Active Spacecraft Research Articles

Online tree-based planning for active spacecraft fault estimation and collision avoidance.

Autonomous robots operating in uncertain or hazardous environments subject to state safety constraints must be able to identify and isolate faulty components in a time-optimal manner. When the underlying fault is ambiguous and intertwined with the robot's state estimation, motion plans that discriminate between simultaneous actuator and sensor faults are necessary. However, the coupled fault mode and physical state uncertainty creates a constrained optimization problem that is challenging to solve with existing methods. We combined belief-space tree search, marginalized filtering, and concentration inequalities in our method, safe fault estimation via active sensing tree search (s-FEAST), a planner that actively diagnoses system faults by selecting actions that give the most informative observations while simultaneously enforcing probabilistic state constraints. We justify this approach with theoretical analysis showing s-FEAST's convergence to optimal policies. Using our robotic spacecraft simulator, we experimentally validated s-FEAST by safely and successfully performing fault estimation while on a collision course with a model comet. These results were further validated through extensive numerical simulations demonstrating s-FEAST's performance.

The ion beam interceptor concept for a space station collision avoidance with space debris

Space debris poses a great threat to space stations. When a collision hazard is detected, the space station's orbit should be adjusted. An alternative approach is to apply collision avoidance measures that consist in creating an external impact on the space debris object to change its orbit and increase the miss distance slightly. Such an impact can be carried out using a ground-based or orbital laser, through the use of nano-tugs thrusters, or by placing a cloud of particles in the path of the object. This study proposes to use the plume of a thruster of an active spacecraft called an Ion Beam Interceptor. When a potential collision is detected, the interceptor undocks from the space station, approaches the space debris object and directs an ion beam, which is generated by the spacecraft's thrusters, at it. Hitting into the surface of the object, the particles of the ion beam exert a force effect on it, which leads to the required change in the orbit. After the completion of the maneuver, the interceptor returns to the space station for refueling and servicing. The objective of the study is a preliminary assessment of the feasibility and effectiveness of the proposed approach of protecting a space station from space debris impact using an active spacecraft. A mission profile of the space station protection is proposed. A series of numerical simulations are performed to determine mission parameters and fuel costs. Based on an analysis of the orbits of space objects located near the space station, the maximum angle between the orbital planes of the dangerous object and the station are determined at which it is advisable to use the Ion Beam Interceptor. The developed concept also allows to implement just-in-time collision avoidance measures for space debris objects whose orbital planes are close to the plane of the space station.

Effects of phase angle and sensor properties on on-orbit debris detection using commercial star trackers

The recent proliferation of resident space objects (RSOs) in low Earth orbit (LEO) threatens the sustainability of space as a resource and requires persistent monitoring to avoid collisions involving valuable space assets. State-of-the-art ground-based space surveillance techniques, due to their susceptibility to atmosphere, weather, and lighting conditions, tend to focus on RSOs with characteristic length greater than 10 cm or 1 dm. Consequently, millions of smaller LEO RSOs remain untracked by ground-based methods, which reduces overall space situational awareness. Onboard satellite sensors offer a space-based method for tracking RSOs. Prior research has investigated the feasibility of using commercial star trackers (CSTs) — optical sensors prevalent on most active spacecraft — to observe, detect, and estimate the position and velocity of RSOs larger than 10 cm. In a recent effort, we expanded on these feasibility studies by assessing the capabilities of CSTs to detect debris particles smaller than 10 cm in characteristic length. We modeled the particles as Lambertian spheres with zero phase angle and ten percent reflectivity and found that typical CSTs can detect properly illuminated debris of characteristic length between 1 cm and 10 cm even at distances of tens of kilometers. More sensitive CSTs can characterize decimeter-scale RSOs hundreds of kilometers away; alternatively, they can track centimeter-scale and smaller RSOs at closer distances. In this paper, we summarize these key results and extend our previous study by relaxing its zero-phase-angle assumption and characterizing the effect of Sun-RSO-CST phase angle on debris detection range. We identify a number of representative CSTs with publicly available optical characteristics and consider the effects of properties such as pixel pitch, focal length, aperture diameter, and field of view (FOV) on the capability of each CST to detect debris at a given distance and relative velocity (in the form of streak speed). We find that, for debris particles modeled as diffuse Lambertian spheres, Sun-RSO-CST phase angles as high as 57° result in no more than 20% reduction to the useful RSO-CST detection range. In addition, we find that pixel pitch and focal length, rather than aperture diameter, tend to determine the capability of a given CST to resolve two distinct RSOs. Furthermore, streak speed may serve as a stronger limiting factor for detection of smaller debris particles than for larger ones. Despite these limitations, the overall results indicate that CSTs have the potential to substantially enhance space-based debris detection capability.

