Orbit Control System Research Articles

Carbon films deposition as protective coating of titanium alloy tube using PIII&D system

The purpose of this work is to deposit carbon films inside titanium alloy tube of small diameter, in order to increase the corrosion resistance of its inner surface. The use of carbon films as protective coating can increase the lifetime of the tubes used in propulsion and thermal control systems of satellites. They can also inhibit the permeation of corrosive species, originated from the fuels and the cooling fluids, between the film-metal interface. Titanium alloy tube with 11 mm diameter and 150 mm length, with one side closed configuration was placed inside a vacuum chamber and pulsed to a high energy voltage. Carbon film deposition was performed by using a Plasma Immersion Ion Implantation and Deposition (PIII&D) system. The hollow cathode discharges (HCD) were produced with pulses of 2 kV, 15 A, 20 μs, and 500 Hz of repetition rate, using a RUP-6 pulser. The maximum temperature measured at the surface of tubes was about 1000 °C. The as-deposited carbon films were analyzed in terms of their microstructure, chemical composition, surface morphology and thickness. The surface characterization of the carbon films were carried out by means of Raman Spectroscopy, Field Emission Gun Scanning Electron Microscopy (FEG-SEM), Atomic Force Microscopy (AFM) and Potenciodynamic Polarization tests. The experimental results show an interesting change on the chemical structure of the as-deposited carbon films, when the deposition time varied from 30 min to 240 min, resulting in thicker and quite adhered films on the inner surface of tubes. In addition, the changes on the microstructure, seen by SEM analysis, as well as the ID/IG variation indicate that nanocrystalline graphite (NC-G) films can be deposited on the inner surface of titanium tubes by using this type of PIII&D system.

The Critical Space Infrastructure and its Importance to Military Operations

Abstract Actual scientific approaches related to critical infrastructure reveal its importance within many of the society’s essential services that are necessary for day-to-day people’s life. Besides other critical infrastructure systems, the emerging development of the space system results in the embedment of it within the critical space infrastructure concept that is based on its description related to the importance of the services provided to population, public or governing institutions, and industry. This paper intends to present the manner in which the space system services influence population’s life and to emphasize the importance of the space infrastructure related to the process of planning and conducting military operations. The reliability and well functioning of the space system from the perspective as a critical infrastructure proves its paramount importance related to essential services provided to the population that results into social well-being. Nowadays, the majority of military operations depend on the development of space infrastructure and its reliability. Moreover, the military space operations are based on satellite control system architectures as well as the interconnected systems on surface and underground systems. From this side, the protection of critical space infrastructure and its resilience level should have a paramount value for all stakeholders that are using the space system services.

Design of Robust Observer-Based Backstepping Control for a Satellite Control System

This paper presents the attitude tracking of a class of satellite control systems under external disturbances. Once the error dynamics of the satellite are obtained, a nonlinear transformation expresses them in a suitable representation for the design of a high-gain observer and backstepping control. The observer-based backstepping controller is then designed to drive the angles of the satellite dynamics to their desired values in the presence of exogenous disturbances. Closed-loop stability of the proposed controller is demonstrated via Lyapunov theory, and its effectiveness is confirmed through numerical simulations.

EVALUATION OF THE RELIABILITY OF COMPLEX TECHNICAL RECONFIGURATION SYSTEMS BY STATISTICAL MODELING METHOD

Goal. Assess the reliability of a complex technical system with periodic reconfiguration and compare the results obtained a similar system, but without reconfiguration. Materials and methods. In this article uses the method of statistical modeling (Monte Carlo) to assess the reliability of complex system. We using the normal and exponential distribution of failure time for modeling failures of system elements. Reconfiguration algorithm is the algorithm proposed for the attitude and orbit control system of spacecraft. Results. A computer program has been developed for assessing reliability on the basis of a statistical modeling method, which makes it possible to evaluate systems of varying complexity with exponential and normal distribution, as well as with and without periodic reconfiguration. A quantitative estimate of the reliability as a function of the probability of system failure is obtained. Conclusion. It has been demonstrated that a system with reconfiguration has the best reliability characteristics, both in the case of exponential and normal distribution of failures.

