Active Spacecraft Research Articles

Genetic Algorithm of Energy Consumption Optimization for Reorientation of the Spacecraft Orbital Plane

The paper is dedicated to the problem of finding optimal spacecraft trajectories. The equations of spacecraft motion are written in quaternion form. The spacecraft moves on its orbit under acceleration from the limited in magnitude jet thrust. It is necessary to minimize the energy costs for the process of reorientation of the spacecraft orbital plane. The equations of spacecraft motion are written in orbital coordinate system. It is assumed that spacecraft orbit is circular and control has constant value on each part of active spacecraft motion. In this case the lengths of the sections of the spacecraft motion are unknown. We need to find the length of each section, their quantity and value of control on each section. The equations of the problem were written in dimensionless form. It simplifies the numerical investigation of the obtained problem. There is a characteristic dimensionless parameter in the phase equations of the problem. This parameter is a combination of dimension variables describing the spacecraft and its orbit. Usually the problems of spaceflight mechanic are solved with the maximum principle. And we have to solve boundary value problems with some kind of shooting method (Newton’s method, gradient descent method etc.) Each shooting method requires initial values of conjugate variables, but we have no analytical formulas to find them. In this paper spacecraft flight trajectories were found with new genetic algorithm. Each gene contains additional parameter which equals to " True" , if the gene forms the control and equals to "False" otherwise. It helps us determine the quantity of spacecraft active motion parts. The input of proposed algorithm does not contain information about conjugate variables. It is well-known that the differential equations of the problem have a partial solution when the spacecraft orbit is circular and control is constant. The genetic algorithm involves this partial solution and its speed is increased. Numerical examples were constructed for two cases: when the difference between angular variables for start and final orientations of the spacecraft orbital plane equals to a few (or tens of) degrees. Final orientation of the spacecraft plane of orbit coincides with GLONASS orbital plane. The graphs of components of the quaternion of orientation of the orbital coordinate system, the longitude of the ascending node, the orbit inclination and optimal control are drawn. Tables were constructed showing the dependence of the value of the quality functional and the time spent on the reorientation of the orbital plane on the maximum length of the active section of motion.

Maneuver Strategy for Active Spacecraft to Avoid Space Debris and Return to the Original Orbit

During normal operation of the on-orbit spacecraft, if some satellite in a nearby orbit suddenly breaks apart, its debris will threat the safe operation of the on-orbit spacecraft. Therefore, it is necessary to study the active spacecraft’s avoidance of the space debris group and returning to the original orbit. In this way, the safe operation of on-orbit spacecraft will be guaranteed. However, as the geometric structure of the space debris group is constantly changing, it is hard to accurately demonstrate the changing shape of the debris group, let alone determine the unreachable domain. Traditional obstacle avoidance problems involve low speed of the vehicle; so, the application of artificial potential field and particle swarm algorithms is suitable for such problems. However, these two methods are not applicable to the maneuver strategy of spacecraft with high initial velocity. Therefore, to help spacecraft avoid the space debris group, a new method is required. This paper has established a simplified model to simulate the unreachable domain of the space debris group. It has modified the artificial potential field (APF) method and particle swarm optimization algorithm, with an aim to help spacecraft avoid the space debris group and return to the original orbit. Based on the method, the paper has proposed a three-stage maneuver strategy for the spacecraft to avoid the debris. To show the effectiveness of the method, this paper has simulated an on-orbit spacecraft’s avoidance of the space debris group nearby and returning to its original orbit. Through simulation, the feasibility of the maneuver strategy for spacecraft in the geosynchronous orbit is evaluated. The simulation results show that the method proposed in this paper can effectively accomplish the task.

Relating Collision Probability and Separation Indicators in Spacecraft Formation Collision Risk Analysis

Active spacecraft formation flying collision avoidance schemes monitor collision risk through indicators such as miss distance and collision probability. This paper compares instantaneous collision probability measures based on planar projections to their three-dimensional counterparts. In this analysis, it is found that the former overestimate the latter. Additionally, this work compares the consistency of risk assessments based on miss distance and instantaneous collision probability in closed Clohessy–Wiltshire trajectories. Certain statistics of relative position are well suited for collision risk assessments because their local minima and collision probability local maxima are anticorrelated. These results are a step toward connecting both types of indicators into a cohesive mathematical framework relating to collision risk.

