Spacecraft Maneuvers Research Articles

Tracking maneuvering spacecraft with filter-through approaches using interacting multiple models

When spacecraft maneuver unknowingly, traditional orbit determination filters diverge. This paper postulates using Interacting Multiple Models and covariance inflation methods to filter-through unknown maneuvers. This work proposes accurate methods for real-time tracking of a non-cooperative, maneuvering spacecraft and compares the performance to traditional techniques. Results show that a filter-through Interacting Multiple Model orbit determination filter can converge on a post-maneuver orbit in real-time with similar performance to off-line Initial Orbit Determination approaches.

Compliance of the Italian satellites in low Earth orbit with the end-of-life disposal guidelines for Space Debris Mitigation and ranking of their long-term criticality for the environment

As of mid-2014, nearly 50 years since the launch of the first satellite, Italy had placed in low Earth orbit 29 objects: 27 payloads, 1 rocket body and 1 mission related object. 19 were yet in space: the IRIS rocket body and 18 payloads, 4 of which, belonging to the COSMO-SkyMed constellation, still operational and maneuverable. 16 objects had been deployed in space before the approval, in 2002, of the IADC Space Debris Mitigation Guidelines and 1 further payload had been launched before the ASI signature of the European Code of Conduct for Space Debris Mitigation, in 2005. While no object had been yet maneuvered to reduce its residual lifetime, due to the operational orbits chosen and area-to-mass ratios, 16 of them had decayed or were predicted to reenter in less than 25 years after mission completion, in agreement with current disposal recommendations. This corresponded to a compliance of 64% over 50 years, of 59% for the objects placed in orbit before the ASI signature of the European Code of Conduct, and of 75% for those launched afterwards, excluding the 4 maneuverable spacecraft still functional. Concerning the risk on the ground associated with uncontrolled reentries, just one satellite decayed in 2003 had a mass greater than 600kg and a casualty expectancy in excess of 10−4. For it, timely reentry predictions and alert time windows had been provided to the countries overflown. In order to evaluate the potential long-term detrimental effects on the environment of the abandoned or unmaneuverable objects, a new ranking index, also useful for active debris removal priority listing, was developed and applied. It is worth noting that all the 14 objects residing or descending below 1000km exhibited an overall ranking index equivalent to just 8% of an average abandoned intact object in a 800km sun-synchronous orbit.

Adaptive Fuzzy PD+ Control for Attitude Maneuver of Rigid Spacecraft

AbstractIn this paper, an adaptive fuzzy PD+ controller is proposed for the attitude maneuver of rigid spacecraft. The novel controller adjusts the gains of the PD+ attitude controller online according to attitude errors and angular velocity errors during the maneuver procedure. Therefore, quick response and avoidance of actuator saturation can be achieved simultaneously. Furthermore, the adaptation mechanism is designed, based on Lyapunov theory, to guarantee the stability of the closed‐loop system. To achieve good performance of the closed‐loop system under the constraint of actuator saturation, controller parameter optimization is developed on the basis of a genetic algorithm. Simulation results show that the transient performance and robustness against parametric uncertainty and environmental disturbance of the adaptive fuzzy PD+ controller are better than those of a constant PD+ controller.

Adaptive Reachability Analysis to Achieve Mission Objectives in Strongly Non-Keplerian Systems

Adaptive numerical reachability analysis is studied as a tool for coping with sensitive, strongly non-Keplerian dynamics when planning impulsive spacecraft maneuvers in pursuit of mission objectives at small-body systems. The concept of an objective reachability map is developed beginning from analytical results for the Kepler problem and is shown to become intricately structured for the environments of asteroid Itokawa and Martian satellite Phobos. An existing method for efficient map computation via heuristic refinement, originally applied to chart safety outcomes on a planar maneuver domain, is extended for application to spherical domains and two types of non-Boolean science objectives defined for target locations on the body surface: cumulative observation time and surface payload deployment accuracy. In addition to its potential as a tool for preliminary mission design and the general study of small-body dynamics, this reachability analysis method is also examined and discussed as a resource for receding-horizon planning onboard an autonomous spacecraft.

