Motion Of Spacecraft Research Articles

Dual-Quaternion Analytic LQR Control Design for Spacecraft Proximity Operations.

Proximity operations offer aggregate capability for a spacecraft operating in close proximity to another spacecraft, to perform on-orbit satellite servicing, or to a space object to perform debris removal. To utilize a spacecraft performing such advanced maneuvering operations and perceiving of the relative motion of a foreign spacecraft, these trajectories must be modeled accurately based on the coupled translational and rotational dynamics models. This paper presents work towards exploiting the dual-quaternion representations of spacecraft relative dynamics for proximity operations and developing a sub-optimal control law for efficient and robust maneuvers. A linearized model using dual-quaternions for the proximity operation was obtained, and its stability was verified using Monte Carlo simulations for the linear quadratic regulator solution. A sub-optimal control law using generalized higher order feedback gains in dual-quaternion form was developed based on small error approximations for the proximity operation and also verified through Monte Carlo simulations. Necessary information needed to understand the theory behind the use of the dual-quaternion is also overviewed within this paper, including the validity of using the dual-quaternions against their Cartesian or quaternion equivalents.

Robust Finite-State Controllers for Uncertain POMDPs

Uncertain partially observable Markov decision processes (uPOMDPs) allow the probabilistic transition and observation functions of standard POMDPs to belong to a so-called uncertainty set. Such uncertainty, referred to as epistemic uncertainty, captures uncountable sets of probability distributions caused by, for instance, a lack of data available. We develop an algorithm to compute finite-memory policies for uPOMDPs that robustly satisfy specifications against any admissible distribution. In general, computing such policies is theoretically and practically intractable. We provide an efficient solution to this problem in four steps. (1) We state the underlying problem as a nonconvex optimization problem with infinitely many constraints. (2) A dedicated dualization scheme yields a dual problem that is still nonconvex but has finitely many constraints. (3) We linearize this dual problem and (4) solve the resulting finite linear program to obtain locally optimal solutions to the original problem. The resulting problem formulation is exponentially smaller than those resulting from existing methods. We demonstrate the applicability of our algorithm using large instances of an aircraft collision-avoidance scenario and a novel spacecraft motion planning case study.

Відносний рух космічного апарата з аеродинамічним компенсатором у перпендикулярному до площини орбіти напрямку при безконтактному видаленні космічного сміття

The article investigates the feasibility of using an aerodynamic compensator for contactless removal of space debris from low Earth orbits, taking into account aerodynamic disturbances in the direction perpendicular to the orbital plane. The object of the research is a modified scheme of the “ion beam shepherd” de-orbiting concept. The modification consists in replacing an additional electric thruster with an aerodynamic compensator designed to compensate the shepherd spacecraft motion caused by the reaction force of the main electric thruster, the ion plume of which creates a “braking” effect on the space debris object. The shepherd spacecraft with the aerodynamic compensator has a relatively large cross-sectional area. In this case, it is necessary to control the relative motion caused by the aerodynamic disturbances in the direction perpendicular to the orbital plane. This control requires additional propellant for the thrusters of the relative motion control system of the shepherd spacecraft. The article presents the calculation of the propellant consumption using a number of simplifying assumptions. The validity of these assumptions is verified by numerical integration of the equations of relative motion. The feasibility of using the aerodynamic compensator for contactless removal of space debris, taking into account aerodynamic disturbances acting in the direction perpendicular to the orbital plane, is shown.

Стратегия развития системы стабилизации космических аппаратов МКРЦ в различных полях

Control of spacecraft motion relative to its center of mass, depending on their specific purpose and intended program of application (for example, reconnaissance and output of target designation (CA)), is generally reduced to provision of specified angular motion of spacecraft relative to its center of mass and its angular orientation with specified accuracy.

