Actuator Fault Tolerant Terminal Sliding Mode Guidance Law with Impact Angle and Acceleration Constraints
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In the guidance literature, various requirements on missile guidance performance have arisen and a significant number of research has been presented on these requirements. Ensuring a successful hit on the target is one of the principal properties that is studied on missile guidance problems. To guarantee high hit probability, the guidance law must be robust against disturbances. Undesired forces created on the missile in case of actuator failure and unknown target acceleration acts as a disturbance on engagement geometry. Thus, a robust guidance law can tolerate the actuator failure and unknown target acceleration effects. Another requirement of guidance problem is achieving the desired impact angle which is mainly important for deactivating the heavily armored ground targets effectively. In this thesis, an actuator fault tolerant terminal sliding mode guidance law is proposed by considering impact angle and acceleration constraints. The sliding mode control method is known to be robust against unknown disturbances and provides an adequate solution for controlling non-linear systems. In this study, the sliding mode control method is adopted for guidance design, since actuator failure and unknown target acceleration behave as a disturbance on the non-linear missile-target engagement kinematics. A first order sliding mode guidance law is designed with equivalent control method. The selected sliding surface ensures achieving a successful hit on the target with the desired impact angle. Bounded target acceleration and actuator failure effects are considered in switching function architecture. Additionally, a guidance law needs necessary data to generate proper commands for the missile. Line-of-sight (LOS) angular rate, one of the commonly used parameters in guidance laws, may not be directly measured on missiles with strapdown seekers. In this study, LOS angular rate is estimated from LOS angle with a second order sliding mode differentiator, if this rate information is not accessible by the missile. The proposed guidance law is used in the terminal flight phase. Slowly moving heavily armored vehicles are considered as a target. The performance of the proposed guidance law is analyzed with a numerical simulation model. The results of the simulation studies prove that the guidance law is robust against actuator failures and unknown target acceleration. Estimation performance of LOS angular rate is interpreted as suitable for the guidance process.
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