Deep operator and SSA collaboration for space sustainability

The space data association (SDA) has been providing reliable flight safety products for approximately 30 spacecraft operators for 14 years now. The service provides conjunction warnings and operator points of contact for around 700 spacecraft occupying all orbital regimes. The SDA's Space Data Center (SDC), built by AGI and maintained and operated by COMSPOC Corporation, has a proven track record of providing high availability space traffic coordination (STC) products since becoming operational on 15 July 2010.Earlier this year, the SDA, and its chief technical consultant, COMSPOC Corporation, supported the U.S. Department of Commerce (DOC) Geosynchronous Earth Orbit (GEO) and Middle Earth Orbit (MEO) Pilot project by providing comprehensively fused orbit solutions, ten-day orbit ephemeris and covariance predictions, and smoothed reference ephemerides for 100 active spacecraft. Most of these spacecraft are operated by SDA members and participants, allowing the operators to collaboratively contribute their maneuver plans, GPS NavSol measurements, active ranging, and passive RF observations, and authoritative spacecraft dimensions to the DOC Pilot project. For its part, COMSPOC employed its Space Situational Awareness (SSA) Software Suite (SSS)to comprehensively fuse this diverse set of spacecraft operator observations with COMSPOC's own observations from its global network of optical sensors.The first phase of this collaborative data fusion required the establishment of accounts, data connectivity, file transfer methods, and sensor calibration. This required about one month of technical interchange, provision of operator sensor locations and specifics, and COMSPOC SSS operator calibration of those sensors for each spacecraft.But once the data flows, sensor calibrations, and maneuver readers were completed, the second phase drew upon a nearly continuous stream of fused observations and maneuvers to yield accurate and timely predictive ephemerides and covariance time histories. These data products are well-suited to the collision avoidance problem.

Fast Spacecraft Charging Mitigation With Plasma Contactors During High‐Power Electron Beam Emission in the GEO Environment

AbstractThe emission of artificial electron beams causes an active spacecraft charging effect that poses challenges for electron beam sounding experiments. Plasma contactors have been suggested as a potential solution to this issue based on theoretical and simulation studies, but the impact of these contactors on spacecraft potentials in complex space environments remains unexplored. This paper investigates the fast active charging effect of spacecraft under the action of plasma contactors in the geostationary Earth orbit environment by using a two‐dimensional Particle‐in‐Cell model. We analyze and compare the effects of the “ion emission” process of the plasma contactor and the “electron collection” process of the background plasma on the spacecraft potential at different beam and environment parameters, as well as the coupling effects when both processes occur simultaneously. Our findings indicate that the “ion emission” and “electron collection” processes are the primary factors in determining the spacecraft potential in relatively thin and dense plasma environments, respectively. When the two processes occur together, the contactor plasma cloud increases the “electron collection” return current by expanding the range of the positive potential area around the spacecraft, while the dense background plasma mediates and provides the return current to ensure that the plasma contactor can effectively suppress the active charging effect even after the ion sheath is present. These results suggest that a reasonable adjustment of the contactor parameters based on the relationship between the electron beam and the background environment can effectively extend the normal emission time of the electron beam.

Planar spacecraft trajectories in the Didymos–Dimorphos binary asteroid system

The dynamics of a spacecraft in the vicinity of the Didymos and Dimorphos binary system is studied, with specific emphasis on the bounded orbits that exist in such a perturbed gravitational environment. The binary environment is approximated with gradually increasing fidelity, starting with a model that simulates Didymos with an oblate spheroid and Dimorphos with a triaxial ellipsoid. Within these simplifying assumptions, we compute all the families of planar periodic orbits and we study the deviations that arise to them under more realistic perturbations that include the actual shape of Didymos, its rotation, the solar gravity and radiation pressure. To establish the robustness of the different orbital configurations, we perform a classification based on planar dynamical maps and connect the stability regions with the periodic motion. A subset of retrograde orbits around Didymos and a particular type of outer orbits with double multiplicity seem to show long-term stability and robustness, so they can be assumed as safe orbits. On the contrary, orbits around Dimorphos, although stable in the basic model, become strongly unstable in our more realistic model. Our approach is based only on natural dynamics and no active spacecraft control is applied, thus providing a useful methodology in finding ‘safe’ orbits for a spacecraft in the vicinity of a binary asteroid system.