Design concept for In-Drag Sail with individually controllable elements

The lifetime of the existing attitude and orbit control systems (AOCS) generally depends on the amount of the fuel carried on a spacecraft. This paper proposes a novel design concept for a low Earth orbit (LEO) spacecraft named In-Drag Sail with Individually Controllable Elements (IDSICE), which can release the fuel's limitation to its attitude control system. IDSICE is designed to have four different working modes, which are Compression and Storage mode (CS mode), Orbital Flight mode (OF mode), Attitude Control mode (AC mode), and De-Orbiting mode (DO mode). By switching between different modes, it can achieve different missions. When IDSICE is in AC mode, the aerodynamic torques can be applied to completed the attitude control phase. When IDSICE is in DO mode, its area-to-mass ratio is enlarged so that the orbit altitude can decrease rapidly to achieve the de-orbiting phase at the end of the life. The decentralized layout is applied in IDSICE and most components have redundant backups, which allows a higher fault tolerance compared with the conventional layout. Finally, some necessary simulations are made, which verified the feasibility of the attitude control and de-orbiting methods proposed based on IDSICE.

AOCS Requirements and Practical Limitations for High-Speed Communications on Small Satellites

In recent years, an increasing number of countries have shown a growing interest in developing their indigenous space capacity building through national small satellite programs. These satellites, which were initially focused on educational and training missions, currently are more scientific and operational-oriented. Thus, small satellite missions are being considered not only as educational tools but also as technological demonstrators or, even, mature enough for commercial and scientific missions, which might generate a huge amount of data to be transmitted to the ground segment. Therefore, an increasing demand on channel capacity will be needed for downloading the generated housekeeping and scientific data for missions based on small satellites. This paper analyses the communication subsystem of a real Cubesat. The influence of geometrical parameters is rigorously calculated both in the signal-to-noise ratio and in the capacity to transmit information. Subsequently, which parameters of the radio link can be modified to increase the transmission capacity, including the pointing requirements and its practical implementation, is studied. Finally, and as a future line, the technical feasibility of using optical links on small satellites that might greatly increase the transmission capacity, including the satellite pointing problems that presents, is presented. In conclusion, this paper presents a rigorous calculation in different frequency bands of the signal-to-noise ratio and the pointing accuracy that is needed to achieve the maximum transmission speed from the satellite to the ground station, and therefore the requirements that the Attitude and Orbital Control Systems (AOCS) must have, as well as the limitations of current systems.

A reference architecture for satellite control systems

Software for satellite control systems (SCS) domain performs a relevant role in space systems, being responsible for ensuring the functioning of the satellites, from the orbit launch to the end of their lifetime. Systems in this domain are complex and are constantly evolving due to technological advancement of satellites, the significant increase in controlled satellites, and the interoperability among space organizations. However, in order to meet such complexity and such evolution, the architectures of these systems have been usually designed in an isolated way by each organization and hence may be prone to recurrent efforts and difficulties of interoperability. In parallel to this scenario, reference architecture, a special type of software architecture that aggregates knowledge of a specific domain, has performed an important role for the success in development, standardization, and evolution of systems in several domains. Nevertheless, the usage of reference architecture has not been explored in the SCS domain. Thus, this article presents a reference architecture for satellite control systems (SCS-RA). Results achieved from usage of SCS-RA in the development of a microsatellite control system for National Institute for Space Research showed a significant reduction of effort, benefits of interoperability, scalability, and sharing of ground resources.

Active Fault-tolerant Gyromoment Control of Satellites and Free-flying Robots

Abstract The methods for modeling and detection of anomalous in the automatic control systems, and also practical methods for active analysis of anomalous situations, are presented. We develop algorithms for consecutive classification of onboard equipment failures and reconfiguration of control loop. The simulation results are presented for the attitude control systems of the information satellites and space robots.

SDRE Control with Nonlinear J2 Perturbations for Nanosatellite Formation Flying

Abstract Nonlinear control for satellite formations is a requirement for challenging missions with large intersatellite distances where spherical gravitational nonlinearities and differential disturbances are significant. This paper presents the design of a formation flying orbit control system for low Earth orbit (LEO) satellites based on the State-Dependent Riccati Equation (SDRE) control approach. An exact nonlinear differential J2 perturbation model is used for the relative orbital dynamics, and SDRE tracking is posed as an error regulation problem. Two State-Dependent Coefficient (SDC) matrices are formulated that avoid singularities such that the controller does not suffer from chattering effects. The additional degree of freedom associated with the SDC parameterizations is exploited using a weighting parameter to tune the control performance. Stable solutions in the domain of interest are obtained and results for thrust and time of formation acquisition for the proposed SDRE controller are presented.