Математическая модель маневрирующего наноспутника с гелиотермической двигательной установкой и маховичной системой ориентации и стабилизации

The paper considers a possibility of using a heliothermal propulsion system for constructing constellations of nanosatellites in the CubeSat format. Such a propulsion system implies direct heating of the working fluid by focused solar radiation. In comparison with electric jet engines, the proposed propulsion has orders of magnitude higher thrust, which makes it possible to build a constellation of nanosatellites in low Earth orbit in less than 6 days. In comparison with electrothermal motors, the presented solution achieves higher efficiency of converting the energy of solar radiation into the thermal energy of the working fluid. The spacecraft motion was simulated taking into account the mutual influence of the propulsion system, the orientation and stabilization system, the power supply system, as well as the passage of the shadow sections of the orbit in which there is a loss of electrical and thermal energy.

Моделі спостережуваного руху космічних об'єктів на основі часових рядів TLE-елементів

Timely detection of changes in the characteristics of space hardware objects during their long-term operation is one of the main tasks in the development and study of onboard systems that maintain the efficiency of their operation. This paper presents a statistical method for simulating the motion of space objects (spacecraft and used launch vehicle stages) in the class of autoregressive models. The method allows one to improve the quality of description and prediction of the motion of space objects based on simulating time series of their TLE-elements (two-line orbital element sets). The purpose of this work is to increase the accuracy of mathematical models of the observed motion of space objects in the problems of deorbit time determination, satellite collision prediction, and space debris cataloging. The paper presents a system for simulating the motion of space objects, which allows one to determine an optimal amount of learning samples in simulating time series of TLE elements, determine the order of autoregression and find an optimal model structure for each variable element, identify model parameters in conditions of unequally spaced observations, identify features of the time behavior of the root-mean-square errors of the constructed autoregressive models on the basis of dividing the initial time series of TLE-elements into successive learning intervals, and obtain predictive estimates of the values of variable elements. The proposed statistical method of space object motion simulation can be recommended to describe and predict the motion of spacecraft and used launch vehicle stages represented as time series of TLE-elements (which are publicly available and regularly updated). The application of the proposed statistical method will increase the accuracy of mathematical models of the observed motion of space objects in the problems of deorbit time determination, satellite collision prediction, and space debris cataloging.

Interactive imitation learning for spacecraft path-planning in binary asteroid systems

Exploration of small body systems poses the problem of designing path planning strategies for possibly uncharted environments. Traditional methods aimed at developing rigorous trajectory baselines may suffer inefficiencies, or turn infeasible when confronted with unknown dynamics. In strongly non-linear dynamics, mapping point design solutions from one dynamical regime to another may be hindered by underlying chaotic behavior. Rather than relying on baseline driven approaches, more generalized strategies may be found by observing human pilots controlling spacecraft motion within varying dynamical environments; the resultant data can then be utilized to initialize machine learning agents to provide more autonomous solutions. A previous numerical experiment resulted in a technical dataset comprising of human-based path planning strategies across a range of binary asteroid systems. This dataset is now used to train various imitation learning agents, and initiate the creation of a framework that integrates human–machine cooperation into the early training phases of artificial intelligent agents; the current application is for spacecraft guidance in binary asteroid systems, as a prototype of complex, potentially unknown, orbit dynamics. An interactive training architecture, based on the DAgger algorithm, is designed and employed to train both original and interactively coached agents, the latter stemming from both corrective and evaluative feedback by a real time human interactor. All agents were interactively trained for a predefined time period. The results from this investigation may provide the first, empirical observations of behavioral cloning within multi-body dynamics with largely randomized parameters, with some notable contributions including early characterization of training time, initial evidence of an autonomous agent learning meaningful policy features via imitation, and early identification of challenges in training fully autonomous agents for a multi-body dynamics path planning problem of this complexity and high dimensional state space.

