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In this paper, the performance of the spacecraft attitude control system is enhanced using model-based disturbance feedback control (DFC) strategy in the presence of disturbance. This control strategy is applied to a single-axis spacecraft attitude control with thruster, reaction wheel, and magnetic torqrod actuators, separately. An anti-windup observer-based modified PI-D is utilized for each actuation system as a main controller. The controller gains are tuned using genetic algorithm when the time average of absolute value of pointing error is chosen as an objective function. The performance of DFC with the modified PI-D controller is investigated under disturbance and model uncertainties. The numerical simulation shows that the DFC strategy can reject disturbance effect and improve the pointing error for the three mentioned actuators; however, for a very large value of external disturbance, a critical value for uncertainty is observed for the thruster lag at which the pointing error is suddenly increased. For this critical value, the control system cannot tolerate any longer lag uncertainty in comparison with the two other actuation systems. Increasing the value of disturbance decreases the tolerable value of uncertainty in the thruster lag.

The main objective of this paper is to enhance the robustness of an on-off attitude control under uncertainties while limiting the probability of failure in attitude control. To do this, the concept of system optimization is utilized for detailed engineering of spacecraft control using reliability-based robust design optimization (RBRDO). The probability of failure of the attitude control is chosen by the system designer as the input of the RBRDO algorithm. The single-axis spacecraft attitude is controlled using a combination of the observer-based anti-windup modified PI-D with pulse-width pulse-frequency modulator in the presence of external disturbance. The on-off thruster is modeled with a delay followed by a second-order transfer function. The input frequency of the thruster is limited to 50 Hz. The uncertain parameters are given as the spacecraft moment of inertia, thrust level, and thruster delay. The controller gains are determined by using traditional, robust, and reliability-based robust design optimizations under uncertainties and disturbance. The simulations are carried out using quasi-normalized equations, along with reducing problem variables and computational burden, to obtain more applicable results for a preliminary design. The traditional optimization gives the highest pointing accuracy without uncertainty, whereas the robust optimization obtains an approximately flat behavior for the mean of absolute pointing error under uncertainties. Under this situation, RBRDO could satisfy the prescribed reliability with a small loss in accuracy for the on-off attitude control of spacecraft, but under system limitations.

PurposeThe purpose of this paper is to develop a novel solution for the predicted error and introduces a systematic method to develop optimal and explicit guidance strategies for different missions.Design/methodology/approachThe predicted error is derived from its basic definition through analytical dynamics. The relations are developed for two classes of systems. First, for systems in which the acceleration commands are truncated at a specified time. Second, for systems in which the corrective maneuvers are cut off at a specified time. The predicted error differential equation is obtained in a way that allows for derivation of several optimal and explicit guidance schemes.FindingsThe effect of tangential acceleration in conjunction with autopilot dynamics can be realized in guidance gain and the predicted error. The differential equation of velocity‐to‐be‐gained is obtained assuming the gravitational acceleration to be given as a vectorial function of time. The relations for different velocity profiles are obtained and discussed including the effective navigation ratio.Research limitations/implicationsThe guidance/control system is modeled as a linear time‐varying dynamic and of arbitrary‐order. The gravitational acceleration is assumed as a given vectorial function of time.Practical implicationsThe presented schemes are applicable to both midcourse and terminal guidance laws with/without velocity constraints.Originality/valueProviding a new analytical solution of predicted errors with final position and velocity constraints and their differential equations considering the thrust/drag acceleration and autopilot dynamics in the presence of gravity.

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