Detumbling of axisymmetric space debris during transportation by ion beam shepherd in 3D case

The paper considers the contactless transportation of space debris by means of an ion beam generated by an active spacecraft electrodynamic engine. The three-dimensional motion of a space debris object relative to its center of mass is studied. Space debris is assumed to be a dynamically symmetric cylindrical rigid body. The aim of the paper is to develop a control of the engine thrust, which ensures the stabilization of spatial oscillations of cylindrical space debris. A simplified mathematical model describing the motion of a dynamically symmetric rigid body in a Keplerian orbit is developed. For the case of a circular orbit, stationary motions relative to the center of mass are found. A feedback control law for the thrust of an electrodynamic engine creating an ion beam, which is aimed at stabilizing the space debris object in a stationary angular position, is proposed. The results of numerical research confirm the effectiveness of this control. The research results can be used in the preparation of space debris removal missions.

Simulation and Uncertainty Quantification of Electron Beams in Active Spacecraft Charging Scenarios

Novel active sensing methods have been recently proposed to measure the electrostatic potential of noncooperative objects in geosynchronous equatorial orbit and deep space. Such approaches make use of electron beams to excite the emission of secondary electrons and X-rays and infer properties of the emitting surface. However, the detectability of secondary electrons is severely complicated in the presence of complex charged bodies, making computationally efficient simulation frameworks necessary for in situ potential estimation. The purpose of this paper is twofold: firstly, to introduce and test a quasi-analytical, uncoupled, and computationally efficient electron beam expansion and deflection model for active charging applications; and secondly, to characterize the uncertainty in the beam–target intersection properties, which condition the measurement of secondary electrons. The results show that a combination of secondary electrons and X-ray methods is highly desirable to yield a robust and accurate measure of the potential of a target spacecraft.

Practical maintenance strategies for teardrop hovering formation relative to elliptical orbit

Hovering technology, as one of the core technologies in on-orbit service (OOS) missions, enables the relative position between an active spacecraft and chief spacecraft to be maintained within a given range, and it has received widespread attention in recent years. In this study, for an elliptical reference orbit, we proposed a series of control strategies to maintain a teardrop hovering formation with a J2 perturbation. First, we formulated a mathematical model of the teardrop hovering formation maintained by the impulse control and presented a rapid method based on a relative-position correction method. Based on the sensitivity matrix, a strategy for maintaining a hovering formation using constant-thrust control was obtained. Furthermore, based on the proposed fast iterative method, an improved strategy was developed to satisfy a magnitude constraint. Finally, we analyzed the influence of the reference orbital elements and formation parameters on the hovering formation using numerical simulations, which demonstrated the efficiency of the proposed method. This provides a reference for the preliminary design of hovering formations.

Fuel costs estimation for ion beam assisted space debris removal mission with and without attitude control

Contactless space debris removal systems are the promising direction for solving the space debris problem. An ion beam transportation method is based on blowing the space debris with a high-speed plasma flow generated by the engine of an active spacecraft. The aim of the work is to evaluate the effect of controlling the attitude motion of space debris on the amount of fuel consumption during the contactless transportation. Four control strategies for an active spacecraft are proposed and compared. The first strategy involves transporting the object without considering its attitude motion. The second strategy involves stabilizing the object in a stable angular equilibrium position. The third strategy assumes transportation the space debris in the oscillation mode with the maximum average force. The fourth strategy involves stabilizing the object in an angular position corresponding to the maximum ion force. The control laws for the spacecraft's thrusters and for the direction of the ion beam axis, implementing these strategies, are proposed. The simulation of the transportation of the rocket stage is carried out and the fuel costs for various strategies are calculated. The most preferable from the point of view of minimizing fuel consumption is the second strategy. The fourth strategy is ineffective. The results obtained in this work can be used in the preparation of contactless space debris removal missions.

Algorithm for solving optimal open-loop terminal control problem for spacecraft rendezvous with account of constraints on the state

The paper proposes an algorithm for solving the optimal open-loop terminal control problem of two spacecraft rendezvous with constraints on their states. A system of nonlinear differential equations that describes the dynamics of the active (maneuvering) spacecraft relative to the passive spacecraft (station) in the central gravitational field of the Earth in the orbital coordinate system of coordinates related to the passive spacecraft center-of-mass is considered as an initial model. The obtained nonlinear model of the active spacecraft dynamics is linearized relative to the specified reference state trajectory of the passive spacecraft, and then it is discretized and reduced to linear recurrence relations. Mathematical formalization of the spacecraft rendezvous problem under consideration is carried out at a specified final moment of time for the obtained discrete-time controlled dynamical system. The quality of solving the problem is estimated by a convex functional taking into account the geometric constraints on the active spacecraft states and the associated control actions in the form of convex polyhedral-compacts in the appropriate finite dimensional vector space. We propose a solution of the problem of optimal terminal control over the approach of the active spacecraft relative to the passive spacecraft in the form of a constructive algorithm on the basis of the general recursive algebraic method for constructing the availability domains of linear discrete controlled dynamic systems, taking into account specified conditions and constraints, as well as using the methods of direct and inverse constructions. In the final part of the paper, the computer modeling results are presented and conclusions about the effectiveness of the proposed algorithm are made.