Cooperative Guidance Strategy for Active Spacecraft Protection from a Homing Interceptor via Deep Reinforcement Learning

The cooperative active defense guidance problem for a spacecraft with active defense is investigated in this paper. An engagement between a spacecraft, an active defense vehicle, and an interceptor is considered, where the target spacecraft with active defense will attempt to evade the interceptor. Prior knowledge uncertainty and observation noise are taken into account simultaneously, which are vital for traditional guidance strategies such as the differential-game-based guidance method. In this set, we propose an intelligent cooperative active defense (ICAAI) guidance strategy based on deep reinforcement learning. ICAAI effectively coordinates defender and target maneuvers to achieve successful evasion with less prior knowledge and observational noise. Furthermore, we introduce an efficient and stable convergence (ESC) training approach employing reward shaping and curriculum learning to tackle the sparse reward problem in ICAAI training. Numerical experiments are included to demonstrate ICAAI’s real-time performance, convergence, adaptiveness, and robustness through the learning process and Monte Carlo simulations. The learning process showcases improved convergence efficiency with ESC, while simulation results illustrate ICAAI’s enhanced robustness and adaptiveness compared to optimal guidance laws.

Active Spacecraft Potential Control in the MMS Mission: Results From Six Years in Orbit

The four spacecraft of the NASA Magnetospheric Multiscale (MMS) mission carry instruments to reduce the positive potential by means of indium ion beams. Since the start of the nominal mission in September 2015 and until the end of 2021, the instruments active spacecraft potential control (ASPOC) have been actively operating for more than 16 000 h at a nominal emission current of 20 <inline-formula xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"> <tex-math notation="LaTeX">$\mu$</tex-math> </inline-formula> A per spacecraft. Based on data from more than six years in orbit with more than 50 000 h in regions of scientific interest, statistical results regarding the potential’s interdependencies with ambient plasma were obtained. This article reports on the derivation of the photo electron energy spectrum from the correlation between the potential and the plasma data obtained by the fast plasma instrument with and without controlled potential. Finally, the time constants during the relaxation of the controlled potential when the active control instrument is turned off, if measured at high time resolution, allow to estimate the electric capacitance of the spacecraft system.

Determining the Effective Space Debris Attitude Motion Modes for Ion-Beam-Assisted Transportation

Contactless space debris transportation systems based on the use of an ion beam are a promising direction among active space debris removal systems. The magnitude of the force generated by the ion beam depends on the orientation of the space debris object in the ion beam created by an active spacecraft. During the ion transport process, the space debris object can oscillate or rotate, which leads to a change in the generated force. The aim of this study is to develop an algorithm for determining the attitude motion mode of space debris, which is most favorable for its contactless transportation. The time-averaged generated ion force is used as the performance index. A simplified system of equations describing the spatial motion of a cylindrical axisymmetric object in a circular orbit is obtained. A generalized energy integral is found. An algorithm for determining the most favorable angular motion mode of a space debris object is developed. A numerical study of a cylindrical space debris object attitude motion under the action of an ion torque is carried out. The favorable attitude motion mode is determined, and the average forces in the most favorable and unfavorable modes are compared.

Precision of Spacecraft Doppler Tracking at Low Signal‐To‐Noise Ratios

AbstractThe signal‐to‐noise ratio received at Earth is typically larger than 10 dB‐Hz for radio tracking of a spacecraft in deep space, allowing a phase‐locked loop to execute reliable carrier tracking. Recently, missions have been proposed to utilize low‐gain antennas for Doppler tracking where the signal‐to‐noise ratio may be at the single‐digit (dB‐Hz) level. In this work, we discuss spacecraft Doppler tracking at these low signal‐to‐noise ratios through an analysis of thermal noise on the radio link, results from ground testing with the Deep Space Network, and demonstrations with an active spacecraft. We show that by utilizing an open‐loop receiver to capture the carrier signal and by applying post‐processing techniques, radio data with the signal‐to‐noise ratio as low as 4 dB‐Hz can be used to derive Doppler data with precision sufficient for tracking a spacecraft in deep space.