In-Orbit Spacecraft Inertia, Attitude, and Orbit Estimation Based on Measurements of Magnetometer, Gyro, Star Sensor and GPS Through Extended Kalman Filter

Spacecraft may exhibit changes of its moments and products of inertia in-orbit due to several reasons. In this case, knowledge of spacecraft moments and products of inertia represents an utmost importance for the spacecraft attitude and orbit control system (AOCS). The contribution of the current research article is to estimate in-orbit spacecraft moments and products of inertia based on measurements of magnetometer, gyro, star sensor, and GPS through extended Kalman filter without test slews while being able to estimate attitude and orbital states at the same time. The proposed solution is also able to converge despite large initial estimation errors.

Model-based deep reinforcement learning with heuristic search for satellite attitude control

PurposeRecently, deep reinforcement learning is developing rapidly and shows its power to solve difficult problems such as robotics and game of GO. Meanwhile, satellite attitude control systems are still using classical control technics such as proportional – integral – derivative and slide mode control as major solutions, facing problems with adaptability and automation.Design/methodology/approachIn this paper, an approach based on deep reinforcement learning is proposed to increase adaptability and autonomy of satellite control system. It is a model-based algorithm which could find solutions with fewer episodes of learning than model-free algorithms.FindingsSimulation experiment shows that when classical control crashed, this approach could find solution and reach the target with hundreds times of explorations and learning.Originality/valueThis approach is a non-gradient method using heuristic search to optimize policy to avoid local optima. Compared with classical control technics, this approach does not need prior knowledge of satellite or its orbit, has the ability to adapt different kinds of situations with data learning and has the ability to adapt different kinds of satellite and different tasks through transfer learning.

Reliability assessment of phased-mission systems under random shocks

Phased-mission systems (PMSs) are widely used, especially in the aerospace industry. As in the outer space there are many kinds of cosmic rays, such as the Galactic Cosmic Rays (GCR), randomly hitting on these systems and causing significant impact on the electronics inside or outside the equipment, a reliability model for PMSs considering both finite and infinite random shocks is proposed in this paper. The modularization method is used to simplify the state space model for each phase and reduce the amount of system states, and the Markov regenerative process (MRGP) is used to describe the hybrid components’ lifetime distributions and the dynamic behaviors within the modules. Then, two kinds of scenarios, finite and infinite random shocks effect, are both integrated into the dynamic modules. For demonstration, a phased altitude and orbit control system (AOCS) subjected to infinite random shocks is illustrated to demonstrate the procedure of the proposed Monte Carlo simulation. Thirdly, the evaluated system reliability under infinite random shocks is compared with the same system without considering random shocks. At last, a sensitivity analysis is also provided for completion.

Design and Control of a Three-Coil Permanent Magnet Spherical Motor

The permanent magnet (PM) spherical motor has been subject to growing interest from the scientific community due to its potential for applications in distinct areas, particularly in robotics, prosthetics, satellite control, sensors or camera systems. Motivated by this movement, the current work presents all the steps for the efficient design and construction of a spherical motor model, using compound deposition technology with the aid of a 3D printer. Furthermore, we report comprehensive studies on the accuracy of the positioning system of the proposed motor using only three stator coils, which jointly act to move the rotor axis toward any point in the hemisphere. Unmodeled nonlinear phenomena, such as friction, impair accurate positioning of the motor actuator, but this is solved by means of a visual servo control system, which allows the user to collect input–output data to train an artificial neural network model. Details on the construction of the proposed motor are reported, in addition to the numerical results of the positioning control in tracking a desired trajectory.