Influence of deviations in manufacturing of electrothermal propulsion system on nanosatellite maneuvering accuracy

A method is proposed for assessing the results of adjustment maneuvers for a nanosatellite (NS) with an electrothermal propulsion system (ETPS). Using the example of the SamSat-M nanosatellite under development, common causes of maneuvering errors associated with deviations in the manufacturing of the propulsion system are revealed. Probabalistic analysis of the NS maneuvering process was carried out. The design parameters of the ETPS are considered as random factors. Statistical models of the distributions of all random factors are assumed to be equally probable, which is the worst-case scenario, since the true distributions of the design parameters of the ETPS are unknown. The methodological basis of the study is the method of statistical modeling (Monte Carlo method) followed by the use of regression and factor analysis, on the basis of which the influence of the scatter of each of the design parameters on the controlled parameters is determined. Requirements for the design parameters of the ETPS that affect the spread of the projections of the velocity growth vector of the NS and the arising angular motion have been formulated. The presented results can be used to assess the influence of production deviations in the design parameters of propulsion systems on the nature of spacecraft motion, as well as to state requirements for the spread of design parameters to ensure the achievement of the objective.

Circulating, eccentric periodic orbits at the Moon

Loosely captured orbits with circulating and pulsating eccentricity vectors have a variety of attractive mission design properties, including low insertion costs, near-circular and highly eccentric phases, mid- to high inclinations, long-term stability, and spatially distributed close approaches. Such orbits are the known result of averaging third-body perturbations over the spacecraft and system motions. Unfortunately, the doubly averaged model does not match long-term unaveraged motion well at high altitudes. Singly averaged dynamics are accurate for orbits with much higher semimajor axis values, providing a mechanism to predict and characterize high-altitude motion in the unaveraged model. In this work, singly averaged circulating, eccentric orbits are surveyed in the specific context of the dimensioned Earth–Moon system, where high-altitude orbits are expected to have the most useful applications. A global search is performed over non-impacting system resonances, and families of periodic orbits are used as a framework to map the feasible space. The resulting database of periodic orbits is provided as an online supplement. The singly averaged families are shown to provide a reliable bridge between the analytic results of the doubly averaged system and initial guesses that converge to periodic orbits in the unaveraged model. Non-impacting circulating, eccentric orbits are demonstrated to maintain their structure for multiple years in a high-fidelity force model. These circulating, eccentric, lunar orbits may be useful for human-crewed space stations, dedicated science orbits, extended missions, loitering orbits, or transfers between the Lagrange points and low-lunar orbits.

Transiting planetary candidates from the K2 Campaign 0 using the PSF method

In the Kepler/K2 data analysis performed to date, aperture photometry has generally been used, but this method has not given any reasonable results when examining nearby sources due to the insufficient number of pixels on Kepler's CCD camera. The point-spread function (PSF) technique prepared by Libralato et al. (2016a), which presents promising results in undersampled pixels even in crowded star groups, was used in this study. Open cluster stars M 35 and NGC 2158 in the K2 Campaign 0 were selected as the target field. By applying the PSF method to the high angular resolution input catalogue prepared using the Asiago Schmidt telescope, M 35 and NGC 2158 provided the opportunity to examine the light curves of faint stars (Kp > 15.5) in crowded regions and gave better results compared to aperture photometry. In this study, I detrended ~50 000 raw light curves from systematic effects and spacecraft motion effects. I ultimately found two candidate host stars that could have possible planetary transits.

Kinetic Modeling of Langmuir Probes in Space and Application to the MAVEN Langmuir Probe and Waves Instrument

AbstractSelf‐consistent kinetic solutions from test‐particle simulations are used to improve the derivation of electron temperature (Te) and density (ne) from Langmuir probes in space. While Langmuir probes are well understood, the non‐ideal characteristics of space‐flight instruments can influence the accuracy of Te and ne and how they are derived. In particular, when in an ionosphere, spacecraft motion often causes a sensor wake and exposes its surface to oxidation. This work has two primary goals. We present kinetic solutions of general interest then apply our findings to the Langmuir Probe and Waves (LPW) instrument on the MAVEN satellite. Of general interest, (1) kinetic solutions show that mechanical mounting of a Langmuir probe sensor and controlling the potential of nearby surfaces is critical for accuracy. (2) An ion wake generated by the sensor can greatly modify how ne must be derived. (3) Interestingly, small voltage variations on the surface of the sensor do not significantly diminish the accuracy of Te and ne. (4) On the other hand, surface resistance on the sensor can appreciably disturb the derivation of Te. The LPW instrument is recalibrated with the aid of kinetic solutions and published results from laboratory experiments. The systematic uncertainty (as opposed to random variations) in Te is improved to as low as ±0.005 eV (±60 K) when ne > ∼3 × 104 cm−3. This recalibration leads to some of the most accurate measurements of Te made in space and can result in improved modeling of Mars' ionosphere.