Визначення сили впливу іонного променя на об’єкт космічного сміття за контурами його зображень з використанням глибокого навчання

The goal of this article is to develop an effective image preprocessing algorithm and a neural network model for determining the force to be transmitted to a space debris object (SDO) for its non-contact deorbit. In the development and study of the algorithm, use was made of methods of theoretical mechanics, machine learning, computer vision, and computer simulation. The force is determined using a photo taken by an onboard camera. To increase the efficiency of the neural network, an algorithm was developed for feature recognition by the SDO edge in the photo. The algorithm, on the one hand, selects a sufficient number of features to describe the properties of the figure and, on the other hand, significantly reduces the amount of data at the neural network input. A dataset with the features and corresponding reference force values was created for model training. A neural network model was developed to determine the force to be exerted on a SDO from the SDO features. The model was tested using a set of eighteen calculated cases to determine the effectiveness, accuracy, and speed of the algorithm. The proposed algorithm was compared with two existing ones: the method of central projections onto an auxiliary plane and the multilayered neural network model that calculates the force using the SDO orientation parameters. The comparison was performed using the root mean square error, the maximum absolute error, and the maximum relative error. The test results are presented as tables and graphs. The proposed approach makes it possible to develop a system of SDO non-contact removal that does not need to determine the exact relative position and orientation with respect to the active spacecraft. Instead, the algorithm uses camera-taken photos, from which the features necessary for calculation are extracted. This makes it possible to reduce the requirements for its computing elements, to abandon sensors for determining the relative position and orientation, and to reduce the cost of the system.

Simultaneous design of passive and active spacecraft attitude control using black-box optimization

This paper investigates the simultaneous design of active attitude control and passive attitude compensation mechanism for a spacecraft to satisfy practically-motivated mission objectives and constraints. The expressions of these fitness-related metrics with respect to the design variables are not analytically available, due to the nontrivial interactions between the spacecraft components and the interactions with the environment. Thus, such functions can only be approximately learned from data derived from simulations. We approach this difficult design problem using a black-box optimization (BBO)-based approach, which combines learning and optimizing the objective and constraint functions by design of experiments. The proposed BBO-based approach is assessed in the context of a 3U CubeSat system design with both a passive magnetic attitude compensation and an active reaction wheel-based control, tested on a simulator considering orbital and environmental dynamics. Simulation results and statistical tests compared to other design methods show the capability of the BBO-based approach to provide a design with the best tracking performance while at the same time satisfying ground station communication requirements and power budget.

KESSYM: A stochastic orbital debris model for evaluation of Kessler Syndrome risks and mitigations

Mankind’s productive use of the low Earth orbit (LEO), from 400-2,000km in altitude, is at risk from increasing counts of debris objects and derelict satellites, which pose collision risks to active spacecraft. Of particular concern to space agencies and industry is the Kessler Syndrome (KS), which is the term for a hypothetical collapse scenario in which collisions between debris and satellites cause more debris, causing a destructive cascade that leaves the orbital environment unusable. In order to better understand this KS tipping point, the KESSYM model has been developed as a stochastic simulation of all the objects in the LEO. This model provides a forecast for the evolution of the orbital environment into the future, including the expected year, if any, that the KS collapse occurs. KESSYM allows for certain risks, such as war or terrorism in space, solar flares, or unconstrained exploitation of the space resources to be analyzed alongside KS mitigation measures, such as the hardening of spacecraft against debris, avoidance of collisions, removal of debris, and effective regulation. The conclusions drawn from the KESSYM simulation are that the KS is almost an inevitability within 200-250 years of today’s date, but can be delayed or avoided altogether if action is taken.

