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The purpose of this paper is to design an integrated guidance and control design for a formation flight of four unmanned aerial vehicles to follow a moving ground target.

The guidance law is based on the line‐of‐sight. The control is optimal. The guidance law is integrated with the optimal control law and is applied to a linear dynamic model.

The theoretical results are supported by the numerical simulations that illustrate a coordinated encirclement of a ground maneuvering target.

A linear dynamic UAV model and a liner engine model were employed.

This is expected to provide efficient coordination technique required in many civilian circular formation UAV applications; also the technique can be used to provide a safe environment required for the civil applications.

The research will facilitate the deployment of autonomous unmanned aircraft systems in various civilian applications such as border monitoring.

The research addresses the challenges of coordination of multiple unmanned aerial vehicles in a circular formation using an integrated optimal control technique with line‐of‐sight guidance.

The paper aims to provide further study on the development and analysis of flight control system for two‐dimensional (2D) differential geometric (DG) guidance and control system based on the application of a set‐point weighting proportional‐integral‐derivative (PID) controller.

The commanded angle‐of‐attack is developed in the time domain using the classical differential geometry theory. Then, a set‐point weighting PID controller is introduced to develop a flight control system so as to form the 2D DG guidance and control system, and the gains of the PID controller are determined by the Ziegler‐Nichols method as well as the Routh‐Hurwitz stability criterion. Finally, the classical frequency method is utilized to study the relative stability and robustness of the designed flight control system.

The results demonstrate that the designed controller yields a fast responding and stable system which is robust to the high frequency parameters variation. Moreover, the DG guidance law is viable and effective in a realistic missile defense engagement.

This paper provides a novel approach on the development of DG guidance and control system associated with its stability analysis.

In this paper, the development of relative guidance and control algorithms for proximity operations to satellite in elliptical orbit are presented. The paper aims to discuss these issues.

The process of autonomous proximity is divided into three phases: proximity manoeuvre, flyaround manoeuvre, and hovering manoeuvre. The characteristics of the three phases are analyzed. Different guidance algorithms are based on using the analytical closed-form solution of the Tschauner-Hempel (TH) equations that is completely explicit in time. Lastly, the linear quadratic regulators control algorithm based on the linearized TH equations is developed to minimize the initial state errors in the last phase.

This paper defines three phases in the satellite proximity operations and develops the guidance and control algorithms. Then, the relative guidance and control algorithms are illustrated through different numerical examples. And the results demonstrate the effectiveness and simplicity of using a TH model in autonomous proximity.

The findings indicate that a TH model is clearly effective at estimating the relative position and velocity and controlling the relative trajectory. In addition, this model is not restricted to a circular orbit, but it can be used as well for an elliptical orbit. Furthermore, by using this model, simple guidance and control algorithms are developed to approach, flyaround and hover from a target satellite.

Based on the guidance algorithms, the manoeuvre-flight period can be set in accordance with the mission requirement. Flyaround with different types of trajectory and a feedback control scheme to achieve stable hovering state are studied. Consequently, this proposed guidance algorithms can effectively implement guidance and control for satellite proximity operations.

The purpose of this paper is to study the closed-loop guidance algorithm for launch vehicles in an atmospheric ascent phase and present a numerical trajectory reconstruction algorithm to satisfy the real-time requirement of generating the guidance commands.

An optimal control model for an atmospheric ascent guidance system is established directly; following that, the detailed process for necessary conditions of the optimal control problem is re-derived based on the calculus of variations. As a result, the trajectory optimization problem can be reduced to a root-finding problem of algebraic equations based on the finite element method (FEM). To obtain an accurate solution, the Newton method is introduced to solve the roots in a guidance update cycle.

The presented approach can accurately and efficiently solve the trajectory optimization problems. A moderate number of unknowns can yield a good optimal solution, which is well suited for the open-loop guidance. To meet the requirements of the rapidity and accuracy for the close-loop guidance, the fewer number of unknowns is artificially chosen to reduce the calculation time, and the on-board trajectory planning strategy can increase the precision of the optimal solution along with the decrease of time-to-go.

The closed-loop guidance algorithm based on an FEM can be found in this paper, which can solve the optimal ascent guidance problems for launch vehicles in the atmospheric flight phase rapidly, accurately and efficiently.

This paper re-derives the necessary conditions of the optimal solution in a different way compared to the previous work, and the closed-loop guidance algorithm combined with the FEM is also a new thought for the optimal atmospheric ascent guidance problems.

The purpose of this paper is to present a novel guidance law that is able to control the impact time while the seeker’s field of view (FOV) is constrained.

The new guidance law is derived from the framework of Lyapunov stability theory to ensure interception at the desired impact time. A time-varying guidance gain scheme is proposed based on the analysis of the convergence time of impact time error, where finite-time stability theory is used. The circular trajectory assumption is adopted for the derivation of accurate analytical estimation of time-to-go. The seeker’s FOV constraint, along with missile acceleration constraint, is considered during guidance law design, and a switching strategy to satisfy it is designed.

The proposed guidance law can drive missile to intercept stationary target at the desired impact time, as well as satisfies seeker’s FOV and missile acceleration constraints during engagement. Simulation results show that the proposed guidance law could provide robustness against different engagement scenarios and autopilot lag.