Optimal propellantless rendez-vous using differential drag

Optimization of fuel consumption is a key driver in the design of spacecraft maneuvers. For this reason, growing interest in propellant-free maneuvers is observed in the literature. Because it allows us to turn the often-undesired drag perturbation into a control force for relative motion, differential drag is among the most promising propellantless techniques for low-Earth orbiting satellites. An optimal control approach to the problem of orbital rendez-vous using differential drag is proposed in this paper. Thanks to the scheduling of a reference maneuver by means of a direct transcription, the method is flexible in terms of cost function and can easily account for constraints of various nature. Considerations on the practical realization of differential-drag-based maneuvers are also provided. The developments are illustrated by means of high-fidelity simulations including coupled 6-degree-of-freedom simulations and an advanced aerodynamic model.

Autonomous star sensor ASTRO APS: flight experience on Alphasat

Jena-Optronik GmbH, located in Jena/Germany, has profound experience in designing and manufacturing star trackers since the early 80s. Today the company has a worldwide leading position in supplying geo-stationary and Earth observation satellites with robust and reliable star tracker systems. In the first decade of the new century Jena-Optronik received a development contract (17317/2003/F/WE) from the European Space Agency to establish the technologically challenging elements for which advanced star tracker technologies as CMOS Active Pixel Sensors were being introduced or were considered strategic. This activity was performed in the frame of the Alphabus large platform pre-development lead by ESA and the industrial Joint Project Team consisting of Astrium (now Airbus Defence and Space), Thales Alenia Space and CNES (Centre national d’etudes spatiales). The new autonomous star tracker, ASTRO APS (Active Pixel Sensor), extends the Jena-Optronik Astro-series CCD-based star tracker products taken the full benefit of the CMOS APS technology. ASTRO APS is a fully autonomous compact star tracker carrying either the space-qualified radiation hard STAR1000 or the HAS2 APS detectors. The star tracker is one of four Technology Demonstration Payloads (TDP6) carried by Alphasat as hosted payload in the frame of a successful Private Public Partnership between ESA and Inmarsat who owns and operates the satellite as part of its geo-stationary communication satellites fleet. TDP6 supports also directly TDP1, a Laser Communication Terminal, for fine pointing tasks. Alphasat was flawlessly brought in orbit at the end of July 2013 by a European Ariane 5 launcher. Only a few hours after launch the star tracker received its switch ON command and acquired nominally within 6 s the inertial 3-axes attitude. In the following days of the early in-orbit operations of Alphasat the TDP6 unit tracked reliably all the spacecraft maneuvers including the 0.1 and 0.2°/s spin stabilization for Sun pointing, all of the apogee engine thrusts, Moon field of view transits and recovered to stable tracking after several Earth and Sun blindings before the spacecraft entered a preliminary Earth pointing in a nominal geo-stationary attitude. The Jena-Optronik TDP6 operation center received daily the star tracker status and attitude data. The huge amount of acquired raw data has been evaluated to characterize the ASTRO APS (STAR1000) star tracker in-orbit performance. The paper will present in detail these data processing activities and will show the extraordinary good results. Due to the diverse transfer orbit satellite operations the key performance star tracker data like attitude random noise, single star noise, star brightness measurement, baffle Sun exclusion angle, temperature control, etc., could be derived and have been compared to the ground based laboratory and field measurements. The ultimate performance parameters achieved and verified as well as the lessons learned from the comparison to the ground test data are summarized in the conclusion of the paper.