Active space debris removal by ion multi-beam shepherd spacecraft

Contactless transportation of a passive object by an ion beam generated by an active spacecraft's thruster is a promising way of space debris removal. The study proposes a modification of the ion beam shepherd concept using an array of impulse transfer thrusters. The multi-beam scheme provides significant increase the generated ion force, which reduces a space debris object deorbit time, thereby reducing the probability of its collision with other orbital objects. The aim of the work is to study the possibility of using several ion beams for effective contactless transportation of a space debris object. Controlling the angular motion of the object permits orientation of the object to maximize the generated ion force. The article proposes a control law for impulse transfer thrusters, which is based on the calculation of the energy of the object's unperturbed motion, and two control schemes that implement this law: (1) relay control, which implies turning the thruster on and off, and (2) ion beam axis direction control for one of the thrusters. For a space debris object of a cylindrical shape, a comparison of the time and fuel costs required for deorbiting the object using one, two, and three impulse transfer thrusters is made. It is shown that the addition of thrusters significantly reduces the descent time, but has little effect on the mass of required fuel. For the case of two engines, the best angle of ion beams axes direction and the most preferable angular motion mode are determined. Numerical simulation of space debris removal is carried out for the case when the space debris object is in the least favorable mode of angular motion and the proposed control schemes are used. The ion beam direction control scheme showed better results than the relay scheme. The use of multiple ion beams opens up new possibilities for creating new ion beam shepherd spacecraft design and developing new control methods and laws.

Detection and characterisation of unknown maneuvers in spacecraft

Space debris detection and characterisation is an important and pressing field, as the amount of space debris is growing due to the increase in spacecraft launches and increasing interests for space uses. Space agencies like NASA and ESA keep track of over 5,000 spacecraft and 20,000 spacecraft-related debris objects currently orbiting the Earth. An active spacecraft may undergo trajectory changes through an impulsive maneuver, or an inactive spacecraft may change its trajectory because of uncontrolled explosions due to residual fuel without warning. These pose risks to operational satellites. In addition, measurement data from tracking and detection consists of uncertainties, causing maneuver detection and characterisation methods to become computationally expensive.Thus, the goal of this work was to develop a quick, reliable, computationally inexpensive method for spacecraft maneuver detection and characterisation to be used by satellite operators and assist in collision avoidance measures. This paper presents such a technique, which is applied to simulated measurement data of an orbiting spacecraft taken from one ground site. With these objectives, a method using a modified Extended Kalman Filter (EKF) with covariance inflation from previous studies was explored and simplified. A new parameter definition is presented for maneuver detection and characterisation, named the Kaderali-parameter, or K, with focus on transverse apogee and perigee maneuvers. A relationship was then found between the parameter and the magnitude of the impulsive artificial maneuver. The method was tested in several different scenarios, varying Δv magnitude and scale, Δv direction, maneuver orbital location, orbit geometry, noise covariance, detection parameter threshold, state covariance inflation threshold, and maneuver duration.In general, it was found that as the Δv magnitude increases, the magnitude of the detection parameter also increases. This relationship was found to be a nonlinear logarithmic relationship through best fit curve testing. Possible extensions of this work in the future include: testing multi-maneuver scenarios, multi-sensor data collection and fusion, break-up or berthing events, rendezvous or mating events, and matlab EKF protocol implementation.

ПРИМЕНЕНИЕ ГЕНЕТИЧЕСКОГО АЛГОРИТМА ДЛЯ ОПТИМИЗАЦИИ ДВУХИМПУЛЬСНОГО ПЕРЕЛЁТА В ЗАДАЧЕ О ВСТРЕЧЕ ДВУХ КОСМИЧЕСКИХ АППАРАТОВ НА ОРБИТЕ МАРСА

In this research, we delve into the application of genetic algorithms for optimizing the twoimpulse rendezvous maneuver of an active spacecraft with a passive spacecraft, situated at orbits with both higher and lower altitudes relative to the passive counterpart. Assumedly, the passive spacecraft resides in a circular Mars orbit at an altitude of 300 km, with defined orbital parameters. Through the employment of genetic algorithms, our primary objective is the optimization of rendezvous parameters, specifically focusing on the rendezvous duration and the initial velocity vector of the active spacecraft within the areocentric equatorial system, while simultaneously striving to minimize the requisite energy expenditure for the rendezvous. In both cases under consideration, we maintain an initial separation of 20 km between the spacecraft. In each case, we track the shift in the relative position vector of the active spacecraft with respect to the passive spacecraft. Our findings reveal that trajectories approximating 75 minutes in duration yield a global minimum energy requirement. Conversely, shorter trajectories necessitate a marginally larger energy expenditure, thereby embodying a trade-off between flight duration and energy expenditure. Interestingly, lengthening the duration of the rendezvous maneuver does not invariably result in a lower energy requirement. Moreover, a thorough numerical analysis uncovers a remarkable similarity in the optimization outcomes across both cases, with the rendezvous trajectory direction serving as the principal differential factor. This investigation furnishes crucial numerical insights into the optimization process of two-impulse rendezvous maneuvers, thereby substantiating the practical utility of genetic algorithms in this domain.

