Missile Autopilot Research Articles

Hybrid Switched Gain-Scheduling Control for Missile Autopilot Design

This paper presents a new hybrid switched gain-scheduling control method for missile autopilot design via dynamic output feedback. For controller design purpose, the nonlinear missile plant model is first converted to a switched linear fractional transformation system. Then, the new hybrid switched gain-scheduling autopilot is designed, which consists of a switching dynamic output–feedback linear fractional transformation controller and a supervisor enforcing a controller state reset at each switching time instant. The proposed hybrid control scheme is shown to provide a systematic yet simple framework for missile autopilot design. Specifically, the control synthesis conditions that guarantee weighted stability performance are formulated in terms of a finite number of linear matrix inequalities, which can be solved effectively via convex optimization without parameter-space gridding. Furthermore, stringent controlled performance and strong robustness against parameter perturbations are achieved using this new control approach, whereas no parameter variation information is required for both controller synthesis and implementation. The advantages of the proposed design approach over existing methods will be shown through nonlinear simulations for the missile autopilot design over a wide range of operating conditions.

Komputerowe wspomaganie analizy i syntezy układów stabilizacji rakiet przeciwlotniczych

Komputerowe wspomaganie analizy i syntezy układów stabilizacji rakiet przeciwlotniczych

Robust output feedback tracking control for a class of perturbed systems using NDOB and nonlinear continuous predictive control

This paper presents a new robust output feedback tracking control scheme for a class of higher-order uncertain systems. Since traditional nonlinear continuous predictive control requires accurate system model as well as full system states to synthesize a controller, a composite control methodology is adopted in the proposed scheme. Specifically, the nonlinear disturbance observer (NDOB) is used to estimate the lumped uncertainty and the unmeasured system states in an integrated manner while the nonlinear continuous predictive control regulates the system states to track the desired reference signal asymptotically. Detailed stability analysis is also presented for the closed-loop nonlinear observer-controller structure through two steps. Then, the obtained results are applied to missile autopilot design to track the desired angle-of-attack signal. Finally, numerical simulations with some comparisons are provided to demonstrate the effectiveness of the proposed formulation.

Design and Comparison of Two-Loop with PI and Three-Loop Autopilot for Static Unstable Missile

Design and Comparison of Two-Loop with PI and Three-Loop Autopilot for Static Unstable Missile

Optimal acceleration autopilot design for non-minimum phase missiles using evolutionary algorithms

The flight control system is a key element to achieve required performance in missiles and aircrafts. First purpose of flight control system is to ensuring the stability of the system, then, it attempts to force it to track the guidance commands. This paper provides a straightforward method using evolutionary optimisation algorithms to design an appropriate autopilot for non-minimum phase missiles. In order to bring the results to the actual conditions, the missile non-minimum phase model and actuator dynamics with time delay is considered. Proper indices such as system speed, overshoot, undershoot, steady state error and control signal effort have been incorporated to propose an innovative cost function. Then, several applicable meta-heuristic techniques are employed to optimise this cost function. Genetic algorithm, particle swarm optimisation, artificial bee colony, imperialist competitive algorithm and cuckoo search techniques have been compared in this optimisation problem. Simulation results on two benchmark problem show that this method has acceptable speed and it can be used in gain scheduling control design method for non-minimum phase systems. This method can be a suitable replacement for the time consuming procedure of gain tuning in gain scheduling method. The superior advantage of this method compared to the other methods is automatic tuning of the autopilot gains.

Adaptive Backstepping Autopilot Design for Missiles of Fast Time-varying Velocity

An adaptive backstepping control scheme is proposed to design a missile autopilot during boost-phase. In the boost-phase, accurately controlled missiles experience fast velocity variations that make the missile dynamics be highly nonlinear. Missile parameters including the mass, the moment of inertia, and a center of gravity, which are usually considered as constants or slowly-varying, change very fast. The time-varying effects combined with nonlinear dynamics yield severe parametric uncertainties in missile dynamics. Therefore, the nonlinear dynamics as well as parametric uncertainty should be carefully treated in autopilot design. To deal with this problem, an adaptive nonlinear controller is designed based on backstepping procedure. Adaptation laws are properly designed so that robustness to the uncertainty as well as the closed-loop stability can be guaranteed using Lyapunov stability theorem. To demonstrate the performance of the proposed controller, numerical simulation is carried out.

Hidden Coupling Terms Inclusion in Gain-Scheduling Control Design: Extension of an Eigenstructure Assignment-Based Technique

This paper tackles the gain-scheduling control design issue when endogenous signals are selected as scheduling parameters. In this case, the gain-scheduled controller exhibits nonlinear dynamics, leading to the occurrence of hidden coupling terms in the linearized dynamics. Omitting hidden coupling terms in the synthesis process may induce performance degradation or even the destabilization of the closed-loop system. The present paper aims at extending a previous work on the design of gain-scheduled controller for hidden coupling terms inclusion via eigenstructure assignment technique. The efficiency of the proposed procedure is illustrated on a pitch-axis missile autopilot benchmark problem and is validated through nonlinear simulations.