Postmaneuver Collision Probability Estimation Using Sparse Polynomial Chaos Expansions

This paper describes the use of polynomial chaos expansions to approximate the probability of a collision between two satellites after at least one performs a translation maneuver. Polynomial chaos provides a computationally efficient means to generate an approximate solution to a stochastic differential equation without introducing any assumptions on the a posteriori distribution. The stochastic solution then allows for orbit state uncertainty propagation. For the maneuvering spacecraft in the presented scenarios, the polynomial chaos expansion is sparse, allowing for the use of compressive sampling methods to improve solution tractability. This paper first demonstrates the use of these techniques for possible intraformation collisions for the Magnetospheric Multi-Scale mission. The techniques are then applied to a potential collision with debris in low Earth orbit. Results demonstrate that these polynomial chaos-based methods provide a Monte Carlo-like estimate of the collision probability, including adjustments for a spacecraft shape model, with only minutes of computation cost required for scenarios with a probability of collision as low as . A graphics processing unit implementation of the polynomial chaos expansion analysis further reduces the computation time for the scenarios presented.

Semi-physical experimental study of adaptive disturbance rejection filter approach for vibration control of a flexible spacecraft

Adaptive disturbance rejection filter (ADRF) approach, an active vibration control for flexible spacecraft, is developed and its parameters are identified using a closed-loop system method. Particularly, semi-physical experimental study of a maneuvering spacecraft equipped with a flexible manipulator is implemented to demonstrate the improvement of the attitude control performance. The control system consists of two main parts: (1) optimal input shaper is applied to the flexible manipulator to suppress the vibration and alleviate the effect of the vibration on the spacecraft; (2) ADRF is presented to mitigate the equivalent periodic disturbance caused by the nonlinear coupling between the spacecraft motion and the manipulator vibration. The disturbance frequency adaptation tracks the frequency of the periodic disturbance using adaptive notch filter, and the obtained frequency is then fed to disturbance rejection filter. The ADRF approach can provide optimal sets of gains based on the characteristics of the system. The semi-physical experimental results show that the frequency tracking method works quite well to be used in a real-time controller. The results also indicate that the ADRF can provide significant vibration reduction and improve the attitude control accuracy of the spacecraft body.

Conceptual design of a low-cost real-time hardware-in-the-loop simulator for satellite attitude control system

Integration of flight hardware with real-time simulation increases satellite attitude control system (ACS) reliability by providing greater test coverage through end-to-end testing in a realistic test environment. In this paper, a compound hardware and software simulator has been designed for evaluation and testing of the spacecraft ACS, placing emphasis on the real-time hardware-in-loop (RTHIL) architecture. The environment comprises both real-time control and data acquisition applications on a network of ACS hardware, the MATLAB Real-Time Workshop, and a PCI device to join the hardware and software units. Moreover, a graphical 3D simulator has been designed, enabling the designers and researchers to intuitively analyze the quality of spacecraft maneuvers. Finally, an ad hoc attitude-stabilizing control law for magnetic actuated satellites has been proposed and implemented in the proposed environment and the efficiency, correctness, and robustness of this control law have been verified using RTHIL simulation results.

Identification of scale factor of a pendulum compensating accelerometer during an orbital spacecraft mission

The paper deals with a crucial task of estimating the metrological characteristics of a pendulum accelerometer which change under the influence of perturbations in the flight of spacecraft. A method is developed that makes it possible to estimate the value of the scale factor of a compensating pendulum accelerometer in the system of spacecraft maneuver control through the use of grapho-analytical and indirect methods of diagnosing first and second order systems. The proposed method makes possible automatic identification of the scale factor in the absence of external forces of non-gravitational origin. The essence of the method is based on the signal-extracting impact on the actuator accelerometer which causes a natural deviation of the sensing element, and the analysis of the transient response of the accelerometer output signal. Numerical studies show that the accuracy of estimating the scale factor in accordance with the proposed method depends on three factors: the output of the measuring noise, the attenuation coefficient of the pendulum accelerometer sensor, and the moments of the external forces acting on the sensitive axis of the accelerometer in the identification process. The results of simulation led to the conclusion about the possibility of applying this method in solving the problem of self-control of the characteristics of a pendulum accelerometer. It is shown that an increase in the accuracy of estimating the scale factor is possible due to the use of statistical methods of smoothing and filtering. Preliminary analysis showed that the proposed method can be implemented in the creation of algorithmic software of the self-control system of smart accelerometers.