SOFTWARE AND MATHEMATICAL SUPPORT FOR THE ANALYSIS OF OPTIONS FOR CONSTRUCTING ORBITAL GROUPINGS OF THE GLONASS SUPPLEMENT

The high accuracy and stability of Global Navigation Satellite System (GLONASS) consumer characteristics allows us to consider it as the basis of coordinate-time and navigation support for regional consumers. At the same time, one of the key factors influencing their choice of the preferred navigation system is to ensure the required levels of accuracy, availability and continuity of the radio navigation field in conditions of complex terrain, urban development and in areas with a high vegetation index. In this sense, the introduction of additional spacecraft into the GLONASS orbital constellation will increase its competitiveness. The article presents a methodology for building software analysis tools options build orbital groups, which allows you to explore the options of structures high-orbit segment of the principles and methods of their construction in order to effectively ensure implementation of navigation services, to analyze the deterioration of the quality of functional capabilities of high-orbit space of the complex due to the deformation under the influence of disturbing factors of various types and build a strategy for correction of motion of spacecraft. The composition and structure of software and mathematical support are proposed. Presents principles of organization and brief description of the software modules comprising: a module to determine the parameters of the perturbed uncontrolled motion of navigation satellites orbital constellation on the interval active life, module to evaluate degradation of functionality for advanced GLONASS and module support decision-making on strategies for correction of navigation satellites by determining the required reserves of the characteristic velocity.

REVEALING CHANGES IN THE SPACE OBJECTS MOTION DESCRIBED BY AUTOREGRESSIVE MODELS

The task of revealing changes in the non-functioning Sich-2 spacecraft motion is completed using autoregressive models and time series of TLE elements. The developed method is recomended to use for revealing changes in the movement of space objects during their long-term operation.

A Big-data Inspired Precision Improvement Algorithm for Autonomous Navigation Based on Period Variable Stars

Periodic variable navigation can realize the integration of positioning, attitude determination, timing and other functions. As a novel autonomous navigation method, autonomous management and autonomous operation of spacecraft are the main direction of great significance to reduce the burden of ground measurement and control, to improve the viability of spacecraft, While, the precision of the navigation is very critical for the use of periodic variable navigation for the long range spacecraft. So, we propose a big-data inspired precision improvement algorithm in this paper. Period variable star phase time measurement is used as observation information, and accomplished by the orbital dynamics equation of spacecraft motion. According to the gathering of the sampling data and the sensor data, a self-learning system is trained with the parameters from the nonlinear filtering methods. Based on the Unscented Kalman Filtering, an autonomous navigation algorithm of period variable star is established to realize the navigation and positioning of spacecraft. Under the measurement conditions of a single sampling interval of 0.01s and the measurement precision of period variable star phase observation time of 10− 5s, It can be find that the navigation position determination precision can reach 400m, the speed determination precision can reach 0.3m/s, and the measurement precision can reach 10− 5s with our proposed algorithm.

