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Trajectory tracking and estimation of hypersonic glide vehicles (HGVs) is a very challenging issue in the defense systems. Insufficient knowledge about the HGV and inaccurate dynamic models for the accelerating HGV are the main challenges in this regard. In the present study, an integrated nonlinear Markov acceleration model is established to formulate the nonlinear dynamic characteristics of HGVs. Since the aerodynamic accelerations of the HGV are dominant and the corresponding aerodynamic coefficients are unknown, a statistics-based aerodynamic model is proposed. The proposed aerodynamic model is capable of providing primary information of the aerodynamic characteristics even without knowing the configuration of the HGV. Then, considering the maneuver mode of the vehicle, the iterative extended Kalman filter (IEKF) is applied to track the trajectory of the HGV by using the proposed model. Obtained results from the numerical simulation for the equilibrium glide mode and skip maneuver mode indicate that the proposed model can effectively improve the velocity estimation accuracy by about 40%-50% and acceleration estimation accuracy by about 20%-50% in the given examples.

This paper has proposed a new robust hybrid nonlinear guidance law, which accounts for a missile’s terminal line-of-sight (LOS) angle constraint, in order to intercept a non-cooperative maneuvering target. The proposed hybrid nonlinear guidance strategy consists of two phases; in the first phase, a guidance law named PIGL is derived from prescribed performance control and the inertial delay control method. In PIGL, a revised prescribed performance function is put forward, and a prescribed performance controller with unknown uncertainties is then derived. The controller smoothly drives both the LOS angle and its rate to a predesigned small region under unknown uncertainties that are induced by target’s maneuvers within a fixed time. Then, a guidance law named SIGL is activated, which is derived from sliding mode control and inertial delay control. By driving the desired sliding mode variable to zero within a finite time, the SIGL guidance law is able to achieve high terminal interception accuracy. The robustness of both of the proposed sub-guidance laws has been proved explicitly in this paper. The hybrid guidance law has the advantage of a tunable convergence rate of the LOS angle and the rate of the LOS angle at the beginning period, by which an excessive large initial maneuver can be avoided. Meanwhile, the hybrid guidance law also has the advantage of lower sensitivity to errors in the estimation of the time-to-go.

This paper derives a new terminal homing guidance law for missiles to intercept non-cooperative maneuvering targets subject to the constraint of terminal line-of-sight angle. The new guidance law avoids abrupt and large changes in guidance commands, while allows the control gains to be determined analytically or chosed loosely to ease the design of guidance law. This is achieved by combining the advantages of sliding mode, prescribed performance, and inertial delay controllers. The prescribed performance controller is designed by a new two-phased prescribed performance function with an improved continuous dynamic property, which drives the predefined state error-based sliding mode variable to a small residual set around zero. The inertial delay controller is adopted to suppress the system uncertainty. Thus, the new guidance law yields continuous guidance commands over the entire terminal homing phase while eliminating abrupt and large changes in the initial phase, resulting from the initial state errors in the prescribed performance function. The newly proposed guidance law improves the prescribed performance control that dominates the gains of the proposed guidance law. Theoretical analyses and numerical simulations are conducted to verify the robustness and preciseness of the proposed guidance law.

A prescribed performance guidance law during the terminal homing phase is proposed to intercept a non-cooperative maneuvering target in this paper. First, on the basis of sliding mode control, a prescribed performance control method based on a revised prescribed performance function is proposed to drive the sliding mode variable to zero. It leads to smooth guidance commands to fulfill the terminal line of sight angle constraint. Then, taking the autopilot's dynamics and its uncertainties into account, a robust controller is designed using dynamic surface control, sliding mode control and inertial delay control, which guarantees the reference guidance commands to be precisely tracked. The highlights of the proposed guidance law are: 1) the transient performance of guidance commands is improved, so that excessive large changes are avoided in the initial phase, 2) parameters that dominant the convergence of the predetermined sliding mode variable can be analytically determined based on the predesigned terminal tolerance, it is more convenient to achieve the prescribed convergence, 3) although uncertainties are existing in the dynamics of missile autopilot, the proposed guidance law could guarantee the terminal interception accuracy. All of the improvements are shown in numerical simulation results.

This paper proposes a new terminal homing guidance law with line-of-sight angle constraint for a missile to intercept a noncooperative maneuvering target by integrating prescribed performance control with sliding mode control. The newly proposed guidance law is derived analytically with a prescribed performance controller in combination with the continuous nonsingular terminal sliding mode manifold and the disturbance observer. The prescribed performance controller is designed with a new prescribed performance function to avoid undesired abrupt change in the derivative of prescribed performance variable. In addition to its robustness and preciseness, this new homing guidance law is able to attenuate a largely abrupt change in guidance commands due to driving the sliding mode variable whose initial derivative is close to zero to a desired residual set at the beginning of the terminal homing phase. Furthermore, the time-varying control gain of the new guidance law is analytically determined by terminal tolerance. Theoretical analysis and numerical simulations are conducted to demonstrate the effectiveness of this new guidance law.

This paper proposed a new fast nonsingular terminal sliding mode to design terminal angle constraint guidance for intercepting maneuvering targets with attenuating the chattering of the command in the guidance law. A fast nonsingular terminal sliding mode (FNTSM) is established for dealing with the relationship between Line of sight (LOS) angle and the rate of LOS angle, the disturbance observer (DOB) is set up to calculate the compensation amount for counteracting the time-varying uncertainty produced by the design of guidance and targets’ maneuver. Combining with FNTSM and DOB, an adaptive law and a compensated parameter are adopted to ease the coupling relationship between the convergence rate of sliding variable and initial parameters design. Furthermore, the chattering amplitude can be reduced, even can be eliminated in some cases. Theoretical analysis and numerical simulations verify the effectiveness of the proposed method.