Adaptive time varying autopilot design

Missile autopilots represent devices for precise following of a target that missiles are sent to intercept. Commands that need to be executed are forwarded to the autopilot in order to achieve demanded performance. Usually, in representation of an autopilot and its performance, a model with constant parameters is used. Constant autopilot parameters and missile coefficients are calculated from time varying values. Model with constant parameters is shown in this paper, as well as the results obtained from this model. These values are then used as the initial ones for calculations in time varying pitch autopilot model design based on gain-scheduling. It is shown how the coupling with other axis such as yaw is cancelled by using time varying parameters. All the models are implemented in Matlab and Simulink toolbox.

Fuzzy control strategy for hypersonic missile autopilot with blended aero-fin and lateral thrust

An aerodynamic force and lateral thrust blended control autopilot is designed for the near-space air-breathing hypersonic missile with low overload cruise phase due to angle of attack restriction. The nonlinear longitudinal short-period motion model of a hypersonic missile is built, and several fuzzy control strategies are applied in overload command allocation, so as to decouple the aerodynamic and direct force subsystems. The adaptive sliding mode control approach is proposed for aerodynamic subsystem to handle the model uncertainties and atmospheric environment disturbances. A two-dimensional fuzzy controller is designed for direct force subsystem to reduce the dependence of the accurate model. The simulations show that the proposed blended controller can overcome the effects of disturbances and improve the delay effect by reason of weak aerodynamic to realize satisfied tracking performance. The simulations are also given to illustrate the validity of the blended control law in long period. The results show that the blended controller can achieve a good tracking performance throughout the cruise phase. The maneuverability of this dual-controlled hypersonic missile, with blended aero-fin and divert-control lateral jets, is enhanced obviously in cruise phase.

Controller Design for a Two-loop Missile Autopilot in Pitch Plane

Controller Design for a Two-loop Missile Autopilot in Pitch Plane

Optimization of three-loop missile autopilot gain under crossover frequency constraint

Abstract The open-loop crossover frequency is pointed as an important parameter for practical autopilot design. Since different gain designs may achieve the same open-loop crossover frequency, it should be neither considered as a performance objective of the optimal autopilot design-schemes nor neglected. Besides, the main assignment of the autopilot is to drive the missile to track the acceleration commands, so the autopilot gain design should be evaluated directly according to the resultant tracking performance. For this purpose, an optimal design methodology of the three-loop missile autopilot is introduced based on constraint optimization technique, where the tracking performance is established analytically as the design objective and the open-loop crossover frequency is formed as inequality constraint function, both are manipulated in terms of stable characteristic parameters of the autopilot closed-loop. The proposed technique is implemented with the assistance of a numerical optimization algorithm which automatically adjusts the design parameters. Finally, numerical simulation results are provided to demonstrate the effectiveness and feasibility of the proposed approach compared with that in some references.

Design of three-dimensional nonlinear guidance law with bounded acceleration command

This paper concentrates on the problem of designing a three-dimensional nonlinear guidance law accounting for saturation nonlinearity and the dynamics of missile autopilot based on a command filtered backstepping (CFBS) scheme. In the design, the nonlinear kinematics of target–missile engagement is considered in the spherical coordinate system and the dynamics of the autopilot is considered as a second-order term. The CFBS scheme is utilized to blend the command filtering approach into standard backstepping method to avoid the complex computation of the analytic derivatives of the intermediate control signals. A command filter with a saturation limiter embedded and some auxiliary filters are combined together to address acceleration command saturation. Simulation results show that even though subject to saturation constraint, the proposed guidance law achieves excellent guidance performance in terms of missed distance.

Two-channel pitch/yaw missile autopilot design using arbitrary order sliding modes based pole placement

AbstractThe paper presents a new missile autopilot system design. The design is achieved through the pole-placement in quasi-continuous high-order sliding mode gains adjustment. Enhanced performance, strong robustness and smooth control are obtained through arbitrary increase of the number of non-oscillatory stable poles. The target application of this technique the two-channel pitch/yaw missile autopilot system is considered. Numerical simulations indicate that the arbitrary-order sliding modes based pole placement’s performance compares favourably against recently proposed high-order pole placement schemes.The proposed arbitrary-order pole placement scheme presents a promising design tool for finite-time stabilisation and control of uncertain multivariable systems.

Finite Time Convergence Guidance Law Accounting for Missile Autopilot

A guidance law with finite time convergence is designed using the sliding mode control method and finite time convergence control theory, accounting for the missile autopilot as second-order dynamics. The high-order derivatives of the line of sight (LOS) angle are avoided in the expression of guidance law such that it can be implemented in practical applications. The designed guidance law is effective in compensating the bad influence of the autopilot dynamics on guidance accuracy. In simulations of intercepting a non maneuvering target or a maneuvering target, respectively, the designed guidance law is compared with the adaptive sliding mode guidance (ASMG) law in the presence of missile autopilot lag. Simulation results show that the designed guidance law is able to guide a missile to accurately intercept a nonmaneuvering target or a maneuvering target with finite time, even if it escapes in a great and fast maneuver and the autopilot has a relatively large lag.