Near-field radiation assisted smart skin for spacecraft thermal control

Thermal control is of critical importance for normal operation of spacecraft. Given thermal radiation is the only means of heat dissipation in space, an efficient thermal control approach for spacecraft is to coat the radiator with a tunable-emittance “skin” that can tune its heat dissipation according to various thermal conditions. The existing schemes solely relying on far-field thermal radiation, which are based on mechanical, electrochromic or thermochromic working principles, are difficult to combine the advantages of all-solid-state structure, actively and accurate tuning, and large tuning range of heat flux. In this work, we propose a near-field radiation assisted (NFRA) smart skin for thermal control which can tune the heat rejection accurately and in a large range. It contains a metal-insulator-semiconductor (MIS) structure, where the carrier distribution in the semiconductor layer can be electrically altered. In this way, the near-field heat flux, and ultimately the skin emission power expressed using effective emittance, can be controlled as a function of the applied voltage. The variation range of the effective emittance can exceed 0.7 when adjusting the applied voltage from −10 V to 100 V with our preliminary design. This work opens a new way of smart skin design for active spacecraft thermal control.

Adaptive neural finite-time control for space circumnavigation mission with uncertain input constraints

This paper explores the design of an anti-saturation adaptive finite-time control strategy with the neural network (NN) technique for the space circumnavigation mission. Before executing the controller design, the analytical solutions of the desired angular velocity and its derivative of the active spacecraft are calculated. Since there are uncertain saturation constraints on control forces and moments in the actual propulsion system, an auxiliary system compensated by an adaptive NN is adopted. The modified auxiliary system no longer needs the precise output values of the actuators. Besides, the hyperbolic tangent function is introduced to design the weight update law for the NN compensator, so that the derivative of the weight estimator will not be amplified by the quadratic of states when the system states are large. It is proved that tracking errors of the system states can converge to a residual set of the origin in finite time. Simulation results show that the maximum amplitudes of the control signals are greatly reduced compared to the classical non-singular terminal sliding-mode control scheme, and that the neural-based compensator can significantly weaken the overshoot and chattering.

Practical Optimization of Low-Thrust Minimum-Time Orbital Rendezvous in Sun-Synchronous Orbits

High-specific-impulse electric propulsion technology is promising for future space robotic debris removal in sun-synchronous orbits. Such a prospect involves solving a class of challenging problems of low-thrust orbital rendezvous between an active spacecraft and a free-flying debris. This study focuses on computing optimal low-thrust minimum-time many-revolution trajectories, considering the effects of the Earth oblateness perturbations and null thrust in Earth shadow. Firstly, a set of mean-element orbital dynamic equations of a chaser (spacecraft) and a target (debris) are derived by using the orbital averaging technique, and specifically a slow-changing state of the mean longitude difference is proposed to accommodate to the rendezvous problem. Subsequently, the corresponding optimal control problem is formulated based on the mean elements and their associated costate variables in terms of Pontryagin’s maximum principle, and a practical optimization procedure is adopted to find the specific initial costate variables, wherein the necessary conditions of the optimal solutions are all satisfied. Afterwards, the optimal control profile obtained in mean elements is then mapped into the counterpart that is employed by the osculating orbital dynamics. A simple correction strategy about the initialization of the mean elements, specifically the differential mean true longitude, is suggested, which is capable of minimizing the terminal orbital rendezvous errors for propagating orbital dynamics expressed by both mean and osculating elements. Finally, numerical examples are presented, and specifically, the terminal orbital rendezvous accuracy is verified by solving hundreds of rendezvous problems, demonstrating the effectiveness of the optimization method proposed in this article.

