Attitude Control in Ascent Phase of Missile Considering Actuator Non-Linearity and Wind Disturbance

Bangsheng Fu,Quanxi Gao,Hui Qi,Shaobo Wang,Jiangtao Xu,Ya Yang

doi:10.3390/app9235113

Abstract

During the ascent phase of a missile, a challenging problem occurs that blocks the construction of a high-precision attitude control scheme, which directly affects accurate modeling including disturbances: non-linearities of an actuator, rapidly time-varying parameters, un-modeled dynamics, etc. In order to improve the control performance, an active disturbance rejection control (ADRC) scheme, considering non-linear dynamics of the actuator and wind disturbance during the ascent phase, is proposed in this paper. An expand state observer (ESO) is planned to estimate and compensate the actuator’s non-linear dynamics, flight model uncertainties, and wind disturbance. Therefore, the complex non-linear time-varying control problem is simplified into a linear time-invariant control problem. The pitch attitude control system is controlled by the cascade method and ADRC controllers are designed for actuator close loop and attitude control loop, respectively. The simulation results show that ADRC has strong robustness under different dead-zones and external disturbances of the actuator. On the other hand, ADRC can effectively suppress the external atmospheric disturbance. Compared with the traditional gain-scheduling control scheme, the ADRC scheme can significantly reduce the overloading of the system and shows remarkable performance for tracking as well as wind resistance.

Highlights

Designing an attitude controller in the ascent phase of a missile is of paramount importance in order to reduce the terminal deviation and enhance the control ability
In the ascent phase of the missile, designing a controller with the traditional modeling concept needs to carry out more modeling of feature points within the range of the ascent phase to achieve better control effect, and carry out dynamic interpolation through the gain-scheduling method to achieve global attitude control [3]
The integral time absolute error (ITAE) parameters tuning method is adopted for proportion integral differential (PID) [25], the control parameters used by the PID method are kp0 = 128, Ti = 1200 and Td = 0.2

Summary

Introduction

Designing an attitude controller in the ascent phase of a missile is of paramount importance in order to reduce the terminal deviation and enhance the control ability. Especially in high-precision attitude control, because of the existence of non-linear factors such as the dead-zone, backlash, saturation, etc., can reduce the dynamic performance of the actuator, which will lead to serious degradation of attitude control effect, it even causes oscillations and risks global stability [9,10]. Due to the limitations of manufacturing technology, the installation error and the debugging error, the dead-zone, the gear backlash and the zero position error, and other non-linearities inevitably exist in the actuator system To solve this problem, Forbes [11] studies the non-linear saturation suppression by adopting a feedback structure, proportional control and dynamic angular velocity control law for attitude control. Non-linearities such as dead-zone, backlash, time-delay, friction, and so on, which can be

Electricfound

Model Simplification and ActivetoDisturbance

Dead-Zone Non-Linearity Influence Analysis

Dead-Zone Non-Linearity Effects

11. Estimation

Command Angle Affects

13. Responses

Modeling and Analysis of Attitude Control of Missile

16. Coordinates

Design of ADRC Attitude Controller

] design the ESO according

Simulation and Analysis