Application of Integrated Guidance and Control Schemes to a Precision Guided Missile

Evers, Johnny; Cloutier, J.R.; Lin, Ching-Fang; Yueh, William R.; Wang, Q.

doi:10.23919/acc.1992.4792745

Cited by 20 publications

(8 citation statements)

References 5 publications

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“…A variety of single-loop IGC algorithms have been proposed in the recent literature. Techniques like linear quadratic optimal control, [4][5][6][7] state-dependent Riccati equation approach, 8,9 suboptimal u-D method, 10 generalized extended state observer (GESO)-enhanced predictive control, 11 feedback linearization, 12,13 and zeroeffort-miss (ZEM)-based sliding mode control [14][15][16] have been applied successfully in the single-loop IGC framework. Among these methods, optimal control is the most intuitive design that can synthesize an overall optimal IGC system subject to certain performance index.…”

Section: Introductionmentioning

confidence: 99%

Partial integrated guidance and control design for supersonic missile based on disturbance rejection

Piao

Sun

Huang

et al. 2019

Measurement and Control

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In this paper, a three-dimensional partial integrated guidance and control law for the supersonic missile impacting the surface target is proposed. In the guidance loop design, a reduced-order extended state observer is integrated with feedback linearization and the pure proportional navigation law to eliminate the effects of the unknown target kinematics and the dynamic coupling between the vertical and horizontal planes. Considering the autopilot lag, two guidance schemes with different observer inputs are compared using the adjoint and frequency domain analyses. For the control loop, the angular velocity feedback is performed first for the pitch and yaw channels as conventionally done in practice, and then first-order linear active disturbance rejection control is designed for the damping-enhanced plant to deal with the fast time-varying, uncertain aerodynamics and the strong coupling effects among the three channels. Both the attitude angle and the acceleration control can be designed in a unified framework. Finally, a three-dimensional engagement scenario is performed using a six-degree-of-freedom supersonic vehicle model, and extensive simulations are carried out to validate the robustness and the high guidance precision of the proposed method.

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Section: Introductionmentioning

confidence: 99%

Partial integrated guidance and control design for supersonic missile based on disturbance rejection

Piao

Sun

Huang

et al. 2019

Measurement and Control

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“…Because the accuracy of sensors in guidance system is much greater than that of autopilot, the IGC can reduce cost and improve the entire guidance and control system performance. In [2], the discussion of IGC scheme goes a step further with three optimal control laws for a tactical missile. In [3,4], new optimal control methods are proposed to effectively design IGC system for missiles.…”

Section: Introductionmentioning

confidence: 99%

Neural Extended State Observer Based Command Filtered Back-stepping Hypersonic Vehicle Integrated Guidance and Autopilot

Wang¹,

Peng²,

Zhang³

et al. 2018

Preprint

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This paper focuses on the integrated guidance and autopilot design with control input saturation in the end-game phase of hypersonic flight. Firstly, uncertain nonlinear integrated guidance and autopilot model is developed with third actuator dynamics, where the control surface deflection has magnitude constraint. Secondly, neural network is implemented in extended state observer (ESO) design, which is used to estimate the complex model uncertainty, nonlinearity and state coupling. Thirdly, a command filtered back-stepping controller is designed with hybrid sliding surfaces to improve the terminal performance. In the process, different command filters are implemented to avoid the influences of disturbances and repetitive derivation, meanwhile solve the problem of unknown control direction caused by saturation. The stability of closed-loop system is proved by Lyapunov theory, and the principles abided by the controller parameters are concluded through the proof. Finally, series of 6-DOF numerical simulations are presented to show the feasibility and validity of the proposed controller.

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“…Linear vehicle model and engagement kinematics were employed. In Lin et al [4] and Evers et al [5], the earlier work [3] was extended to include a target acceleration response model. Integrated steering/control design was also reported in [6], where flight in a 2-D plane with simplified linear dynamics was considered.…”

Section: Introductionmentioning

confidence: 99%

Nonlinear integrated steering and control of flight vehicles

Poh

1998

Guidance, Navigation, and Control Conference and Exhibit

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In this paper, we present an approach to integrated steering and control design for flight vehicles. In this approach, input-output linearization is applied to derive the integrated controller, which performs the function of both autopilot controller and steering law. The nonlinear six degree-of-freedom dynamics together with aerodynamics model of vehicle is considered. Simulation shows that the design has excellent performance and is robust against unmodelled actuator dynamics and uncertainties. Nomenclature b = wing span c = wing chord 8 a = aileron deflection 8 e = elevator deflection 8 r = rudder deflection m = vehicle mass p = body-axis roll rate q = body-axis pitch rate r = body-axis yaw rate; norm of r r = \x y z J vehicle position vector in inertial coordinate C ; = aerodyn. moment coeff. about body x axis C m = aerodyn. moment coeff. about body y axis C n = aerodyn. moment coeff. about body z axis C x = aerodynamic force coeff. in body x axis C y = aerodynamic force coeff. in body y axis C z = aerodynamic force coeff. in body z axis F x = aerodynamic force in body x axis F y = aerodynamic force in body y axisF z = aerodynamic force in body z axis I = vehicle's moment of inertia matrix / =Jacobian matrix of aerodyn force coefficients L = aerodynamic roll moment M = aerodynamic pitch moment N = aerodynamic yaw moment Q = dynamic pressure S = reference area T = engine thrust Ty, = Transformation matrix from body frame to inertial frame V = magnitude of velocity CC = angle of attack /? = sideslip angle \j/ = yaw angle 6 = pitch angle (j) = roll angle C = LOS vector <7 z ,C7 y ,ff z = components of LOS vector in inertial coordinate p =air density Q = skew-symmetric matrix of body rate vector A = Transpose of matrix/vector A

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Application of Integrated Guidance and Control Schemes to a Precision Guided Missile

Cited by 20 publications

References 5 publications

Partial integrated guidance and control design for supersonic missile based on disturbance rejection

Partial integrated guidance and control design for supersonic missile based on disturbance rejection

Neural Extended State Observer Based Command Filtered Back-stepping Hypersonic Vehicle Integrated Guidance and Autopilot

Nonlinear integrated steering and control of flight vehicles

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