Robust terminal angle constraint guidance law with autopilot lag for intercepting maneuvering targets

In this paper, a robust continuous guidance law with terminal angle constraint for intercepting maneuvering targets in the presence of autopilot lag is proposed and its finite-time stability is proved. First, assuming that the missile autopilot is sufficiently fast, a composite fast nonsingular terminal sliding mode guidance (CFNTSMG) law is presented. The proposed guidance law is obtained through a combination of fast nonsingular terminal sliding mode control theory and generalized disturbance observer (GDOB). The presented guidance law requires no priori information on target maneuver, which is estimated and compensated online by GDOB. Moreover, no discontinuous term exists in CFNTSMG law and therefore chattering is eliminated effectively. Next, viewing the missile autopilot as an uncertain second-order system, an integrated guidance and control dynamics is formulated and the systematic step-by-step backstepping technique is used to derive a new robust guidance law, which not only holds the advantages of CFNTSMG law, but also is insensitive to autopilot lag. At each backstepping step, novel continuous virtual/real control laws using finite-time control approach are designed and tracking differentiator is used to overcome the ‘explosion of complexity’ problem encountered with traditional backstepping method. Theoretical analysis and numerical simulations demonstrate the effectiveness of the proposed method.

Performance investigation of extended-state-observer-based roll autopilot design

In this work, performance analysis of the recently appeared extended-state-observer (ESO)-based roll autopilot design is carried out. Firstly, considering a more realistic roll dynamics, performance comparison of the ESO-based design with several classical roll autopilot structures is carried out and the related results are presented. Next, the better performing designs are implemented and validated through a high fidelity nonlinear 6-DOF environment for a typical tactical missile in a realistic engagement scenario. Lastly, performance of the ESO-based design is investigated using Monte Carlo simulations under several subsystem parameter uncertainties. The results show that the ESO-based design offers satisfactory performance in controlling roll angle and its rate in the presence of significant disturbances, parameter uncertainties and cross-coupling effects and thus offers a viable approach towards design of roll autopilot for high performance tactical missiles.

Agile Missile Autopilot Design for High Angle of Attack Maneuvering with Aerodynamic Uncertainty

Agile Missile Autopilot Design for High Angle of Attack Maneuvering with Aerodynamic Uncertainty

Robust Predictive Control of a Class of Nonlinear Systems

In this work, a continuous-time predictive controller designed for a class of nonlinear systems is robustified using an extended state observer. The predictive controller formulation requires the availability of an accurate mathematical model of the plant and the complete state vector. In this work, the extended state observer is employed to address these issues. The extended state observer simultaneously estimates the state and the effect of uncertainty and disturbances in an integrated manner. The estimated states and disturbances are used to robustify the predictive controller designed for nominal system. Closed-loop stability for the overall system is established. As an illustrative example, missile autopilot design using the proposed approach is presented along with simulation results. Finally, comparison results of the proposed approach with existing techniques appeared in the literature toward the robustification of the predictive controller are presented. The study clearly brings out the efficacy and effectiveness of the extended state observer-based predictive controller in offering improved robustness and performance.

Gain‐Scheduled H∞ Control for Tensor Product Type Polytopic Plants

AbstractA tensor product (TP) model transformation is a recently proposed technique for transforming a given linear parameter‐varying (LPV) model into polytopic model form, for which there are many methods that can be used for controller design. This paper proposes an alternative approach to the design of a gain‐scheduled output feedback H∞ controller with guaranteed L2‐gain parameter‐dependent performance for a class of TP type polytopic models using parameter‐dependent Lyapunov functions where the linear matrix inequalities (LMIs) need only to be evaluated at all vertices of the system state‐space model matrices and the variation rate of the scheduled parameters. In addition, a construction technique of the intermediate controller variables is also proposed as a matrix‐valued function in the polytopic coordinates of the scheduled parameters. The performance of the proposed approach is tested on a missile autopilot design problem. Furthermore, nonlinear simulation results show the effectiveness of these proposed techniques.

Nonlinear autopilot design for interceptors with tail fins and pulse thrusters via θ-D approach

The missile autopilot for an interceptor with tail fins and pulse thrusters is designed via the θ-D approach. The nonlinear dynamic model of the pitch and yaw motion of the missile is transformed into a linear-like structure with state-dependent coefficient (SDC) matrices. Based on the linear-like structure, a θ-D feedback controller is designed to steer the missile to track reference acceleration commands. A sufficient condition that ensures the asymptotic stability of the tracking system is given based on Lyapunov's theorem. Numerical results show that the proposed autopilot achieves good tracking performance and the closed-loop tracking system is asymptotically stable.