Indirect Optimization of Finite-Thrust Cooperative Rendezvous

A time-constrained finite-thrust rendezvous between two cooperating spacecraft is investigated in detail in this paper. To ensure high numerical accuracy, the optimization is carried out by means of an indirect method that exploits an “a priori” subdivision of the trajectory into burn and coast arcs, whose time lengths become additional unknown parameters. Issues related to simultaneous presence of two switching control structures, one for each maneuvering spacecraft, are analyzed. A peculiar approach, which relies on the adoption of a different timescale for each spacecraft, is proposed; the introduction of multiple decision vectors allows for an easy management of the switching control structures for missions involving two or more maneuvering spacecraft. As an example, a coplanar rendezvous is thoroughly analyzed; necessary conditions for optimality are comprehensively derived for both target/chaser and cooperative maneuvers. Numerical results are provided and solutions are discussed.

Navigation Solution to Maneuver a Spacecraft Relative to a Sphere Centered on a Cooperative Satellite

A differential correction algorithm is presented to deliver an impulsive maneuver to a satellite to place it within a sphere, with a user defined radius, centered around a non-maneuvering satellite within a constrained time. The differential correction algorithm develops and utilizes the State Transition Matrix along with the Equations of Motion and multiple satellite׳s state information to determine the optimum trajectory to achieve the desired results. The results from the differential correction algorithm are very accurate for prograde orbits, as presented. The results allow for orbit design trade-offs, including satellites׳ initial inclinations, semi-major axes, as well as the ballistic coefficients. The results also provide an empirical method to determine the optimum ΔV solution for the provided problem. Understanding that the minimum fuel solution lies with a semi-major axis ratio of 1, a very accurate empirical approximation is presented for semi-major axis ratio values less than and greater than 1. This work ultimately provides the generalized framework for applying the algorithm to a unique user defined maneuvering spacecraft scenario.

Navigation solution to maneuver a spacecraft relative to multiple satellites and ground locations

A dynamic differential correction algorithm is presented to deliver an impulsive maneuver to a satellite to place it within overlapping spheres, with user defined radii, centered around multiple non-maneuvering satellites and ground location׳s user defined line-of-sight within a constrained time. The solution is further constrained by elevation considerations for the ground locations. The differential correction algorithm develops and utilizes the state transition matrix, along with the equations of motion and multiple satellite׳s state information and ground location state information to determine the optimum trajectory to achieve the desired results. The results from the differential correction algorithm are very accurate for prograde orbits, as presented. The results allow for orbit design trade-offs, including the maneuvering satellite׳s initial inclination. The results deliver an analytical method to determine the optimum ΔV solution for the provided problem. This work ultimately establishes the generalized framework for applying the algorithm to a unique maneuvering spacecraft scenario which includes multiple satellites and ground locations.

Nonlinear Back Stepping Adaptive Controller Design for Spacecraft Attitude Maneuver

The nonlinear backstepping adaptive controller is designed for the large angle fast attitude maneuver of spacecraft with uncertainties in the moment of inertia. Deviation between the actual position and command position is depicted by error quaternion. The nonlinear backstepping tracking function is employed to design the adaptive controller both of command torque and estimation of unknown moment inertia. And the stability is analyzed based on Lyapunovo function. By Matlab/Simulink programming, the simulation of spacecraft attitude manoeuver control is discussed, and its results demonstrate the effectiveness and feasibility of the proposed controller.