The concept of a long-term service station to increase the life duration of some satellites or to remove space debris

Every year the concentration of space debris is steadily growing, which significantly complicates the conduct of both modern and future space missions using automatic, and especially manned space vehicles. To date, over 18,000 artificial objects and fragments recorded in near-Earth space. Therefore, the issues of cleansing outer space of space debris of various sizes are quite relevant. In this article, the concept of a multirole station is studied. For the successful design of such a station that could serve spacecraft in space and successfully cope with space debris, it is necessary to consider and necessarily take into account in its further work the ballistic aspects of the orbital motion of spacecraft and the logistics of building such a station. This station could be a maintenance hub for satellites or a space debris collector. Because some satellites have very precious instruments onboard, it could be a loose of money to launch a new satellite if the first one is out of service because of a system failure or a lack of fuel. In this perspective, an orbital maintenance hub would enable to fix the failure (batteries replacement, refueling, mechanical maintenance) and to give to the satellites a new life. In this article, we also study the possibility of using such a hub to remove space debris from their orbit and to collect them until they can be deorbiting.

Problematic issues of spacecraft development for contactless removal of space debris by ion beam

Application of the method of removal of space debris objects by an ion beam - the IBS method - seems to be the most effective for cleaning the GEO region. This is due to the low energy consumption for transfer one object in a safe orbit delta-V≃11 m/s and the possibility of removal several objects by one spacecraft. The service spacecraft for this purpose can be developed on the basis of telsat technology, taking into account its functions. Such a spacecraft should contain an ion beam injector to act on the removed object and an electric propulsion system, which serves to compensate for the ion injector thrust and to control the spacecraft motion both during transportation of the removed object and during interorbital flights. The development of a weakly diverging ion beam injector, determination of the composition and structure of an electric propulsion system, ballistics of the optimal sequence of interorbital flights, and development of an algorithm for controlling a virtual cluster: spacecraft-removed object are discussed in the article. In the majority of works on IBS, commercial ion thrusters are considered as an ion injector and engines. However, ion beams from ion thrusters have a half angle of divergence of 120-150. When propagating in outer space, the ion beam expands irresistibly due to electron thermal effects. As a result, the distance at which the spacecraft must follow the removed object during its transportation for effective impulse transmission to it should be small (estimated, about 7 m). Taking into account the dimensions of both objects and the uncontrolled movement of the removed object relative to its center of mass, the object transportation operation carries noticeable risks. The results of an experimental study of a prototype of an ion injector with an initial half angle of divergence of 20-40 are presented. The use of such an injector would provide a distance between objects of 15–20 m and reduce the risk of their collision. In connection with IBS, various aspects of spacecraft control during transportation of an object with one, two or more engines as part of the electric propulsion system are considered. In the conceptual design of the spacecraft presented in this work, the electric propulsion system contains two stationary plasma thrusters, each of which is mounted with the possibility of independent rotation respect to two mutually perpendicular directions. Spacecraft motion control, including thrust vector control, is carried out by means of the thrusters coordinated rotation. The characteristics of the spacecraft and its subsystems intended for the removal of several objects are conceptually considered. For example, parameters of the mission for the removal of up to 10 objects from the NORAD catalog from the GEO region are given. In the future, using the result obtained, the control of SC of this scheme when IBS implementing in the low-Earth orbit will be considered.

Coupled Relative Orbit and Attitude Control Augmented by the Geomagnetic Lorentz Propulsions

An electrostatically charged spacecraft is subject to the Lorentz force and torque when moving in the geomagnetic field. By active modulation of the surface charge, the induced Lorentz force and torque can be used for orbit and attitude control, respectively. Because of the electromagnetic mechanism, the orbit and attitude motion of such spacecraft are naturally coupled, and the induced Lorentz force and torque are both instantaneously underactuated. Thus, other kinds of actuation are required to render the system fully actuated and controllable. A hybrid control system is then proposed in this paper for Lorentz-augmented spacecraft relative orbit and attitude control. The relative orbit motion is controlled by the combination of the Lorentz force and the thruster-generated control force, and the relative attitude motion is governed by the combination of the Lorentz torque and the magnetic torque. By using the adaptive backstepping control method, a closed-loop control scheme is designed for the hybrid control system to deal with the dynamic coupling, underactuation, unknown external disturbances, and model approximation errors simultaneously. The parameter adaptation laws are derived via the Lyapunov theory to guarantee the stability of the closed-loop system. Optimal distribution laws of the hybrid control inputs are then analytically solved. Finally, numerical examples are simulated in a -perturbed environment to verify the feasibility of the proposed hybrid control system and the validity of the closed-loop control scheme.