Computational modeling of nonlinear propellant sloshing for spacecraft AOCS applications

During all operational phases, propellant sloshing may have a significant influence on spacecraft performance and stability. The proper description of propellant sloshing effects is then essential for the verification of the Attitude and Orbit Control System (AOCS). As soon as nonlinear effects predominate (e.g. for high-agility or in a micro-gravity environment), classical approaches as mechanical analog models (e.g. pendulums) fail and computational models need to be utilized to describe the propellant dynamics. These computational models currently lack the ability to accurately describe nonlinear effects like high-velocity impacts, as well as cohesion and adhesion forces in a micro-gravity environment. Additionally these models are computationally very expensive, so that they are commonly not directly used in AOCS verification campaigns, where a large number of simulations are performed. The scope of this work is the development of a computational model based on smoothed particle hydrodynamics, which is able to accurately describe nonlinear propellant sloshing effects in gravity-dominated regimes, which is of relevance for example for high-agility spacecraft missions. The newly developed code is validated using analytical expressions, experimental data as well as other numerical codes. Subsequently, a study is performed showing that the global variables like forces acting on the tank wall or the center of mass are still approximated reasonably well when using lower spatial resolutions, resulting in much faster simulation runs and making it feasible to use the computational model directly in AOCS verification campaigns. Finally, the coupling of the propellant sloshing code to an AOCS and Flight Dynamics rigid-body simulator is demonstrated by simulating a nonlinear forced-roll motion of a partially-filled tank under Earth-gravity.

AOCS design for the ATHENA X-ray telescope: challenges and solutions

The ATHENA—Advanced Telescope for High-ENergy Astrophysics—mission is currently assessed in a phase A feasibility study as L-class mission in ESA’s Cosmic Vision 2015–2025 plan, with launch foreseen in 2028. Primary mission goal is the mapping of hot gas structures and the determination of their physical properties to search for supermassive black holes. ATHENA is an X-ray telescope with a focal length of 12 m. It has a mass of ~ 7000 kg and it is ~ 15 m high with a diameter of ~ 3 m. The main mass is distributed to the mirror on the one side of the spacecraft and to the science instrument module on the other side of the spacecraft. To achieve its science goals, ATHENA performs a sky survey with precision line-of-sight pointing requirements in the order of arc seconds for absolute pointing and sub-arc seconds for relative pointing in time windows > 1 ks, all at 95% confidence level. That is very demanding for large X-ray telescopes. In addition to the precision pointing requirements, ATHENA cannot violate a certain sun exclusion zone. This is a hard constraint to prevent any stray-light falling onto the instruments, as it would immediately destroy them. The sky survey is defined by an observation plan that is demanding in terms of availability and thus spacecraft agility. The pointing and agility requirements and the fact that ATHENA is a spacecraft with high mass and volume introduce several design challenges for the attitude and orbit control system. This paper presents those challenges, corresponding solutions, and preliminary results, which have been achieved during the phase A study led by Airbus in Friedrichshafen, Germany. The main focus and contribution of this paper are the identification of research and development needs for attitude and orbit control systems to enable the ATHENA mission. In this respect, the ATHENA design challenges are discussed and addressed with the state-of-the-art design methods. This paper concludes with the main identified technology development needs and formulates specific research questions related to practical design problems. In particular, the following attitude and orbit control system design challenges are addressed: autonomous and agile large angle slew manoeuvres with exclusion zones, availability for science observations, precision line-of-sight determination as well as analysis during the design process using the ESA Pointing Error Engineering Tool and pointing control with a hexapod as line-of-sight actuator in the control loop. The last challenge, namely, the hexapod in the control loop, is without precedence in Europe and to the best knowledge of the authors in the world.

Evaluation of calibration accuracy of magnetometer sensors of Aist small spacecraft

In the paper the technique of estimation of calibration accuracy of magnetometer gauges by the example of an Aist small spacecraft is stated. According to the measurement of the Earth's magnetic field in the orbital flight of a small spacecraft, the parameters of its rotational motion around the center of mass are estimated and primary information is generated for the magnetic actuators of the orbital motion control system. Therefore, calibration of the magnetometer sensors at the ground test stage is essential for the successful execution of the flight program. The technique can be used at the stages of ground and flight tests of magnetic field measuring instruments.