Attitude and Vibration Control of Flexible Spacecraft Using Singular Perturbation Approach

This paper addresses a composite two-time-scale control system for simultaneous three-axis attitude maneuvering and elastic mode stabilization of flexible spacecraft. By choosing an appropriate time coordinates transformation system, the spacecraft dynamics can be divided into double time-scale subsystems using singular perturbation theory (SPT). Attitude and vibration control laws are successively designed by considering a time bandwidths separation between the oscillatory flexible parts motion describing a fast subsystem and rigid body attitude dynamics as a slow subsystem. A nonlinear quaternion feedback control, based on modified sliding mode (MSM), is chosen for attitude control design and a strain rate feedback (SRF) scheme is developed for suppression of vibrational modes. In the attitude control law, the modification to sliding manifold for slow subsystem ensures that the spacecraft follows the shortest possible path to the sliding manifold and highly reduces the switching action. Stability proof of the overall closed-loop system is given via Lyapunov analysis. The proposed design approach is demonstrated to combine excellent performance in the compensation of residual flexible vibrations for the fully nonlinear system under consideration, as well as computational simplicity.

Comparison of Preshaping Profilers for Reducing Residual Vibrations After Rest-to-Rest Maneuvers

Generally, spacecraft with large flexible structures have multiple flexible modes with large effective masses and various mode shapes; therefore, the presence of uncertain high-order flexible modes should be considered in high-accuracy controller design. However, most researchers have proposed input preshaping algorithms that are valid only for a few flexible modes. Consequently, when these algorithms are applied to a flexible spacecraft having uncertain high-order flexible modes, residual vibrations might not be reduced sufficiently. This paper presents a smooth feedforward control algorithm for rest-to-rest maneuvers that significantly reduces residual vibrations at the endpoint of a maneuver. The new preshaping profiler is based on a filtering technique that combines a pair of sinc functions. The new profiler is compared to several conventional preshaping profilers. In addition, residual vibrations at the endpoint of maneuvers in the presence of unknown higher modes are found to be much smaller than those in other techniques. Only a hybrid-type profiler shows similar level of small residual vibrations comparable with the new profiler. The new profiler is suitable for rest-to-rest rotational maneuvers of a flexible spacecraft, especially when fine pointing accuracy is required at the endpoint of maneuvers; such accuracy is sometimes required for missions involving scientific observations.

Review of past on-orbit collisions among cataloged objects and examination of the catastrophic fragmentation concept

By mid-2013, four accidental hypervelocity collisions among cataloged objects have been recorded in low Earth orbit. Three were debris–intact impacts, generating very few cataloged fragments (less than 10, in total), while one involved the collision between two intact spacecraft, one of which maneuverable, with the consequent production of more than 2000 cataloged fragments. In order to evaluate the past, current and future average collision rates, the results of debris evolution simulations carried out in 1999–2000 were carefully re-analyzed and the corresponding collision rates, among objects larger than 10cm, were estimated from 1980 to 2040. The value predicted for mid-2013, 0.240 events per year, was further validated by analyzing the current unclassified debris catalog and assessing its qualitative and quantitative incompleteness. It was also found that the intact–intact average collision rate is approximately 30% of the total, while debris–debris impacts account for less than 2%, the majority of the events involving debris–intact collisions. The potential benefits of a wide adoption of collision avoidance practices for maneuverable spacecraft was analyzed as well. If applied to 50% of the current operational spacecraft below 2000km, it would lead to an 11% reduction of the overall collision rate and to a 16% reduction of the intact–intact collision rate. The model estimates from 1980 to mid-2013 were then used to compute the probability of collision occurrence as a function of the time interval considered. It was found that the number of accidental collisions observed so far is in extremely good agreement with the model predictions, corresponding to the maximum likelihood of the probability distributions. Moreover, even the proportion between debris–intact and intact–intact events reflects well what should be expected. Concerning the projections up to 2040, the probability of avoiding further collisions is less than 1%, while there is a probability of more than 50% of having 8 additional collisions, or more. Focusing the attention on the intact–intact events, the probability of avoiding another intact–intact collision is around 8%. At least 1 further collision will have therefore a probability to occur higher than 90%, 2 or more collisions will have a probability higher than 70%, and 3 or more collisions will have a probability higher than 45%, even though the broad adoption of collision avoidance for maneuverable spacecraft might reduce these figures. Finally, the concept of catastrophic collision was re-analyzed, taking into account the observational record and the basic assumptions usually adopted by the long-term debris evolution models. Potential sources of collision fragments over- and underestimation were identified and discussed, coming to the preliminary conclusion that after one century the number of collision fragments currently predicted by the evolutionary models might be overestimated.