Processing of laser altimeter time-of-flight measurements to geodetic coordinates

Laser altimeters are commonly used in planetary research for their high geodetic accuracy. A key procedure in processing of laser altimeter data is the geolocation. In this process, the time-of-flight measurements are converted to coordinates of laser pulse footprints on the surface of the target body. Here, we present a consistent and systematic formulation of three commonly used geolocation models with increasing complexity: static model, spacecraft motion model, pointing aberration model and special relativity model. We show that for small velocities of the spacecraft relative to the target the special relativity model can be reduced to the pointing aberration model without significant loss in the geolocation accuracy. We then discuss the respective accuracies of the proposed models and apply them to time-of-flight measurements from the Mars Orbiter Laser Altimeter (MOLA) onboard the Mars Global Surveyor (MGS) spacecraft and the Mercury Laser Altimeter (MLA) onboard the MErcury Surface, Space ENvironment, GEochemistry and Ranging spacecraft (MESSENGER). While, the archived datasets had not considered the effect of pointing aberration, we demonstrate that a correction due to pointing aberration makes insignificant improvements of 4–5 m laterally and up to ± 3 cm radially for MOLA profiles, these figures enormously increase to up to about 150 m laterally and ± 25 m radially when applied to the MLA orbital profiles.

Periodic attitude motions along planar orbits in the elliptic restricted three-body problem

A way to improve the accuracy of the three-body problem model is taking into account the eccentricity of primary attractors. Elliptic Restricted Three-Body Problem (ER3BP) is a model for studying spacecraft trajectory within the three-body problem such that the orbital eccentricity of primaries is reflected in it. As the principal cause of perturbation in the employed dynamical model, the primaries eccentricity changes the structure of orbits compared to the ideal Circular Restricted Three-Body Problem (CR3BP). It also changes the attitude behavior of a spacecraft revolving along periodic orbits in this regime. In this paper, the coupled orbit-attitude dynamics of a spacecraft in the ER3BP are exploited to find precise periodic solutions as the spacecraft is considered to be in planar orbits around Lagrangian points and Distant Retrograde Orbits (DRO). Periodic solutions are repetitious behaviors in which spacecraft whole dynamics are repeated periodically, these periodic behaviors are the main interest of this study because they are beneficial for future mission designs and allow delineation of the system’s governing dynamics. Previous studies laid the foundation for spacecraft stability analysis or studying pitch motion of spacecraft in the ER3BP regime. While in this paper, at first, initial guesses for correction algorithms were derived through verified search methods, then correction algorithms were used to refine calculated orbit-attitude periodic behaviors. Periodic orbits and full periodic solutions are portrayed and compared to previous studies and simpler models. Natural periodic solutions are valuable information eventuate in the longer functional lifetime of spacecraft. Since the problem assumption considered in this paper is much closer to real mission conditions, these results may be the means to use natural bounded motions in the actual operational environment.

Расчет аэродинамического момента в задачах математического моделирования вращательного движения транспортных грузовых кораблей Прогресс

We describe the method of representation the aerodynamic torque in problems of mathematical modeling uncontrolled attitude motion of spacecraft Progress. The components of the torque are calculated in the nodes of the grid on the two-dimensional sphere. The calculations were based on the detailed geometric model of the spacecraft external surface. The found arrays of torque data were approximated by the segments of the Fourier series by spherical functions up to the order (8,8) inclusive. These approximations are used in the equations of attitude motion of the spacecraft. We describe the use of such equations in reconstruction of the motion of spacecraft Progress MS-07, MS-08 in the gravitational orientation mode by measurements of their angular rates and in the prognosis of electric charge from the solar arrays of the spacecraft in the mode of its single-axis solar orientation.