Adaptive and Supertwisting Adaptive Spacecraft Orbit Control Around Asteroids

This paper addresses the development of control systems for the orbit control of spacecraft around irregularly shaped rotating asteroids with uncertain parameters. The objective is to steer the spacecraft along prescribed orbits. First, a nonlinear adaptive law for orbit control was designed. This was followed by the design of a supertwisting adaptive (STWA) control system. In the closed-loop system, which includes the adaptive law or the STWA law, all the signals remain bounded, and the trajectory tracking error asymptotically converges to zero for any initial condition. Finally, under the assumption of boundedness of the derivative of the uncertain functions of the model in a region of the state space, a supertwisting control (STW) law for finite-time convergence of the trajectory was obtained. Based on the Lyapunov theory, stability properties of the closed-loop systems were analyzed. Simulation results for 433 Eros and Ida asteroids were presented for illustration. The results showed that control of spacecraft along closed orbits or to a fixed point is accomplished using each of these controllers, despite uncertainties in the parameters of the asteroid models

Asteroid Deflection: How, where and when?

To deflect impact-trajectory of massive km^3 and spinning asteroid by a few terrestrial radius one need a large momentum exchange. The dragging of huge spinning bodies in space by external engine seems difficult or impossible. Our solution is based on the landing of multi screw-rockets, powered by mini-nuclear engines, on the body, that dig a small fraction of the soil surface, to use as an exhaust propeller, ejecting it vertically in phase among themselves. Such a mass ejection increases the momentum exchange, their number redundancy guarantees the stability of the system. The slow landing (below 40 cm s^-1) of each engine-unity at those lowest gravity field, may be achieved by save rolling and bouncing along the surface. The engine array tuned activity, overcomes the asteroid angular velocity. Coherent turning of the jet heads increases the deflection efficiency. A procession along its surface may compensate at best the asteroid spin. A small skin-mass (about 2 10^4 tons) may be ejected by mini nuclear engines. Such prototypes may also build first save galleries for humans on the Moon. Conclusive deflecting tests might be performed on remote asteroids. The incoming asteroid 99942 Apophis (just 2% of km^3) may be deflected safely a few Earth radius. How to tag its trajectory is described. Its encounter maybe not just a hazard but an opportunity, learning how to land, dig, build and also to nest save human station inside. Asteroids amplified deflections by gravity swing maybe driven into longest planetary journeys. Mars journey may benefict by Phobos natural asteroid parking and fueling role.Comment: 14 pages, 5 figures; editorial corrections and answer to referee open questions on project time scal

Optimal landing site selection based on safety index during planetary descent

AbstractLanding safety is the prior concern in planetary exploration missions. With the development of precise landing technology, future missions require vehicles to land on places of great scientific interest which are usually surrounded by rocks and craters. In order to perform a safe landing, the vehicle should be capable of detecting hazards, estimating its fuel consumption as well as touchdown performance, and locating a safe spot to land. The landing site selection process can be treated as an optimization problem which, however, cannot be efficiently solved through traditional optimization methods due to its complexity. Hence, the paper proposes a synthetic landing area assessment criterion, safety index, as a solution of the problem, which selects the best landing site by assessing terrain safety, fuel consumption and touchdown performance during descent. The computation effort is cut down after reducing the selection scope and the optimal landing site is found through a quick one-dimensional search. A typical example based on the Mars Science Laboratory mission is simulated to demonstrate the capability of the method. It is proved that the proposed strategy manages to pick out a safe landing site for the mission effectively. The safety index can be applied in various planetary descent phases and provides reference for future mission designs

Development of non-linear guidance algorithms for asteroids close-proximity operations

In this paper, we discuss non-linear methodologies that can be employed to devise real-time algorithms suitable for guidance and control of spacecrafts during asteroid close-proximity operations. Combination of optimal and sliding control theory provide the theoretical framework for the development of guidance laws that generates thrust commands as function of the estimated spacecraft state. Using a Lyapunov second theorem one can design non-linear guidance laws that are proven to be globally stable against unknown perturbations with known upper bound. Such algorithms can be employed for autonomous targeting of points of the asteroid surface (soft landing , Touch-And-Go (TAG) maneuvers). Here, we theoretically derived and tested the Optimal Sliding Guidance (OSG) for close-proximity operations. The guidance algorithm has its root in the generalized ZEM/ZEV feedback guidance and its mathematical equations are naturally derived by a proper definition of a sliding surface as function of Zero-Effort-Miss and Zero-Effort-Velocity. Thus, the sliding surface allows a natural augmentation of the energy-optimalguidance via a sliding mode that ensures global stability for the proposed algorithm. A set of Monte Carlo simulations in realistic environment are executed to assess the guidance performance in typical operational scenarios found during asteroids close-proximity operations. OSG is shown to satisfy stringent requirements for asteroid pinpoint landing and sampling accuracy

Adaptive-Optimal Control of Spacecraft near Asteroids

Spacecraft dynamics and control in the vicinity of an asteroid is a challenging and exciting problem. Currently, trajectory tracking near asteroid requires extensive knowledge about the asteroid and constant human intervention to successfully plan and execute proximity operation. This work aims to reduce human dependency of these missions from a guidance and controls perspective. In this work, adaptive control and model predictive control are implemented to generating and tracking obstacle avoidance trajectories in asteroid’s vicinity. Specifically, direct adaptive control derived from simple adaptive control is designed with e modification to track user-generated trajectories in the presence of unknown system and sensor noise. This adaptive control methodology assumes no information on the system dynamics, and it is shown to track trajectories successfully in the vicinity of the asteroid. Then a nonlinear model predictive control methodology is implemented to generate obstacle avoidance trajectories with minimal system information namely mass and angular velocity of the asteroid. Ultimately, the adaptive control system is modified to include feed-forward control input from the nonlinear model predictive control. It is shown through simulations that the new control methodology names direct adaptive model predictive control (DAMPC), is able to generate sub-optimal trajectories. A comparative study is done with Asteroid Bennu, Kleopatra and Eros to show the benefits of DAMPC over adaptive control and MPC. A study on effect of noisy measurements and model is also conducted on adaptive control and direct adaptive model predictive control

Mars Powered Descent Phase Guidance Design Based on Fixed-Time Stabilization Technique

This paper proposes a guidance scheme to achieve an autonomous precision landing on Mars and proposes a practical fixed-time stabilization theorem to analyze the robustness of the guidance. The proposed guidance is mainly based on the fixed-time stabilization method, and it can achieve the precision landing within a pre-defined time. This property enables the proposed guidance to outperform the finite-time stabilization technique which cannot handle uncertainties well and whose convergence time is dependent on initial states. Compared with the existing fixed-time stabilization theorem, the proposed practical fixed-time stabilization theorem can achieve a shorter convergence time and cope with unknown disturbances. When the Mars landing guidance is designed by this proposed theorem, the upper bound of the landing time and the maximum landing error subject to unknown disturbances can be calculated in advance. Theoretical proofs and Monte Carlo simulation results confirm the effectiveness of the proposed theorem and the proposed guidance. Furthermore, the efficacy of the proposed guidance with thrust limitations is also demonstrated by testing of 50 cases with a range of initial positions and velocities