Cassini Thruster Calibration Algorithm Using Reaction Wheel Biasing Data

Thrust force estimates for the reaction control thrusters onboard the Cassini spacecraft are presented in this paper. Cassini consists of two thruster branches, each with eight thrusters. The eight thrusters in each branch control the three body-fixed axes of the spacecraft. It is important to track the thrust force estimates in order to detect any thruster degradation and for supporting various activities in spacecraft operations (Titan flyby, spacecraft maneuvers, etc.). The Euler equation, which describes the rotational motion of the spacecraft during a reaction wheel bias event (wheel momentum change), is used to develop the algorithm. The thrust estimates are obtained from the pseudoinverse solution using flight telemetry during the bias. These results are similar to thrust estimates from the Kalman filtering algorithm. Cassini operations used the first thruster branch from launch until November 2008. Results show that two thrusters in this branch exhibited degraded thrust in November 2008. Because of the degraded thrust performance, the first branch usage was discontinued, and the prime branch was swapped to the second thruster branch in March 2009. The thrust estimates from this branch do not show any degradation to date. The algorithm is used to trend the thruster branch thrust force estimates as the mission continues.

Validation of a Compression Mass Gauge using ground tests for liquid propellant mass measurements

To properly estimate orbital lifetimes and predict the maneuverability of spacecraft, the remaining liquid propellant mass must be accurately known at every moment of a space mission. This paper studies the Compression Mass Gauge (CMG) method to determine the mass of liquid contained in a tank in a low-gravity environment with high accuracy. CMG is a thermodynamic method used to determine the quantity of liquid by measuring the gas pressure change when the tank volume changes, and has been previously theoretically and experimentally studied by researchers. The primary objective of this investigation is to explore the effects of attitude disturbance and the spacecraft thermal environment on the accuracy of the method. A ground test system, consisting of several test apparatuses, was fabricated and described as part of this study. The test results and analyses indicate that the CMG performs well and has an accuracy of ±1%. Additionally, demonstrations were performed to show that measurement errors do not increase drastically or exceed ±1% when the test system is vibrated to simulate the tank being perturbed as a result of an attitude disturbance. Liquid sloshing resonance was found to have a significant effect on the gauging accuracy. Measurements in a real thermal environment in which heat transfers into and out of the propellant tank were also conducted. The results show that the gauging accuracy is acceptable for normal liquid propellant. Furthermore, theoretical research shows that heat leakage has a significant influence on cryogenic propellant mass gauging.

Responsiveness in Low Orbits Using Electric Propulsion

One promising option for space operational responsiveness is orbital maneuvering. In low orbit, a maneuverable spacecraft can provide valuable benefits such as better coverage properties, reduced revisit times, selectable targets, and local overflight times. Such maneuvers are not common due to the high cost of chemical propulsion. The more recent paradigm of operationally responsive space is to rapidly launch a small inexpensive asset and use it in a short disposable fashion. This concept relies on drastically reducing the cost of launch, yet it remains the most expensive piece, and so additional cost savings can be realized by minimizing the need for launches. Electric propulsion has been considered as an efficient alternative to chemical propulsion. With technological advances, electric propulsion can provide responsiveness in a timely propellant-efficient manner without requiring repeated launches to satisfy multiple missions. This study shows the control algorithm for a single low-Earth satellite equipped with the proper electric propulsion to overfly any target inside its coverage area in as little as 34 h for 1.8% of its propellant budget. A comprehensive survey to quantify global reach requirements is provided, and the optimal time and propellant solutions are explored. The results strongly support the argument that electric propulsion could be a key enabler in responsive operations.