Reliability assessment of multi-state phased mission system with non-repairable multi-state components

• A MMDD algorithm for PMS and a PMS-MMDD model are proposed to model the non-repairable multi-state phased-mission system. • A Markov renewal equation-based method is proposed to evaluate system probabilities with non-exponential distributions. • The comparison to the Monte Carlo simulation method illustrate the computation accuracy and efficiency of the proposed methods. • An attitude and orbit control system in spacecraft is studied as a practical example to illustrate the proposed methods. Phased mission systems (PMSs) like satellites and spacecraft perform their functions over non-overlapping mission periods, called phases. One of the challenges in assessing reliability of PMSs comes from considering the s -dependence among phases, and the consideration on the multi-state behavior of components and systems makes the reliability analysis even more difficult. To effectively address this problem, a multi-state multivalued decision diagram algorithm for PMS and a multi-state multi-valued decision diagram model for phased mission system (PMS-MMDD) method is developed for the reliability modelling of non-repairable multi-state components. Based on the Semi-Markov process, a Markov renewal equation-based method is developed to deal with non-exponential multi-state components and a numerical method, the trapezoidal integration method, is used to compute the complex integrals in the path probability evaluation. A case study of a multi-state attitude and orbit control system in a spacecraft is analyzed to illustrate the proposed PMS-MMDD model and the Markov renewal equation-based evaluation method. The accuracy and computation efficiency of the proposed method are verified by the Monte Carlo simulation method.

Reliability analysis of phased mission system with non-exponential and partially repairable components

Phased mission systems (PMSs) have wide applications in engineering practices, especially in aerospace industry such as man-made satellite and spacecraft. To achieve high reliability in a PMS, certain critical parts in the system are designed to have a redundant architecture, such as cold standby (structural or functional). State-space models such as Markov processes have been widely used in previous studies to evaluate the reliabilities of these systems. But in practice, many real systems consist of mechanical components or mechatronics whose lifetime follow non-exponential distributions like the Weibull distribution. In this type of system, the Markov process is not capable of modeling the system behavior. In this paper, the SMP (Semi-Markov Process) is applied to solve the problem that the components’ lifetime in dynamic systems follows non-exponential distributions. An approximation algorithm for the SMP is proposed to assess the reliability of the PMSs consisting of non-exponential components. Furthermore, the accuracy and calculation efficiency of the approximation algorithm are explored. At last, the reliability assessment of a complex multi-phased altitude and orbit control system (AOCS) in a man-made satellite is presented to illustrate the method.

A data-driven method of health monitoring for spacecraft

PurposeThe purpose of this paper is to detect the occurrence of anomaly and fault in a spacecraft, investigate various tendencies of telemetry parameters and evaluate the operation state of the spacecraft to monitor the health of the spacecraft.Design/methodology/approachThis paper proposes a data-driven method (empirical mode decomposition-sample entropy-principal component analysis [EMD-SE-PCA]) for monitoring the health of the spacecraft, where EMD is used to decompose telemetry data and obtain the trend items, SE is utilised to calculate the sample entropies of trend items and extract the characteristic data and squared prediction error and statistic contribution rate are analysed using PCA to monitor the health of the spacecraft.FindingsExperimental results indicate that the EMD-SE-PCA method could detect characteristic parameters that appear abnormally before the anomaly or fault occurring, could provide an abnormal early warning time before anomaly or fault appearing and summarise the contribution of each parameter more accurately than other fault detection methods.Practical implicationsThe proposed EMD-SE-PCA method has high level of accuracy and efficiency. It can be used in monitoring the health of a spacecraft, detecting the anomaly and fault, avoiding them timely and efficiently. Also, the EMD-SE-PCA method could be further applied for monitoring the health of other equipment (e.g. attitude control and orbit control system) in spacecraft and satellites.Originality/valueThe paper provides a data-driven method EMD-SE-PCA to be applied in the field of practical health monitoring, which could discover the occurrence of anomaly or fault timely and efficiently and is very useful for spacecraft health diagnosis.

Mathematical model of a flexible asymmetrical rotor for active magnetic bearing reaction wheel

The paper introduces the mathematical model of rotor for active magnetic bearing reaction/momentum wheels, used as actuator in spacecraft attitude and orbit control system. Developed model is used for estimation of critical speeds and forced oscillation magnitudes with a glance of the rotor modes. Rotor is considered as a two-mass system, consisting of a shaft and a rim, active magnetic bearings are assumed to be a linear elastic springs. The equations of the rotor motion are derived using the Lagrange equation. Developed model is verified by comparing the calculated Campbell diagrams with the results of the finite-element modal analysis, performed in the ANSYS software.