The presented guidance law lays a foundation for using cooperative strategies, such as simultaneous attack.

This paper presents further study on the impact time control problem considering the seeker’s FOV constraint, which conforms better to reality.

This paper aims to investigate a novel impact time control guidance (ITCG) law based on the sliding mode control (SMC) for a nonmaneuvering target using the predicted interception point (PIP).

To intercept the target with the minimal miss distance and desired impact time, an estimation of time-to-go is introduced. This estimation results in a precise impact time for multimissiles salvo attack the target at the same time. Even for a large lead angle, the desired impact time is achieved by using the sliding mode and Lyapunov stability theory. The singularity issue of the proposed impact time guidance laws is also analyzed to achieve an arbitrary lead angle with the desired impact time.

Numerical scenarios with desired impact time are presented to illustrate the performance of the proposed ITCG law. Comparison with the state-of-art impact time guidance laws proves that the guidance law in this paper can enable the missile to intercept the target with minimal miss distance and final impact time error. This method enables multiple missiles to attack the target simultaneously with different distances and arbitrary lead angles.

An ITCG law based on sliding mode and Lyapunov stability theory is proposed, and the switching surface is designed based on a novel estimation time-to-go for the missile to intercept the target with minimal miss distance. To intercept the target with initial arbitrary lead angles and desired impact time, the authors analysis the singular issue in SMC to ensure that the missile can intercept the target with arbitrary lead angle. The proposed approach for a nonmaneuvering target using the PIP has simple forms, and therefore, they have the superiority of being implemented easily.

This study examines the connection between different digital-twin characteristics and organizational control. Specifically, the study aims to examine whether the digital-twin characteristics exploration, guidance and gamification will affect formal and social control.

The study is based on an analysis of survey results from 139 respondents comprising applied university students who use digital twins.

The results offer an interesting contribution to the literature. The authors consider the digital-twin characteristics exploration, guidance and gamification and investigate their contribution to two types of organizational controls: formal and social. The results show that two characteristics, exploration and gamification, affect the extent to which digital twins can be utilized for social control. Exploration and guidance’s role is significant concerning the extent to which digital twins can be utilized for formal control.

This study contributes to literature by considering multiple digital-twin characteristics and their contribution to two different control outcomes. First, it diverges from previous technical-oriented research by investigating digital twins in a human context. Second, the study is the first to examine digital twins’ effects from an organizational control perspective systematically.

The aim of this paper is the implementation and validation of control and guidance algorithms for unmanned aerial vehicle (UAV) autopilots.

The path-following control of the UAV can be separated into different layers: inner loop for pitch and roll attitude control, outer loop on heading, altitude and airspeed control for the waypoints tracking and waypoint navigation. Two control laws are defined: one based on proportional integrative derivative (PID) controllers both for inner and outer loops and one based on the combination of PIDs and an adaptive controller.

Good results can be obtained in terms of trajectory tracking (based on waypoints) and of parameter variations. The adaptive control law guarantees smoothing responses and less oscillations and glitches on the control deflections.

The proposed controllers are easily implementable on-board and are computationally efficient.

The algorithm validation via hardware in the loop simulations can be used to reduce the platform set-up time and the risk of losing the prototype during the flight tests.

This paper seeks to examine the development of the on‐board guidance law for multi‐revolutions orbit transfer spacecraft with low‐thrust propulsion systems.

In the research, first, a set of equinoctial elements is utilized to avoid the singularities in dynamical equation of classical orbit elements. A thruster switch law is derived by analyzing the efficiency of the changing of each orbit elements. Second, by using the theory of Lyapunov feedback control, analytic expressions of thrust angles are derived. Finally, the weights of the Lyapunov function are adjusted by hybrid genetic algorithm to improve the performance of the guidance law.

First, the dynamical equations of classical orbit elements are always singularity during the orbit transfer. By using modified equinoctial elements, these singularities could be avoided. Second, the trajectory is sensitive to the weights in Lyapunov function. With reasonable weights, the key parameters under the control of the guidance law presented in this paper are very close to that of optimal trajectory.

In further research, some dynamical weights methods will be used in the control law to improve the performance index, and approach the optimal solution.

The guidance law presented in this paper could be easily used as an on‐board algorithm for the multi‐revolutions orbit transfer or stationkeeping. Furthermore, it could also be utilized as an initial design method for low‐thrust orbit transfer.

Providing a low‐thrust guidance law by combining the concept of Lyapunov feedback control with hybrid genetic algorithm. This method has a super convergence and a low‐computational cost.

The purpose of this paper is to present a novel guidance law for hypervelocity descent to a stationary target such that the impact angle and impact velocity can be constrained.

The proposed method is based on inverse dynamics and is designed using a third-order Bézier curve approximation to the reference trajectory.

Simulations indicate that the proposed law is able to satisfy impact angle and impact velocity constraints as well as follow control and path limitations in the case of guidance under perturbations. Comparisons with other methods also indicate better performance.

The onboard implementation requires an offline selection of Bézier parameters.

The presented scheme could be extremely important for further research on automated onboard control of impact angle and velocity for both re-entry and terminal guidance laws.

This paper presents an innovative method for the solution of an inverse dynamics-based guidance law using Bézier curve approximation.