Attitude motion of a space object during its contactless ion beam transportation

The development of systems for transporting space objects by an active spacecraft is an urgent task of modern astronautics. One of the promising areas in this field is contactless ion beam transport systems. The goals of this work are to study a space object attitude motion dynamics during its contactless transportation and the influence of this motion on the magnitude of the force transmitted by the ion beam. A mathematical model of a mechanical system consisting of the space object and the active spacecraft was developed using the Lagrange formalism. A simplified equation of the object's angular motion in Kepler's orbit was obtained. The influence of the object's center of mass position and the ratio of ion beam and gravity gradient torques on the phase portrait view in the case of cylindrical object in a circular orbit was studied. The most favorable areas on phase portrait from the point of view of the magnitude of the force transferred by the ion beam to the object were found. It is shown that in the case of an elliptical orbit chaos exists. Entering a phase trajectory in a chaotic layer can lead to a noticeable decrease in the ion beam force.

Feasibility analysis of LEO and GEO large space debris de/re-orbiting taking into account launch mass of spacecraft-collector and its configuration layout

Analysis of the efficiency of two basic strategies for de/re-orbiting large space debris objects to disposal orbits (DO) is given. Large objects in LEO are classified into groups with similar orbital inclinations and comprise primarily last stages of launch vehicles, in GEO vicinity the paper studies upper stages. Under the first de/re-orbiting variant, it is assumed a spacecraft-collector is equipped with several thruster de/re-orbiting kits (TDKs); one of them can be fixed on an object and is capable of de/re-orbiting an object to a DO independently of the collector. In the second variant, a collector operates as a space tug: transfers objects to a DO and then returns to the next objects in line. The authors study possible configuration layouts of collectors in LEO and near GEO. The available analogous projects are analyzed. The efficiency of both de/re-orbiting variants can be properly compared using the estimations of collector's dry mass and having at one's disposal the parameters of the maneuvers required for transfers between all objects in the group. As reasonable criteria of effectiveness, one can consider (separately or jointly) the launch mass of an equipped collector, its ΔV budget, and the required number of such active spacecraft. Two de/re-orbiting variants are compared in terms of these criteria via mass-energy diagrams constructed for each group of objects in both altitude regions. Analysis of these diagrams shows that low Earth orbits can be more efficiently cleaned under the first de-orbiting variant by using a two-stage space system consisting of an active spacecraft carrying TDKs. For GEO, it is expedient to choose the second re-orbiting variant using a single-stage spacecraft. Our analysis shows that LEO cleaning is an order of magnitude more expensive than that for GEO, hence the problem of LEO population should be given increased attention.

Teardrop hovering formation for elliptical orbit considering J2 perturbation

On-orbit service (OOS) technologies, applied to provide essential maintenance to prolong the lifetime of spacecraft and to extend their function, have received widespread attention recently. The hovering formation, as one of the core technologies, enables the relative motion state between an active spacecraft and chief spacecraft within a specified range. In this study, we proposed a teardrop hovering formation, which is suitable for elliptical reference orbit with a J2 perturbation. We formulated a mathematical model of the teardrop hovering formation maintained by the impulse control and presented a rapid method based on the state transition matrix. With the triangles of the three tangent lines solved by the rapid method, we estimated the envelope of the multi-loop hovering formation. Furthermore, we analyzed the influence of the reference orbital elements and formation parameters on the hovering formation by numerical simulations, which demonstrated the efficiency of the proposed method and provided a reference for the preliminary design of hovering formations.

Chaotic motion of a cylindrical body during contactless transportation from MEO to LEO by ion beam

The attitude dynamics of a cylindrical body under the action of an ion beam generated by the low-thrust engine of an active spacecraft is studied. The aim of the work is the analysis of chaotic and regular attitude motion of the body during contactless transportation. The eccentricity of the orbit could cause chaotic motion of the system. A simplified mathematical model, which describes the motion of a cylinder under the action of an ion beam, was developed. A bifurcation diagram was constructed. It allows judgment of the number and type of singular points of the phase portrait of an unperturbed system depending on the ratio between gravitational and ionic influences. An analysis of the chaotic motion was carried out using Poincare sections and Lyapunov exponents. It shows that in the process of contactless transportation from MEO to LEO, the cylinder can become uncontrollable from the oscillation motion mode to the rotation mode. The orientation of the body in the ion beam affects the magnitude and direction of the ion beam force vector; thus, the body angular motion mode affects the time duration of the performed orbital maneuver and the magnitude of the required fuel.

