the adaptivecontrol law. The results show that the maximum yaw angle is2°, the maximum yawdistance is1.5m, the crosswind disturbance is well inhibited.