Development and coupling test of active spacecraft potential control – Next generation (ASPOC-NG)

The active spacecraft potential control (ASPOC) system developed in the 1990s emits positive ions to neutralise the spacecraft potential, such as used in several missions like Geotail, Equator-S, Cluster, Doublestar and MMS. With the experience gained, the next generation of the active spacecraft potential control (ASPOC-NG) instrument has been developed over the last three years. Thereby, three emission technologies were tested including Liquid Metal Ion Source (LMIS), Liquid Metal Electron Source (LMES) and Solid Metal Electron Source (SMES). The development of the emitter module by FOTEC and the corresponding electronics control unit by IWF is presented. Optimisations were carried out with the focus on the reduction of mass and power consumption to comply with the requirements of future scientific missions. Coupling tests of the modules and the electronics control unit were performed including range, accuracy and lifetime tests. Both ASPOC-NG instruments for positive and negative charge compensation and their performance values show excellent results.

The Rendezvous Mission of Two in-Orbit Sattelites Using Numerical Simulation

In this work, a rendezvous of two satellites is considered. The active satellite approaches the passive satellite for a remote partial recharge of the passive spacecraft’s dead battery. In the first case, the active and passive spacecraft are in geocentric circular orbits, having an altitude of 499.9 km and 500 km, respectively. It is assumed that the rendezvous is unperturbed and the rendezvous plane coincides with the plane of the passive satellite’s orbit. Close-range guidance starts as soon as the sensor on the active spacecraft recognizes the passive spacecraft. A method implementing the line of sight is used for the close-range guidance process. A mathematical model of the relative motion is formed and numerically solved to investigate the rendezvous parameters against time. For solving the model, different values of 499.9, 499.8, 499.7 km are considered as the initial orbit altitude of the active spacecraft, assuming that the angle between state vectors of the spacecraft has a small value of 0.01 degrees. The results show that the active spacecraft with an initial altitude of 499.7 km approaches the passive spacecraft to a distance of 80 meters in less than 50 seconds, where a final velocity impulse is needed to maintain this distance.

Finite-Time Prescribed Performance Control for Space Circumnavigation Mission With Input Constraints and Measurement Uncertainties

This article presents a 6-DOF attitude-orbit synchronous control problem for the space circumnavigation (SCN) mission with parameter uncertainties, time-varying uncertainties, and input constraints. In particular, from the perspective of engineering application, time-varying measurement uncertainties are taken into account of the 6-DOF attitude–orbit coupling kinematics and dynamics, and the analytical solution of the desired attitude is derived based on the measured relative orbit information with measurement uncertainties. To drive the active spacecraft approach to the faulty target safely, a time-varying exponential prescribed convergence boundary is introduced into the sliding surface. A finite-time disturbance observer is involved in equivalent tracking errors for compensating the mismatched uncertainties. In addition, an auxiliary system is designed to overcome the instability danger caused by input constraints. The stability of the controlled system is discussed in the nonautonomous finite-time stable framework, which is proved via Lyapunov analysis that the attitude-orbit tracking errors converge to the equilibrium within finite time. The simulation experiment with mismatched uncertainties and prescribed constraints shows the superiority of the designed control scheme.

An analytical review for the power system of a spacecraft

Electrical bus is needed for the operation of all active spacecraft systems and subsystems. In the spacecraft mechanism, the battery charging system and the bus storage continued with bus control, the conditioning and bus distribution is an essential part for keeping the spacecraft in working mode. In this paper, the discussion for electrical power system (EPS) analysis in the spacecraft body will be elaborated from the previous research papers on this subject covering the topics of solar array, solar array drive, over current and over voltage protection, battery charging and discharging, bus control, conversion.