Evaluation of debris mitigation options for a large constellation

Large constellations of satellites in Low Earth Orbit (LEO) may adversely impact the sustainable use of the space environment over the long-term unless appropriate plans for orbital debris mitigation measures are incorporated into the design and operation. In particular, recent computer modelling studies have shown that comprehensive observance of the Inter-Agency Space Debris Coordination Committee (IADC) guideline on post-mission disposal is a vital element of debris mitigation plans intended for large constellations in LEO. Several operators have proposed that the region of LEO below 600 km altitude could be used to accommodate large numbers of spacecraft without unduly heightening orbital debris concerns. The objective of this study, using the DAMAGE computer model, was to evaluate the implications for the orbital debris population of orbit design options for a large constellation. The build-up, replenishment and disposal of a large, complex constellation comprising 10,440 active spacecraft was simulated. A subset of the constellation, a component comprising 1800 spacecraft in a single band, was deployed at either 1100 km or 550 km altitude. The positioning of this component provided the two fundamental study conditions. The results showed that positioning the 1800-spacecraft component at 550 km altitude reduced the overall impact of the constellation substantially, whilst also reducing the need for a high post-mission disposal success rate.

Оптимальное управление тягой буксира при развертывании троса после захвата космического мусора гарпуном

The paper considers the process of removing a passive object by an active spacecraft, consisting of three stages: harpooning the object, deploying the tether, and towing. The shock pulse from the harpoon is used to reduce the angular velocity of the object and transfer it to the towing state. An algorithm for determining the position of the capture point on the object surface is proposed. The equations of relative motion of the transfer vehicle at the stage of tether deployment are given in the dimensionless form, which allows studying the movement in any parameter space. The law of thrust control at the stage of cable deployment is proposed and optimal control parameters that ensure safe towing are determined. The limits of applicability of the considered control law are found taking into account the structural and strength limitations of the system. As an example of using the proposed approach, a numerical simulation of withdrawing the Ariane 4 rocket upper stage was performed.

Autonomous relay control system for satellite formation operating in low Earth orbit

A relay control algorithm is proposed for the satellite formation consisting of a «passive» (virtual) spacecraft moving in an unperturbed circular orbit, and several “active” spacecraft maneuvering relative to it. The purpose of control is to keep each active spacecraft on its finite path relative to the virtual spacecraft defining the center of formation with as low as possible propellant consumption. To describe the motion of maneuvering spacecraft relative to the center of formation, modified Hill–Clohessy–Wiltshire equations are used accounting for the earth's flattening and aerodynamic drag. The main emphasis of this paper is on the study of the existence of stable limit cycles and determination of their domains of attraction, as well as the search for values of the relay control system parameters ensuring a minimum propellant consumption for maintaining a proper dynamic formation behavior. Key words: relay control, point transformation method, formation flying, limit cycle, phase plane.

SATELLITE SEGMENT OF NATIONAL INFORMATION INFRASTRUCTURE. APPROACHES TO IMPLEMENTATION

It is proved that the satellite segment is a necessary component of the national information infrastructure. From this position, the state and development trends of satellite telecommunication systems and satellite channel organization technologies are analyzed. It is shown that there are two ways to build a satellite segment: based on the use of leased resources of satellite operators operating in the satellite communications market; based on the resources of the national satellite communications system, which still needs to be created. It is shown that satellite telecommunication systems have a steady tendency to increase the number of spacecraft, until the launch of “heavyˮ satellites with a payload of more than 70 transponders and to increase the frequency resource. The diagram of satellites location density of in orbit is given. In the case of orientation to rent the resources of active spacecraft, the criteria for their selection is indicated. The arc of the geostationary orbit, on which the satellite should be selected, is calculated. For satellites located on this arc, diagrams of equivalent isotropically radiated power values in the C and Ku bands, which they provide in Ukraine, are shown. A list of satellites that you can focus on, recommendations for selection and a list of recommended for the satellite segment is indicated. The technology standards used in payload and methods for organizing satellite communications channels is noticed. To select the parameters of the satellite segment, the results of calculations of the effectiveness of signal-code structures are provided. Approaches to creating a satellite segment of the national information infrastructure are proposed. It is shown that the orientation to the Lybid national satellite does not make sense; a new frequency-orbit resource is required. The types of platforms that can be used on a national satellite are considered.

Control of the Rendezvous of Two Spacecraft Using a Tether System

The rendezvous of two spacecraft (passive and active) using a tether system is studied. Two laws of the tether control are proposed to ensure the safe rendezvous of bodies. A minimum allowable thrust of the active spacecraft and the parameters of the tether control laws are defined, under which the rendezvous is implemented safe. Based on the results of numerical simulation, the ability to control the tether using the proposed laws is showed.